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"Launch into Space",1,0,0,0
Our age-old fascination with the stars and planets has put us on a path of discovery that continues today. Investigate our history in space, from the early Soviet satellites to the Mars Pathfinder Mission and read about upcoming space projects. You can also visit the space photo gallery, which contains over 100 color photographs taken on American and Russian space missions.
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"Space Table of Contents",2,0,0,0
This section of \IWebster's World Encyclopedia\i contains in-depth text and pictorial information on space history as well as upcoming space exploration and projects.
It is divided into eight parts:
\JEarly Space Missions\j
\JMajor Moon Missions\j
\JNew and Upcoming Space Projects\j
\JSoviet/Russian Space Programs\j
\JSpace Photo Gallery\j
\JSpace Shuttle Information\j
\JYear in Space\j
\JMonth in Space\j
Click on the entry of your choice to move to that topic.
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"Early Space Missions",3,0,0,0
This part covers information on the following space missions:
\JClementine Project\j
\JGalileo Project Information\j
\JGiotto Mission\j
\JHubble Space Telescope (HST)\j
\JMagellan Program\j
\JMariner Program\j
\JMars Observer Project\j
\JPioneer Venus Program\j
\JSakigake and Suisei Projects\j
\JSkylab Program\j
\JUlysses Project Information\j
\JViking Mission to Mars\j
\JVoyager Project Information\j
Click on the entry of your choice to move to that part.
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"Major Moon Missions",4,0,0,0
This part covers information on the following space missions:
\JNASA Mercury Project\j
\JNASA Gemini Project\j
\JRanger (1964 - 1965)\j
\JSurveyor (1966 - 1968)\j
\JLunar Orbiter (1966 - 1967)\j
\JApollo 11, Twenty-Five Years On\j
\JApollo Program (1968 - 1972)\j
\JApollo Missions Contents\j
\JApollo 13 Mission\j
\JKey Documents from the Apollo Space Program\j
\JRemarks by the President at the 25th Anniversary of Apollo 11\j
Click on the entry of your choice to move to that part.
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"New and Upcoming Space Projects",5,0,0,0
\JInternational Space Station\j
\JMars Global Surveyor Updates\j
\JMars Pathfinder Status\j
\JMir 23 Status Report\j
\JShuttle Columbia Mission\j
\JSpace Calendar June97 - July98\j
\JSpace Launches 1997 (Worldwide)\j
\JMir Space Station\j
\JShuttle Missions (Upcoming)\j
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"Soviet/Russian Space Programs",6,0,0,0
\JSoviet/Russian Space Programs continued \j
\JSoviet/Russian Space Programs continued 2\j
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"Space Photo Gallery",7,0,0,0
This part provides over 100 color (plus some black and white) pictures of the different planets taken on various American and Russian space missions. It includes photographs of the following planets and planetary bodies:
\JAsteroid Photo Gallery\j
\JComets\j
\JEarth and Moon Photo Gallery\j
\JEarth Photo Gallery\j
\JJupiter Photo Gallery\j
\JMars Photo Gallery\j
\JMercury Photo Gallery\j
\JMoon Photo Gallery\j
\JNeptune Photo Gallery\j
\JPluto Photo Gallery\j
\JSaturn Photo Gallery\j
\JUranus Photo Gallery\j
\JVenus Photo Gallery\j
Click on the entry of your choice to move to that part.
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"Space Shuttle Information",8,0,0,0
\JShuttle CHALLENGER (10) 51-L (25)\j
\JSpace Shuttle, Living in a\j
\JShuttle Location\j
\JNASA Facilities\j
\JShuttle Flights To Date\j
\JSpace Shuttle Launch Team\j
\JSpace Shuttle Mission Summary (1972-1988)\j
\JSpace Shuttle System\j
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"Year in Space",9,0,0,0
\JYear in Space 1994\j
\JYear in Space 1995\j
\JYear in Space 1996\j
\JSpace News Update 1997\j
\JSpace News Update 1998\j
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"Month in Space",10,0,0,0
\JMonth in Space - July\j
\JMonth in Space - June\j
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"International Space Station",11,0,0,0
April 9, 1997
"International Space Station: \JEngineering\j the Future"
The International Space Station program promises a new era of space exploration and space-based scientific research, and will come to fruition through an unprecedented level of international cooperation. When complete, the Space Station will be permanently occupied by a crew of six, and it is anticipated that it will remain fully operational for ten years following its planned completion in June, 2002.
The International Space Station program is divided into three phases, each defined by the inception of new capabilities:
PHASE I (1994-1997) is currently underway, and utilizes existing resources, primarily the Space Shuttle and Russian Mir space station to build experience in the technical aspects of space station construction and occupation, and to usher in a new area of international cooperation in space.
Phase I began with the launch of Discovery for Shuttle mission STS-60, when Russian \Jcosmonaut\j Sergei Krikalev became the first Russian to fly aboard a U.S. \Jspacecraft\j. Phase I will be completed following the ninth Shuttle/Mir docking mission, currently envisioned to be STS-91, in May, 1998, which will carry the Alpha Magnetic Spectrometer (AMS) and other scientific experiments specially selected to maximize the utilization of the space station environment.
PHASE II (November, 1997 - February, 1999) will mark the beginning of the assembly phase of the International Space Station, and will see the first "production" experiments conducted aboard the Station. Phase II will commence upon launch, aboard a Russian Proton launch vehicle, of the Russian-built "FGB" module, a propulsion and attitude control module of a design proven years ago aboard Russian military flights.
In May, 1998, permanent occupation of the station will commence with a crew of three, and the Canadian Mobile Servicing System will be launched in late 1998. The end of Phase II will be marked by the delivery of the U.S. Laboratory Module aboard Shuttle flight STS-94. (Refer to Table)
PHASE III (February, 1999 - June, 2002) will see permanent occupation of the partially completed International Space Station by a crew of six. Construction will continue along with flights tailored to Station utilization. The first Phase III flight will be Shuttle mission STS-96, which will also be the first Space Station Utilization flight.
After the permanent docking of a second Soyuz craft to serve as an escape "lifeboat" and the installation of the U.S. Habitation Module aboard flight 19-A, STS-122, in June, 2002, the Station will be able to support a full-time six-person crew. Phase III will include the addition of the Japanese Experiment Module (JEM) in 2000 and the European Columbus Orbiting Facility (COF) in late 2002 or early 2003.
Phase III will draw to a close with the flight of Shuttle mission STS-121 and the addition of the Station's largest module, the U.S. Habitation Module. With assembly now complete, full operational use of the International Space Station will begin. (Refer to Table)
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"Mars Global Surveyor Updates",12,0,0,0
\JMars Surveyor Report (January)\j
\JMars Surveyor Report (February)\j
\JMars Surveyor Report (March)\j
\JMars Surveyor Report (April)\j
\JMars Surveyor Report (May)\j
\JMars Surveyor Report (June)\j
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"Mars Surveyor Report (January)",13,0,0,0
Friday, 10 January 1997
This week marked the transition from the inner-cruise to the outer- cruise phase of the mission. One of the first transition tasks occurred early Monday morning when the flight team sent a set of commands to change Surveyor's pointing orientation. The commands turned the \Jspacecraft\j from its previous orientation of +X axis pointed 60 degrees away from the Sun to a position where the +X axis is pointed directly at the Earth.
One of the benefits of this new pointing orientation is that Surveyor can now use its high-gain antenna to communicate with the Earth. This antenna is mounted on the \Jspacecraft\j's +X axis and its narrow-beam signal requires that the \Jspacecraft\j point directly at Earth.
Until now, Surveyor was utilizing its wide-beam, low-gain antenna for communications. The high-gain antenna broadcasts with greater power and will allow the \Jspacecraft\j to transmit data at higher data rates.
Before January, it was impossible to use the high-gain antenna because an Earth-pointed orientation would have placed Surveyor at an unfavorable angle with respect to the Sun. The switch from the low-gain to the high-gain antenna occurred early Thursday morning.
The flight team is continuing to diagnose the position discrepancy in Surveyor's -Y solar panel which is deployed, but 20.5 degrees from its proper position. \JEngineering\j data transmitted to Earth during the five solar array "wiggle tests" conducted in December supports the current model regarding the nature of the obstruction keeping the array out of position.
The model suggests that a damper shaft in the solar array's deployment mechanism broke shortly after launch, approximately 43 seconds after the start of the array's deployment. This damper is a device that was installed to minimize the mechanical shock of deployment by slowing the motion of the array during deployment.
The flight team theorizes that the broken shaft caused the damper arm to wedge into the hinge joint connecting the solar panel to the \Jspacecraft\j. Attitude-control telemetry recorded by the \Jspacecraft\j during solar array deployment corroborates this theory.
Plans are currently being developed for three more solar array "wiggle tests" during the week of January 20th. Data from these upcoming tests and the five previous tests in December will assist the flight team in determining the best method to attempt to free the damper arm from the hinge joint.
Today, the flight team transmitted the C4 sequence to Surveyor. C4 contains commands that will control Surveyor for the next five weeks. The first activities in C4 will start on January 13th and will involve using the Mars Orbiter Camera to image stars over four consecutive days. These star images will allow the camera team to refine the camera's focusing capability.
After a mission elapsed time of 64 days from launch, Surveyor is 14.79 million kilometers from the Earth and is moving in an orbit around the Sun with a velocity of 31.32 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
Friday, 17 January 1997
On Monday of this week, Surveyor's flight team activated the Mars Orbiter Camera in preparation for four days of star imaging. Once per afternoon from Tuesday through Friday, the \Jspacecraft\j turned to point the camera at a cluster of stars called the Pleiades. Over the course of one hour on each imaging day, the camera observed stars within the cluster in order to perform focus checks.
Communications with the \Jspacecraft\j during star imaging was not possible because the star-pointed orientation resulted in pointing the high-gain antenna away from the Earth. Consequently, all of the data from the camera was stored on Surveyor's solid-state recorders.
This data was transmitted back to Earth approximately three hours after the conclusion of each day's imaging. The daily playback of camera data required 49 minutes. During that time, Surveyor transmitted 250 megabits of data at a downlink rate of 85,333 bits per second.
Next week, the onboard flight computer will activate heaters in the camera that will bake the epoxy structure of the camera to remove residual moisture. A set of four more star images will be taken after the bakeout period ends in late March. The star images taken this week will serve as a reference to assess the focusing capability of the camera after the bakeout.
Other activities this week included a two-hour radio-science \Jcalibration\j that occurred late in the evening on Wednesday. This test involved using the \Jspacecraft\j's ultra-stable \Joscillator\j to control the frequency or "tone" of Surveyor's radio transmissions to the Earth.
Normally, the \Jspacecraft\j listens to a signal transmitted from the Earth as a reference to set the tone of the signal transmitted to Earth. The \Joscillator\j functions as an electronic clock that can precisely control the tone of Surveyor's signal without listening to the Earth-based reference signal.
Future tests of the \Joscillator\j will occur approximately every other week until the \Jspacecraft\j reaches Mars. These tests are important because a stable radio signal as controlled by the \Joscillator\j will be critical toward the collection of scientific data at Mars.
After a mission elapsed time of 71 days from launch, Surveyor is 16.05 million kilometers from the Earth, 136.00 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 30.85 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
Friday, 31 January 1997
Early Monday morning, flight controllers sent several commands to Surveyor that deactivated the Mars Orbiter Camera's 53-Watt bakeout heater. This heater was activated on Wednesday, January 22nd to remove residual moisture in the camera's \Jgraphite\j epoxy structure. If the bakeout had not been performed, the moisture in the camera's tube-like structure would have slowly leaked into space and caused its length to gradually change.
As a consequence, this tiny, slow-rate change in the structure's length would have resulted in a gradual shift in the focus of the camera during science operations. The goal of the bakeout was to remove all of the moisture at once in order to stabilize the focus of the camera.
Originally, the bakeout was scheduled to last for 60 days. This duration was subsequently reduced to 14 days last Wednesday when data from the camera suggested that the structure contained significantly less moisture than predicted.
Upon request from the camera team, the flight operations manager made the decision to terminate the bakeout after only six days. The concern is that baking the camera for longer than necessary would be detrimental to the camera's focusing capability.
In several weeks, the camera will image stars over a one-week period for the purpose of acquiring focus \Jcalibration\j images. These images will be compared to the star images taken before bakeout in order to assess the best focus settings for the camera.
Other activities this week included a two-hour radio-science \Jcalibration\j that occurred Thursday morning, just after midnight. This test involved using the \Jspacecraft\j's ultra-stable \Joscillator\j to control the frequency or "tone" of Surveyor's radio transmissions to the Earth.
Later on Thursday, flight controllers sent a command that activated a flange heater located near Surveyor's main rocket engine. The heater will gradually increase the pressure of the \Jnitrogen\j tetroxide inside the oxidizer tank.
As a consequence, the increase in oxidizer pressure will improve the efficiency of the propellant during the second trajectory correction maneuver. This maneuver is currently scheduled for March 20th.
After a mission elapsed time of 85 days from launch, Surveyor is 19.29 million kilometers from the Earth, 116.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 29.83 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
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"Mars Surveyor Report (February)",14,0,0,0
Friday, 7 February 1997
Today, the flight team sent a command to Surveyor to activate the Mars Orbiter Camera. Over the weekend, the camera team will collect temperature data from the instrument in order to determine the best focus setting for a focus check test that will be performed on Tuesday, February 11th.
Earlier in the week, the flight team completed \Jcalibration\j activities on the gyroscopes in the inertial measurement unit. These gyroscopes are devices that provide critical data to the flight computers regarding Surveyor's pointing orientation in space. Each one of the three gyroscopes on the \Jspacecraft\j has a primary and backup data channel.
Over the course of a several day period, the \Jspacecraft\j team examined data from the backup \Jgyroscope\j channels in order to understand the slight variations between the in-flight performance and the performance as specified by the manufacturer.
The knowledge of these minor variations were incorporated into Surveyor's flight software. This activity was performed to improve the \Jspacecraft\j's ability to maintain a proper orientation in the event that the backup \Jgyroscope\j channels are used.
Throughout this past week, the Magnetometer science instrument has also been active. The data collected during the week will provide the Magnetometer team with an opportunity to conduct further calibrations on the instrument. In addition, the data will provide the team with an opportunity to study the solar wind.
This "wind" is a stream of protons and electrons that are constantly blown out from the Sun at a speed of 100,000 kilometers per second.
After a mission elapsed time of 92 days from launch, Surveyor is 21.51 million kilometers from the Earth, 107.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 29.31 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C4 command sequence, and all systems continue to be in excellent condition.
Friday, 14 February 1997
On Tuesday, the Surveyor \Jspacecraft\j rotated to a position that pointed the Mars Orbiter Camera at a cluster of stars called the Pleiades. Over the course of an hour, the camera imaged stars within the cluster. These images were used by the camera team to determine the focus of the narrow-angle camera following the bakeout period that ended two weeks ago.
During that five-day bakeout period, a 53-Watt heater was used to remove residual moisture from the camera's \Jgraphite\j epoxy structure. This moisture affects the camera's focus. Preliminary results from this week's activity indicates that additional bakeout will not be necessary.
Over the next two weeks, the camera will image the Pleiades on four separate opportunities to allow the camera team to make adjustments to the focus settings.
On Wednesday, the \Jspacecraft\j was commanded to spin in the opposite direction for a period of three hours. Under normal conditions during the journey to Mars, Surveyor's high-gain antenna is pointed at the Earth, and the \Jspacecraft\j slowly spins in the clockwise direction as seen from the Earth.
During the three hours, the \Jspacecraft\j spun in a counter- clockwise direction to allow the \Jspacecraft\j team to calibrate the gyroscopes. These devices provide information to Surveyor's flight computers regarding the \Jspacecraft\j's pointing orientation in space.
Today, the flight team transmitted the C5 sequence to Surveyor. C5 contains commands that will control the \Jspacecraft\j for the next four weeks. The first activities in C5 will start on Monday, February 17th.
After a mission elapsed time of 99 days from launch, Surveyor is 24.30 million kilometers from the Earth, 98.95 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 28.78 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C4 command sequence, and all systems continue to be in excellent condition.
Friday, 21 February 1997
Today, in an activity similar to the one that occurred last week, the Surveyor \Jspacecraft\j rotated to a position that pointed the Mars Orbiter Camera at a cluster of stars called the Pleiades. Over the course of an hour, the camera imaged stars within the cluster.
Images from today's opportunity, combined with three image sets that will be taken between February 24th and February 28th, will allow the camera team to determine settings to control the instrument's focus.
Other major events this week included a complete memory read-out of Surveyor's on-board flight computers on Monday. During this activity, the flight team commanded the \Jspacecraft\j's computers to transmit the contents of its memory banks back to Earth.
The read-out was performed to allow the flight team to verify the values of critical flight software parameters that control the \Jspacecraft\j. Because some of these parameters are periodically updated, the results of the memory read-out were entered into a tracking system that provides a historical record of the changes. Monday's activity was only the second time during the mission that the memory has been completely read out.
After a mission elapsed time of 106 days from launch, Surveyor is 27.71 million kilometers from the Earth, 90.93 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 28.25 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 28 February 1997
On Monday, Wednesday, and Friday of the week that began on February 24th, the Surveyor \Jspacecraft\j rotated to a position that allowed the Mars Orbiter Camera to obtain images within a cluster of stars called the Pleiades. Images were gathered over the course of one hour on each day's opportunity. These images, combined with the images obtained on February 21st, will allow the camera team to determine settings to control the instrument's focus.
Late in the afternoon on Friday, the \Jspacecraft\j experienced a minor glitch with the star scanner. Normally, this device constantly scans a set of reference stars in deep space. These distant stars serve as fixed reference points that allow the \Jspacecraft\j to determine its proper pointing orientation relative to the Earth and Sun. This process is called attitude control and is not related to the camera's star imaging for focus determination purposes.
This glitch occurred during Friday's playback of Mars Orbiter Camera data from Surveyor's recorders. At that time, the star scanner began misidentifying stars. As a consequence, the flight team transmitted a command to the flight software to reset the portion of the attitude control software that controls the star scanner. After several hours, all conditions returned to normal.
Although the cause of the glitch has not yet been determined, the flight team suspects that the star scanner was fooled by sunlight reflecting off of dust particles in the vicinity of the \Jspacecraft\j. In order to further investigate this event, a playback of \Jspacecraft\j \Jengineering\j data recorded during the glitch will occur later this week.
After a mission elapsed time of 113 days from launch, Surveyor is 31.76 million kilometers from the Earth, 83.40 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 27.74 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
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"Mars Surveyor Report (March)",15,0,0,0
Friday, 7 March 1997
On Monday, the on-board command sequence controlling Surveyor executed a test called the "Solar Array Feather." During the several-hour test, the solar arrays were rotated back and forth several times in a similar fashion to the motion that a person makes when rotating the wrist joint.
This activity was performed for the benefit of the Magnetometer science team. The test simulated the rotation of the solar arrays that will occur as the arrays automatically track the Sun during Mars mapping operations. Because the Magnetometer sensors sit at the end of the solar arrays, the data collected from the test will allow the science team to determine the effect of the solar array rotation on the quality of their data.
On Tuesday, the flight team loaded new parameters to Surveyor's attitude control software. These parameters deal with the performance of the star scanner that controls the \Jspacecraft\j's ability to point at targets in space. With this parameter update, the \Jspacecraft\j will be able to point its science instruments at objects with better accuracy than previously possible.
Later on Tuesday, the Ka-band communications team accomplished a major milestone in their experiment. Over a several hour time period, an antenna at the Goldstone tracking station recorded data transmitted simultaneously from Surveyor's X-band and Ka-band transmitters.
Normally, the \Jspacecraft\j utilizes the 25-Watt, X-band transmitter for communicating with the Earth. The main difference between the two signals is that the 1-Watt, Ka-band transmitter operates at a frequency near 32 gigaHertz versus 8 gigaHertz for X-band.
An analysis of the experiment indicated that no disagreements existed between the X-band and Ka-band data for all 12 million data bits observed on Tuesday. This positive result marks the first verified data transmission by an interplanetary \Jspacecraft\j using a Ka-band signal.
The result affirms a long-held belief that the use of Ka-band signals can allow a \Jspacecraft\j to transmit information at faster data rates with transmitters that consume much less power.
After a mission elapsed time of 120 days from launch, Surveyor is 36.46 million kilometers from the Earth, 76.39 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 27.23 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 14 March 1997
On Monday of this week, the flight team loaded new parameters to Surveyor's attitude control software. These parameters deal with the alignment of the Inertial Measurement Unit. This device contains three gyroscopes that provide the flight computers with critical information regarding the \Jspacecraft\j's pointing orientation in space.
The new parameters, combined with the new parameters for the star scanner that were loaded last week, will enable Surveyor to point its science instruments at objects with better accuracy than previously possible.
Today marked the first day since the launch of both Mars Pathfinder and Mars Global Surveyor that Pathfinder's distance to Mars was less than Surveyor's. However, because the two \Jspacecraft\j are on different types of flight paths to Mars, they did not physically fly past each other.
At the time of closest approach, Pathfinder and Surveyor were separated by 4.7 million kilometers. Pathfinder was launched after Surveyor, but will reach Mars first because it is traveling on a shorter, more direct flight path.
This week was a relatively quiet week as the flight team prepared for next week's trajectory correction maneuver. This engine firing will refine Surveyor's flight path to Mars and will take place on Thursday, March 20th at 10:00 a.m. PST.
After a mission elapsed time of 127 days from launch, Surveyor is 41.78 million kilometers from the Earth, 69.86 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 26.74 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 21 March 1997
On Wednesday, the flight team transmitted the C6 sequence to Surveyor. This sequence contains commands that will control the \Jspacecraft\j for the next four weeks. C6 became active on Thursday at 6:00 a.m. PST.
The first major event in C6 occurred at 10:00 a.m. PST on Thursday. At that time, the onboard flight computer commanded the \Jspacecraft\j's main rocket engine to fire for six seconds in order to make minor corrections to Surveyor's flight path.
During this trajectory correction maneuver, the main engine burned a propellant combination of hydrazine fuel and \Jnitrogen\j tetroxide oxidizer. In total, the \Jspacecraft\j expended approximately 1.4 kilograms of propellant.
Immediately before the six-second burn was performed, Surveyor ignited eight of its 12 attitude-control thrusters for 20 seconds. These tiny thruster rockets are normally used to stabilize the \Jspacecraft\j during main engine firings.
The initial, 20-second thruster firing settled the liquid in the \Jspacecraft\j's tanks to ensure a smooth flow of propellant to the more powerful main rocket engine that was used to perform the correction maneuver.
At this time, the navigation team is busy analyzing the accuracy of yesterday's trajectory correction maneuver. However, preliminary results from the accelerometer onboard the \Jspacecraft\j show that the engine firing provided a velocity change of 3.875 meters per second. This value was within 0.5% of the predicted change of 3.857 meters per second.
Yesterday's maneuver was the second in a series of four trajectory correction maneuvers that are designed to refine the \Jspacecraft\j's flight path to Mars. The first maneuver occurred shortly after launch last November. The third and fourth are currently scheduled for April 21st and August 25th, respectively.
After a mission elapsed time of 134 days from launch, Surveyor is 47.69 million kilometers from the Earth, 63.84 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 26.27 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 28 March 1997
No major activities occurred onboard the Mars Global Surveyor \Jspacecraft\j this week. Meanwhile, at the Jet Propulsion Laboratory in Pasadena, Surveyor's navigation team has completed their preliminary assessment of the trajectory correction maneuver that took place on March 20th. This short firing of the \Jspacecraft\j's main rocket engine resulted in a velocity change of 3.875 meters per second and refined Surveyor's flight path to Mars.
Initial analysis provided by the navigation team indicates that the \Jspacecraft\j performed the maneuver with an accuracy of greater than 99%. Consequently, the \Jspacecraft\j is now on a flight path that will come within 630 kilometers of the Martian surface at the point of closest approach on September 12th. Additional trajectory correction maneuvers scheduled for April 21st and August 25th will reduce this approach altitude to 500 and 380 kilometers, respectively.
After a mission elapsed time of 141 days from launch, Surveyor is 54.12 million kilometers from the Earth, 58.29 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 25.82 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
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"Mars Surveyor Report (April)",16,0,0,0
Friday, 4 April 1997
On Saturday, March 29th, the flight team performed a several-hour communications test to measure low-level interference between Surveyor's ultra-stable-oscillator-generated X-band signal and the Ka-band signal. Normally, the \Jspacecraft\j utilizes the 25-Watt, X-band transmitter for communicating with the Earth.
The main differences between the two are that the 1-Watt, Ka-band transmitter is experimental and operates at a frequency near 32 gigaHertz versus 8 gigaHertz for X-band.
During last Saturday's test, the \Jspacecraft\j simultaneously activated both the X- and Ka-band signal sources. The test was designed to determine the effect of the Ka-band signal on the purity of the X-band signal as generated by the ultra-stable \Joscillator\j.
Understanding the performance of the \Joscillator\j under potential interference conditions is important because this device functions as an electronic clock that precisely controls the tone of Surveyor's radio signal. Precision control of the signal's tone is important for gathering data regarding the Martian atmosphere.
On Monday, March 31st, the \Jspacecraft\j passed through a major milestone on its way to Mars. For the first time in the mission, Surveyor was closer to Mars than to Earth. This equidistant point was approximately 57 million kilometers between the two planets. The half-way point measured in terms of days to reach Mars will occur on April 10th.
After a mission elapsed time of 148 days from launch, Surveyor is 61.05 million kilometers from the Earth, 53.20 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 25.36 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 11 April 1997
This week, the Mars Global Surveyor science team received an unexpected bonus from the Sun due to a solar flare eruption that took place on Monday. Eruptions of solar flares occur when disturbances deep within the Sun's interior cause streams of electrically charged atomic particles to be ejected from the solar surface. These charged particles move through the solar system at speeds in excess of 1,000,000 kilometers per hour.
In order to allow the science team to study this event, the flight team sent commands to Surveyor that enabled the \Jspacecraft\j to record solar flare data gathered from the Magnetometer science instrument. These commands activated the \Jspacecraft\j's data recorders late Wednesday afternoon, about half a day before the stream of charged particles from Monday's eruption reached Surveyor.
Although past occurrences of solar flares have both disrupted space communications and damaged \Jspacecraft\j, Monday's eruption was relatively mild in comparison. The Mars-bound Surveyor \Jspacecraft\j sustained no damage from the solar flare.
Late Thursday afternoon, the navigators on the project canceled the trajectory correction maneuver that was planned for later this month. This maneuver would have refined the flight path to Mars by slightly altering the \Jspacecraft\j's speed and velocity.
However, analysis showed that this month's maneuver involves a velocity change of only 40 millimeters per second (less than one-tenth of a mile per hour). The maneuver was canceled because with such a small velocity change, the errors in executing the maneuver are comparable to the size of the maneuver.
This canceled maneuver would have been the third of four planned maneuvers during the journey to Mars. The first two occurred in November 1996 and March 1997. The fourth trajectory correction maneuver will take place on August 25th, 1997.
Yesterday marked the halfway point in the journey to Mars with respect to time of flight. As of April 10th, Surveyor has completed 154 of the 308 days required to reach the red planet. The halfway point in terms of distance between the Earth and Mars occurred last week on Monday, March 31st. This difference in halfway dates arises from the fact that the positions of the two planets constantly change during the \Jspacecraft\j's journey to Mars.
After a mission elapsed time of 155 days from launch, Surveyor is 68.43 million kilometers from the Earth, 48.55 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.98 kilometers per second.
This orbit will intercept Mars 153 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 18 April 1997
No major mission activities occurred this week onboard the Mars Global Surveyor \Jspacecraft\j. Back at the Jet Propulsion Laboratory in Pasadena, \JCalifornia\j, the project management has made a decision not to attempt any more efforts to free debris that is currently keeping the -Y-side solar array slightly out of position. This solar panel is currently deployed and fully functional, but is 20.5 degrees from its proper position.
The flight team believes that the position discrepancy was caused when a damper shaft in the array's deployment mechanism broke shortly after launch. This damper is a device that was installed to minimize the mechanical shock of deployment by slowing the motion of the array during deployment. The flight team theorizes that the broken shaft caused the damper arm to wedge into the hinge joint connecting the solar panel to the \Jspacecraft\j.
An important aspect of this position discrepancy is that the solar panels will be used at Mars not only to produce electrical power, but also to help the \Jspacecraft\j attain its final mapping orbit. Over the course of a four-month period following Mars orbit insertion, Surveyor will be dipped into the upper Martian atmosphere on every orbit.
During these atmospheric passes, air resistance generated by the solar panels will slow the \Jspacecraft\j and gradually lower its orbit. Surveyor will use this "aerobraking" technique to lower the high point of its orbit from an initial 56,000 kilometer altitude to just under 400 kilometers.
For the last few months, the flight team has been considering several options to free the debris and allow the panel to latch and lock into its proper position. One idea involved a short firing of Surveyor's main rocket engine to provide a small force to dislodge the damper arm.
However, such efforts will not be necessary because an extensive analysis has indicated that aerobraking with the -Y solar panel slightly out of position is feasible with a few minor modifications to the original plan.
One of the minor changes involves rotating the panel into a position where the front side will face into the air flow instead of the back side. This orientation will keep the unlatched panel from folding up on itself when it encounters the air flow during aerobraking.
Because the front side contains the silicon cells that produce electricity, it is more fragile than the back side and cannot tolerate as much heating from the air flow. As a result, the flight plan will be modified so that Surveyor aerobrakes at a slightly slow pace than previously planned.
After a mission elapsed time of 162 days from launch, Surveyor is 76.20 million kilometers from the Earth, 44.32 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.59 kilometers per second.
This orbit will intercept Mars 146 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 25 April 1997
Last Friday afternoon, the flight team transmitted the C7 sequence to Surveyor. This sequence became active at 7:00 a.m. PDT on Monday, April 21st and contains commands that will control the \Jspacecraft\j for the next 28 days.
Late in the evening on Monday, Surveyor transmitted 1.5 gigabytes of recorded data back to Earth. This data was collected by the Magnetometer science instrument two weeks ago during the solar flare eruption. The playback of the data took five hours to complete and represents nearly 52 hours of recorded science. On Tuesday, the \Jspacecraft\j repeated the five-hour data transmission for redundancy purposes.
No other major activities occurred this week. After a mission elapsed time of 169 days from launch, Surveyor is 84.32 million kilometers from the Earth, 40.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.23 kilometers per second.
This orbit will intercept Mars 139 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C7 command sequence, and all systems continue to be in excellent condition.
#
"Mars Surveyor Report (May)",17,0,0,0
Friday, 2 May 1997
No major activities took place this week. For the past three weeks, few activities have occurred because the Surveyor \Jspacecraft\j has been configured in a quiet state for a search campaign to detect gravity waves. According to theoretical physics, these waves are gravitational disturbances emitted by all objects in the universe.
However, because gravity is a relatively weak force, detection of these waves is almost impossible unless they are generated by massive objects such as black holes and matter at the center of the Milky Way Galaxy.
To date, nobody has ever detected a gravity wave. If Surveyor encountered these waves, the \Jspacecraft\j would experience an extremely small jolt. This tiny bumping motion would cause a tiny shift in the frequency of the \Jspacecraft\j's radio signal transmitted to Earth. Analysis of the data generated by this experiment will take six months or more.
After a mission elapsed time of 176 days from launch, Surveyor is 92.74 million kilometers from the Earth, 37.03 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.89 kilometers per second.
This orbit will intercept Mars 132 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C7 command sequence, and all systems continue to be in excellent condition.
Friday, 9 May 1997
At 4:30 a.m. PDT on Thursday, the flight software onboard Mars Global Surveyor commanded the \Jspacecraft\j into safe mode. Entry into this operational mode placed the \Jspacecraft\j in a safe power, thermal, and communications configuration. This precautionary measure is taken if the \Jspacecraft\j detects an unexpected event in one or more of its subsystems.
The chain of events that resulted in safe mode began Wednesday night. At that time, the flight team was finishing the second of two calibrations of Surveyor's gyroscopes. These calibrations involved commanding the \Jspacecraft\j to rotate in various directions in order to ascertain the performance of the gyroscopes.
Surveyor had just completed the \Jcalibration\j that involved a +Z-axis rotation when the flight software commanded the \Jspacecraft\j into contingency mode. This mode is similar to safe mode, but involves fewer precautionary measures taken to safe the \Jspacecraft\j.
Entry into contingency mode was triggered when the direction to the Sun as measured by Surveyor's Sun sensors disagreed with the predicted direction to the Sun as calculated by the onboard flight software. This discrepancy in Sun position was approximately 5 degrees.
Entry into safe mode occurred about five hours later when a flight software task timed out and failed to report back Surveyor's central processor. At this time, the flight team is identifying the software task that timed out.
The entry into contingency and safe mode resulted in the flight software terminating the execution of the current command sequence, powering off the science payload and non-essential components, and turning the \Jspacecraft\j toward the Sun to guarantee adequate power. Analysis of telemetry transmitted from Surveyor over the last 24 hours indicates that all systems are healthy.
After the exact cause of safe- mode entry is identified and resolved, the flight team will command the \Jspacecraft\j back into its normal operational mode. This process will consume at least the next few days.
Late Thursday night, the flight team transmitted a series of commands to Surveyor for thermal maintenance purposes. One set of commands shut off the secondary set of heaters to avoid overheating the \Jspacecraft\j's 12 attitude-control thruster rockets.
The other set of commands changed Surveyor's pointing orientation from high-gain antenna pointed directly toward the Sun to antenna pointed 10 degrees away from the Sun. This orientation change allowed for more sunlight to maintain warm temperatures on the science instruments.
After a mission elapsed time of 183 days from launch, Surveyor is 101.43 million kilometers from the Earth, 33.90 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.58 kilometers per second. This orbit will intercept Mars 125 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th).
Although the \Jspacecraft\j is currently operating in safe mode, all systems are functioning properly, there are no \Jspacecraft\j hardware problems, and there is no threat to the mission.
Friday, 16 May 1997
This week, flight team members concentrated their efforts on determining what event caused the Mars Global Surveyor \Jspacecraft\j to enter safe mode early in the morning on May 8th. Since then, Surveyor has been operating in a configuration that ensures that the \Jspacecraft\j has adequate power, thermal, and communications margins.
Flight software on the \Jspacecraft\j automatically commands entry into this safe mode if it detects an unexpected event in one or more of Surveyor's subsystems.
One of the major diagnostic activities involved commanding the \Jspacecraft\j to transmit portions of its computer memory back to Earth for analysis. An examination of a region of memory called the Audit Queue revealed that entry into safe mode occurred when a flight software task timed out and failed to report back to Surveyor's central processor.
Each software task executed by Surveyor's computer is allocated a certain amount of time to complete. Timeouts occur when a task fails to complete in the allocated time. Members of the flight team at the Lockheed Martin facility in \JDenver\j traced the source of this timeout to an infinite loop that occurred in flight software. A timeout resulted because infinite loops are impossible to complete.
The infinite loop resulted from the corruption of an area of computer memory called the Active Script Table. This table contains a list of programs executed by Surveyor's central processor, and corresponding links to the locations in computer memory where those programs are stored.
A software task that was executing prior to safe- mode entry caused the infinite loop when it incorrectly updated one of the entries in the table by linking that entry back to itself.
Over the last few days, engineers on the flight team reproduced the safe-mode entry conditions in the \Jspacecraft\j simulator. Subsequent analysis indicates that the action that created the infinite loop is uncommon, but predictable. Consequently, the Flight Operations Manager has decided to allow the flight team to begin procedures that will return the \Jspacecraft\j back to its normal operating state.
Commands to perform this recovery will be sent starting in the afternoon on Monday, May 19th. Once recovery is complete, the flight team will transmit modifications to Surveyor's flight software that will prevent this infinite loop condition from occurring again. Normal operations should be restored by mid-week.
In other news not related to safe-mode operations, the flight computer powered down \Jgyroscope\j #2 on Tuesday, May 13th. This power down occurred automatically when the electrical current used by the \Jgyroscope\j exceeded a preset limit.
Gyro #2's functions were automatically assumed by Gyro #1 and #3, the transition was smooth, and there is no performance degradation with respect to Surveyor's ability to point at targets in space. The powered-down \Jgyroscope\j will be reactivated after normal operations recommence.
After a mission elapsed time of 190 days from launch, Surveyor is 110.33 million kilometers from the Earth, 31.07 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.29 kilometers per second.
This orbit will intercept Mars 118 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). Although the \Jspacecraft\j is currently operating in safe mode, all systems are functioning properly, and no \Jspacecraft\j hardware problems exist that pose a threat to the mission.
Tuesday, 27 May 1997
Shortly after 9:00 p.m. PDT last Saturday, operators staffing the Goldstone antenna complex in the Mojave desert announced that they had locked up on a signal transmitted from Surveyor at a data rate of 2,000 bits per second.
This milestone marked the transition out of safe-mode and back to normal operating conditions. Since early in the month, Surveyor's safe-mode orientation had limited the maximum data transmission rate to 250 bits per second or less.
The \Jspacecraft\j automatically entered safe-mode on the morning on Thursday, May 7th when the onboard computer encountered an infinite loop in flight software. Entry into safe mode placed Surveyor in a configuration that guaranteed adequate power, thermal, and communications margins. This mode is intended to be a benign operating state favorable for diagnostic and recovery activities if an unexpected event occurs in one or more of the \Jspacecraft\j's systems.
Recovery operations involved a multi-step process that began on Friday. First, the flight team sent a series of instructions to Surveyor's backup flight computer. These instructions initialized the backup computer to begin using its normal flight software rather than the limited software set utilized in safe mode. Then, the flight team commanded the backup computer to control the \Jspacecraft\j while performing the same software initialization procedure on the Surveyor's primary computer.
The next step required reestablishing the \Jspacecraft\j's ability to point at targets in space. In safe mode, the flight computer assumes that its ability to find and point at targets other than the Sun has been compromised.
Restoration of pointing capability involved commanding the \Jspacecraft\j to rotate in a cone-shaped pattern around the Sun for several hours. This action allowed Surveyor's star scanner to lock-up on distant guide stars in space. The \Jspacecraft\j determines its orientation in space by using these stars as reference points.
Pointing capability was restored early Saturday evening. At that time, the flight team commanded Surveyor to rotate from its safe-mode, Sun-pointed orientation to an Earth-pointed orientation.
Aiming the \Jspacecraft\j's antenna directly at the Earth enabled Surveyor to begin transmitting data using any one of its standard rates of 2,000 bits per second or faster. Early next week, the flight team will transmit modifications to Surveyor's flight software to prevent the infinite-loop condition from occurring again.
Surveyor would have been stable in safing for an indefinite period of time even if no corrective action had been taken. However, the flight team worked on restoring standard operations as quickly as possible because normal command sequences, such as those that control science \Jcalibration\j activities, are prohibited from executing in safe mode.
After a mission elapsed time of 201 days from launch, Surveyor is 124.64 million kilometers from the Earth, 27.15 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 22.89 kilometers per second. This orbit will intercept Mars 107 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). All systems continue to be in excellent condition.
#
"Mars Surveyor Report (June)",18,0,0,0
Friday, 6 June 1997
Two weeks after recovery from safe mode and the restoration of standard operations, the Mars Global Surveyor \Jspacecraft\j continues to perform excellently as it cruises toward an encounter with the red planet later this summer. Currently, Surveyor is operating in a quiet state with no major activity sequences programmed in the onboard computer. The flight team will transmit the next major sequence load toward the end of the month.
On Tuesday of this week, the flight team sent a few commands to Surveyor that activated \Jgyroscope\j #2 for a period of one hour. Several weeks ago, this gyro was automatically powered down when its usage of electrical current exceeded a preset limit.
Although gyros help the \Jspacecraft\j keep track of its pointing orientation in space, there was no loss of control because Surveyor's #1 and #3 gyros seamlessly assumed the function of the powered down unit.
During the one hour of operation, the amount of electrical current used by gyro #2 was well below the level that would have resulted in an automatic power down. Although the gyro is functional, the project management has decided to leave it powered off. Flight software code is being developed that will autonomously activate gyro #2 in the unlikely event that an anomalous condition precludes the usage of either the #1 or #3 gyro. This new software will be transmitted to Surveyor in a few weeks.
The only other notable activity this week occurred late Thursday. That evening, the flight team transmitted a short series of commands to Surveyor that modified the onboard software. These minor changes will ensure that the infinite-loop condition that resulted in safe mode entry will never happen again.
After a mission elapsed time of 211 days from launch, Surveyor is 137.88 million kilometers from the Earth, 24.04 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 22.57 kilometers per second. This orbit will intercept Mars 97 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). All systems continue to be in excellent condition.
Due to the dynamic nature of space flight operations, please keep in mind that dates of events can change at a moment's notice. This calendar was last updated on 18 May 1997.
Short-Term Event Schedule
19-May-97
Begin transition out of safe mode and back to normal operations
30-May-97
Uplink command sequence C8 to \Jspacecraft\j
9-May-97
Uplink software patch for Payload Data Subsystem
High-Level Master Schedule
Launch
7-Nov-96
Lift-off occurred at 12:00:50 EST
TCM121-
Nov-96
Optimized maneuver to correct for launch injection errors (1st of 2 burn sequence)
TCM220-
Mar-97
Flight path correction (2nd of 2 burn sequence)
TCM321-
Apr-97
Maneuver to correct for execution errors from TCM2
TCM425-
Aug-97
Maneuver for final adjustment for orbit insertion aim point
MOI12-
Sep-97
Mars orbit insertion burn will last for about 20 minutes, closest approach will occur at about 1:26 a.m.
UTCOrbit Insertion Phase
12-Sep-97 to 14-Mar-98
Lasts for 5 months to reach mapping orbit using aerobraking and propulsive maneuvers
Mapping Phase
15-Mar-98 to 31-Jan-00
Mapping operations will last 1 Mars year, about 687 Earth days in duration
Relay Phase
1-Feb-00 to 1-Jan-03
Communications support for future Mars missions
#
"Mars Pathfinder Status",19,0,0,0
\JPathfinder Status (August96)\j
\JPathfinder Status (September96)\j
\JPathfinder Status (October96)\j
\JPathfinder Status (December96)\j
\JPathfinder Status (January97)\j
\JPathfinder Status (February97)\j
\JPathfinder Status (March97)\j
\JPathfinder Status (April97)\j
\JPathfinder Status (May97)\j
\JPathfinder Status (June97)\j
\JPathfinder: Entry, Descent and Landing\j
#
"Pathfinder Status (August96)",20,0,0,0
Week of August 19, 1996
Once the \Jspacecraft\j arrived and unpacked at KSC, one of the first things we did was re-run several "EDL runs" in several different conditions on the \Jspacecraft\j (EDL stands for Entry Descent and Landing). We wanted to re-verify some minor changes in the flight software that automatically controls the detailed series of events in this critical mission phase. In these tests, the software starts out in the late cruise mode of operation.
Using support equipment that can "stimulate" accelerometers to simulate the effects of the deceleration of entry and landing, as well as other equipment that can simulate the data that comes from the radar \Jaltimeter\j, the entire EDL process can be artificially created without having anything really happen (other than petal and airbag retraction actuators moving).
These tests were quite productive. In several cases we even "pulled the plug" on the computer and let it re-boot, with great success. This was the next-to-last time that the EDL software will be run on the real lander, the next time will be on the Fourth of July, 1997!
#
"Pathfinder Status (September96)",21,0,0,0
Week of September 2, 1996
EDL testing done, the lander was dissembled so that the interior \Jelectronics\j could be accessed. We had two fuses, some relays, and a waveguide transfer switch that needed to be replaced. Also we installed a fresh, fully charged, flight battery. This battery is the best one that we have used to date and will now have to survive to do its job over the next year. Once the \Jelectronics\j were updated, we performed a full functional test to confirm that the fixes worked and that we didn't disturb anything in the process.
Week of September 9, 1996
Early this week, for the final time, we reinstalled the ISA (Integrated Structure Assembly) - that's the white thermal and structural box that surrounds the lander \Jelectronics\j and has the red "JPL" letters on it. We then calibrated the stop positions on the HGA (High Gain Antenna) - that's the lollipop-shaped articulated antenna that sits nest to the camera on the ISA.
We used special theodolites (similar to those used by construction surveyors) to verify that the HGA mechanically points in the direction we want it to point with respect to the lander's base petal. We then reinstalled and checked out the pyro switching \Jelectronics\j and installed the lander thermal batteries.
We use "thermal" batteries to provide power (current) to ignite explosive initiators in the EDL pyrotechnic devices (things like the parachute mortar, separation nuts, cable cutters and rocket ignitors). The batteries are called "thermal" because they get their electrical energy from self-generated chemical heat.
Similar to the pyrotechnic initiators to which they provide, these batteries themselves need to be "lit" seconds before they are used on the \Jspacecraft\j during EDL. Once "lit", these batteries will operate for only a few minutes - plenty of time to do their jobs.
Once the ISA was installed, we performed some radio communication tests between the rover and the lander. We had been uncertain whether or not we needed to launch with an RF (radio frequency) attenuator in series between the lander's rover antenna and the lander's RFD modem used to talk with the rover.
This attenuator was thought some time ago to be needed to allow communication with the lander at close distances. These tests and some others coming up have nearly convinced us that we can live without it. We think that it would be good if we did not use it because the attenuator might reduce the communication range if we ever decided to drive the rover a long way away from the lander in its "extended" mission. Either way, the primary mission is unaffected.
We successfully performed other radio tests as well. Until this week, we had not yet tried to uplink the large software patch files using the real X-band radio and a ground station. The patch files are used in the unlikely event we have to reload large portions of the flight software into the EEPROM memory during the mission. Using the MIL-71 ground station at the Kennedy Space Center, we found that the process works fine.
We also took a few last verification images from each eye of the IMP camera on the lander. This is the last time the IMP camera will be taking interesting pictures until we land on Mars.
There has also been much work on the cruise stage. The HRS (Heat Rejection System) freon pumps have been installed and checked out. The HRS is the system used to flow freon inside the lander and around the perimeter of the cruise stage to keep the lander \Jelectronics\j cool.
We need to keep the battery, the digital \Jelectronics\j, the rover, and the big X-band radio transmitter we call the SSPA (Solid State Power Amplifier) cool during the especially warm early part of the "cruise" phase of the mission as we leave Earth.
We had to replace the pumps that had been installed during this summer's thermal tests because we think that it may have been damaged during one of our electrical tests. Because it is so hard to take apart, we can't tell for sure that it is broken. So just in case it was, we decided to replace it with the flight spare unit.
Later this week we will reattach the petals (the Sojourner Rover is already mounted on its "Y" petal). Next week we begin the long process of installing the flight airbag as well as the many pyrotechnic devices on the lander.
Week of September 16, 1996
As you can see from the live image, the mechanical team has completed installation of the three lander petals. Each of these petals are mounted on a hinge that allows them to open and close by more than 110 degrees of travel. The petals are moved under software command via three actuators mounted along the "hinge line" of each of the three petals.
One or more of these petals open up, after the airbags are retracted, about an hour and fifteen minutes after landing. They are the mechanism for automatically righting the tetrahedral-shaped lander so that no matter how the lander comes to rest inside the inflated airbag \Jcocoon\j, the lander will always end up "top side up" and not the reverse!
Once the petals were installed, the "rock membrane" was placed on each of the petals (except for the rover petal which will be done next week). This is a layer of aluminum sheet metal placed on the outside of the petals (but under the airbags).
The petals are about two inches thick, but for the most part are hollow. The inside of the petals have the lander's solar arrays attached on an aluminum substrate. This layer will provide protection from any tall sharp rocks that might be inclined to try to penetrate a side (or base) petal and damage the delicate solar arrays from below.
When the lander rolls onto its base petal from a side petal during the righting phase, the 270 kg lander (only 223 lbs on Mars) can roll quite hard. The airbags will help some, but aluminum will really do the trick.
The three solid rockets used to stop the lander just before landing are now installed. First they had to be prepared with heaters and thermal blanketing before mounting to the inside of the backshell. The backshell, the heatshield and the cruise stage are not far from the lander, but are out of sight of the camera.
Progress on the cruise stage is also being made. We decided some time ago to separate the fuel tank heater circuits from the propellant line heaters so that we can maintain the option of turning off the tank heaters during battery charging while also not allowing the fuel in the propellant lines to freeze.
Turning the tank heaters on for the few hours of battery charging allows us to gain a few extra precious watts when we need it most during "cruise". The necessary cabling changes on the cruise stage were done this past week and verified.
#
"Pathfinder Status (October96)",22,0,0,0
Week of October 8, 1996
A lot has been accomplished in the last two weeks. For those of you who have been watching the images from KSC regularly, you will have noticed that the airbags have been installed for nearly two weeks.
The airbag and gas generator installation operation is quite complex, but the talented folks (Skip Wilson) from the airbag manufacturer (ILC Dover Inc. of Frederica, DE) and the JPL airbag cognizant engineer (Tom Rivellini) have done this so many times before on other test landers, that they make it look easy.
Once the bags were installed, a detailed "walk through" of the whole lander was performed early last week by some the best \Jspacecraft\j mechanical engineers around. They looked for anything that might appear to be amiss.
For example, they pointed out that the exposed edge of the solar panel honeycomb substrates really ought to be taped closed rather than provide the opportunity for manufacturing particulates to escape and potentially contaminate sensitive surfaces such as the camera \Joptics\j. The edges were taped and bonded closed within two of hours of being mentioned by our very talented mechanical team.
One other item was reviewed and commented upon by the mechanical review team. We all noticed for the first time, that with the fully loaded mass of the lander and the airbags, the petal latches were no longer perfectly aligned when closed.
This meant that when the six petal latch separation bolts are torqued down (they get released after landing released via pyrotechnic separation nuts) there may be some rebound in the structure that would "race" the separation nuts at the moment of release. The engineers were concerned that there was adequate time for the bolts to exit the holes before the structure moved to the point that the bolts could not clear the holes and hang up.
Also we knew that the petal actuators and the mechanical petal stops used to prevent the lander petals from closing in on and making unwanted contact with the internal structures on the base petal, would also contribute some outward \Jtorque\j and angular acceleration at the moment of petal latch release.
So just to be on the safe side, with the help of Jim Baughman, the lander structural engineer and designer, we made the decision to slightly widen the separation bolt holes so that there will be plenty of time for the bolt to exit the hole before the structure moved to the point of impeding bolt motion (the bolts move out of the holes from the energy of the released tension in the bolt plus the bolt has a spring at the head of the bolt that moves it quite quickly out when the nut splits open).
This delayed the final petal closing a day or two and by last Saturday night we had the petals closed for the last time. Fortunately we still have plenty of time built into our schedule to get all of the work still to go completed before our first launch opportunity on Dec. 2.
This coming week, we will lift and weigh the lander and we will install it in to the backshell by raising it from below. Shortly thereafter we will install the heatshield for the last time. Later in the week we will take this "entry" assembly and place it upside down (heatshield up) onto a spin table in the high bay and we will balance the vehicle (much like an \Jautomobile\j tire) so that there is virtually no wobble when it spins at 2 rpm during Mars entry.
For week of October 17, 1996
Spacecraft Status at KSC:
As you can see from the live KSC image, quite a lot has been accomplished over the last two weeks. The petals have been closed for the final time, the lander bolted (using separation nuts) to the "backshell interface plate" inside the backshell. Most recently, the heatshield has been bolted to the aeroshell.
These accomplishments come as a great relief to the entire team. As we and other \Jspacecraft\j builders have discovered, the flight hardware is often more in danger of being damaged inadvertently by human hands than from anything the rigors of outer space could dish out!
For example we had to repair a broken wind sensor in the MET mast after a ground wrist strap brushed past it, damaging a very tiny element. During the mission, nothing will come near this sensor. Fortunately, it was fixed without much trouble. Now that the bulk of the system is neatly packaged inside the backshell, at least we can breathe a partial sigh of relief.
The next step in this delicate process will be to turn the entry vehicle upside down and spin balance it at 70 rpm using a spin table in the facility. This process will remove any wobble from the mass properties. (For the technically inclined, this process aligns the principle axis of \Jinertia\j with the axis of symmetry.) Once done later this week, we will flip the vehicle over again (a tricky process all by itself) and then mount the cruise stage. We will then do yet another spin balance in the launch configuration.
When we started this assembly sequence at KSC two months ago we had about 14 days of schedule "pad" in case anything slowed us down. Well of course, things did slow us down and we now have about 5 days of schedule margin left until we place the \Jspacecraft\j on top of the upper stage in November. However that isn't bad considering the complexity of this operation. And we had expected to use the margin up. We are all very happy with the progress so far.
Mission Operations Status at JPL:
For the past two weeks those of us on the team working at JPL have been focused on operations. We have a "testbed" at in Pasadena that has a complete duplicate set of all of the flight \Jelectronics\j that are on the \Jspacecraft\j (including a "sim" rover when needed).
In addition to testing flight software on this testbed, we also use it in an operational mode: essentially "flying" the testbed in a simulated mission to Mars! Early last week we "launched" the testbed using the same people, procedures and software that we will use on launch day. We even used the launch team at the Cape who have been doing the \Jelectronics\j \Jintegration\j testing.
Afterward we performed the instrument and rover health checks that we will perform in the weeks after launch. This week, we performed two TCMs (trajectory correction maneuvers - or "burns"). Later this week and this weekend we will "land" the testbed: actually running the "EDL" (entry descent and landing) flight software in the "testbed spacecraft". Saturday we will simulate the first day on Mars.
We are currently in the process of building up a "Mars Room" adjacent to the testbed near our operations area. This room is filled with sand and rocks and will eventually be the home of a test lander and the sim rover.
Although most of the \Jelectronics\j will reside in the testbed a few feet away (so we can more easily access it), the lander's sensors and actuators will reside on the lander in the room. This lander will include the petals, camera, high gain antenna, airbags and rover communication antenna.
The rover ramps will also be installed. In the coming months, we will use this room to simulate the first days and weeks on Mars. These simulations help us learn the finer points of operating this complex little lander and rover.
#
"Pathfinder Status (December96)",23,0,0,0
December 4, 1996
10:30 a.m. Pacific Standard Time
NASA's Mars Pathfinder is reported to be performing well on the first day of a seven-month journey to the red planet following a perfect launch today from Cape Canaveral, FL at 1:58 a.m. Eastern time.
"The \Jspacecraft\j team was ecstatic at seeing good \Jspacecraft\j data," said Brian Muirhead, Pathfinder Deputy Project Manager at NASA's Jet Propulsion Laboratory. "The command and data telecommunications subsystems are working perfectly, sending down data at 1,183 bits per second.
The temperature control and propulsion subsystems reported all temperatures and pressures are within expected ranges. All systems are healthy," he said. Pathfinder is traveling away from Earth at a speed of 3.9 kilometers (2.4 miles) per second.
The Delta II launch vehicle performed flawlessly, placing the \Jspacecraft\j on its trajectory to Mars well within acceptable limits. NASA's Deep Space Network acquired the \Jspacecraft\j telemetry signal on schedule, about five minutes after separation of the Delta's third stage.
When Pathfinder came out of Earth's shadow at one hour and 38 minutes after launch, the solar arrays took over powering the \Jspacecraft\j as planned. "Power from the array looks to be about 10% better than initially predicted," said Muirhead.
Pathfinder engineers continue to analyze data from the \Jspacecraft\j's sun sensor, an instrument that helps the \Jspacecraft\j determine its orientation with respect to the Sun. "The sensor's voltage output is below expected levels but it does appear to be giving good data," said Muirhead.
Navigation data and the sun sensor data agree and show the \Jspacecraft\j to be properly oriented, spinning at the expected 12 rpm and pointed 26 degrees off the Sun. Later today, Muirhead said the \Jspacecraft\j will be commanded to switch to a redundant sensor head to see if it is also performing at a low voltage. "Should the problem persist, we have a number of work around options and there is no risk to the continuation of the mission."
Carried inside the cone-shaped \Jspacecraft\j is Sojourner, the small robotic rover that will roll out to traverse the surface of Mars when the \Jspacecraft\j makes its Martian landing on July 4, 1997.
It will be the first \Jspacecraft\j to land on Mars since NASA's Viking mission soft-landed two \Jspacecraft\j there in 1976. Pathfinder is the second mission in NASA's Discovery program, which is designed to send low-cost \Jspacecraft\j with highly focused mission objectives to explore space.
December 6, 1996
12:00 p.m. Pacific Standard Time
The Mars Pathfinder \Jspacecraft\j continues to perform well in the early part of its cruise to Mars, which is about 209 million kilometers (130 million miles) away today.
Currently the \Jspacecraft\j is 750,000 kilometers (0.5 million miles) from Earth, or about two times the distance that the Moon is from Earth, traveling at a speed of 3.3 kilometers per second (7,400 miles per hour).
The \Jspacecraft\j is performing just as expected, with the exception of the sun sensor. The temperatures of the lander and its \Jelectronics\j are at their predicted levels for this phase of the mission. The cruise stage solar array, propulsion module and \Jelectronics\j are also at their predicted temperatures.
Two of the four segments of the solar array are currently in use, producing approximately 250 watts of power, about 10 percent more power than the original predicts. The battery is charged at 75 percent of its full capacity, and is showing a temperature of 9 \JCelsius\j (48 degrees Fahrenheit), which is approaching the desired steady state of 8 \JCelsius\j (46 degrees Fahrenheit).
The telecommunications system is performing well within its predicted range, indicating that it will be able to maintain higher data rates throughout the mission.
The JPL flight team is continuing its investigation of a lower than expected voltage reading on the sun sensor. However, since the sensor data are good, flight controllers have decided to implement a software update to compensate for this low voltage condition.
The software modification has already been coded and validated in the project's testbed and will be sent to the \Jspacecraft\j this weekend. The software modification will allow Pathfinder's on-board attitude control system to use the sun sensor data in its normal calculations of the \Jspacecraft\j's orientation. Once the attitude control calculations are verified, the planned spin-down maneuver to 2 rpm will be performed, probably early next week.
The \Jspacecraft\j is pointed approximately 55 degrees from Earth and 25 degrees off the Sun. Doppler and ranging data continue to look very good. Because the \Jspacecraft\j is not pointed directly at Earth, flight controllers are able to observe the motion of the antenna as Pathfinder spins about its axis and have confirmed a spin rate of 12.3 rpm.
The latest orbital data from tracking operations at all three Deep Space Network stations around the world indicate that the magnitude of the first trajectory correction maneuver, if performed as scheduled on Jan 4, 1997, would be 29.5 meters per second (96 feet per second).
Mars Pathfinder, the second in NASA's Discovery program of low-cost, highly focused spaceflight missions, is scheduled to land on the surface of Mars on July 4, 1997, and deploy a small rover, called Sojourner, to explore the Martian landscape.
December 10, 1996
12:00 p.m. Pacific Standard Time
The Mars Pathfinder \Jspacecraft\j continues to perform nearly flawlessly on its 203 million kilometer (126 million mile) flight path to Mars. Currently the \Jspacecraft\j is 1.8 million kilometers (1.1 million miles) from Earth, traveling at a speed of 3.2 kilometers per second (7,155 miles per hour). Temperatures and power utilization of the lander and cruise stage remain at predicted levels for this early phase of the mission.
The \Jspacecraft\j's sun sensors are the only issue being watched closely on an otherwise beautifully performing \Jspacecraft\j, the flight team reported. There are five sun sensor heads on board the \Jspacecraft\j, two pointed along the craft's spin axis and three that are equally spaced around the circular cruise stage that look out at about 105 degrees from the spin axis.
Of the five sensor heads, unit #4 on the spin axis is obscured or contaminated to the point of not being useful. Sensor #5, which is also on the spin axis, is providing good sun orientation data, but at a lower voltage than was expected. The other three sensor heads are working fine.
The flight team at JPL uploaded a software modification to the \Jspacecraft\j on Saturday, December 7, which allowed the on-board attitude control system to use the sun sensor data from sensor #5 in its normal calculations of the \Jspacecraft\j's orientation. The software patch was successful and the team was exuberant to see the \Jspacecraft\j's attitude control estimators operating properly.
The team then began to prepare for turning the \Jspacecraft\j more toward Earth to improve the telecommunications link. At the time, Pathfinder was about 58 degrees from the Earth, which is near the edge of the antenna's performance.
Since this was to be the first time flight controllers used the propulsion module, they planned a small turn of two degrees to verify that everything was working properly. Thirty minutes later, they planned to turn the \Jspacecraft\j an additional 20 degrees.
"The turn maneuvers were conducted successfully on Monday morning [December 12]," said Brian Muirhead, Pathfinder flight system manager. "The propulsion and attitude control systems worked properly and the \Jspacecraft\j's spin axis is currently pointed about 44 degrees from the Sun and 37 degrees from Earth. The downlink performance improved as expected and we continue to communicate with Pathfinder at 1,185 bits per second."
The flight team is planning its next maneuver to spin the \Jspacecraft\j down from 12.3 rpm to 2 rpm. The maneuver will be performed in the next few days, Muirhead said. Pathfinder's first trajectory correction maneuver remains on schedule, to take place on January 4, 1997.
December 18, 1996
12:00 p.m. Pacific Standard Time
Sojourner, a 10-kilogram (22-pound) rover tucked away on a petal of the Mars Pathfinder \Jspacecraft\j, got a 'wake up' call on Dec. 17 from flight controllers at NASA's Jet Propulsion Laboratory. After waking up, Sojourner conducted an internal health check and sent data back to the flight team that all was well.
The Pathfinder flight team was ecstatic with the rover data, which showed that all systems within the rover were operating normally. In addition, data from the rover's main science instrument -- the alpha proton x-ray spectrometer -- showed that it was operating properly.
"The rover woke up, did its internal health check, sent the lander its status data and went back to sleep, all as planned," said Art Thompson, rover operations team member. "All subsystems were verified as being in good health."
Pathfinder continues to perform very well on its 500 million-kilometer (310 million-mile) journey to Mars, the team reported. Currently the \Jspacecraft\j is 4 million kilometers (2.5 million miles) from Earth, traveling at a speed of 3.1 kilometers per second (7,000 miles per hour).
Its destination, Mars, is currently about 190 million kilometers (118 million miles) away. All temperatures and power utilization of the lander and cruise stage remain at their predicted levels for this phase of the mission.
The \Jspacecraft\j was spun down from 12.3 rpm to 2 rpm on Dec. 11. Flight controllers first instructed the \Jspacecraft\j to turn to a Sun angle of 50 degrees and an Earth angle of 32 degrees. This allowed them to use all four operating Sun sensors. The \Jspacecraft\j executed the commanded spin down to the normal cruise spin rate of 2 rpm in steps of 2 rpm at a time.
Once the normal spin rate was established, the team turned on the \Jspacecraft\j's star scanner on Dec. 12. Star scanner data allows the \Jspacecraft\j to establish full, three-axis knowledge of its orientation in space. This is the normal cruise attitude control mode and the one in which all trajectory correction maneuvers will be performed.
While Sun sensor #5 continues to work well after a software fix, the flight team continues to investigate the cause of the loss of Sun sensor head #4. The team expects to reach a likely conclusion on the cause of the problem within the next month or two.
Dave Gruel, Pathfinder flight director at JPL, conducted the Dec.16 health check of the lander science instruments, including the atmospheric sensor instrument and \Jmeteorology\j (ASI/MET) package and the imager.
Temperature, pressure and accelerometer readings from the atmospheric/meteorology instrument verified it was in normal working order. Power and dark current measurements received from the imager while it was imaging the darkness around it, confirmed that the instrument was working properly, Gruel said.
Richard Cook, Pathfinder mission operations manager at JPL, reported today that Pathfinder has been fully checked out for this phase of the mission and that all subsystems are "go" for a successful seven-month cruise to Mars.
The next major in-flight event will be Pathfinder's first trajectory correction maneuver, which is scheduled for Jan. 4, 1997.
31 December 1996
The \Jspacecraft\j is currently 7.2 million kilometers from Earth and traveling at 32.6 kilometers per second. All \Jspacecraft\j subsystems continue to operate as expected.
On 30 December 1996 we performed another successful ASI/MET science instrument health check. These are intended to monitor the performance of the pressure \Jtransducer\j that will measure Martian air pressure.
Last Friday, 27 December 1996, we successfully performed our first celestial mode attitude turn. In this turn, the spin axis was turned about 43 degrees, mostly out of the plane of the \Jecliptic\j. The \Jspacecraft\j used the sun sensors and the star scanner to precisely orient the \Jspacecraft\j's spin axis to the new direction, which is about 35 degrees off the Sun and the Earth.
This direction was picked because it is the direction we want the \Jspacecraft\j's thrusters to be in when we perform our first and largest Trajectory Correction Maneuver.
This maneuver is now scheduled for the evening of 3 January 1997 (Important: see update of 2 January 1997). It has been fully designed and tested using the Flight System Testbed \Jspacecraft\j simulation at JPL. Because of the superb performance of the Delta II rocket in getting Mars Pathfinder into a Martian trajectory, this maneuver is expected to use less than 25% of the 93 kilograms of hydrazine on-board the \Jspacecraft\j. It will change the velocity of the \Jspacecraft\j by about 30 m/s over a burn time of about two hours.
#
"Pathfinder Status (January97)",24,0,0,0
2 January 1997
The \Jspacecraft\j is currently 8 million kilometers from Earth and traveling at 32.5 kilometers per second. All \Jspacecraft\j subsystems continue to operate as expected.
Last Friday, we successfully turned the \Jspacecraft\j's spin axis about 43 degrees to the attitude we want the \Jspacecraft\j to be in when we perform our first and largest TCM (Trajectory Correction Maneuver). This had been planned for the evening of 3 January 1997.
However, in the course of testing this TCM using detailed models of the \Jspacecraft\j and the celestial sensors, we discovered that due to the partial obscuration of the sun sensors that occurred shortly after launch, the attitude control software would have unnecessarily fired spin thrusters.
Although in no way dangerous to the \Jspacecraft\j, we would rather not have the thrusters fire any more than absolutely necessary. This problem can be easily solved by changing a parameter in the flight software that will partially reject the bad sun sensor data. Because of the short time before Friday's planned TCM, we decided to postpone it.
Early next week we will send commands that will change the flight software parameters and later in the week we will do the TCM (now tentatively scheduled for 9 January 1997).
Starting tomorrow the team will resume its normal schedule.
10 January 1997
The Mars Pathfinder \Jspacecraft\j is currently 10 million km (6.2 million miles) from Earth, traveling at 32 km/s on its trajectory to Mars. All \Jspacecraft\j subsystems continue to operate as expected.
Earlier in the week we successfully updated attitude control software to further compensate for the apparent obscuration of the sun sensors. Miguel San Martin, lead Attitude Control Subsystem Engineer, stated: "This software update should allow the \Jspacecraft\j to perform all turns and maneuvers needed to get it to Mars. We are very pleased with the subsystem's performance in spite of the sun sensor obscuration."
At 7:40 PM PST on January 9, we successfully performed our first and largest Trajectory Correction Maneuver (TCM). This maneuver, planned by lead Flight Engineer Rob Manning, used two of the \Jspacecraft\j's eight 1 pound thrusters. These thrusters were fired continuously for an hour and a half, and changed the velocity of the \Jspacecraft\j by 31 meters/second.
The purpose of this maneuver was to correct small launch vehicle targeting errors and reduce the planetary protection trajectory bias. Later we turned the \Jspacecraft\j's spin axis 35 degrees back toward Earth so that we can perform radio navigation more effectively. We will leave the \Jspacecraft\j in this attitude until the next TCM scheduled for early February.
From the Doppler radio signature, it appears that last night's TCM performance was well within our expectations. In the next few days, radio ranging data will allow us to precisely gauge the quality of the maneuver.
Next week we will turn on the backup heat rejection system pump and allow it to operate for a while in parallel with the primary pump. These pumps circulate Freon around the perimeter of the cruise stage and down into the lander to keep the lander and rover \Jelectronics\j cool during the 7 month cruise to Mars.
24 January 1997
The \Jspacecraft\j continues to function well and is currently 14 million kilometers from Earth. The major activity for last week was starting the performance testing of the K=15 R=1/6 convolutional code. Early tests results indicate that the Block III MCD is operating as expected and that the expected telecom link improvements match predicts.
The project is investigating some minor anomalies which occurred this week involving the \Jspacecraft\j Command Detector Unit. The most serious of these occured on Monday, January 20 when the CDU transitioned to a lock state during a period when we were not uplinking to the \Jspacecraft\j.
The CDU lost lock as expected when we transmitted an uplink signal, but this "self-lock" behavior is not expected. In addition, we experienced two other episodes where commands were rejected by the \Jspacecraft\j uplink hardware for unexplained reasons.
The project has started a tiger team activity to further investigate these problems and determine potential causes. The possibility of solar flare induced SEUs is being assessed, but we are not ruling out other potential causes.
The flight team is also investigating an attitude control fault that we experienced on Sunday, January 19. This fault occured when estimates of the attitude \Jcovariance\j matrix inexplicably jumped by several orders of magnitude. The resulting fault response reinitialized attitude control flight software and turned off the Propulsion Drive \JElectronics\j.
Analysis of telemetry before and after the fault did not show a definitive cause. We are currently planning to perform additional diagnostic tests to assess memory and software integrity in case SEUs or numerical divergence problems may have caused this problem.
Continuing EDL and surface operations planning. The Rover team completed a successful Rover Operations Readiness Test, and planning is nearly complete for the first full team Surface Operations Readiness Test on January 27-28.
31 January 1997
The \Jspacecraft\j continues to be in excellent health, and is now about 16 million km from Earth. Key activities completed this week include successful completion of the K=15, R=1/6 convolutional code tests and resolution of the attitude control software glitch detected last week.
Attitude control software has been re-enabled and is currently operating nominally. In addition, we verified that the noise seen during ASI/MET health checks is due to the Propulsion Drive \JElectronics\j. This noise appears to be radiative in nature, and will not be an issue for surface operations because the PDE is located on the cruise stage.
The Uplink Problem Tiger team has developed a plausible explanation for the majority of the command rejections and the CDU In Lock conditions. It involves harmonics from the uplink sweep locking up the CDU and pulling it away from the nominal command frequency.
The team is developing a test plan to confirm this hypothesis and is also gathering information about the incidents where the CDU went into lock while we were not tracking.
An Operational Readiness Test (ORT) of the Sol 1-2 sequences was run on Jan 27 and 28. The sequences used were identical to the last pre-launch surface ORT. The ORT was successful in that all of the sequences were executed properly by the simulated lander and rover.
However, a number of relatively minor problems were logged during the test. These problems were reviewed and action has been assigned in all cases for problem resolution.
Nineteen investigators have been selected by NASA Headquarters in response to the Announcement of Opportunity for selection of Mars Pathfinder Participating Scientists and a Facility Instrument Science Team for the Atmospheric Structure Instrument/Meteorology Package.
An "All Hands" Pathfinder Science Team meeting has been set up for Feb. 5-7, 1997 at JPL to begin integrating the new investigators into the Experiment Operations Team.
#
"Pathfinder Status (February97)",25,0,0,0
4 February 1997
The Mars Pathfinder \Jspacecraft\j is currently 19 million km (11 million miles) from Earth traveling at 30 km/s on its trajectory to Mars . All \Jspacecraft\j subsystems continue to operate as expected.
At 5:00 PM PST on February 3, we successfully completed our second Trajectory Correction Maneuver. This maneuver was designed to correct errors in the first TCM performed on January 9, and move us closer to our final trajectory.
The \Jspacecraft\j will not be placed on a Mars atmospheric entry trajectory until after TCM-3 (currently scheduled for May 5) because of planetary quarantine requirements. The TCM-2 design team, led by Flight Engineer Guy Beutelschies, developed a two part approach to perform the maneuver. In the first part, the \Jspacecraft\j fired two of its forward facing thrusters continuously for five minutes. The change in velocity for this "axial" component was about 1.5 m/s.
The second part of the maneuver was a smaller velocity correction of 0.1 m/s performed in the "lateral" mode. In this mode, the \Jspacecraft\j pulses all four thrusters on one side of the \Jspacecraft\j for five seconds. This pulse causes a small change in the \Jspacecraft\j velocity in the direction perpendicular to the \Jspacecraft\j spin axis.
This mode will be used for all future maneuvers, so TCM-2 was a good proof-of-concept test. Early analysis of tracking data from NASA's Deep Space Network indicates that both components were completed successfully.
Upon completing the maneuver, the \Jspacecraft\j's spin axis was turned 15 degrees back toward Earth so that we can perform radio navigation more effectively. The \Jspacecraft\j is currently pointed about 5 degrees from Earth and 2 degrees from the Sun. We will remain in this attitude until late March.
The \Jspacecraft\j will remain in a relatively quiescent mode for the next two to three months. The flight team is currently working hard to complete planning for Mars entry and surface operations.
7 February 1997
Successfully completed Trajectory Correction Maneuver #2 on February 3. The purpose of this maneuver was to clean up TCM-1 execution errors and had a magnitude of about 1.6 m/s. The maneuver consisted of two parts, an axial component of 1.5 m/s and a lateral component of 0.1 m/s.
All \Jspacecraft\j subsystems performed as expected, and the resulting maneuver execution error was less than 2%. The \Jspacecraft\j was turned back to Earth after the maneuver, and will remain in this attitude until late March.
The \Jspacecraft\j computer reset during an off-track period on Wednesday, February 5. Analysis of post reset telemetry indicates that it was caused by a divide-by-zero fault in the attitude control flight software (ACS). The \Jspacecraft\j responded as expected to this anomaly, and correctly enforced the early cruise boot configuration.
This configuration idles ACS and reduces the uplink and downlink data rates from nominal cruise values. Several commands were sent on February 6 to increase the data rates and return diagnostic data, but attitude control is still idle. The \Jspacecraft\j is in a safe attitude, however, so there is no compelling reason to restart ACS.
An all-hands science team meeting was held on February 5-7. All existing investigators plus the newly selected ASI/MET FIST scientists and participating scientists were invited. The purpose of this meeting was to organize the science teams into a set of science operations groups and to review the current baseline plans for surface operations. A number of useful suggestions were made for modifying the nominal Sol 1-2 plans.
21 February 1997
The \Jspacecraft\j remains in excellent health and is currently about 25 million kilometers from Earth. No significant operational activities were conducted this week, and we have adopted a policy of unattended operations for most of our tracking sessions.
The DSN Radiometric Tiger Team reported the cause of the ~500 Range Unit bias seen in DSS 15 range data between TCM-1 and TCM-2 was caused by an incorrect value for the signal inversion parameter in the SRA configuration table. The error has been fixed and subsequent DSS 15 data looks good.
A set of improvements to the Flight System Testbed Ground Support Equipment Software have been completed which will allow higher fidelity testing of the EDL flight software. Robustness testing of the EDL software has now been started and will continue for the next two months.
28 February 1997
The \Jspacecraft\j remains in excellent health and is currently about 32 million kilometers from Earth. No significant operational activities were conducted this week. Continued investigation of the recent reset and attitude control software problems now indicates that they are related to the Command Detector Unit erroneous lock problem.
A bug was found in the codeblock error detection software which causes a corruption to the floating point registers. The reset and attitude control problems were caused by floating point error conditions and can be tied directly to this corruption.
We are currently in the process of developing a patch to correct this problem. Congratulations to Steve Stolper and Glenn Reeves for quickly identifying the problem and developing the required fix.
A combined Project Science Group meeting involving Mars Pathfinder, Mars Global Surveyor, and Mars Surveyor '98 was held on Thursday and Friday, February 27-28. Although there is not a great deal of overlap between Pathfinder and these other projects, there are some synergistic investigations that can be performed.
The Rover Operations Team completed a Rover Operational Readiness Test in the Mars Yard this week. Preparations are proceeding for next week's project wide surface Operations Readiness Test in the Pathfinder Sandbox. The current plan is to conduct nominal Sol 1 and 2 operations.
#
"Pathfinder Status (March97)",26,0,0,0
7 March 1997
The \Jspacecraft\j is currently about 37 million kilometers from Earth and continues to function as expected. The total travel distance covered since launch is 248 million kilometers, which means that the \Jspacecraft\j has reached the halfway point to Mars.
A set of Entry, Descent, and Landing communications tests were started this week using the \Jspacecraft\j and the Deep Space Network \JGalileo\j Telemetry recorders at Goldstone. These tests are meant to simulate the open loop strategy that we intend to use during entry to record significant events. The first test was successfully completed on March 3, and three additional tests will be performed during the next week.
The project completed Surface Operational Readiness Test #3 on March 7-8. This test was the first formal operations test after launch, and was designed to test the nominal Sol 1 and 2 sequences. Although there were a few start up problems, the test was generally successful. All elements of the project worked well together to complete the critical Sol 1 operations and re-plan Sol 2.
The Rover Operations Team performed remote field testing on Monday and Tuesday. With the SIM Rover at Amboy Crater, the Operations Team ran four Martian sol sequences from JPL. The sequences included navigation and traverse activities, and science and technology experiments. The Pathfinder Science Team also participated in the testing.
14 March 1997
The \Jspacecraft\j remains in good health and is currently about 44 million kilometers from Earth. We experienced a minor command error this week when a command was sent to activate a sequence which was not on-board. The \Jspacecraft\j rejected the command as expected, but we have tightened up our command approval process to prevent future incidents.
The set of Entry, Descent, and Landing (EDL) communications tests started last week were completed this week. All of the Deep Space Network hardware and software elements required to support EDL communications are performing as expected.
EDL Flight software testing is progressing well, and has been the primary focus of testbed activities this week. A full scale airbag retraction test is planned for Friday, March 14 in the Mars sandbox in Building 230.
21 March 1997
The \Jspacecraft\j remains in good health and is currently about 49 million kilometers from Earth. No significant \Jspacecraft\j operations were performed this week.
Entry, Descent, and Landing Flight Software Testing is progressing well. We completed an airbag retraction test in the Building 230 Sandbox last week, and have finished a set of parachute deploy and rocket ignition \Jalgorithm\j robustness tests this week.
A couple of significant issues have come up in this testing which are likely to cause us to change flight software. A flight software change board meeting will be held on April 1 to determine what changes to make and what regression tests to perform.
28 March 1997
The \Jspacecraft\j remains in good health and is currently about 55 million kilometers from Earth. The most significant \Jspacecraft\j activity performed this week was to turn the \Jspacecraft\j to a 5 degree Earth leading attitude. Regular attitude turns will be required for the remainder of cruise to keep the \Jspacecraft\j pointed within 5 degrees of Earth. The propulsion and attitude control subsystems functioned flawlessly after a 7 week hiatus.
Successfully completed a set of sun recognition tests using the Prototype IMP at the University of \JArizona\j. These tests have boosted our confidence that we can successfully perform sun search and point the High Gain Antenna after landing.
An IMP Science Team meeting was held at the University of \JArizona\j on March 25, 26. The sequence changes for Sols 1 and 2 were reviewed. The mission plan for sols 3-5 was discussed and a scenario for the imaging observations was adopted. Discussions of image processing plans and data distribution were held and a number of contentious issues were resolved.
#
"Pathfinder Status (April97)",27,0,0,0
4 April 1997
The \Jspacecraft\j remains in good health and is currently about 64 million kilometers from Earth. The most significant \Jspacecraft\j activity performed this week was to switch to a new convolutional code, K=15, Rate 1/6, on our downlink.
This code gives us a significant increase in our downlink capability. Mars Pathfinder is the first \Jspacecraft\j to operationally use this code with the DSN and we are very pleased with the results.
The project conducted a meeting to discuss Flight Software changes for Entry, Descent and Landing (EDL) and Surface phases of the mission. These changes are to correct bugs found during testing performed since launch. We expect to finalize these changes and prepare for loading them onboard within the next two months.
11 April 1997
The \Jspacecraft\j remains in good health and is currently about 71 million kilometers from Earth. The only \Jspacecraft\j activities performed this week were a regular ASI/MET health check, a Heat Rejection System Pump B cycle, and some modification to Attitude Control Subsystem fault protection parameters. The total flight time since launch is now 128 days, and we have 85 days until Mars arrival.
We successfully completed an operational readiness test of all activities from Mars entry -2 days through Sol 2. This test was performed using the Pathfinder Testbed and Mars Sandbox. The Mars approach phase included periodic updates to the Entry, Descent, and Landing flight software parameter set and execution of a contingency Trajectory Correction Maneuver #5.
We are investigating a minor problem which occurred during airbag retraction, which did not significantly effect the test. The surface operations test included a planned failure of the High Gain Antenna and subsequent Low Gain Antenna operations.
John Wellman and Matt Golombek attended a special Project Science Group meeting at NASA Headquarters to discuss data rights issues and the role of the participating scientists. A number of useful discussions were held and a draft policy on these issues was developed.
18 April 1997
The \Jspacecraft\j remains in good health and is currently about 80 million kilometers from Earth. The only activity performed this week was to continue gathering sun sensor data to characterize its performance at large sun angles. No additional degradation of the sun sensor has been observed since launch. The total flight time since launch is now 135 days, and we have 78 days until Mars arrival.
The flight team is completing preparations for next week's Operational Readiness Test #5. This test is a five day simulation of surface operations using the flight system testbed and Mars sandbox. A significant number of science team members and participating scientists will be in attendance, and are already here this week conducting test and training activities.
The project completed the first of two sessions on lessons learned during the Mars Pathfinder development effort. The first session focused on system level issues and was well received by a lab-wide audience. The second session is scheduled for April 28, and will cover subsystem lessons. Both sessions are being videotaped, and the presentation material will be gathered into a single "book".
25 April 1997
The \Jspacecraft\j remains in good health and is currently about 88 million kilometers from Earth. Major activities performed this week included a regularly scheduled attitude turn to maintain Earth point. We also transitioned to the Late Cruise mission phase and switched in the fourth and final solar panel quadrant. The total flight time since launch is now 142 days, and we have 71 days until Mars arrival.
We successfully completed a week long surface Operational Readiness Test (ORT #5). The purpose of this test was to train team members on operational processes and procedures and verify nominal surface operations plans. Although we had some early difficulties deploying the rover, the test was a great learning experience and an overall success. We did discover several issues with our tools and processes that we will correct prior to ORT #6 (scheduled for May 19).
#
"Pathfinder Status (May97)",28,0,0,0
2 May 1997
The \Jspacecraft\j remains in good health and is currently about 98 million kilometers from Earth. No significant \Jspacecraft\j activities were performed this week. The total flight time since launch is now 149 days, and we have 63 days until Mars arrival.
The EDL planning team completed a very successful Sequence of Events Peer Review. This review covered the detailed flight software sequence used during EDL, recent changes in the EDL software, robustness and regression testing performed since launch, and detailed operational plans for TCM-5 and EDL approach.
The review board, led by Jim Marr, agreed that no significant holes exist in the EDL sequence and that we are well on our way to being prepared for EDL operations.
9 May 1997
The \Jspacecraft\j remains in good health and is currently about 107 million kilometers from Earth. We performed our third Trajectory Correction Maneuver as scheduled on May 6. The purpose of this maneuver was to make a small correction in the trajectory and to test our implementation approach for a potential contingency maneuver just before Mars entry.
The maneuver consisted of three components: a 0.4 m/s lateral burn away from Mars, a 0.1 m/s axial mode burn to correct arrival time, and a 0.5 m/s burn back towards Mars. All three burns were performed without incident, and placed us on a trajectory that is closed to our desired Mars entry trajectory.
We will perform the final planned maneuver (TCM-4) on June 24 to clean up TCM-3 execution errors. The \Jspacecraft\j is now back into quiescent cruise mode. The total flight time since launch is now 156 days, and we have 56 days until Mars arrival.
16 May 1997
The \Jspacecraft\j remains in good health and is currently about 115 million kilometers from Earth (26 million km from Mars). The only major \Jspacecraft\j activity performed this week was a battery heating and solar array characterization test. The total flight time since launch is now 163 days, and we have 49 days until Mars arrival.
Successfully completed the Entry, Descent, and Landing (EDL) and Surface Operations Readiness Review. The review board, led by Mike Sander, asked many useful questions and generated several advisories, but agreed that the project will be ready for pre-entry and surface operations by July 4.
Completed the second of three EDL Operations Readiness Tests (ORT). This test used the testbed to simulate all pre-EDL operations, including TCM-5a. In addition, we conducted a simultaneous EDL communications ORT using the actual \Jspacecraft\j to simulate the telecom behavior during EDL.
Several members of the EDL data acquisition team traveled to Madrid to support this test. Although a number of lessons were learned during the ORT, both the data acquisition and pre-EDL operations teams completed the test successfully.
Completed a set of surface operations mini-ORTs which tested our strategy for petal movements after landing, end-to-end image processing and rover target designation, and the ramp deployment decision process. These three issues were among the most significant concerns in ORT #5, and appear to be resolved at this point.
23 May 1997
The \Jspacecraft\j remains in good health and is currently about 129 million kilometers from Earth. The only significant \Jspacecraft\j activity performed this week was a turn to maintain Earth point attitude. In addition, we have resumed nearly continuous DSN coverage. The total flight time since launch is now 170 days, and we have 42 days until Mars arrival.
Completed the sixth Surface Operations Readiness Test (ORT). The purpose of this test was to validate our low power and no battery contingency scenarios and correct problems from ORT #5. The test was hampered by a recurring testbed hardware problem involving the Imager for Mars Pathfinder (IMP) and the power support equipment.
The problem caused the testbed flight computer to reset several times and caused the flight team to invoke reset recovery procedures instead of performing normal operations. The problem was fixed after Sol 3, at which time we restarted the test and ran through Sol 1-2 operations. In spite of this problem, the test was generally a success in that we exercised all of our core operational processes. Several other minor problems occurred, and all will be corrected by ORT #7.
30 May 1997
The \Jspacecraft\j remains in good health and is currently about 140 million kilometers from Earth (17 million km from Mars). Major \Jspacecraft\j activities performed this week included a turn to maintain Earth point attitude and starting battery charge. The total flight time since launch is now 175 days, and we have 35 days until Mars arrival.
The most significant project activity completed this week was to begin charging the flight battery. Approximately 22 amp-hours of capacity has been taken out of the battery since installation, and the objective of charging was to replace as much as possible.
Degradation of the total battery capacity has occured over the six months since launch, but a total capacity of at least 40 amp-hours should still be possible (compared to a pre-launch capacity of 56 amp-hours). A total of 7 amp-hours has been added after two days of charging, so we are already above the 40 amp-hour level. The flight team is currently assessing whether additional charging is warranted.
Completed updates to the final flight software load. We are currently performing a final set of regression tests prior to patching the software next week.
Deep Space Network personnel corrected the final open ranging data accuracy problem which has been a concern for the last several months. This problem involved larger than expected range jitter in the data acquired at Goldstone. The problem is related to a faulty board in the Sequential Ranging Assembly. All ranging data acquired for the project now meets the pre-launch accuracy specifications.
#
"Pathfinder Status (June97)",29,0,0,0
6 June 1997
The \Jspacecraft\j remains in good health and is currently about 148 million kilometers from Earth (14 million km from Mars). The total flight time since launch is now 182 days, and we have 28 days until Mars arrival.
The flight team completed uploading the final flight software patch to the \Jspacecraft\j on June 5-6. The patch corrects a number of problems in the Entry, Descent, and Landing control software. The total volume of the patch was approximately 400 Kbytes.
The decision was made to delay using the new software until after Operational Readiness Test #7 (ORT #7) because it will be a good final regression test. The \Jspacecraft\j will begin using the new software on June 16.
Completed two mini-ORTs which tested our petal move and low gain antenna contingency plans for Sol 1. All flight sequences for pre-entry and surface operations have now been completed (except for some minor tweaks to the backup mission load), and will be loaded on the \Jspacecraft\j on June 18. We have also completed all preparations for the mission dress rehearsal (ORT #7), which will start on June 9.
Completed a detailed review of Public Outreach plans involving project, program office, and JPL PIO personnel. No major issues or concerns exist, and the detailed implementation activities are proceeding well.
13 June 1997
The \Jspacecraft\j remains in good health and is currently about 160 million kilometers from Earth (10 million km from Mars). The total flight time since launch is now 190 days, and we have 21 days until Mars arrival.
The Operations Team completed the final Operational Readiness Test in preparation for Entry, Descent, and Landing (EDL) and surface operations. This test simulated activities from two days before landing through two days after landing and was very successful.
Major activities completed during this ORT included a simulated Trajectory Correction Maneuver (TCM)-5, a nominal EDL, a series of simulated press conferences for Sol 1 & Sol 2, a successful rover deployment of Sol-1, and extended rover traverses over the course of the following two days.
20 June 1997
The \Jspacecraft\j remains in good health and is currently about 168 million kilometers from Earth (7 million km from Mars). The total flight time since launch is now 196 days, and we have 14 days until Mars arrival.
Completed the final flight software load process including resetting the flight computer. The \Jspacecraft\j is now operating using the new software, including two small changes made as a result of Operational Readiness Test #7. The patch process worked flawlessly and we recovered from the reset within two hours.
Completed the rover flight software patch and health check. A number of changes to the rover software have been identified as a result of post-launch testing, so a small patch was performed this week in conjunction with the planned rover health check. The rover woke up as expected based on lander command, the code patch was accepted, and all \Jengineering\j telemetry measurements were normal.
Completed validating all sequences required for pre-EDL and initial surface operations. Approximately 370 sequences will be loaded on the \Jspacecraft\j starting this weekend.
TMOD conducted a readiness review for Mars Pathfinder EDL and Surface Operations on June 13. No major issues were identified, and all DSN and MGSO elements will be ready to support project operations.
27 June 1997
Mars Pathfinder, now eight days away from landing on the surface of Mars, performed the last of its scheduled trajectory correction maneuvers at 10 a.m. Pacific Daylight Time on Wednesday, June 25.
The correction maneuver was performed in two phases occurring 45 minutes apart. The first burn, lasting just 1.6 seconds, involved firing four thruster engines on one side of the vehicle. The second burn lasted 2.2 seconds and involved firing two thrusters closest to the heat shield.
The combined effect of both burns changed Pathfinder's velocity by 0.018 meters per second (0.04 miles per hour), which places the \Jspacecraft\j on target for a July 4 landing in an ancient flood basin called \JAres\j Vallis. Pathfinder is scheduled to land at 10:07 a.m. PDT (in Earth-received time). The one-way light time from Mars to Earth is 10 minutes, 35 seconds, so in actuality, Pathfinder lands at 9:57 a.m. PDT.
If necessary, a fifth trajectory correction maneuver may be performed just before Pathfinder hits the upper atmosphere of Mars. The maneuver would be carried out either 12 hours or six hours before Pathfinder reaches the atmosphere at 10 a.m. PDT in Earth-received time. The flight team will make a decision to proceed with the final correction maneuver the evening before landing.
A final health check of the \Jspacecraft\j and rover was performed on June 20. All \Jspacecraft\j systems, including science instruments and the critical radar \Jaltimeter\j, remain in excellent health from the last check about six months ago.
The rover received a "wake up" call, woke up on command from the lander, then accepted a software upgrade. Flight controllers next loaded the 370 command sequences that will be required by Pathfinder to carry out its surface operations mission.
The \Jspacecraft\j is now ready to begin its entry, descent and landing phase. It will be commanded into that mode at 1:42 p.m. PDT on June 30 by an onboard sequence.
Mars Pathfinder is currently about 180 million kilometers (111 million miles) from Earth and about 3.5 million kilometers (2.2 million miles) from Mars. After 202 days in flight, the \Jspacecraft\j is traveling at about 18,000 kilometers per hour (12,000 miles per hour) with respect to Mars.
#
"Pathfinder: Entry, Descent and Landing",30,0,0,0
The entry, descent and landing (EDL) process for Mars Pathfinder will begin days before landing when controllers at JPL will send commands to the \Jspacecraft\j to tell it precisely when and how to begin the complex autonomous series of steps necessary to safely land on the surface of Mars. These commands are sent periodically right up to a few hours before landing, when controllers on the Earth will have the most precise knowledge of where the \Jspacecraft\j is relative to Mars (the effect of Mars' gravity well is not felt until the \Jspacecraft\j is less than 48 hours away).
Landing occurs at about 3:00 am local time on Mars, which will be about 10:00 am PDT on Friday, July 4, 1997. From an hour and a half before landing until about 3 and a half hours later, the \Jspacecraft\j is under control of autonomous on-board software that precisely controls the many events that must occur.
The fast-paced approach of Pathfinder at Mars begins with venting of the heat rejection system's cooling fluid about 90 minutes prior to landing. This fluid is circulated around the cruise stage perimeter and into the lander to keep the lander and rover cool during the 7 month cruise phase of the mission.
Its mission fulfilled, the cruise stage is then jettisoned from the entry vehicle about one-half hour prior to landing at a distance of 8500 km from the surface of Mars.
Several minutes before landing, the \Jspacecraft\j begins to enter the outer fringes of the atmosphere about 125 km. (80 mi.) above the surface. Spin stabilized at 2 rpm, and traveling at 7.5 km/sec, the vehicle enters the atmosphere at a shallow 14.8 deg angle. A shallower entry angle would result in the vehicle skipping off the atmosphere, while a steeper entry would not provide sufficient time to accomplish all of the entry, descent and landing tasks.
A Viking-derived aeroshell (including the heatshield) protects the lander from the intense heat of entry. At the point of peak heating the heatshield absorbs more than 100 megawatts of thermal energy. The Martian atmosphere slows the vehicle from 7.5 km/sec to only 400 m/sec (900 mph).
Then entry deceleration of up to 20 gees, detected by on-board accelerometers, sets in motion a sequence of preprogrammed events that are completed in relatively quick succession.
Deployment of the single, 24-ft. diameter parachute occurs 2-3 min. after atmospheric entry at an altitude of 5-11 km. (3-7 mi.) above the surface, eventually slowing the vehicle down to 65 meters/sec. The parachute is similar in design to those used for the Viking program but has a wider band around the perimeter which helps minimize swinging.
The heatshield is pyrotechnically separated from the lander 20 sec. later and drops away at an altitude of 2-9 km. (3-6 mi.). The lander soon begins to separate from the backshell and "rappels" down a metal tape on a centrifugal braking system built into one of the lander petals.
The slow descent down the metal tape places the lander into position at the end of a braided Kevlar tether, or bridle, without off-loading the parachute or placing excessive loads on the backshell. The 20 m bridle provides space for airbag deployment, distance from the solid rocket motor exhaust stream and increased stability.
Once the lander has been lowered into position at the end of the bridle, the radar \Jaltimeter\j is activated and aids in the timing sequence for airbag \Jinflation\j, backshell rocket firing and the cutting of the Kevlar bridle.
The lander's Honeywell radar \Jaltimeter\j is expected to acquire the surface about 32 sec. prior to landing at an altitude of about 1.5 km. The airbags are inflated about 8 sec. before landing at an altitude of 300 meters above the surface.
The airbags have two pyro firings, the first of which cuts the tie cords and loosens the bags. The second, 0.25 sec. later, and 4 sec. before the rockets fire, ignites three gas generators that inflate the three 5.2 m (17-ft) dia. bags to a little less than 1 psi. in less than 0.3 sec.
The conical backshell above the lander contains three solid rocket motors each providing about a ton of force for over 2 seconds. They are activated by the computer in the lander. Electrical wires that run up the bridle close relays in the backshell which ignite the three rockets at the same instant.
The brief firing of the solid rocket motors at an altitude of 80-100 meters is intended to essentially bring the downward movement of the lander to a halt some 12 meters (â–’10 m) above the surface. The bridle separating the lander and heatshield is then cut in the lander, resulting in the backshell driving up and into the parachute under the residual impulse of the rockets, while the lander, encased in airbags, falls to the surface.
Because it is possible that the backshell could be at a small angle at the moment that the rockets fire, the rocket impulse may impart a large lateral velocity to the lander/airbag combination. In fact the impact could be as high as 25 m/sec (56 mph) at a 30 deg grazing angle with the terrain.
It is expected that the lander may bounce at least 12 m about the ground and soar 100-200 m between bounces. (Tests of the airbag system verified that it was capable of much higher impacts and longer bounces.)
Once the lander has settled on the surface, pyrotechnic devices in the lander petal latches are blown to allow the petals to be opened. The latches locking the sturdy side petals in place are necessary because of the pulling forces exerted on the lander petals by the deployed airbag system.
In parallel with the petal latch release, a retraction system will begin slowly dragging the airbags toward the lander, breaching vent ports on the side of each bag, in the process deflating the bags through a cloth filter. The airbags are drawn toward the petals by internal lines extending between attachments within the airbags and small winches on each of the lander sides. It takes about 64 minutes to deflate and fully retract the bags.
There is one high-torque motor on each of the three petal hinges. If the lander comes to rest on its side, it will be righted by opening a side petal with a motor drive to place the lander in an upright position. Once upright, the other two petals are opened.
About 3 hours is allotted to retract the airbags and deploy the lander petals. In the meantime, the lander's X-band radio transmitter will be turned off for the first time since before launch on December 4, 1996. This saves battery power and will allow the transmitter \Jelectronics\j to cool down from being warmed up during entry without the cooling system.
It also allows time for the Earth to rise well above the local horizon and be in a better position for communications with the lander's low-gain antenna later in the morning.
Normal digital data transmissions will cease near the time of cruise stage separation due to the dynamics of EDL. Instead, the transmitter's carrier signal and sidebands will be recorded by the Deep Space Network's Madrid station so that the effects of the many events on the signal may be discerned. The digital data downlink will automatically resume 3.5 hours after landing, long after the airbags have been retracted and the petals opened.
#
"Mir 23 Status Report",31,0,0,0
Mission Control Center, Korolev
\BApril 29, 1997\b
U.S. \Jastronaut\j Jerry Linenger and Mir 23 Commander Vasily Tsibliev conducted a successful five-hour spacewalk today, the first joint U.S.-Russian spacewalk ever undertaken, to attach and retrieve several experiments designed to collect data on the environment around the orbiting space complex.
Linenger and Tsibliev opened the airlock hatch on the Kvant-2 module at 12:10 a.m. Central time this morning and the two spacewalkers went right to work, testing the mobility and design of new Orlan-M spacesuits earmarked for eventual use in the assembly of the International Space Station.
Linenger and Tsibliev reported that new visors in the spacesuit helmets to protect them from the harsh effect of the Sun worked to perfection, and prevented their visors from fogging during the most strenuous periods of activity. Linenger was congratulated by Russian ground controllers at the start of his first spacewalk as he and Tsibliev used a telescoping cargo crane to move themselves and their equipment from the Kvant-2 module to the Mir's Docking Module for the installation of the Optical Properties Monitor (OPM).
The device, which is designed to collect data on the environment around the Mir, was installed near a pair of similar experiments attached to the Docking Module by STS-76 spacewalkers Linda Godwin and Rich Clifford 13 months ago. The so-called Mir Environmental Experiment Packages (MEEPS) will be retrieved by veteran \Jcosmonaut\j Vladimir Titov and \Jastronaut\j Scott Parazynski during a spacewalk outside Atlantis during the STS-86 mission to the Mir in September.
Titov will become the first Russian to conduct a spacewalk wearing a U.S. suit during that excursion. A short time after its installation, the OPM was activated and was reported to be in good working order.
With their first task completed, Linenger and Tsibliev returned to the cargo crane and slowly swung back to the Kvant-2 module, where they installed a meter to monitor radiation levels around the Mir. Video of the spacewalk, downlinked to the Russian Mission Control Center by Flight Engineer Alexander Lazutkin inside Mir, showed the two spacewalkers operating with care near the Mir's delicate solar arrays as they worked to the timeline crafted over the past year.
Linenger and Tsibliev then retrieved a pair of micrometeorite and debris particle collection experiments from the exterior of Kvant-2 which had been left outside by Mir 21 cosmonauts Yuri Onufrienko and Yuri Usachev last year. The experiments were returned to the Kvant-2 airlock where they will be stowed before being brought back to Earth.
Finally, at 5:08 a.m. Central time, after five hours outside Mir, Linenger and Tsibliev returned to Kvant-2 and repressurized the airlock to complete the spacewalk. It was Tsibliev's sixth excursion outside Mir in his two flights dating back to 1993.
Gen. Yuri Glazkov, the Deputy Director of the Gagarin \JCosmonaut\j Training Center in Star City, outside Moscow, congratulated Linenger and Tsibliev for their performance following the completion of the spacewalk for which they had trained for over a year. In a post-spacewalk debriefing with flight controllers, Tsibliev again praised the new spacesuits, particularly the helmet visors and the flexibility of the shoulders, arms and knees, which enabled him and Linenger to move with relative ease outside the station.
Tsibliev plans two more spacewalks outside Mir with Lazutkin in late June and early July to erect an experiment platform on the Spektr module and to prepare valves on the outside of the Core Module for later work designed to add a second carbon dioxide removal system to the outpost.
The crewmembers plan to relax on Wednesday before resuming their scientific agenda and their search and repair of a small cooling loop leak in the Kvant-1 module. Otherwise, the Mir's systems continue to operate normally as plans proceed for the launch of Atlantis in mid-May to deliver U.S. \Jastronaut\j Mike Foale to the complex to replace Linenger.
\BJuly 21 1997\b
Mir Spacewalk Officially Rescheduled to August
Top Russian space officials met Monday and officially announced that the next Mir crew, Commander Anatoly Solovyev and Flight Engineer Pavel Vinogradov, will perform the internal spacewalk to try to restore power from solar arrays outside the damaged Spektr module. The pair is scheduled to be launched Aug. 5, arriving two days later at the Russian station.
Mir 23 cosmonauts Vasily Tsibliev and Alexander Lazutkin are expected to return to Earth on Aug. 14, after the completion of a one-week handover with Solovyev and Vinogradov. The internal spacewalk into Spektr will occur no earlier than Aug. 20.
#
"Shuttle Columbia Mission",32,0,0,0
Space Shuttle Columbia
Last Updated: June 27, 1997, 11:50 a.m. CDT
Preparations for the fifth space shuttle mission of 1997 are on schedule at the Kennedy Space Center in \JFlorida\j, where the shuttle Columbia was rolled out to launch pad 39-A last week.
Ground crews remain on track for a targeted July 1st launch of mission STS-94, the reflight of April's \Jmicrogravity\j science laboratory mission, which was curtailed after four days due to indications of a faulty power-producing fuel cell. Mission managers plan to start Columbia's fuel cells earlier than usual on this mission to allow for additional monitoring prior to liftoff. An official launch date for STS-94 should be set at next Thursday's shuttle program flight readiness review at the Kennedy Space Center.
Why STS-94?
A shuttle flight takes approximately two years to prepare, including all the safety reviews of the planned operations, \Jengineering\j \Jintegration\j to connect the payloads up properly in the payload bay, and development of the flight plan to meet the various payload requirements.
When a flight gets a number, all the documentation that has been prepared carries that designation. If a flight slips for some reason, it would take a lot of time and money to change all the documentation just to keep the order right. Those resources are better spent preparing to go to work in space.
STS-83 landed early in April because of indications of a fuel cell problem. That same flight will fly again reusing all the STS-83 products, with their original designation. The same crew will fly again. But because STS-83 has already flown, NASA will designate the July reflight STS-94, which was the next available number.
The Crew:
James D. Halsell Jr., Commander
Susan Still, Pilot
Janice E. Voss , Mission Specialist
Michael L. Gernhardt, Mission Specialist
Donald Thomas, Mission Specialist
Roger K. Crouch, Payload Specialist
Gregory T. Linteris, Payload Specialist
Columbia Launches
Last Update: 1 July 1997, 6:30 p.m. CDT (-0500 GMT)
After a short delay due to bad weather, Columbia launched at 1:02 p.m. CDT.
The Space Shuttle Columbia and seven astronauts are back in space, resuming a mission of scientific research in \Jmicrogravity\j that was cut short earlier this year by a suspect fuel cell. Less than three months after returning to Earth, the same crew and the \JMicrogravity\j Science Laboratory-1 payloads are once again on orbit to study how certain materials behave and change in the absence of gravity.
This is the twenty-third flight of the oldest orbiter in the NASA fleet, the fifth shuttle mission of this year. The turnaround took only 82 days after the premature landing of Columbia on April 8th. Mission managers decided at that time to bring Columbia home after just four days on orbit due to indications of a problem with that power-producing fuel cell.
After extensive analysis of the fuel cell, the shuttle program concluded that an undetermined and isolated incident caused a slight voltage change in about one-fourth of the 96 cells that make up each fuel cell. To ensure the pre- launch health of those cells this time, the fuel cells were activated earlier than usual in the countdown to provide additional time for monitoring prior to lift-off.
Today's launch, at 1:02 p.m. CDT, was slightly delayed due to weather concerns at the Kennedy Space Center. Shortly after reaching orbit pilot Susan Still reported an unusual reading during shutdown of one of the orbiter's three auxiliary power units, but space shuttle launch \Jintegration\j manager Loren Shriller says mission managers don't believe that reading indicates any problem.
The shuttle's auxiliary power units will be used to supply the hydraulic power to run the orbiter's flight control surfaces during entry and landing. Mission managers also report no indication of anything unusual with Columbia's power-producing fuel cells.
This 85th mission in space shuttle program history is the first time that an entire crew from one mission has flown together a second time. These seven astronauts, operating in two teams to support around-the-clock operations in the Scapulae module, were already on their separate schedules when they started launch day activities very early today.
Columbia is orbiting at 160 nautical, with all systems operating in good shape and the red team astronauts already sent off to bed.
Leading the red team is commander Jim Hales, a 40-year-old Lieutenant Colonel in the U.S. Air Force on his fourth trip to space. Also on the red team is the pilot on this mission, Susan Still. STS-94 is the second spaceflight for the 35-year-old Lieutenant Commander in the U.S navy.
The other two red team members are mission specialist-3 Dr. Don Thomas, a 42-year-old materials scientist on his fourth trip to space, and payload specialist-2 Dr. Greg Linnets, a 39-year-old engineer on his second spaceflight. Linnets is the crew specialist on the combustion experiments in the Scapulae module.
Leading the blue team of astronauts is payload commander Dr. Janice Voss. The 40-year-old aeronautical engineer, making her fourth shuttle flight, has entered the Scapulae module to begin activating its systems. Working with her is payload specialist-1 Dr. Roger Crouch, a 56-year-old physicist on his second spaceflight. Crouch is the chief scientist of NASA's \JMicrogravity\j Science and Applications Division.
The final blue team member is mission specialist-2 Dr. Mike Gernhardt, a 41- year-old bioengineer on his third spaceflight. Gernhardt is seeing to the orbiter systems as flight engineer on this mission.
Orbiter operations for Columbia are being overseen from the Mission Control Center in Houston, and the science operations for the \JMicrogravity\j Science Laboratory are being managed from the Scapulae Operations Control Center, located at the Marshall Spaceflight center in \JHuntsville\j, \JAlabama\j.
A few minutes after 6:00 CDT this evening payload commander Janice Voss will go to work in the spacelab module activating payload systems. At 11:32 p.m. CDT the red team astronauts will get their wake-up call from Mission Control, and half an hour later they'll begin taking a handover from the blue team astronauts. Voss, mission specialist Mike Gernhardt, and payload specialist Roger Crouch begin an eight-hour sleep period shortly after 1:00 a.m.
At 2:17 a.m. Thomas and Linnets will begin their day's work with the experiments in the Scapulae module. At 5:32 a.m. Still will set up the wireless data acquisition system, a risk mitigation experiment for the international space station that will gather remote data on the health of payload hardware. At 9:02 a.m. the wake-up call will sound for the three blue team astronauts.
#
"Space Calendar June97 - July98",33,0,0,0
\JSpace Calendar (June97)\j
\JSpace Calendar (July97)\j
\JSpace Calendar (August97)\j
\JSpace Calendar (September97)\j
\JSpace Calendar (October97)\j
\JSpace Calendar (November97)\j
\JSpace Calendar (December97)\j
\JSpace Calendar (January98)\j
\JSpace Calendar (February98)\j
\JSpace Calendar (March98)\j
\JSpace Calendar (April98)\j
\JSpace Calendar (May98)\j
\JSpace Calendar (June98)\j
\JSpace Calendar (July98)\j
#
"Space Calendar (June97)",34,0,0,0
Jun 13 - \JComet\j C/1997 L3 (SOHO) Perihelion (0.00844 AU)
Jun 13 - Moon Occults Mars
Jun 13-18 - 13th North American Workshop on Cataclysmic Variables, Jackson Hole, Wyoming
Mir represents a unique capability -- an operational space station that can be permanently staffed by two or three cosmonauts. Visiting crews have raised Mir's population to six for up to a month.
#
"Mir Station - Progress-M",51,0,0,0
Progress-M is the unmanned supply ship used to send food and other supplies to the Astronauts and Cosmonauts aboard Mir. It is launched atop a Russian Soyuz SL-4 rocket.
#
"Mir Station - Core Module",52,0,0,0
Mir is the first space station designed for expansion. The 20.4-ton Core Module, Mir's first building block, was launched in February 1986. The Core Module provides basic services (living quarters, life support, power) and scientific research capabilities. It has two axial docking ports, fore and aft, for Soyuz-TM manned transports and automated Progress-M supply ships, plus four radial berthing ports for expansion modules.
The 20.4-ton Core Module, Mir's first building block, was launched in February 1986. The Core Module provides basic services (living quarters, life support, power) and scientific research capabilities. It has two axial docking ports, fore and aft, for Soyuz-TM manned transports and automated Progress-M supply ships, plus four radial berthing ports for expansion modules.
The working compartment is the main habitable volume on Mir and is made up of two concentric cylinders connected by a tapered conical section. The interior of the working compartment is divided into an operations zone and a living area.
Mir crews prefer a spatial orientation of floor and ceiling with the sides arranged in a bottom-to-top orientation despite the formal irrelevance of the terms in the absence of gravity. The floor of the operations area is covered with dark green carpet, the walls are light green and the ceiling is white with fluorescent lamps.
The arrangement of equipment and the interior finish of the working compartment are designed to reinforce this bottom-to-top orientation. The living area uses the same spatial orientation concepts, but soft pastel colors are used to imply a home-like atmosphere.
The living area of the working compartment provides the necessities for long-term human missions. The living area contains a galley area with a table, cooking elements, and trash storage. Individual crew cabins, which include a porthole, hinged chairs and a sleeping bag are found next as one moves axially through the working compartment. The aft end of the working compartment contains the personal hygiene area with toilet, sink, and shower.
#
"Mir Station - Priroda",53,0,0,0
On April 23, 1996, \JRussia\j launched the Mir module Priroda (Nature) on a Proton launcher for rendezvous and docking with the space station on April 26. Weighing nearly 20 tons, the unit carries more than a ton of U.S. cargo for \Jastronaut\j Shannon Lucid aboard the space station. Other Priroda equipment includes optical systems to survey Earth's resources.
When docked, the new module completes the Mir construction complex started ten years earlier; four other modules -- Kristall, Kvant, Kvant-2, and Spektr -- have been launched and attached to the core unit before. Unlike them, Priroda has no solar power arrays but must rely on its on-board batteries as long as it is not docked to Mir.
Priroda was the last module to be added to the Mir. After its launch from Baikonur on April 23, 1996, it docked to the space station as scheduled on April 26. Its primary purpose is to add Earth remote sensing capability to Mir. It also contains the hardware and supplies for several joint U.S.-Russian science experiments.
Its Earth remote sensing capabilities include:
ò monitoring the ecological situation of large industrial areas, estimation of anthropogenous effects on ecological systems
ò measuring concentration and spacial distribution of small gaseous components in atmosphere of ozone and anthropogenous impurities
ò determining temperature fields on the ocean surface and researching the process of energy and mass exchange between ocean and atmosphere affecting the weather
ò receiving data on classification, structure, and moisture of clouds, including their optical characteristics
ò receiving data for plotting geological structure maps on refinement of mineral reserves, water reserves, erosion of soil and conditions of forests and crops
ò acquiring emergency information from buoys in areas of nuclear power stations, seismically dangerous and other zones to create an integrated monitoring and warning system (Kentavr)
ò performing measurements in order to obtain data for working out ecological and economic theory of natural resources utilization
#
"Mir Station - Spektr",54,0,0,0
Launched on a Russian Proton rocket from the Baikonur launch center in central Asia, Spektr was lofted into orbit on May 20, 1995. The module was berthed at the radial port opposite Kvant 2 after Kristall was moved out of the way. Spektr carries four solar arrays and scientific equipment (including more than 1600 pounds of U.S. equipment).
Spektr was badly damaged at 5:18 a.m. EDT on June 25, 1997, when Progress M-34, an unmanned resupply vessel, crashed into the module during tests of the new TORU Progress guidance system. The module lost pressure and electricity and had to be shut completely down and sealed off from the remainder of the Mir complex. It had served as the living quarters for U.S. astronauts aboard Mir. The possibility of repairing the module is under consideration by the Russian Mir team.
The focus of scientific study for this module is Earth observation, specifically natural resources and atmosphere. The equipment onboard is supplied by both Russian and the United States.
Scientific equipment on Spektr includes:
ò Pion, Lira and Buton equipment for atmospheric research.
ò Faza and Feniks equipment for surface studies.
ò Astra-2 equipment for atmospheric trace constituent monitoring.
ò Taurus and Grif equipment for monitoring Mir's induced x-ray and gamma-ray background.
#
"Mir Station - Soyuz-TM",55,0,0,0
Soyuz-TM is the Russian manned \Jspacecraft\j that ferries Cosmonauts and Astronauts to and from Mir. It also serves as an escape "lifeboat" in the event Mir should experience any life-threatening condition.
#
"Mir Station - Kristall",56,0,0,0
Berthed opposite Kvant 2 in 1990, Kristall weighs 19.6 tons and carries two stowable solar arrays, science and technology equipment, and a docking port equipped with a special androgynous docking mechanism designed to receive heavy (up to about 100 tons) \Jspacecraft\j equipped with the same kind of docking unit.
The androgynous unit was originally developed for the Russian Buran Shuttle program. Atlantis used the androgynous docking unit on Kristall during mission STS-71.
Added in 1990, Kristall carries scientific equipment, retractable solar arrays, and a docking node equipped with a special androgynous docking mechanism designed to receive \Jspacecraft\j weighing up to 100 tons.
The purpose of the Kristall module is to develop biological and materials production technologies in the space environment. One component of the Kristall is a radial docking port. Originally designed as a potential means of docking the Russian Buran reusable shuttle orbiter, this port is now attached to the Docking Module.
#
"Mir Station - Docking Module",57,0,0,0
The Docking Module was launched in the payload bay of Atlantis and berthed at Kristall's androgynous docking port during the STS-74 mission. The Docking Module will provide clearance for future Shuttle dockings with Mir and will carry two solar arrays -- one Russian and one jointly developed by the U.S. and \JRussia\j -- to augment Mir's power supply.
#
"Mir Station - Kvant-2",58,0,0,0
Berthed at a radial port since 1989, the module weighs 19.6 tons and carries an EVA airlock, two solar arrays, and science and life support equipment.
Kvant 2, added in 1989, carries an EVA airlock, solar arrays, and life support equipment. The 19.6-ton module is based on the transport logistics \Jspacecraft\j originally intended for the Almaz military space station program of the early 1970s.
The purpose of Kvant-2 is to provide biological research data, Earth observation data, and EVA capability. It adds additional system capability to Mir. Kvant-2 includes additional life support system, drinking water, and oxygen provisions, motion control systems, and power distribution, as well as shower and washing facilities.
Kvant-2 is divided into three pressurized compartments: instrumentation/cargo, science instrument, and airlock.
The airlock not only provides EVA capability, but also contains a self-sustained \Jcosmonaut\j maneuvering unit that increases the range and complexity of tasks that can be attempted via EVA. For instance, various construction materials and electronic components can be placed on the outside of the Mir Complex modules via EVA. The effects of space environment exposure on these construction materials can later be investigated.
#
"Mir Station - Kvant",59,0,0,0
Berthed at the core module's aft axial port in 1987, the module weighs 11 tons and carries telescopes and equipment for attitude control and life support.
Kvant was added to the Mir core's aft port in 1987. This small, 11-ton module contains astrophysics instruments, life support and attitude control equipment.
The purpose of the Kvant-1 module is to provide data and observations for research into the physics of active galaxies, quasars, and neutron stars. This data is gathered with devices which measure electromagnetic spectra and x-ray emissions. The Kvant-1 also supports biotechnology experiments in the areas of antiviral preparations and fractions.
The Kvant-1 module is divided into a pressurized laboratory compartment and a nonpressurized equipment compartment. The laboratory compartment is further divided into an instrumentation area and a living area, which are separated by an interior partition. A pressurized transfer chamber connects the Passive Docking Unit with the laboratory chamber. The nonpressurized equipment compartment contains power stabilizers.
#
"Shuttle Missions (Upcoming)",60,0,0,0
(Refer to Table)
#
"Apollo Missions Contents",61,0,0,0
\JApollo Missions, The\j
\JApollo Flights, The\j
\JSkylab Missions\j
\JApollo-Soyuz Mission Outline\j
\JApollo 7\j
\JApollo 8\j
\JApollo 9\j
\JApollo 10\j
\JApollo 11 - 'One Giant Leap'\j
\JApollo 12\j
\JFlight Of Apollo 13, The\j
\JApollo 13 Accident, Chronology of Events\j
\JApollo 14\j
\JApollo 16\j
\JApollo 17 - Farewell\j
\JFlight Of Apollo-Soyuz, The\j
\JApollo-Soyuz\j
#
"Apollo Missions, The",62,0,0,0
The Apollo Program began before the first American was launched into space. In July, 1960, NASA announced that a program to fly Astronauts around the moon would follow the planned Mercury program, but with President Kennedy's famous speech on May 25, 1961, the focus on the Apollo missions shifted to a lunar landing and came into sharper focus with the concrete goal of achieving this before the decade's end.
Many people feel that the Apollo program stands as mankind's greatest technological achievement. In all, six missions landed on the surface of the moon, and three others orbited the moon without landing, including the ill-fated Apollo 13.
The \Jspacecraft\j was in three parts: The conical Command Module where the crew ate and slept on its way to the moon and home; the Service Module, supplying electricity, maneuvering power and thrust to get home from lunar orbit, and water to the \Jspacecraft\j; and the Lunar Module, or LM, a two-part, totally self-contained \Jspacecraft\j that used its own rockets to land on and take off from the surface of the moon, and even served as its own launch pad.
Apollo missions were launched atop two different boosters, the Saturn 1B used for the Earth orbiting missions (including Skylab and Apollo-Soyuz), and the mighty Saturn V, the rocket to the moon.
Apollo started in tragedy, when a fire on the launch pad in the Command Module of Apollo 1 claimed the lives of our second man in space and first Gemini \Jastronaut\j, Virgil I. "Gus" Grissom, our first space walker, Edward White, and rookie \Jastronaut\j Roger Chaffee on January 27, 1967, in a routine training exercise for what had been scheduled to be the first Apollo mission.
A detailed description of the ill-fated Apollo 1 mission is available from KSC, and images are available from JSC's image library.
Although the \Jspacecraft\j had to be modified to prevent any chance of a recurrence, Apollo 7 was readied for flight by October, 1968, after an unoccupied test (named Apollo 4, the first flight of the Saturn V). Following its success, Apollo 8 , the first human flight of the Saturn V, was launched around the moon in December, and by the following July, Apollo 11 actually placed a man on another celestial body and brought him home again.
Twelve men in all walked on the moon before Apollo was done. The last three missions featured the Lunar Rover, which permitted the astronauts to drive about and explore various terrains too rough for the LM to attempt to land upon. On the last Apollo mission to the moon, the astronauts spent 22 hours in moon walks and camped out on the moon for three days total.
Sadly, Apollo 18, 19 and 20 were canceled due to budget limitations. One of these missions had been scheduled to explore the scientifically intriguing crater Aristarchus, where astronomers through the ages had witnessed geological (or, more properly, "lunalogical") activity through their telescopes and wondered whether or not it might be volcanism. We are still wondering.
We have never yet returned to the moon. How sad. The Apollo \Jspacecraft\j was used for four later missions, the three long-duration Skylab missions and the final Apollo flight, the Apollo-Soyuz linkup with the Soviet Soyuz 19.
#
"Apollo Flights, The",63,0,0,0
(Refer to Table)
#
"Skylab Missions",64,0,0,0
(Refer to Table)
#
"Apollo-Soyuz Mission Outline",65,0,0,0
(Refer to Table)
#
"Apollo 7",66,0,0,0
\BLaunched:\b October 11, 1968\b
\BSplashed Down:\b October 22, 1968\b
\BCrew:\b
ò Walter M. Schirra. Jr.
ò Donn F. Eisele
ò Walter Cunningham
Following the tragedy of the January, 1967, Apollo 1 launchpad fire, the success of this mission was a badly needed confidence builder. The Command Module had been extensively redesigned to eliminate the risk of another conflagration, and the program had fallen behind schedule.
This mission was primarily intended as a shakedown of the new Command Module and the Service Module, including the Service Propulsion System (SPS) that would be relied upon to place Apollo into and out of Lunar orbit.
The mission was a roaring success. Because it was launched without the Lunar Module, it was able to be placed into orbit by the smaller, more extensively tested Saturn 1B rather than the magnificent but still brand-new Saturn V. The SPS performed almost flawlessly on eight separate firings, and the larger and more commodious cabin was significantly more comfortable than the cramped quarters of the Gemini flights.
The eleven days in orbit, though, took their toll: The food was bad, and all three astronauts caught colds. But Apollo 7 demonstrated the spaceworthiness of the basic Apollo vehicle, and truly great deeds were very near in its future.
On this flight, Wally Schirra became the only \Jastronaut\j to fly on Mercury, Gemini and Apollo missions, and will forever remain the one and only man with that honor.
#
"Apollo 8",67,0,0,0
\BLaunched:\b December 21, 1968\b
\BSplashed Down:\b December 27, 1968\b
\BCrew:\b
ò Frank Borman
ò James A. Lovell, Jr.
ò William A. Anders
For the first time in human existence, man finally broke the bounds of the earth when the three Apollo 8 astronauts took man's first trip to the moon. This flight was initially planned as another earth orbiting checkout of the Apollo hardware, but rumors that the Soviets were plotting to beat us into orbit around the moon caused a little last-minute change in plans.
This was the first human spaceflight atop the wondrous Saturn V launch vehicle, which had flown only twice before. There could be no thought of a lunar landing; like Apollo 7, no Luner Module had been launched with them.
The booster worked flawlessly, as did the Service Propulsion Module. Who can forget Christmas Eve, when the astronauts passed behind the moon, and fired their engines while outside communications range?
The whole nation, the whole world, sat and waited on the edge of our seats, waiting for the craft to come out from behind the moon. Did the engines fire properly? Could they come home, or would they remain, forever trapped in the icy cold of space?
It all worked, and Apollo 8 proved at last that our horizons are truly limitless. Man had, at last, seen the Earth rise above the horizon of another world, and brought back the pictures to prove it!
#
"Apollo 9",68,0,0,0
\BLaunched:\b March 3, 1969\b
\BSplashed Down:\b March 13, 1969\b
\BCrew:\b
ò James A. McDivitt
ò David R. Scott
ò Russell L. Schweickart
This mission was the first flight test of the Lunar Module, the third critical piece of Apollo hardware. During ten days in earth orbit, the crew undocked, maneuvered and docked the LM and the Command Modules, simulating as closely as possible the conditions that would be encountered when men finally would land on the moon itself.
This mission also tested the Apollo spacesuit, the first to carry self-contained life support rather than being dependent on an umbilical connection.
This was the first Apollo mission where the astronauts were granted the honor, often exercised rather whimsically, of naming their \Jspacecraft\j. The Apollo 9 astronauts named the Command Module "Gumdrop" and the Lunar Module "Spider." Oh, well, they can't all be Eagles, now, can they?.
After Apollo 8, Apollo 9 was a bit of an anticlimax, but it was another absolutely necessary step that we needed to take if we were going to place a man onto the moon.
#
"Apollo 10",69,0,0,0
\BLaunched:\b May 18, 1969\b
\BSplashed Down:\b May 26, 1969\b
\BCrew:\b
ò Thomas P. Stafford
ò John W. Young
ò Eugene A. Cernan
This was the "dress rehearsal," the penultimate in space exploration. Apollo 10 entered actual orbit around the moon, separated from the Lunar Module, and the LM (named "Snoopy") with Stafford and Cernan aboard descended to within nine miles of the lunar surface before returning and redocking with the waiting John Young in the Command Module (named "Charlie Brown").
The astronauts tested the LM's radar and ascent engine and surveyed Apollo 11's eventual landing site, the Sea of Tranquility. This mission also served as a test of the extensive new Apollo tracking and control network on earth.
Another first for this mission: For the first time, live color TV pictures were broadcast into our homes from space. What a country!
#
"Apollo 11 - 'One Giant Leap'",70,0,0,0
\BLaunched:\b July 16, 1969\b
\Blanded:\b July 20, 1969, Sea of Tranquility\b
\BSplashed Down:\b July 24, 1969\b
\BCrew:\b
ò Neil A. Armstrong
ò Michael Collins
ò Edwin E. "Buzz" Aldrin, Jr.
On July 20, 1969, the human race accomplished its single greatest technological achievement of all time when a man first set foot on another celestial body. We entered a new era, no longer bound by the circles of the earth that had held us so jealously so close to its surface for so long.
On that day we evolved from lowly, ape-like homo sapiens to homo universalis, Man of the Universe, through the power of our minds and the strength of our indomitable will.
Six hours after landing at 4:17 p.m. Eastern Standard Time (with less than 30 seconds of fuel remaining), Neil A. Armstrong took the "Small Step" into our greater future when he stepped off the Lunar Module, named "Eagle", onto the surface of the moon, from which he could look up and see the earth in the heavens as no man had done before him.
He was shortly joined by "Buzz" Aldrin, and the two astronauts spent 21 hours on the lunar surface and returned 46 pounds of lunar rocks. Their liftoff from the surface of the moon was (partially) captured on a TV camera they left behind, and they successfully docked with Michael Collins, patiently orbiting the cold but no longer lifeless moon alone in the Command Module "Columbia."
The moon walkers left behind a plaque on the lunar surface that read:
"Here Men From \JPlanet\j Earth First Set Foot Upon The Moon. July 1969 A.D. We Came In Peace For All Mankind."
Since that day the world's wars have all wound down, implacable foes have become allies, and we have known the longest period of near universal peace and prosperity in recorded history. Coincidence? Perhaps. But then perhaps we all could now see, along with the astronauts, the world as it truly exists, nestled in the heavens, without borders or boundaries, and with wealth beyond the reckoning of kings. And perhaps we became more cognizant of the value of life beyond the need for such petty strife.
In any event, we all grew that day, that fateful day, that greatest of days, back in 1969.
#
"Apollo 12",71,0,0,0
\BLaunched:\b November 14, 1969\b
\BLanded:\b November 18, 1969, Ocean of Storms\b
\BSplashed Down:\b November 24, 1969\b
\BCrew:\b
ò Charles "Pete" Conrad, Jr.
ò Richard F. Gordon
ò Alan L. Bean
This was an exercise of precision targeting. The Lunar Module Intrepid was brought to the surface of the moon automatically by radar and computer, needing only a few manual corrections by the pilot Pete Conrad. It landed only 183 meters from its target in the Ocean of Storms, where the old Surveyor 3 robot \Jspacecraft\j had soft-landed on the surface of the moon back in 1967.
Conrad and Bean brought back pieces of the old Surveyor and took two moon walks, each lasting almost four hours, during which they set up seismic and magnetism experiments and some experiments to test the effects of the solar wind. When the crew, back aboard the Command Module "Yankee Clipper", launched the LM's ascent stage into the moon, the seismometers they had left behind recorded the vibrations of its impact for over an hour.
They had such a good time that they stayed an extra day in lunar orbit taking pictures.
#
"Flight Of Apollo 13, The",72,0,0,0
\BMission Data:\b
ò Pad 39-A (7)
ò Saturn-V AS-508 ()
ò High Bay 1
ò MLP 3
ò Firing Room 1
\BFlight Crew:\b
ò James A. Lovell, Jr.
ò John L. Swigert, Jr.
ò Fred W. Haise, Jr.
\BBackup Crew:\b
ò John W. Young
ò Thomas K. Mattingly II
ò Charles M. Duke, Jr.
\BMilestones:\b
ò 06/13/69 - S-IVB ondock at Kennedy Space Center (KSC)
ò 06/16/69 - S-1C Stage ondock at KSC
ò 06/29/69 - S-II Stage ondock at KSC
ò 07/07/69 - S-IU ondock at KSC
ò 04/11/70 - Launch
\BPayload:\b
ò Odyssey (CM-109) Block II Command and Service Module
ò Aquarius (LM-7) Lunar Module
\BLaunch:\b
ò April 11, 1970 , 19:13:00 Greenwich Mean Time (GMT)
\BEarth Orbit:\b
ò Translunar injection performed over Pacific Ocean before completion of second revolution
\BMission Duration:\b
ò 05 Days, 22 hours, 54 minutes
\BLanding:\b
ò April 17, 1970
\BMission Highlights:\b
Third lunar landing attempt. Mission aborted after rupture of service module oxygen tank. Classed as "successful failure" because of experience in rescuing crew. Spent upper stage successfully impacted on the Moon.
\BMission Narrative:\b
Apollo 13, the third human lunar landing and exploration mission, had been tentatively scheduled in July 1969 for launch in March 1970, but by the end of the year the launch date had been shifted to April. In August 1969 crew assignments for Apollo 13 were announced, eventually James A. Lovell commanded the mission, with Fred Haise and John Swigert.
The target for the mission was the Fra Mauro Formation, a site of major interest to scientists, specifically a spot just north of the crater Fra Mauro, some 550 kilometers (340 miles) west-southwest of the center of the Moon's near side.
On March 24, 1970, during the countdown demonstration test for Apollo 13, Kennedy Space Center test engineers encountered a problem with an oxygen tank in the service module. The \Jspacecraft\j carried two such tanks, each holding 320 pounds (145 kilograms) of supercritical oxygen. They provided the oxygen for the command module atmosphere and (along with two tanks of hydrogen) three fuel cells, which were the \Jspacecraft\j's primary source of electrical power.
Besides power, the chemical reaction in the cells produced water, which not only supplied the crew's drinking water but was circulated through cooling plates to remove heat from certain critical electronic components. The tanks were designed to operate at pressures of 865 to 935 pounds per square inch (psi) (6,000 to 6,450 kilopascals) and temperatures between-340 degrees F and +80 degrees F (-207 degrees C to +27 degrees C). Inside each spherical tank were a quantity gauge, a thermostatically controlled heating element, and two stirring fans driven by electric motors.
The fans were occasionally operated to homogenize the fluid in the tank; it tended to stratify, leading to erroneous quantity readings. All wiring inside the tank was insulated with Teflon, a fluorocarbon plastic that is ordinarily noncombustible.
Each tank was fitted with a relief valve designed to open when the pressure rose above 1,000 psi (6,900 kilopascals); the tanks themselves would rupture at pressures above 2,200 psi (15,169 kilopascals). Both tanks were mounted on a shelf in the service module between the fuel cells and the \Jhydrogen\j tanks.
The countdown demonstration test called for the tanks to be filled, tested, and then partially emptied by applying pressure to the vent line, thus forcing oxygen out through the fill line. Number one tank behaved normally in this test, but number two released only 8 percent of its contents, not 50 percent as required.
Test engineers decided to proceed with the rest of the test and investigate the problem later. The next day, after KSC engineers had discussed the problem with colleagues at the Manned \JSpacecraft\j Center (MSC, the Johnson Space Center after 1973), North American Rockwell (builders of the service module), and Beech \JAircraft\j (manufacturers of the oxygen tanks), they tried emptying the tank again, with no success.
Further talks led to the conclusion that the tank probably contained a loose-fitting fill tube, which could allow pressure to escape without emptying the tank.
When normal procedures again failed to empty the tank, engineers decided to use its internal heaters to boil off the contents and applied direct-current power at 65 volts to the heaters. This was successful but slow, requiring eight hours of heating. It was then decided that if the tank could then be filled normally it would not cause a problem in flight. A third test gave the same result as the second, requiring heating to empty the tank.
In view of the difficulty of replacing the oxygen shelf a job that would take at least 45 hours and the possibility that other components might be damaged in the process and the launch delayed for a month, NASA and contractor officials decided not to replace the tanks.
The \Jspacecraft\j was launched on April 11, 1970, and the mission was quite routine for the first two days. At 30 hours and 40 minutes after launch (30:40 ground elapsed time), the crew ignited their main engine to put the \Jspacecraft\j on a hybrid trajectory, a flight path that saved fuel in reaching the desired lunar landing point.
At 46:40 the crew routinely switched on the fans in the oxygen tanks briefly. A few seconds later the quantity indicator for tank number two went off the high end of the scale, where it stayed. The tanks were stirred twice more during the next few hours; and at 55:53, after a master alarm had indicated low pressure in a \Jhydrogen\j tank, the Mission Control Center (MCC) directed the crew to switch on all tank stirrers and heaters.
Shortly thereafter the crew heard a loud banging and felt unusual vibrations in the \Jspacecraft\j. Mission controllers noticed that all telemetry readings from the \Jspacecraft\j dropped out for 1.8 seconds. In the command module (CM), the caution and warning system alerted the crew to low voltage on d.c. main bus B, one of two power distribution systems in the \Jspacecraft\j. At this point command module pilot Jack Swigert told Houston, "Hey, we've had a problem here."
Because of the interruption of telemetry that had just occurred, flight controllers in the MCC had difficulty for the next few minutes determining whether they were getting true readings from the \Jspacecraft\j sensors or whether the sensors had somehow lost power.
Before long, however, both MCC and the crew realized that oxygen tank number two had lost all of its contents, oxygen tank number one was slowly losing its contents, and the CM would soon be out of oxygen and without electrical power. Among the first actions taken were shutting down one fuel cell and switching off nonessential systems in the CM to minimize power consumption; shortly after, the second fuel cell was shut down as well. When the remaining oxygen ran out, the CM would be dead; its only other power source was three reentry batteries providing 120 ampere-hours, and these had to be reserved for the critical reentry period.
An hour and a half after the "bang," MCC notified the crew that "we're starting to think about the lifeboat" using the lunar module (LM) and its limited supplies to sustain the crew for the rest of the mission. Plans for such a contingency had been studied for several years, although none had anticipated a situation as grave as that of Apollo 13. Many of these studies were retrieved and their results were adapted to the situation as it developed.
Shortly after the accident, mission commander James Lovell reported seeing a swarm of particles surrounding the \Jspacecraft\j, which meant trouble. Particles could easily be confused with stars, and the sole means of determining the \Jspacecraft\j's attitude was by locating certain key stars in the onboard \Jsextant\j. Navigational sightings from the lunar module (LM) were difficult in any case as long as it was attached to the command module, and this would only complicate matters.
Flight controllers decided to align the lunar module's guidance system with that in the command module while the CM still had power. That done, the last fuel cell and all systems in the command module were shut down, and the crew moved into the lunar module. Their survival depended on this craft's oxygen and water supplies, guidance system, and descent propulsion engine (DPS).
Normally all course corrections were made using the service propulsion system (SPS) on the service module, but flight controllers ruled out using it, partly because it required more electrical power than was available and partly because no one knew whether the service module had been structurally weakened by the explosion. If it had, an SPS burn might be dangerous. The DPS would have to serve in its place.
When word got out that Apollo 13 was in trouble, off-duty flight controllers and \Jspacecraft\j systems experts began to gather at MSC, to be available if needed. Others stood by at NASA centers and contractor plants around the country, in touch with Houston by \Jtelephone\j. Flight directors Eugene Kranz, Glynn Lunney, and Gerald Griffin soon had a large pool of talent to help them solve problems as they arose, provide information that might not be at their fingertips, and work on solutions to problems they could anticipate further along in the mission.
Astronauts occupied the CM and LM training simulators at Houston and at Kennedy Space Center, testing new procedures as they were devised and modifying them as necessary. MSC director Robert R. Gilruth, Dale D. Myers, director of manned space flight, and NASA administrator Thomas O. Paine were all on hand at Mission Control to provide high-level authority for changes.
Soon after the explosion, the assessment of life-support systems determined that although oxygen supplies were adequate, the system for removing carbon dioxide (CO2) in the lunar module was not. The system used canisters filled with \Jlithium\j \Jhydroxide\j to absorb CO2 as did the system in the command module.
Unfortunately the canisters were not interchangeable between the two systems, so the astronauts were faced with plenty of capacity for removing CO2 but no way of using it. A team in Houston immediately set about improvising a way to use the CM canisters, using materials available in the \Jspacecraft\j.
Flight controllers, meanwhile, were addressing operational problems. Their first critical decision was to put the crippled \Jspacecraft\j back on a free-return trajectory, which was accomplished by firing the LM descent engine at 61:30.
Mission Control then had some 18 hours to consider the remaining problems; the next was a possible adjustment to change the \Jspacecraft\j's landing point on Earth. If this was to be done, it was scheduled for "PC + 2" two hours after pericynthion (closest approach to the Moon), after the \Jspacecraft\j emerged from behind the Moon. In the interval, Houston worked out a new flight plan that would minimize the consumption of oxygen, water, and electricity while keeping vital systems operating.
The alternatives for the PC +2 maneuver were worked out by about 64 hours into the flight. A major consideration was the total time to splashdown. Left on its free-return course the command module would return at about 155 hours after launch to a landing in the Indian Ocean. Three options would bring it back in the mid-Pacific and could reduce the total mission time to as little as 118 hours.
The fifth possibility returned the \Jspacecraft\j in 133 hours, but to the South Atlantic.
For one reason or another, all but one of these choices were discarded. The free-return (no course correction) choice was abandoned, since there was no known reason not to use the LM descent propulsion system. Recovery in either the Atlantic or the Indian Ocean was far from ideal; the main recovery force was deployed in the mid-Pacific and there was not enough time to move it or to make adequate arrangements elsewhere.
Two options giving the shortest return time (118 hours) had other drawbacks. Both would require using virtually all of the available propellant, and it was not prudent to assume that no additional course corrections would be required. One of them involved jettisoning the service module, which would expose the CM heat shield to the cold of space for 40 hours and raise questions about its integrity on reentry.
After five and a half hours of weighing the choices and their consequences, flight directors met with NASA and contractor officials and presented their findings and recommendations. The decision, made some ten hours before the scheduled engine burn, was to go for mid-Pacific recovery at 143 hours.
During all of these deliberations the atmosphere in the lunar module was gradually accumulating carbon dioxide as the absorbers in the environmental control system became saturated. Members of MSC's Crew Systems Division devised a makeshift air purifier by taping a plastic bag around one end of a CM \Jlithium\j \Jhydroxide\j cartridge and attaching a hose from the portable life-support system, allowing air from the cabin to be circulated through it.
After verifying that this jury rig would function, they prepared detailed instructions for building it from materials available in the \Jspacecraft\j and read them up to the crew. For the rest of the mission the improvised system kept the CO2 content of the atmosphere well below hazardous levels.
The decision to recover in the Pacific fixed the time line for the remainder of the mission and imposed some rigid constraints on preparations for reentry. The final course correction had to be made with the LM engine; command module systems had to be turned on and the guidance system aligned; the service module had to be discarded; and when all preparations had been made, the lunar module would be cut loose.
In all these preparations the power available from the CM's reentry batteries was a limiting factor. From the PC + 2 burn until about 35 hours before reentry the sequence of activation of CM systems was worked out, checked in the simulators, and modified. Fifteen hours before beginning reentry the revised sequence of activities was read to the crew, to give them time to review and practice it.
The husbanding of expendable resources, particularly electrical power, paid off on the morning of landing, when it was discovered that power reserves in the LM were adequate to allow use of it in the CM. Some of the early CM activities could then be done at a less hurried pace.
The Apollo 13 command module splashed down within a mile of the recovery carrier with about 20 percent of its battery power remaining. Three weary, chilled astronauts came aboard the U.S.S. Iwo Jima on April 17 and were flown to Hawaii for an emotional reunion with their families.
Mission Control teams and their hundreds of helpers were no less drained. The usual cigars were lit up after recovery, but the splashdown parties that evening were subdued: most of those who went quit early and went home to bed. Their efforts were recognized the next day when President Richard M. Nixon, on his way to Hawaii, stopped in Houston to present the Presidential Medal of Freedom, the nation's highest civilian award, to the entire team.
NASA immediately convened an investigation board to determine the cause of the accident and postponed Apollo 14 until its results were in. Lacking the \Jspacecraft\j itself, the service module had been jettisoned before reentry, and the crew had been able to take only a few rather poor photographs of it, the board initially had only the data from inflight telemetry to work with.
When it became clear that the fault lay in oxygen tank number two, the board carefully reviewed its entire history, from fabrication to launch, as recorded in the detailed documentation that followed every piece of equipment from plant to launch pad. Under the board's direction, MSC and other NASA centers conducted tests under simulated mission conditions to verify its findings. The investigation, which concluded in a few weeks, turned up a highly improbable sequence of human error and oversight that led inexorably to the failure in flight.
Board Chairman Edgar M. Cortright, director of Langley Research Center, explained the board's findings to congressional committees in June 1971. The accident, he reported, was not a random malfunction but resulted from an unusual combination of mistakes as well as "a somewhat deficient and unforgiving design."
As the board's report reconstructed the events leading up to the accident, the tank left Beech \JAircraft\j's plant on May 3, 1967, after passing all acceptance tests. It was installed as part of a shelf assembly in service module no. 106 on June 4, 1963, having passed all tests conducted at North American Rockwell during assembly. Design changes in the service module, however, necessitated removing the entire shelf from SM 106 for modification.
During removal, which was accomplished by use of a special fixture that fit under the shelf to lift it upward, workmen overlooked one bolt that held down the back of the shelf, with the result that the removal fixture broke, dropping the shelf two inches. The board concluded that this incident might have jarred loose a poorly fitting fill tube. Subsequent tests did not detect any flaws, and after modification the shelf was shipped to Kennedy Space Center for installation in SM 109, the Apollo 13 \Jspacecraft\j.
What was not known was that this oxygen tank was fitted with obsolete thermostatic switches protecting its heating elements. Original specifications for the switches called for operation on 28 volts d.c.; in 1965 this was changed to 65 volts d.c. to match the test and checkout equipment at the Cape. Later tanks conformed to the new specifications, but this one, which should have been modified, was not, and the discrepancy was overlooked at all stages thereafter.
In a normal checkout of a normal tank, this would not have mattered, because the switches would not have opened during normal operation. But the improvised procedure used when this tank failed to empty (the result of a loose fitting, as noted above) raised the temperature in the tank above 80 degrees F (27 degrees C), at which point the switches opened.
Tests conducted during the investigation showed that the higher current produced by the 65-volt power source caused an arc between the contact points as they separated, welding them together and preventing their opening when the temperature dropped. This went undetected during the detanking procedure at the Cape; it could have been noticed if anyone had monitored the heater current, which would have shown that the heaters were operating when they should not have been.
But all attention was on the specific malfunction, and no one was aware that the heaters were on continuously for eight hours on two separate occasions. The result, as tests showed, was that the heater tube reached 1,000 degrees F (538 degrees C) in spots, damaging the Teflon \Jinsulation\j on the adjacent fan-motor wiring and exposing bare wire. From that point on, the board concluded, the tank was hazardous when filled with oxygen and electrically powered. Teflon can be ignited at a high enough temperature in the presence of pure oxygen, and the tank contained small amounts of other combustibles as well.
Unfortunately for Apollo 13, the tank functioned normally for the first 56 hours of the mission, when the heaters and the fans were energized during routine operations. At that point an arc from a short circuit probably ignited the Teflon, and the rapid pressure rise that followed either ruptured the tank or damaged the conduit carrying wiring into the tank, expelling high-pressure oxygen.
The board could not determine exactly how the tank failed or whether additional combustion occurred outside the tank, but the pressure increase blew off the panel covering that sector of the service module and damaged the directional antenna, causing the interruption of telemetry observed in Houston. It also evidently damaged the oxygen distribution system, or the other oxygen tank, as well, leading to the loss of all oxygen supplies and aborting the mission.
The board pointed out that although the circumstances of the tank failure were highly unusual and that the system had worked flawlessly on six successful missions, Apollo 13 was a failure whose causes had to be eliminated as completely as possible. It recommended that the oxygen tanks be modified to remove all combustible material from contact with oxygen and that all test procedures be thoroughly reviewed for adequacy.
Compared to the AS-204 fire in 1967, Apollo 13 was only a frightening near-miss, and because its cause was localized and comparatively easy to discover, it had fewer adverse effects on the program. Only the skill and dedication of hundreds of members of the often-celebrated "manned space flight team" saved it, however, and the accident served to remind NASA and the public that human flight in space, no matter how commonplace it seemed to the casual observer, was not a routine operation.
#
"Apollo 13 Accident, Chronology of Events",73,0,0,0
The following includes events from 2.5 minutes before the accident to about 5 minutes after. Times given are in Ground Elapsed Time (G.E.T.), that is, the time elapsed since liftoff of Apollo 13 on April 11, 1970, at 2:13 PM Eastern Standard Time (EST). 55:52:00 G.E.T. is equal to 10:05 PM EST on April 13, 1970.
55:52:31 - Master caution and warning triggered by low \Jhydrogen\j pressure in tank no. 1
55:52:58 - CapCom (Charlie Duke): "13, we've got one more item for you, when you get a chance. We'd like you to stir up the cryo tanks. In addition, I have shaft and trunnion .....
55:53:06 - Swigert: "Okay."
55:53:07 - CapCom: ".... for looking at \JComet\j Bennett, if you need it."
55:53:12 - Swigert: "Okay. Stand by."
55:53:18 - Oxygen tank No. 1 fans on.
55:53:19 - Oxygen tank No. 2 pressure decreases 8 psi.
55:53:20 - Oxygen tank No. 2 fans turned on.
55:53:20 - Stabilization control system electrical disturbance indicates a power transient.
55:53:21 - Oxygen tank No. 2 pressure decreases 4 psi.
55:53:22.718 - Stabilization control system electrical disturbance indicates a power transient.
55:53:22.757 - 1.2 Volt decrease in ac bus 2 voltage.
55:53:22.772 - 11.1 \Jamp\j rise in fuel cell 3 current for one sample.
55:53:26 - Oxygen tank No. 2 pressure begins rise lasting for 24 seconds.
55:53:38.057 - 11 volt decrease in ac bus 2 voltage for one sample.
55:53:38.085 - Stabilization control system electrical disturbance indicates a power transient.
55:53:41.172 - 22.9 \Jamp\j rise in fuel cell 3 current for one sample
55:53:41.192 - Stabilization control system electrical disturbance indicates a power transient.
55:54:00 - Oxygen tank No. 2 pressure rise ends at a pressure of 953.8 psia.
55:54:15 - Oxygen tank No. 2 pressure begins to rise.
55:54:30 - Oxygen tank No. 2 quantity drops from full scale for 2 seconds and then reads 75.3 percent.
55:54:31 - Oxygen tank No. 2 temperature begins to rise rapidly.
55:54:43 - Flow rate of oxygen to all three fuel cells begins to decrease.
55:54:45 - Oxygen tank No. 2 pressure reaches maximum value of 1008.3 psia.
55:54:51 - Oxygen tank No. 2 quantity jumps to off-scale high and then begins to drop until the time of telemetry loss, indicating failed sensor.
55:54:52 - Oxygen tank No. 2 temperature sensor reads -151.3 F.
55:54:52.703 - Oxygen tank No. 2 temperature suddenly goes off-scale low, indicating failed sensor.
55:54:52.763 - Last telemetered pressure from oxygen tank No. 2 before telemetry loss is 995.7 psia.
55:54:53.182 - Sudden accelerometer activity on X, Y, Z axes.
55:54:53.220 - Stabilization control system rate changes begin.
55:54:53.323 - Oxygen tank No. 1 pressure drops 4.2 psi.
55:54:53.500 - 2.8 \Jamp\j rise in total fuel cell current.
55:54:53.542 - X, Y, and Z accelerations in CM indicate 1.17g, 0.65g, and 0.65g.
55:54:53.555 - Master caution and warning triggered by DC main bus B undervoltage. Alarm is turned off in 6 seconds. All indications are that the cryogenic oxygen tank No. 2 lost pressure in this time period and the panel separated.
55:54:54.741 - \JNitrogen\j pressure in fuel cell 1 is off-scale low indicating failed sensor.
55:54:55.350 - Telemetry recovered.
55:54:56 - Service propulsion system engine valve body temperature begins a rise of 1.65 F in 7 seconds. DC main A decreases 0.9 volts to 28.5 volts and DC main bus B 0.9 volts to 29.0 volts. Total fuel cell current is 15 amps higher than the final value before telemetry loss. High current continues for 19 seconds. Oxygen tank No. 2 temperature reads off-scale high after telemetry recovery, probably indicating failed sensors. Oxygen tank No. 2 pressure reads off-scale low following telemetry recovery, indicating a broken supply line, a tank pressure below 19 psi, or a failed sensor. Oxygen tank No. 1 pressure reads 781.9 psia and begins to drop.
55:54:57 - Oxygen tank No. 2 quantity reads off-scale high following telemetry recovery indicating failed sensor.
55:55:01 - Oxygen flow rates to fuel cells 1 and 3 approached zero after decreasing for 7 seconds.
55:55:02 - The surface temperature of the service module oxidizer tank in bay 3 begins a 3.8 F increase in a 15 second period. The service propulsion system \Jhelium\j tank temperature begins a 3.8 F increase in a 32 second period.
55:55:09 - DC main bus A voltage recovers to 29.0 volts, DC main bus B recovers to 28.8.
55:55:20 - Swigert: "Okay, Houston, we've had a problem here."
55:55:28 - Duke: "This is Houston. Say again please."
55:55:35 - Lovell: "Houston, we've had a problem. We've had a main B bus undervolt."
55:55:42 - Duke: "Roger. Main B undervolt."
55:55:49 - Oxygen tank No. 2 temperature begins steady drop lasting 59 seconds indicating a failed sensor.
55:56:10 - Haise: "Okay. Right now, Houston, the voltage is--is looking good. And we had a pretty large bang associated with the caution and warning there. And as I recall, main B was the one that had an \Jamp\j spike on it once before.
55:56:30 - Duke: "Roger, Fred."
55:56:38 - Oxygen tank No. 2 quantity becomes erratic for 69 seconds before assuming an off-scale low state, indicating a failed sensor.
55:56:54 - Haise: "In the interim here, we're starting to go ahead and button up the tunnel again."
55:57:04 - Haise: "That jolt must have rocked the sensor on -- see now -- oxygen quantity 2. It was oscillating down around 20 to 60 percent. Now it's full-scale high."
55:57:39 - Master caution and warning triggered by DC main bus B undervoltage. Alarm is turned off in 6 seconds.
55:57:40 - DC main bus B drops below 26.25 volts and continues to fall rapidly.
55:57:44 - Lovell: "Okay. And we're looking at our service module RCS \Jhelium\j 1. We have -- B is barber poled and D is barber poled, \Jhelium\j 2, D is barber pole, and secondary propellants, I have A and C barber pole." AC bus fails within 2 seconds.
55:57:45 - Fuel cell 3 fails.
55:57:59 - Fuel cell current begins to decrease.
55:58:02 - Master caution and warning caused by AC bus 2 being reset.
55:58:06 - Master caution and warning triggered by DC main bus undervoltage.
55:58:07 - DC main bus A drops below 26.25 volts and in the next few seconds levels off at 25.5 volts.
55:58:07 - Haise: "AC 2 is showing zip."
55:58:25 - Haise: "Yes, we got a main bus A undervolt now, too, showing. It's reading about 25 and a half. Main B is reading zip right now."
56:00:06 - Master caution and warning triggered by high \Jhydrogen\j flow rate to fuel cell 2.
#
"Apollo 14",74,0,0,0
\BLaunched:\b January 31, 1971\b
\BLanded:\b February 3, 1971, Fra Mauro Region\b
\BSplashed Down:\b February 9, 1971\b
\BCrew:\b
ò Alan B. Shepard, Jr.
ò Stuart A. Roosa
ò Edgar D. Mitchell
Following the frightening problems of Apollo 13, almost ten months elapsed before we returned to the moon with Apollo 14. This flight marked the return to space of America's first spaceman, Alan B. Shepard, who had first flown aboard Freedom 7 a decade earlier.
Shepard and Mitchell had to scrap a planned rock-collecting trip to the 1,000 foot wide Cone Crater when they became disoriented and almost got lost. Interestingly, it was later discovered that they were only a little over 30 yards from the crater's rim when they gave up the search.
After their return aboard the Command Module "Kitty Hawk,", the three Apollo 14 astronauts became the last to be required to undergo a period of quarantine.
#
"Apollo 16",75,0,0,0
\BLaunched:\b April 16, 1972\b
\BLanded:\b April 20, 1972, Descartes Highlands\b
\BSplashed Down:\b April 27, 1972\b
\BCrew:\b
ò John W. Young
ò Thomas K. Mattingly II
ò Charles M. Duke, Jr.
A malfunction in the main propulsion system of the Lunar Module "Orion" almost scrubbed the landing, but after its success Young and Duke spent three days exploring the geologically interesting Descartes Highlands region while Mattingly circled overhead in the Command Module "Casper."
It was thought that Descartes may be an area of active volcanism, but this proved not to be the case. Among other specimens, the astronauts returned the largest moon rock ever, a 23-pound chunk that turned out to contain not even a single gram of green cheese.
During this flight the moon racers also had a bit of fun testing out the capabilities of the Lunar Rover, at one point getting it up to almost 11 miles per hour!
#
"Apollo 17 - Farewell",76,0,0,0
\BLaunched:\b December 7, 1972\b
\BLanded:\b December 11, 1972, Taurus-Littrow Valley\b
\BSplashed Down:\b December 19, 1972\b
\BCrew:\b
ò Eugene A. Cernan
ò Ronald E. Evans
ò Harrison H. "Jack" Schmitt
Schmitt and Cernan left behind the Lunar Module "Challenger" and drove around almost 34 kilometers of lunar ground during this, the last mission to the moon. Schmitt was the first scientist, a geologist by trade, to visit the Moon, and became the last man (so far) to set foot on another celestial body.
The astronauts collected a record 108.86 kilograms of lunar rocks, including some that were orange in color, during their three moon walks and drives through the Taurus-Littrow Valley.
The last three scheduled Apollo missions had already been cancelled because of budgetary shortfalls by the time of this flight. The astronauts left behind a plaque that read:
"Here Man completed his first exploration of the Moon, December 1972 A.D. May the spirit of peace in which we came be reflected in the lives of all mankind."
#
"Flight Of Apollo-Soyuz, The",77,0,0,0
\BFlight Crew:\b
ò Apollo:
ò Thomas P. Stafford
ò Vance D. Brand
ò Donald K. Slayton
ò Soyuz:
ò Alexey A. Leonov
ò Valery N. Kubasov
\BLaunch:\b
ò Apollo: July 15, 1975
ò Soyuz: July 15, 1975
\BLanding:\b
ò Apollo: July 24, 1975
ò Soyuz: July 21, 1975
\BMission Duration:\b
ò 09 days, 07 hours, 28 minutes
ò July 15-24, 1975
\BMission Highlights:\b
The Soyuz was launched just over seven hours prior to the launch of the Apollo CSM. Apollo then maneuvered to rendezvous and docking 52 hours after the Soyuz launch. The Apollo and Soyuz crews conducted a variety of experiments over a two-day period. After separation, Apollo remained in space an additional 06 days. Soyuz returned to Earth approximately 30 hours after separation.
\BMission Narrative:\b
The final flight of the Apollo program was the first spaceflight in which \Jspacecraft\j from different nations docked in space. In July 1975, a U.S. Apollo \Jspacecraft\j carrying a crew of three docked with a Russian Soyuz \Jspacecraft\j with its crew of two.
For the Apollo-Soyuz Test Project (ASTP), the United States used an Apollo Command and Service Module (CSM) modified to provide for experiments to be conducted during the mission, extra propellant tanks and the addition of controls and equipment related to the Docking Module. Launch was accomplished with a Saturn IB.
The Docking Module was designed jointly by the United States and Soviet Union, and built in the United States. Its purpose was to enable a docking between the dissimilar Soyuz \Jspacecraft\j and the U.S. Apollo. It was a three meter long cylinder 1.5 meters in diameter, and in addition to serving as a docking device, also served as an airlock module between the different atmospheres of the two ships (the U.S. ship with 100% oxygen at 260 millimeters of mercury; the Soyuz with a mixed oxygen-nitrogen atmosphere at 520 mm HG--lowered from its usual 760 mm Hg for this mission).
Prior to the conduct of ASTP, the astronauts and cosmonauts visited each other's space centers and became familiar with the \Jspacecraft\j of the other country. The first visit was by the Russians to Johnson Space Center in July 1973, followed by a U.S. visit to Moscow in November 1973. In late April and early May 1974, the Russian flight crews returned to Johnson Space Center, and the U.S. crews went to Moscow in June and July 1974.
The Russian crew made a third trip to the United States in September 1973 and came for the fourth and last time in February 1975. The U.S. crew visited the Soviet Union in late April and early May 1975 and became the first Americans to see the Russian launch facilities at Tyuratam on April 28, 1975.
Three simulation sessions were conducted between flight controllers and the ASTP crew in Houston and Moscow on May 13, 15 and 18, 1975 involving communications links between the two control centers, and fully occupied control center facilities. A final simulation was conducted from June 30-July 1, 1975. Additionally, in December 1974, the Russians made a human flight of the modified version of the Soyuz spaceship for system tests (Soyuz 16).
One of the most difficult problems to overcome was that of language differences. To alleviate this problem as much as possible, the Americans learned Russian and the Russians learned English. It was found that the best scenario was for the Russians to speak English and for the Americans to speak Russian.
Soyuz Launch: Soyuz 19, carrying cosmonauts Aleksey A. Leonov and Valery N. Kubasov, was launched into sunny skies from Baykonur Cosmodrome at 5:20 pm local time (8:20 am EDT) July 15, 1975. The \Jspacecraft\j entered orbit with a 221.9-km apogee, 186.3-km perigee, 88.5-min period, and 51.8 inclination.
Foreign correspondents, barred from the launch site, watched the launch on color TV sets in a Moscow press center. The first Soviet launch to be televised live, it was transmitted to viewers throughout the Soviet Union, the U.S., and eastern and western Europe. President Ford watched from a U.S. State Dept. auditorium with Soviet Ambassador to the U.S, Anatoly P. Dobrynin and NASA Administrator James C. Fletcher, before Dr. Fletcher and Ambassador Dobrynin flew to Kennedy Space Center to watch the Apollo launch.
On the third orbit the Soyuz 19 crew established contact with U.S. mission control in Houston, putting into operation the global Moscow and Houston Soyuz-Apollo communications system. On the fifth orbit the cosmonauts made the first of two maneuvers to place Soyuz 19 into a circular docking orbit. New orbital parameters were 231.7-km apogee and 192.4-km perigee. The \Jspacecraft\j was spin-stabilized at 3 per sec with all systems operating normally.
Apollo Launch: At 3:50 pm EDT July 15, 1975, 7 hr, 30 min, after the Soyuz launch-a Saturn IB flawlessly lifted the Apollo \Jspacecraft\j from Kennedy Space Center's launch complex 39, carrying Apollo commander Thomas P. Stafford, command-module pilot Vance D. Brand, and docking-module pilot Donald K. Slayton. The \Jspacecraft\j entered orbit with a 173.3-km apogee, 154.7-km perigee, 87.6-min period, and 51.8 inclination.
The \Jspacecraft\j's launch-vehicle adapter was jettisoned at 9 hr 4 min ground elapsed time (9:04 GET, counted from the Soyuz 19 launch) and the crew maneuvered the Apollo 180 to dock with the adapter and extract the docking module. These events were videotaped and transmitted to earth later via ATS 6 (NASA's Applications Technology Satellite launched 30 May 1974). A maneuver 2 hr later at 7:35 pm circularized the orbit at 172 km. The Saturn S-IVB stage was deorbited into the Pacific Ocean 1 hr 30 min later.
A second Soyuz 19 circularization burn of 18.5 sec at 8:43 am EDT July 16 placed that \Jspacecraft\j in a circular orbit of 229 km, with all systems functioning normally.
Rendezvous and Docking: A series of Apollo maneuvers, with the final braking maneuver at 8:51 am EDT July 17, put the Apollo \Jspacecraft\j in a 229.4-km circular orbit matching the orbit of Soyuz 19. A few minutes later Brand reported, "We've got Soyuz in the sextant." Voice contact was made soon after. Hello. Soyuz, Apollo," Stafford said in Russian. Kubasov replied in English, "Hello everybody. Hi to you, Tom and Deke. Hello there, Vance."
All communications among the five crew members during the mission were made in the language of the listener, with the Americans speaking Russian to the Soviet crew and the Soviet crew speaking English to the Americans. Contact of the two \Jspacecraft\j, 51 hr, 49 min, into the mission (12:09 pm July 17) was transmitted live on TV to the earth, and Stafford commented, "We have succeeded. Everything is excellent." "Soyuz and Apollo are shaking hands now," the cosmonauts answered. Hard docking was completed over the Atlantic Ocean at 12:12 pm, 6 min earlier than the prelaunch flight plan watched by millions of TV viewers worldwide.
"Perfect. Beautiful. Well done, Tom. It was a good show. We're looking forward to shaking hands with you in board [sic] Soyuz," Leonov said. Tass later reported that Kubasov told Moscow ground controllers that "we felt a slight jolt at the moment of docking" but that all went according to plan.
Joint Activities: At 3:17 pm hatch 3 opened; Apollo commander Stafford and Soyuz commander Leonov shook hands 2 min later. "Glad to see you," Stafford told Leonov in Russian. "Glad to see you. Very, very happy to see you," Leonov responded in English. "This is Soyuz and the United States," Slayton told TV viewers around the world. Both Soviet Communist Party General Secretary Leonid I. Brezhnev and President Ford congratulated the crews and expressed their confidence in the success of the mission.
Stafford then presented Leonov with "five flags for your government and the people of the Soviet Union" with the wish that "our joint work in space serves for the benefit of all countries and peoples on the earth." Leonov presented the U.S. crew with Soviet flags and plaques. The men signed international certificates and exchanged other commemorative items. After nearly 4 hrs of joint activities, including a meal aboard the Soyuz, the Americans returned to the Apollo and the hatch was closed at 6:51 pm.
An integrity check of the hatches indicated an atmospheric leak on the Soviet side. Ground controllers later attributed the indication to temperature changes in the sealed docking module that were detected by the sensitive Soviet instrumentation. Future integrity checks of the hatches would be more rigorous, however.
Following a sleep period, the crews prepared for another day of joint activity. Kubasov described the mission to Soviet TV viewers while the rest of the crews performed experiments in their respective \Jspacecraft\j. At 5:05 am July 18, 1975, Brand entered the Soviet \Jspacecraft\j; Leonov joined Stafford and Slayton in Apollo, greeting them with "Howdy partner."
Kubasov gave American TV viewers a tour of his Soyuz, and Stafford followed with a tour of the Apollo. Then both Kubasov and Brand videotaped scientific demonstrations for transmission to earth later. Kubasov and Brand ate lunch in the Soyuz while Leonov ate with Stafford and Slayton in Apollo.
During a third transfer, Stafford and Leonov went into the Soyuz and Kubasov and Brand joined Slayton in Apollo. Brand gave Soviet viewers a Russian-language tour of the eastern U.S. as seen from space. Further speeches and exchanges of commemorative items were made for both U.S. and Soviet viewers before the final handshakes at 4:49 pm EDT July 18, when the crews returned to their respective \Jspacecraft\j.
The hatches were closed after Brand told Leonov and Kubasov, "We wish you the host of success. I'm sure that we've opened up a new era in history. Our next meeting will be on the ground." Total time for all transfers and joint activities was 19 hr 55 min. Stafford had spent 7 hr 10 min aboard Soyuz; Brand, 6 hr 30 min; Slayton, 1 hr 35 min. Leonov spent 5 hr 43 min in the Apollo, Kubasov 4 hr 57 min. During nearly 2 days of joint activities, the five men carried out five joint experiments.
Undocking and Separation: The Apollo and Soyuz \Jspacecraft\j undocked at 95:42 GET (8:02 am EDT July 19, 1975). While the \Jspacecraft\j were in station-keeping mode, the crews photographed them and the docking apparatus, transmitting the pictures live on TV to earth.
The Apollo \Jspacecraft\j then served as an occulting disk, blocking the sun from the Soyuz and simulating a solar eclipse the first man-made eclipse. Leonov and Kubasov photographed the solar corona as the Apollo backed away from the Soyuz and toward the sun.
The two \Jspacecraft\j then redocked at 8:34 am EDT with the Apollo maneuvering and the Soyuz docking system active while good quality TV was transmitted to earth. The second docking was not as smooth as the first because a slight misalignment of the two \Jspacecraft\j caused both to pitch excessively at contact.
Final undocking also with the Soyuz active went smoothly and was completed at 11:26 am. As the \Jspacecraft\j separated, the two crews performed the ultraviolet atmospheric absorption experiment, making unsuccessful data measurements at 150 m and then moving to a distance of 500 and 1,000 m, where data were successfully collected.
The Apollo maneuvered to within 50 m of Soyuz and took intensive still photography of the Soyuz. Separation maneuvers to put the two \Jspacecraft\j on separate trajectories began at 2:42 pm with a reaction-control system burn. With the maneuvers completed, Leonov told the Apollo crew, "Thank you very much for your very big job....It was a very good show." Brand answered, "Thank you, also. This was a very good job."
Soyuz Orbit and Landing: Soyuz 19 remained in orbit nearly 30 hrs after the undocking. The cosmonauts conducted biological experiments with microorganisms and zone-forming fungi. At 2:39 am EDT 21 July the Soyuz crew closed hatch 5 between their orbital vehicle and descent module and began depressurizing the orbital module.
Braking burns of the descent engines began at 6:06 am when the \Jspacecraft\j was 772 km from the Apollo. The 194.9-sec burn slowed the \Jspacecraft\j to 120 km per sec. After another burn to stabilize the \Jspacecraft\j the orbital and descent modules separated over Central Africa.
While Soviet viewers watched the first landing of a Soviet \Jspacecraft\j televised in real time, the main parachute deployed at 7 km and jettisoned before the soft-landing engines fired. Soyuz 19 landed about 11 km from the target point northeast of Baykonur Cosmodrome at 6:51 am EDT July 21, after a 142-hr 31-min mission.
The rescue \Jhelicopter\j approached the capsule immediately and specialists opened hatch 5. Kubasov stepped out waving to rescue-team members, followed by Leonov, both cosmonauts in apparent good health and spirits. The cosmonauts returned to Baykonur for medical checks and debriefings.
Apollo Postdocking Orbital Activities: Apollo remained in orbit while its crew continued U.S science experiments begun during predocking. Searching for extreme ultraviolet radiation, the ASTP crew marked the birth of a new branch of \Jastronomy\j when they found, for the first time, extreme ultraviolet sources outside the solar system; some scientists had believed that such sources could never be found.
One of the newly discovered sources turned out to be the hottest known white dwarf star. The Apollo detector also revealed the existence of the first pulsar discovered outside the Milky Way. About 200 000 light years from earth's galaxy, in the Small Magellanic Cloud, it was the most luminous pulsar known to astronomers, 10 times brighter than any discovered so far. After repairing some malfunctioning equipment, the astronauts also mapped x-ray sources throughout the Milky Way.
The crew completed nearly all the 110 earth-observation tasks assigned. Coordinated investigations had been made simultaneously by six groups of scientists on the ground, on ships at sea and in \Jaircraft\j. The astronauts looked at ocean currents, ocean \Jpollution\j, desert \Jgeography\j, shoreline erosion, volcanoes, \Jiceberg\j movements, and vegetation patterns.
On 23 July the command-module tunnel was vented and the crew put on spacesuits to jettison the docking module. The command and service module unlocked from the DM at 3:43 pm EDT, and a 1-sec engine firing put the CSM into a higher orbit (232.2-km apogee, 219.0 km perigee) so that the DM could move ahead. A second maneuver put the CSM in a 223.2-km by 219.0-km orbit. Deorbit began at 4:38 p.m.
The command module and service module separated, the drogue and main parachutes deployed normally, and the Apollo splashed down at 224:58 GET (5:18 p.m. EDT July 24) in the Pacific Ocean 163W and 22N, 500 km west of Hawaii. This was the last ocean landing planned for U.S. human space flights; future flights on the Space Shuttle would be wheeled touchdowns at land bases.
The CM landed in "stable 2" position (upside down 7.4) km from the prime recovery ship, U.S.S. New Orleans. After swimmers from the rescue \Jhelicopter\j righted the \Jspacecraft\j and attached a flotation collar, the Apollo was lifted by crane on to the deck of the recovery ship and Stafford, Brand, and Slayton stepped out to the cheers of the ship's crew. President Ford telephoned congratulations.
During the welcome, the crew was evidently experiencing eye and lung discomfort; subsequent conversations and \Jspacecraft\j data revealed that, during reentry, the earth landing system had failed to jettison the apex cover and drogues as scheduled and had to be fired manually, without first disabling the reaction-control system thrusters.
With the CM oscillating, the thrusters began firing rapidly to compensate, and combustion products including a small amount of \Jnitrogen\j tetroxide entered through the cabin-pressure relief valves. As soon as the RCS system had been disabled, fresh air was once again drawn into the cabin. The crew members told flight officials that they had put on oxygen masks once the \Jspacecraft\j had landed, and then activated the postlanding vent system.
Because of the crew's discomfort, further shipboard ceremonies had been canceled and the crew had been sent to sick bay and then to Tripler \JHospital\j in Hawaii for observation until August 8, 1975.
Primary ASTP mission objectives were to evaluate the docking and undocking of an Apollo \Jspacecraft\j with a Soyuz, and determine the adequacy of the onboard orientation lights and docking target; evaluate the ability of astronauts and cosmonauts to make inter-vehicular crew transfers and the ability of \Jspacecraft\j systems to support the transfers: evaluate the Apollo's capability of maintaining attitude-hold control of the docked vehicles and performing attitude maneuvers; measure quantitatively the effect of weightlessness on the crews' height and lower limb volume, according to length of exposure to zero-g; and obtain relay and direct synchronous-satellite navigation tracking data to determine their accuracy for application to Space Shuttle navigation-system design. The objectives were successfully completed, and the mission was adjudged successful on August 15, 1975.
#
"Apollo-Soyuz",78,0,0,0
\BApollo 18 and Soyuz 19 Launched:\b July 15, 1975\b
\BMeeting in Space:\b July 17, 1975\b
\BSoyuz 19 Landed:\b July 21, 1975\b
\BApollo 18 Splashed Down:\b July 24, 1975\b
\BDuration:\b
ò Apollo 18: 217 hours, 30 minutes
ò Soyuz 19: 143 hours, 31 minutes
\BOrbits:\b (Apollo 18) 136; (Soyuz 19) 96\b
\BAstronaut Crew:\b
ò Thomas P. Stafford
ò Vance D. Brand
ò Donald K. "Deke" Slayton
\BCosmonaut Crew:\b
ò Alexei Leonov
ò Valeri Kubasov
ò The Flight of Apollo-Soyuz from the NASA History Office.
This, the final flight of the Apollo \Jspacecraft\j, was the first docking of \Jspacecraft\j built by different nations and presaged the era of cooperation between the Russians and the Americans that is now such an essential part of our efforts to build a permanently occupied space station.
The American crew included three-flight veteran Thomas P. Stafford, rookie Vance Brand, and the last of the original seven Mercury astronauts to make it into orbit, Donald K. "Deke" Slayton, whose heart murmur had previously kept him grounded. The Soviet crew included the first space walker, Alexei Leonov, and rookie Valeri Kubasov.
While this mission is generally remembered as a political/public relations venture, it resulted in some major technological advancements necessitated by the requirement to dock the two extremely variant \Jspacecraft\j, neither of which had been built for the purpose, together.
The two \Jspacecraft\j were launched within seven and a half hours of one another, and, three hours after they docked two days later, the Astronauts and Cosmonauts met in the middle and shook hands in orbit, exchanged flags and gifts (including the seeds of trees that were later planted in each others' countries) and conversed haltingly with one another in each other's native tongues.
It would be six long years before another American \Jastronaut\j would fly in space, this time aboard the reusable Space Shuttle. The Apollo era, an era of the greatest achievements in mankind's history, had ended.
#
"NASA Mercury Project",79,0,0,0
\JNASA Mercury Project Summary\j
\JFlights of Project Mercury\j
\JMercury-Redstone 4: Liberty Bell 7\j
\JMercury-Atlas 6: Friendship 7\j
\JMercury-Atlas 8: Sigma 7\j
\JMercury-Atlas 9: Faith 7\j
#
"NASA Mercury Project Summary",80,0,0,0
At the time of the announcement of Project Apollo by President Kennedy in May 1961, NASA was still consumed with the task of placing an American in orbit through Project Mercury. Stubborn problems arose, however, at seemingly every turn. The first space flight of an \Jastronaut\j, made by Alan B. Shepard, had been postponed for weeks so NASA engineers could resolve numerous details and only took place on 5 May 1961, less than three weeks before the Apollo announcement.
The second flight, a suborbital mission like Shepard's, launched on 21 July 1961, also had problems. The hatch blew off prematurely from the Mercury capsule, Liberty Bell 7, and it sank into the Atlantic Ocean before it could be recovered. In the process the \Jastronaut\j, "Gus" Grissom, nearly drowned before being hoisted to safety in a \Jhelicopter\j. These suborbital flights, however, proved valuable for NASA technicians who found ways to solve or work around literally thousands of obstacles to successful space flight.
As these issues were being resolved, NASA engineers began final preparations for the orbital aspects of Project Mercury. In this phase NASA planned to use a Mercury capsule capable of supporting a human in space for not just minutes, but eventually for as much as three days. As a launch vehicle for this Mercury capsule, NASA used the more powerful Atlas instead of the Redstone. But this decision was not without controversy. There were technical difficulties to be overcome in mating it to the Mercury capsule to be sure, but the biggest complication was a debate among NASA engineers over its propriety for human spaceflight.
When first conceived in the 1950s many believed Atlas was a high-risk proposition because to reduce its weight Convair Corp. engineers under the direction of Karel J. Bossart, a pre-World War II immigrant from \JBelgium\j, designed the booster with a very thin, internally pressurized fuselage instead of massive struts and a thick metal skin. The "steel balloon," as it was sometimes called, employed \Jengineering\j techniques that ran counter to a conservative \Jengineering\j approach used by Wernher von Braun for the \JV-2\j and the Redstone at \JHuntsville\j, \JAlabama\j.
Von Braun, according to Bossart, needlessly designed his boosters like "bridges," to withstand any possible shock. For his part, von Braun thought the Atlas too flimsy to hold up during launch. He considered Bossart's approach much too dangerous for human spaceflight, remarking that the \Jastronaut\j using the "contraption," as he called the Atlas booster, "should be getting a medal just for sitting on top of it before he takes off!" The reservations began to melt away, however, when Bossart's team pressurized one of the boosters and dared one of von Braun's engineers to knock a hole in it with a sledge hammer. The blow left the booster unharmed, but the recoil from the hammer nearly clubbed the engineer.
Most of the differences had been resolved by the first successful orbital flight of an unoccupied Mercury-Atlas combination in September 1961. On 29 November the final test flight took place, this time with the chimpanzee Enos occupying the capsule for a two-orbit ride before being successfully recovered in an ocean landing. Not until 20 February 1962, however, could NASA get ready for an orbital flight with an \Jastronaut\j.
On that date John Glenn became the first American to circle the Earth, making three orbits in his Friendship 7 Mercury \Jspacecraft\j. The flight was not without problems, however; Glenn flew parts of the last two orbits manually because of an autopilot failure, and left his normally jettisoned retrorocket pack attached to his capsule during reentry because of a loose heat shield.
Glenn's flight provided a healthy increase in national pride, making up for at least some of the earlier Soviet successes. The public, more than celebrating the technological success, embraced Glenn as a personification of heroism and dignity. Hundreds of requests for personal appearances by Glenn poured into NASA headquarters, and NASA learned much about the power of the astronauts to sway public opinion.
The NASA leadership made Glenn available to speak at some events, but more often substituted other astronauts and declined many other invitations. Among other engagements, Glenn did address a joint session of Congress and participated in several ticker-tape parades around the country. NASA discovered in the process of this hoopla a powerful public relations tool that it has employed ever since.
Three more successful Mercury flights took place during 1962 and 1963. Scott Carpenter made three orbits on 20 May 1962, and on 3 October 1962 Walter Schirra flew six orbits. The capstone of Project Mercury was the 15-16 May 1963 flight of Gordon Cooper, who circled the Earth 22 times in 34 hours. The program had succeeded in accomplishing its purpose: to successfully orbit a human in space, explore aspects of tracking and control, and to learn about \Jmicrogravity\j and other biomedical issues associated with spaceflight.
#
"Flights of Project Mercury",81,0,0,0
Project Mercury began on October 7, 1958, one year and three days after the launch of \JSputnik\j 1 by the Soviet Union heralded the beginning of the Space Age. The challenges were seemingly insurmountable; to devise a vehicle light enough to be launched into Earth orbit yet sturdy enough to withstand the forces of liftoff and splashdown, while protecting the pilot from the vacuum of space and the intense heat of re-entry into the atmosphere.
To meet these goals, they designed the \Jspacecraft\j as a wingless capsule, complete with an ablative heat shield that would literally be burned off during re-entry. This capsule was to sit atop a liquid-fueled ballistic missile.
The first, for suborbital flights, was the Redstone, designed by Wernher von Braun's \JHuntsville\j team. The second vehicle, designed to actually rocket the capsule into orbit, was the Atlas-D, with steel skin so thin that it would collapse like a plastic bag if it were not kept under constant pressure from within.
Three weeks after Alan Shepard's first manned suborbital flight, on May 5, 1961, President John F. Kennedy announced the goal of landing a man on the moon within the decade. On July 20, 1969, Neil Armstrong became that man when he stepped off Apollo 11's LEM onto the Moon's surface, taking the last "small step" on that Giant Leap begun a decade earlier by the Mercury astronauts.
The six Mercury flights totaled only two days, six hours in space. (Refer to Table)
#
"Mercury-Redstone 4: Liberty Bell 7",82,0,0,0
Astronaut: Virgil I. "Gus" Grissom
Launched: July 21, 1961
Landed: July 21, 1961 (15 Minutes later)
Orbits: 0
This mission was basically a repeat of Freedom 7, with some improvements, including a window and hand controls, added to the capsule design. Also added was an explosive escape hatch, which almost proved disastrous.
While in the ocean off the \JBahamas\j waiting to be picked up, the side hatch blew, causing the capsule to fill with water and sink, almost taking the \JAstronaut\j with it. A soaking wet Gus Grissom was safely rescued, although the Liberty Bell 7 sank to the bottom and was lost.
#
"Mercury-Atlas 6: Friendship 7",83,0,0,0
Astronaut: John H. Glenn, Jr.
Launched: February 20, 1962
Landed: February 20, 1962 (4 Hours, 55 Minutes later)
Orbits: 3
Apogee: 161.75 Miles
John Glenn became the first American to orbit the Earth aboard Friendship 7, circling the Earth three times. He became the first American to see a sunrise and sunset from space, and became the the first space photographer, using a Minolta camera he bought in a \JCocoa\j Beach drugstore to take some holiday snaps through the window.
There were some exciting times on this flight, particularly when a malfunctioning sensor indicated that the heat shield had come loose. During reentry, Glenn ran out of fuel trying to compensate for the capsule's bucking motions, but he nevertheless came to a safe landing about 40 miles from the target, a miss attributed to having overlooked the loss of consumable items when computing the \Jspacecraft\j's weight.
Those were simpler times!
#
"Mercury-Atlas 8: Sigma 7",84,0,0,0
Astronaut: Walter M. Schirra, Jr.
Launched: October 3, 1962
Landed: October 3, 1962 (9 Hours, 13 Minutes later)
Orbits: 6 Apogee: 175.73 Miles
On this flight, Wally Schirra became the first man to conduct a live TV broadcast from space. After Carpenter's overshoot of the landing site, this flight was intended to improve the fine-tuning of the \Jspacecraft\j's controls.
Much of the flight was spent with the craft on autopilot, although Schirra attempted to do some steering by the stars, a task he found difficult. He observed \Jlightning\j bolts in the atmosphere and took photographs with a Hasselblad camera he brought on-board.
This, at a little over 175 miles, was the highest flight of the Mercury program.
#
"Mercury-Atlas 9: Faith 7",85,0,0,0
Astronaut: L. Gordon Cooper, Jr.
Launched: May 15, 1963
Landed: May 16, 1963 (34 Hours, 19 Minutes later)
Orbits: 22.5
By far the longest of the Mercury flights, it was during this flight that the first satellite was released from a \Jspacecraft\j when Gordon Cooper detached a 152.4 mm sphere containing a beacon as part of a test of the \Jastronaut\j's visual tracking capabilities. Cooper also spotted a 44,000-watt \Jxenon\j lamp shining up from the ground, and he claimed to be able to discern individual buildings and smoke rising from chimneys.
He became the first American to sleep in space, and he took the best space pictures up to that time. His mission was such a roaring success that the planned seventh Mercury flight was canceled and we went instead directly into the somewhat more ambitious Gemini program.
#
"NASA Gemini Project",86,0,0,0
\JNASA Gemini Project Summary\j
\JFlights of Project Gemini\j
\JGemini 3: The Unsinkable Molly Brown\j
\JGemini IV\j
\JGemini V\j
\JGemini VI\j
\JGemini VII\j
\JGemini VIII\j
\JGemini IX\j
\JGemini XI\j
\JGemini XII\j
#
"NASA Gemini Project Summary",87,0,0,0
Even as the Mercury program was underway, and work took place developing Apollo hardware, NASA program managers perceived a huge gap in the capability for human spaceflight between that acquired with Mercury and what would be required for a Lunar landing. They closed most of the gap by experimenting and training on the ground, but some issues required experience in space.
Three major areas immediately arose where this was the case. The first was the ability in space to locate, maneuver toward, and rendezvous and dock with another \Jspacecraft\j. The second was closely related, the ability of astronauts to work outside a \Jspacecraft\j. The third involved the collection of more sophisticated physiological data about the human response to extended spaceflight.
To gain experience in these areas before Apollo could be readied for flight, NASA devised Project Gemini. Hatched in the fall of 1961 by engineers at Robert Gilruth's Space Task Group in cooperation with McDonnell \JAircraft\j Corp. technicians, builders of the Mercury \Jspacecraft\j, Gemini started as a larger Mercury Mark II capsule but soon became a totally different proposition. It could accommodate two astronauts for extended flights of more than two weeks. It pioneered the use of fuel cells instead of batteries to power the ship, and incorporated a series of modifications to hardware.
Its designers also toyed with the possibility of using a paraglider being developed at Langley Research Center for "dry" landings instead of a "splashdown" in water and recovery by the Navy. The whole system was to be powered by the newly developed Titan II launch vehicle, another ballistic missile developed for the Air Force. A central reason for this program was to perfect techniques for rendezvous and docking, so NASA appropriated from the military some Agena rocket upper stages and fitted them with docking adapters.
Problems with the Gemini program abounded from the start. The Titan II had longitudinal oscillations, called the "pogo" effect because it resembled the behavior of a child on a pogo stick. Overcoming this problem required \Jengineering\j imagination and long hours of overtime to stabilize fuel flow and maintain vehicle control. The fuel cells leaked and had to be redesigned, and the Agena reconfiguration also suffered costly delays.
NASA engineers never did get the paraglider to work properly and eventually dropped it from the program in favor of a parachute system the one used for Mercury. All of these difficulties shot an estimated $350 million program to over $1 billion. The overruns were successfully justified by the space agency, however, as necessities to meet the Apollo landing commitment.
By the end of 1963 most of the difficulties with Gemini had been resolved, albeit at great expense, and the program was ready for flight. Following two unoccupied orbital test flights, the first operational mission took place on 23 March 1965. Mercury \Jastronaut\j Grissom commanded the mission, with John W. Young, a Naval aviator chosen as an \Jastronaut\j in 1962, accompanying him.
The next mission, flown in June 1965 stayed aloft for four days and \Jastronaut\j Edward H. White II performed the first extra-vehicular activity (EVA) or spacewalk. Eight more missions followed through November 1966. Despite problems great and small encountered on virtually all of them, the program achieved its goals.
Additionally, as a technological learning program, Gemini had been a success, with 52 different experiments performed on the ten missions. The bank of data acquired from Gemini helped to bridge the gap between Mercury and what would be required to complete Apollo within the time constraints directed by the president.
#
"Flights of Project Gemini",88,0,0,0
Somewhat like the current Space Shuttle is a transition between our earliest space exploration and actually living and working in space, Project Gemini was a transitional project between the initial, pioneering Mercury Program and the actual space travel accomplished by the Apollo Program. And also akin to the Shuttle, its success was absolutely critical to achieving our goal of reaching the Moon.
The primary purpose of the Gemini missions was to learn how to "fly" a space vehicle, to maneuver in orbit, to rendezvous and dock with another vehicle, which were essential to the later Apollo missions. Project Gemini also demonstrated that astronauts could endure conditions of weightlessness for the length of time necessary for a lunar mission.
There were ten Gemini missions spanning a period of 20 months. It was during this period that Mission Control was transferred to the Johnson Space Center in Houston. Sixteen new astronauts joined the original seven, and space flight began to become routine.(Refer to Table)
The capsule was larger, almost twice as heavy as the Mercury capsule, but was a much tighter fit for the astronauts, having an interior space only 50% larger than Mercury's. There were many technological improvements, including fuel cells in lieu of batteries, complex maneuvering jets, on-board computers, a modular construction that permitted easy replacement of malfunctioning parts, and ejection seats to replace Mercury's escape rockets.
The launch vehicle, the Titan 2, was far more powerful than the old Atlas-D that had launched Mercury into orbit, and it also served to launch the "Agena" upper stage that contained the docking mechanisms used to train the astronauts for Apollo.
By the time of Gemini's final flight, Lunar Orbiter 2 was already mapping out Apollo landing sites. The Gemini Program provided another rather large step toward our Giant Leap to the moon.
#
"Gemini 3: The Unsinkable Molly Brown",89,0,0,0
Launched: March 23, 1965
Splashed Down: March 23, 1965
Orbits: 3
Duration: 4 hours, 53 minutes
Crew:
ò Virgil I. "Gus" Grissom
ò John W. Young
In a joking reference to the sinking of Liberty Bell 7 on the second suborbital Mercury mission, Gemini 3 became the only one of the Gemini missions to get a "nickname." Subsequent Gemini missions only received numbers, and were numbered with Roman numerals.
This was primarily a testing shakedown for the new, maneuverable Gemini capsule. During the flight, the astronauts used the thrusters to change the shape of their orbit and drop to a lower altitude. Because the aerodynamic behavior of the craft did not match wind tunnel predictions, the splashdown missed the target point by over 80 kilometers, and the shift of the capsule to landing position was so abrupt that Gus Grissom broke the faceplate of his helmet.
#
"Gemini IV",90,0,0,0
Launched: June 3, 1965
Splashed Down: June 7, 1965
Orbits: 62
Duration: 97 hours, 56 minutes
Crew:
ò James A. McDivitt
ò Edward H. White II
This mission was a real learning experience. The plan was to fly in formation with the spent Titan 2 second stage, but the dynamics of maneuvering toward another object in orbit proved to be far different from what flight engineers expected. When the astronauts tried to fly toward the target, the craft got farther and farther away. They had to give up the effort after burning half their fuel.
They discovered that, to catch up with an object ahead of you, you must drop down, and then rise back up after you catch up, rather than speed up, because speeding up puts you into a higher, and therefore slower, orbit.
The highlight of the mission was Ed White's 22-minute space walk, during which he used a hand-held "zip gun" to maneuver at the end of a tether. Gemini IV also marked the first flight controlled out of the new Mission Control Facility in Houston.
#
"Gemini V",91,0,0,0
Launched: August 21, 1965
Splashed Down: August 29, 1965
Orbits: 120
Duration: 190 hours, 55 minutes
Crew:
ò L. Gordon Cooper, Jr.
ò Charles "Pete" Conrad. Jr.
Some mark the Gemini V mission as the point at which America finally took the lead in the Space Race with the Soviets. By far the longest mission to date (surpassing the record of 119 hours, 6 minutes and 81 orbits set by Valery A. Bykovsky on Vostok 5 back in June, 1963), Gemini V, its duration was made possible by the use of new fuel cells for power. Of course, as with so many NEW! IMPROVED!! systems, the fuel cells proved a little bit troublesome at first, and their malfunctions prevented a rendezvous with a pod released from the craft and forced the cancellation of several other experiments, which gave the astronauts so much "free time" in orbit that Conrad complained about the fact that he had not brought along a book.
Medical tests were very positive, indicating that long-duration spaceflight is indeed very feasible, and this was essential in paving the way for the later Apollo lunar missions.
#
"Gemini VI",92,0,0,0
Launched: December 15, 1965
Splashed Down: December 16, 1965
Orbits: 16
Duration: 25 hours, 51 minutes
Crew:
ò Walter M. Schirra, Jr.
ò Thomas P. Stafford
This mission was originally intended to dock with an unoccupied Agena vehicle, but failure of an earlier test of the Agena prompted a rather spectacular change in plans; instead of the Agena, Gemini VI would rendezvous instead with Gemini VII.
This mission was originally scheduled to blast off on December 12, but the Titan 2 launch vehicle shut down right at T-minus-0. The astronauts could have ejected, but Schirra made the command decision to stay; he felt no motion and trusted his senses. Three days later, the lift off was successful.
Schirra used information from his on-board computer (as well as a bit of "seat-of-the-pants" flying) to rendezvous with Gemini VII on the afternoon of December 15. The two craft flew in formation and around one another, at one point coming within a foot without actually touching, for five hours. The astronauts returned the first pictures of another human-occupied craft in space.
#
"Gemini VII",93,0,0,0
Launched: December 4, 1965
Splashed Down: December 18, 1965
Orbits: 206
Duration: 330 hours, 35 minutes
Crew:
ò Frank Borman
ò James A. Lovell, Jr.
This mission set a duration record that remained unbroken until Cosmonauts Georgi Dobrovolsky, Vladislav Volkov and Viktor Patsayev stayed in orbit for 360 revolutions aboard Soyuz 11 in June, 1971, long after American astronauts had walked on the surface of the Moon. In the cramped Gemini quarters it proved to be a real endurance test, during which the astronauts tested a new, lighter space suit (which proved quite uncomfortable if worn for long periods).
This crew conducted the most experiments, 20, of any Gemini mission, inculding studies of \Jnutrition\j in space. The highlight of this mission was, of course, the successful rendezvous with Gemini VI, which was launched after Gemini VII had already been in orbit eleven days. Following the rendezvous, the astronauts just marked time for three more days; fortunately, they heeded the advice of Pete Conrad and brought their books along.
#
"Gemini VIII",94,0,0,0
Launched: March 16, 1966
Splashed Down: March 16, 1966
Orbits: 16
Duration: 10 hours, 11 minutes
Crew:
ò Neil A. Armstrong
ò David R. Scott
This mission was the first to actually dock with another craft in space, linking up to an unmanned Agena target vehicle that had been launched earlier. What followed, though, was one of the most frightening events in the history of the space program. The two craft, still linked together, began to roll continuously.
When the astronauts succeeded in undocking from the Agena, the spinning got faster and faster because the problem was a stuck thruster on the Gemini capsule itself. The dizzying spin got up to one revolution per second, and Armstrong and Scott had to use the reentry control thrusters to stabilize the craft.
The mission had to be cut short; a planned spacewalk by Scott had to be cancelled. The still-nauseous astronauts returned only ten hours after launch, but at least they came back alive, and Neil Armstrong survived to take his "giant leap for mankind" aboard Apollo 11.
#
"Gemini IX",95,0,0,0
Launched: June 3, 1966
Splashed Down: June 6, 1966
Orbits: 45
Duration: 72 hours, 21 minutes
Crew:
ò Thomas P. Stafford
ò Eugene A. Cernan
This mission began on a very sad note; the crew originally scheduled for this flight, Elliott See and Charles Bassett, had died in a plane crash four months earlier. Stafford and Cernan were the first backup crew to fly in space.
Gemini IX was supposed to dock with with the new ATDA (Augmented Target Docking Adapter), a shortened version of the Agena, but the docking had to be cancelled when, after the successful rendezvous, it was discovered that the protective shroud failed to disengage from the ATDA properly, resulting in what came to be known as the "Angry Alligator."
Cernan was also supposed to have test-flown the AMU (Astronaut Maneuvering Unit) during a space walk, but so many hassles ensued, including a fogged visor in his space helmet, that the test had to be cancelled. The device was never tested until Skylab, seven years later.
#
"Gemini XI",96,0,0,0
Launched: September 12, 1966
Splashed Down: September 15, 1966
Orbits: 44
Duration: 71 hours, 17 minutes
Crew:
ò Charles "Pete" Conrad
ò Richard F. Gordon, Jr.
Gemini XI was a race to dock quickly with an already orbiting Agena, to closely simulate the conditions that would soon be encountered by LEM astronauts attempting to dock with the Apollo Command Module in orbit around the Moon. Only 85 minutes after launch the rendezvous was successful and GEMINI XI docked with the Agena target vehicle several times successfully.
There was a chance that this mission would be given a go-ahead for an orbit around the moon, but this was scrapped and the crew had to be satisfied with reaching the highest orbit ever by a human spaceflight, 848.7 miles up.
This mission also succeeded in tethering the Gemini capsule and the Agena vehicle together, and a partially successful attempt to rotate the pair resulted in the first "artificial gravity," although there were problems keeping the tether taut.
This mission also featured the first fully automatic, computer controlled landing, and splashed down only 2.8 miles from the recovery ship.
#
"Gemini XII",97,0,0,0
Launched: November 11, 1966
Splashed Down: November 15, 1966
Orbits: 59
Duration: 94 hours, 35 minutes
Crew:
ò James A. Lovell, Jr.
ò Edwin E. "Buzz" Aldrin, Jr.
Gemini XII finally demonstrated what had proven elusive on all previous Gemini flights; that it WAS indeed possible for man to work effectively outside the protected environment of a \Jspacecraft\j in 0G. Prior to this flight, the new technique of underwater training had prepared Aldrin far better for the rigors of weightlessness, and during his two hour, 20 minute space walk he photographed star fields, retrieved a micrometeorite collector and performed various other tasks.
The docking with the Agena was routine; we had that little chore down pat now. Problems with the Agena vehicle's rockets prevented attaining a high orbit, but two additional space walks were successful and the remaining planned Gemini missions were cancelled; the Gemini program had done its job, preparing us well for the challenges of Apollo.
#
"Ranger (1964 - 1965)",98,0,0,0
The Ranger series was the first US attempt to obtain close-up images of the Lunar surface. The Ranger \Jspacecraft\j were designed to fly straight down towards the Moon and send images back until the moment of impact. This image has a resolution of 5 meters. Ranger 7 impacted in mare terrain modified by crater rays. Ranger 8 also impacted in mare terrain, but this area contained a complex system of ridges. Ranger 9 impacted in a large crater in the lunar highlands.
\BRanger 7
Launched 28 July 1964\b
Impacted Moon 31 July 1964 at 13:25:49 UT
Latitude 10.35 S, Longitude 339.42 E - Mare Cognitum (Sea of Clouds)
\BRanger 8
Launched 17 February 1965\b
Impacted Moon 20 February 1965 at 09:57:37 UT
Latitude 2.67 N, Longitude 24.65 E - Mare Tranquillitatis (Sea of Tranquility)
\BRanger 9
Launched 21 March 1965\b
Impacted Moon 24 March 1965 at 14:08:20 UT
Latitude 12.83 S, Longitude 357.63 E - Alphonsus
Each Ranger \Jspacecraft\j had 6 cameras on board. The cameras were fundamentally the same with differences in exposure times, fields of view, lenses, and scan rates. The camera system was divided into two channels, P (partial) and F (full). Each channel was self-contained with separate power supplies, timers, and transmitters. The F-channel had 2 cameras: the wide-angle A-camera and the narrow angle B-camera. The P-channel had four cameras: P1 and P2 (narrow angle) and P3 and P4 (wide angle).
The final F-channel image was taken between 2.5 and 5 sec before impact (altitude about 5 km) and the last P-channel image 0.2 to 0.4 sec before impact (altitude about 600 m). The images provided better resolution than was available from Earth-based views by a factor of 1,000. These highly detailed images showed Apollo planners that finding a smooth landing site was not going to be easy.
#
"Lunar Orbiter (1966 - 1967)",99,0,0,0
Five Lunar Orbiter missions were launched in 1966 through 1967 with the purpose of mapping the lunar surface before the Apollo landings. All five missions were successful, and 99% of the Moon was photographed with a resolution of 60 m or better. The first three missions were dedicated to imaging 20 potential lunar landing sites, selected based on Earth-based observations. These were flown at low inclination orbits. The fourth and fifth missions were devoted to broader scientific objectives and were flown in high altitude polar orbits. Lunar Orbiter 4 photographed the entire nearside and 95% of the farside, and Lunar Orbiter 5 completed the farside coverage and acquired medium (20 m) and high (2 m) resolution images of 36 pre-selected areas.
\BLunar Orbiter 1
Launched 10 August 1966\b
Imaged Moon: 18-29 August 1966
Apollo landing site survey mission
\BLunar Orbiter 2
Launched 06 November 1966\b
Imaged Moon: 18-25 November 1966
Apollo landing site survey mission
\BLunar Orbiter 3
Launched 05 February 1967\b
Imaged Moon: 15-23 February 1967
Apollo landing site survey mission
\BLunar Orbiter 4
Launched 04 May 1967\b
Imaged Moon: 11-26 May 1967
Lunar mapping mission
\BLunar Orbiter 5
Launched 01 August 1967\b
Imaged Moon: 06-18 August 1967
Lunar mapping and hi-res survey mission
The Lunar Orbiters had an ingenious imaging system, which consisted of a dual-lens camera, a film processing unit, a readout scanner, and a film handling apparatus. Both lenses, a 610-mm narrow angle high-resolution (HR) lens and an 80-mm wide-angle medium resolution (MR) lens, placed their frame exposures on a single roll of 70 mm film. The axes of the two cameras were coincident so the area imaged in the HR frames were centered within the MR frame areas. The film was moved during exposure to compensate for the \Jspacecraft\j velocity, which was estimated by an electric-optical sensor. The film was then processed, scanned, and the images transmitted back to Earth.
#
"Surveyor (1966 - 1968)",100,0,0,0
The Surveyor probes were the first US \Jspacecraft\j to land safely on the Moon. The main objectives of the Surveyors were to obtain close-up images of the lunar surface and to determine if the terrain was safe for manned landings. Each Surveyor was equipped with a \Jtelevision\j camera. In addition, Surveyors 3 and 7 each carried a soil mechanics surface sampler scoop which dug trenches and was used for soil mechanics tests and Surveyors 5, 6, and 7 had magnets attached to the footpads and an alpha scattering instrument for chemical analysis of the lunar material.
\BSurveyor 1
Launched 30 May 1966\b
Landed 02 June 1966, 06:17:37 UT
Latitude 2.45 S, Longitude 316.79 E - Flamsteed P
\BSurveyor 2
Launched 20 September 1966\b
Crashed on Moon 22 September 1966
Vernier engine failed to ignite - southeast of Copernicus Crater
\BSurveyor 3
Launched 17 April 1967\b
Landed 20 April 1967, 00:04:53 UT
Latitude 2.94 S, Longitude 336.66 E - \JOceanus\j Procellarum (Ocean of Storms)
\BSurveyor 4
Launched 14 July 1967\b
Radio contact lost 17 July 1967
2.5 minutes from touchdown - Sinus Medii
\BSurveyor 5
Launched 08 September 1967\b
Landed 11 September 1967, 00:46:44 UT
Latitude 1.41 N, Longitude 23.18 E - Mare Tranquillitatus (Sea of Tranquility)
\BSurveyor 6
Launched 07 November 1967\b
Landed 10 November 1967, 01:01:06 UT
Latitude 0.46 N, Longitude 358.63 E - Sinus Medii
\BSurveyor 7
Launched 07 January 1968\b
Landed 10 January 1968, 01:05:36 UT
Latitude 41.01 S, Longitude 348.59 E - Tycho North Rim
#
"Apollo Program (1968 - 1972)",101,0,0,0
The Apollo program was designed to land humans on the Moon and bring them safely back to Earth. Six of the missions (Apollos 11, 12, 14, 15, 16, and 17) achieved this goal. Apollos 7 and 9 were Earth orbiting missions to test the Command and Lunar Modules, and did not return lunar data. Apollos 8 and 10 tested various components while orbiting the Moon, and returned photography of the lunar surface.
Apollo 13 did not land on the Moon due to a malfunction, but also returned photographs. The six missions that landed on the Moon returned a wealth of scientific data and almost 400 kilograms of lunar samples. Experiments included soil mechanics, meteoroids, seismic, heat flow, lunar ranging, magnetic fields, and solar wind experiments.
\BApollo lunar missions
Apollo 8
Launched 21 December 1968\b
Lunar Orbit and Return
Returned to Earth 27 December 1968
\BApollo 10
Launched 18 May 1969\b
Lunar Orbit and Return
Returned to Earth 26 May 1969
\BApollo 11
Launched 16 July 1969\b
Landed on Moon 20 July 1969
Sea of Tranquility
Returned to Earth 24 July 1969
\BApollo 12
Launched 14 November 1969\b
Landed on Moon 19 November 1969
Sea of Storms
Returned to Earth 24 November 1969
\BApollo 13
Launched 11 April 1970\b
Lunar Flyby and Return
Malfunction forced cancellation of lunar landing
Returned to Earth 17 April 1970
\BApollo 14
Launched 31 January 1971\b
Landed on Moon 5 February 1971
Fra Mauro
Returned to Earth 9 February 1971
\BApollo 15
Launched 26 July 1971\b
Landed on Moon 30 July 1971
Hadley Rille
Returned to Earth 7 August 1971
\BApollo 16
Launched 16 April 1972\b
Landed on Moon 20 April 1972
Descartes
Returned to Earth 27 April 1972
\BApollo 17
Launched 07 December 1972\b
Landed on Moon 11 December 1972
Taurus-Littrow
Returned to Earth 19 December 1972
The Apollo mission consisted of a Command Module (CM) and a Lunar Module (LM). The CM and LM would separate after lunar orbit insertion. One crew member would stay in the CM, which would orbit the Moon, while the other two astronauts would take the LM down to the lunar surface. After exploring the surface, setting up experiments, taking pictures, collecting rock samples, etc., the astronauts would return to the CM for the journey back to Earth.
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"Key Documents from the Apollo Space Program",102,0,0,0
This section looks at five key documents from the Apollo Project. They include:
\B\IPresident Kennedy asks the Vice President for Recommendations on the US Space Program
The Vice President Answers
President Kennedy Announces the Apollo Decision
The Kennedy Administration Defends Apollo
Neil Armstrong Sets Foot on the Moon\b\i
\BPresident Kennedy asks the Vice President for Recommendations on the US Space Program\b
The memorandum that follows led directly to the Apollo program. By posing the question:"Is there any . . . space program which promises dramatic results in which we could win?" President Kennedy set in motion a review that concluded that only a crash effort to send Americans to the Moon met the criteria Kennedy had laid out. This memorandum followed a week of discussion within the White House on how best to respond to the challenge to US interests posed by the 12 April 1961 orbital flight of Soviet \JCosmonaut\j Yuri Gagarin.
\BThe White House
April 20, 1961
Memorandum for Vice President\b
In accordance with our conversation I would like for you as Chairman of the Space Council to be in charge of making an overall survey of where we stand in space.
1. Do we have a chance of beating the Soviets by putting a laboratory in space, or by a trip round the moon, or by a rocket to land on the moon, or by a rocket to go to the moon and back with a man. Is there any other space program which promises dramatic results in which we could win?
2. How much additional would it cost?
3. Are we working 24 hours a day on existing programs. If not, why not? If not, will you make recommendations to me as to how work can be speeded up.
4. In building large boosters should we put our emphasis on nuclear, chemical or liquid fuel, or a combination of these three?
5. Are we making maximum effort? Are we achieving necessary results?
I have asked Jim Webb, Dr. Weisner, Secretary McNamara and other responsible officials to cooperate with you fully. I would appreciate a report on this at the earliest possible moment.
[Signed] John F. Kennedy
\BThe Vice President Answers\b
This memorandum, prepared by Edward Welsh, new Executive Secretary of the National Aeronautics and Space Council, and signed by Vice President Johnson, was the first report to President Kennedy on the results of the review he had ordered of the space program on 20 April 1961. It identified a human Lunar landing by 1966 or 1967 as the first dramatic space project in which the United States could beat the Soviet Union. Vice President Johnson identified US "leadership" in the world arena as sufficient justification of this undertaking in space.
\BOffice of the Vice President
April 28, 1961
MEMORANDUM FOR PRESIDENT
Subject: Evaluation of Space Program.\b
Reference is to your April 20 memorandum asking certain questions regarding this country's space program.
A detailed survey has not been completed in this time period. The examination will continue. However, what we have obtained so far from knowledgeable and responsible persons makes this summary reply possible.
Among those who have participated in our deliberations have been the Secretary and Deputy Secretary of Defense; General Schriever (AF); Admiral Hayward (Navy); Dr. von Braun (NASA); the Administrator, Deputy Administrator, and other top officials of NASA; the Special Assistant to the President on Science and Technology; representatives of the Director of the Bureau of the Budget; and three outstanding non-Government citizens of the general public: Mr. George Brown (Brown & Root, Houston, Texas); Mr. Donald Cook (American Electric Power Service, New York, NY); and Mr. Frank Stanton (Columbia Broadcasting System, New York, N.Y.).
The following general conclusions can be reported:
a. Largely due to their concentrated efforts and their earlier emphasis upon the development of large rocket engines, the Soviets are ahead of the United States in world prestige attained through impressive technological accomplishments in space.
b. The US has greater resources than the USSR for attaining space leadership but has failed to make the necessary hard decisions and to marshal those resources to achieve such leadership.
c. This country should be realistic and recognize that other nations, regardless of their appreciation of our idealistic values, will tend to align themselves with the country which they believe will be the world leader -- the winner in the long run. Dramatic accomplishments in space are being increasingly identified as a major indicator of world leadership.
d. The US can, if it will, firm up its objectives and employ its resources with a reasonable chance of attaining world leadership in space during this decade. This will be difficult but can be made probable even recognizing the head start of the Soviets and the likelihood that they will continue to move forward with impressive successes. In certain areas, such as communications, navigation, weather, and mapping, the US can and should exploit its existing advance position.
e. If we do not make the strong effort now, the time will soon be reached when the margin of control over space and over men's minds through space accomplishments will have swung so far on the Russian side that we will not be able to catch up, let alone assume leadership.
f. Even in those areas in which the Soviets already have the capability to be first and are likely to improve upon such capability, the United States should make aggressive efforts as the technological gains as well as the international rewards are essential steps in eventually gaining leadership. The danger of long lags or outright omissions by this country is substantial in view of the possibility of great technological breakthroughs obtained from space exploration.
g. Manned exploration of the moon, for example, is not only an achievement with great propaganda value, but it is essential as an objective whether or not we are first in its accomplishment -- and we may be able to be first. We cannot leapfrog such accomplishments, as they are essential sources of knowledge and experience for even greater successes in space. We cannot expect the Russians to transfer the benefits of their experiences or the advantages of their capabilities to us. We must do these things ourselves.
h. The American public should be given the facts as to how we stand in the space race, told of our determination to lead in that race, and advised of the importance of such leadership to our future.
i. More resources and more effort need to be put into our space program as soon as possible. We should move forward with a bold program, while at the same time taking every practical precaution for the safety of the persons actively participating in space flights.
As for the specific questions posed in your memorandum, the following brief answers develop from the studies made during the past few days. These conclusions are subject to expansion and more detailed examination as our survey continues.
Q.1- Do we have a chance of beating the Soviets by putting a laboratory in space, or by a trip around the moon, or by a rocket to land on the moon, or by a rocket to go to the moon and back with a man. Is there any other space program which promises dramatic results in which we could win?
A.1- The Soviets now have a rocket capability for putting a multi-manned laboratory into space and have already crash-landed a rocket on the moon. They also have the booster capability of making a soft landing on the moon with a payload of instruments, although we do not know how much preparation they have made for such a project. As for a manned trip around the moon or a safe landing and return by a man to the moon, neither the US nor the USSR has such capability at this time, so far as we know. The Russians have had more experience with large boosters and with flights of dogs and man. Hence they might be conceded a time advantage in circumnavigation of the moon and also in a manned trip to the moon. However, with a strong effort, the United States could conceivably be first in those two accomplishments by 1966 or 1967.
There are a number of programs which the United States could pursue immediately and which promise significant world-wide advantage over the Soviets. Among these are communications satellites, and navigation and mapping satellites. These are all areas in which we have already developed some competence. We have such programs and believe that the Soviets do not. Moreover, they are programs which could be made operational and effective within reasonably short periods of time and could, if properly programmed with the interests of other nations, make useful strides toward world leadership.
Q.2- How much additional would it cost?
A.2- To start upon an accelerated program with the aforementioned objectives clearly in mind, NASA has submitted an analysis indicating that about $500 million would be needed for FY 1962 over and above the amount currently requested of the Congress. A program based upon NASA's analysis would, over a ten-year period, average approximately $1 billion a year above the current estimates of the existing NASA program.
While the Department of Defense plans to make a more detailed submission to me within a few days, the Secretary has taken the position that there is a need for a strong effort to develop a large solid-propellant booster and that his Department is interested in undertaking such a project. It was understood that this would be programmed in accord with the existing arrangement for close cooperation with NASA, which Agency is undertaking some research in this field. He estimated they would need to employ approximately $50 million during FY 1962 for this work but that this could be financed through management of funds already requested in the FY 1962 budget. Future defense budgets would include requests for additional funding for this purpose; a preliminary estimate indicates that about $500 million would be needed in total.
Q.3- Are we working 24 hours a day on existing programs? If not, why not? If not, will you make recommendations to me as to how work can be speeded up?
A.3- There is not a 24-hour-a-day work schedule on existing NASA space programs except for selected areas in Project Mercury, the Saturn-C-1 booster, the \JCentaur\j engines and the final launching phases of most flight missions. They advise that their schedules have been geared to the availability of facilities and financial resources, and that hence their overtime and 3-shift arrangements exist only in those activities in which there are particular bottlenecks or which are holding up operations in other parts of the programs. For example, they have a 3-shift 7-day-week operation in certain work at Cape Canaveral; the contractor for Project Mercury has averaged a 54-hour week and employs two or three shifts in some areas; Saturn C-1 at \JHuntsville\j is working around the clock during critical test periods while the remaining work on this project averages a 47-hour week; the \JCentaur\j \Jhydrogen\j engine is on a 3-shift basis in some portions of the contractor's plants.
This work can be speeded up through firm decisions to go ahead faster if accompanied by additional funds needed for the acceleration.
Q.4- In building large boosters should we put our emphasis on nuclear, chemical or liquid fuel, or a combination of these three?
A.4- It was the consensus that liquid, solid and nuclear boosters should all be accelerated. This conclusion is based not only upon the necessity for back-up methods, but also because of the advantages of the different types of boosters for different missions. A program of such emphasis would meet both so-called civilian needs and defense requirements.
Q.5- Are we making maximum effort? Are we achieving necessary results?
A.5- We are neither making maximum effort nor achieving results necessary if this country is to reach a position of leadership.
[signed] Lyndon B. Johnson
\BPresident Kennedy Announces the Apollo Decision\b
John F. Kennedy unveiled the commitment to execute Project Apollo before Congress on 25 May 1961 in a speech on "Urgent National Needs," billed as a second State of the Union message. In the speech he asked for support to accomplish four basic goals in space exploration, only the Lunar landing is usually remembered. In addition, he asked for congressional appropriations for weather satellites, communications satellites, and the Rover nuclear propulsion rocket. Congress agreed to all of them with barely any comment. As seen in this excerpt from the speech, Kennedy couched the space program in the context of the cold war rivalry with the Soviet Union:
. . . Finally, if we are to win the battle that is now going on around the world between freedom and tyranny, the dramatic achievements in space which occurred in recent weeks should have made clear to us all, as did the \JSputnik\j in 1957, the impact of this adventure on the minds of men everywhere, who are attempting to make a determination of which road they should take. Since early in my term, our efforts in space have been under review. With the advice of the Vice President, who is Chairman of the National Space Council, we have examined where we are strong and where we are not, where we may succeed and where we may not. Now it is time to take longer strides -- time for this nation to take a clearly leading role in space achievement, which in many ways may hold the key to our future on earth.
I believe we possess all the resources and talents necessary. But the facts of the matter are that we have never made the national decisions or marshaled the national resources required for such leadership. We have never specified long-range goals on an urgent time schedule, or managed our resources and our time so as to insure their fulfillment.
Recognizing the head start obtained by the Soviets with their large rocket engines, which gives them many months of lead-time, and recognizing the likelihood that they will exploit this lead for some time to come in still more impressive successes, we nevertheless are required to make new efforts on our own. For while we cannot guarantee that we shall one day be first, we can guarantee that any failure to make this effort will make us last. We take an additional risk by making it in full view of the world, but as shown by the feat of \Jastronaut\j Shepard, this very risk enhances our stature when we are successful. But this is not merely a race. Space is open to us now; and our eagerness to share its meaning is not governed by the efforts of others. We go into space because whatever mankind must undertake, free men must fully share.
I therefore ask the Congress, above and beyond the increases I have earlier requested for space activities, to provide the funds which are needed to meet the following national goals:
First, I believe that this nation should commit itself to achieving the goal, before this decade is out, of landing a man on the moon and returning him safely to the earth. No single space project in this period will be more impressive to mankind, or more important for the long-range exploration of space; and none will be so difficult or expensive to accomplish. We propose to accelerate the development of the appropriate lunar space craft. We propose to develop alternate liquid and solid fuel boosters, much larger than any now being developed, until certain which is superior. We propose additional funds for other engine development and for unmanned explorations -- explorations which are particularly important for one purpose which this nation will never overlook; the survival of the man who first makes this daring flight. But in a very real sense, it will not be one man going to the moon -- if we make this judgment affirmatively, it will be an entire nation. For all of us must work to put him there.
Secondly, an additional 23 million dollars, together with 7 million dollars already available, will accelerate development of the Rover nuclear rocket. This gives promise of some day providing a means for even more exciting and ambitious exploration of space, perhaps beyond the moon, perhaps to the very end of the solar system itself.
Third, an additional 50 million dollars will make the most of our present leadership, by accelerating the use of space satellites for world-wide communications.
Fourth, an additional 75 million dollars -- of which 53 million dollars is for the Weather Bureau -- will help give us at the earliest possible time a satellite system for world-wide weather observation.
Let it be clear -- and this is a judgment which the Members of the Congress must finally make -- let it be clear that I am asking the Congress and the country to accept a firm commitment to a new course of action -- a course which will last for many years and carry very heavy costs: 531 million dollars in fiscal '62 -- and estimated seven to nine billion dollars additional over the next five years. If we are to go only half way, or reduce our sights in the face of difficulty, in my judgment it would be better not to go at all.
Now this is a choice which this country must make, and I am confident that under the leadership of the Space Committees of the Congress, and the Appropriating Committees, that you will consider the matter carefully.
It is a most important decision that we make as a nation. But all of you have lived through the last four years and have seen the significance of space and the adventure in space, and no one can predict with certainty what the ultimate meaning will be of mastery of space.
I believe we should go to the moon. But I think every citizen of this country as well as the Members of the Congress should consider the matter carefully in making their judgment, to which we have given attention over many weeks and months, because it is a heavy burden, and there is no sense in agreeing or desiring that the United States take an affirmative position in outer space, unless we are prepared to do the work and bear the burdens to make it successful. If we are not, we should decide today and this year.
This decision demands a major national commitment of scientific and technical manpower, material and facilities, and the possibility of their diversion from other important activities where they are already thinly spread. It means a degree of dedication, organization and discipline which have not always characterized our research and development efforts. It means we cannot afford undue work stoppages, inflated costs of material or talent, wasteful interagency rivalries, or a high turnover of key personnel.
New objectives and new money cannot solve these problems. They could in fact, aggravate them further -- unless every scientist, every engineer, every serviceman, every technician, contractor, and civil servant gives his personal pledge that this nation of freedom, in the exciting adventure of space. . . .
\B\ITo continue please click\b\i \JKey Documents from the Apollo Space Program continued\j
#
"Key Documents from the Apollo Space Program continued",103,0,0,0
\BThe Kennedy Administration Defends Apollo\b
Criticism of the priority assigned to the space program, and particularly Project Apollo, increased in 1963. Congress tried unsuccessfully, for example, to delete $700 million from the NASA appropriation for Apollo. As a result, on 9 April 1963 Kennedy asked Lyndon Johnson as head of the National Aeronautics and Space Council for a careful review of the program. Johnson replied on 13 May with a lengthy report that emphasized the positive results of the space program and noted challenges that it posed. In the end, as these excerpts suggest, Johnson's report reflected the administration's continued commitment to an aggressive Lunar landing program for international prestige, scientific, and cost benefit reasons.
\BII. Benefits to National Economy from NASA Space Programs\b
1. It cannot be questioned that billions of dollars directed into research and development in an orderly and thoughtful manner will have a significant effect upon our national economy. No formula has been found which attributes specific dollar values to each of the areas of anticipated developments, however, the "multiplier" of space research and development will augment our economic strength, our peaceful posture, and our standard of living.
2. Even though specific dollar values cannot be set for these benefits, a mere listing of the fields which will be affected is convincing evidence that the benefits will be substantial. The benefits include:
a. Additional knowledge about the earth and the Sun's influence on the earth, the nature of interplanetary space environment, and the origin of the solar system as well as of life itself.
b. Increased ability and experience in managing major research and development efforts, expansion of capital facilities, encouragement of higher standards of quality production,
c. Accelerated use of liquid oxygen in steelmaking, coatings for temperature control of housing, efficient transfer of chemical energy into electrical energy, and wide-range advances in \Jelectronics\j.
d. Development of effective filters against detergents; increased accuracy (and therefore reduced costs) in measuring hot steel rods; improved medical equipment in human care; stimulation of the use of fiberglass refractory welding tape, high energy metal forming processes; development of new coatings for plywood and furniture; use of frangible tube energy absorption systems that can be adapted to absorbing shocks of failing elevators and emergency \Jaircraft\j landings.
e. Improved communications, improved weather forecasting, improved forest fire detection, and improved navigation.
f. Development of high temperature gas-cooled \Jgraphite\j moderated reactors and liquid metal cooled reactors; development of \Jradioisotope\j power sources for both military and civilian uses; development of instruments for monitoring degrees of radiation; and application of thermoelectric and thermionic conversion of heat to electric energy.
g. Improvements in metals, alloys, and \Jceramics\j.
h. An augmentation of the supply of highly trained technical manpower.
i. Greater strength for the educational system both through direct grants, facilities and scholarships and through setting goals that will encourage young people.
j. An expansion of the base for peaceful cooperation among nations.
k. Military competence. (It is estimated that between $600 and $675 million of NASA's FY 1964 budget would be needed for military space projects and would be budgeted by the Defense Department, if they were not already provided for in the NASA budget.)
\BIII. Problems Resulting from the Space Program\b
1. The introduction of a vital new element into an economy always creates new problems but, otherwise, the nation's space program creates no major complications. The program has, to a lesser magnitude, the same problems which Defense budgets and programs have been creating for several years.
2. Despite claims to the contrary, there is no solid evidence that research and development in industry is suffering significantly from a diversion of technical manpower to the space program. NASA estimates that:
a. The nation's pool of scientists and engineers was 1,400,000 as of January l, 1963.
b. NASA programs employed 42,000 of these scientists and engineers -- only 9,000 directly on NASA payrolls.
c. On this basis, the NASA space program currently draws upon only 3% of the national pool of scientists and engineers.
d. Taking into account anticipated expansion, NASA programs are not expected to absorb more than 7% of our country's total supply of scientists and engineers.
3. The majority of the technical people working for NASA fall in the category of \Jengineering\j. However, NASA's education programs are designed to help the universities train additions to the nation's technical manpower needs.
4. NASA has undertaken to support the annual graduate training of 1,000 Ph.Ds, 1/4 of the estimated overall shortage of 4,000 per year. This program would more than replace those drawn upon by the agency.
5. In overall terms, NASA finds that diversion of manpower and resources is not a major problem arising from the space program. A major problem, however, is the need to minimize waste and inefficiency. To help meet this challenge, turnover of top level Government talent should be reduced and compensation more in line with responsibilities would contribute to this objective.
\BConclusion\b
There is one further point to be borne in mind. The space program is not solely a question of prestige, of advancing scientific knowledge, of economic benefit or of military development, although all of these factors are involved. Basically, a much more fundamental issue is at stake -- whether a dimension that can well dominate history for the next few centuries will be devoted to the social system of freedom or controlled by the social system of \Jcommunism\j.
The United States has made clear that it does not seek to "dominate" space and, in fact, has led the way in securing international cooperation in this field. But we cannot close our eyes as to what would happen if we permitted totalitarian systems to dominate the environment of the earth itself. For this reason our space program has an overriding urgency that cannot be calculated solely in terms of industrial, scientific, or military development. The future of society is at stake.
\BNeil Armstrong Sets Foot on the Moon\b
After eight years of all-out effort, nearly $20 billion expended, and three astronauts' deaths, on 20 July 1969 Apollo 11 landed on the Moon. The two astronauts who set foot on the surface, Neil A. Armstrong and Edwin E. Aldrin, called it in what later astronauts thought of as an understatement, "magnificent desolation." This document contains the radio transmissions of the landing and Armstrong's first venture out onto the Lunar surface. The "CC" in the transcript is Houston Mission Control, CDR is Neil Armstrong, and LMP is Buzz Aldrin.
04 06 45 52....CC....We copy you down, Eagle [the name of the Lunar Module].
04 06 45 57....CDR....Houston, Tranquility Base here.
04 06 45 59....CDR....THE EAGLE HAS LANDED.
04 06 46 04....CC....Roger, Tranquility. We copy you on the ground. You got a bunch of guys about to turn blue. We're breathing again. Thanks a lot.
04 06 46 16....CDR....Thank you.
04 06 46 18....CC....You're looking good here.
04 06 46 23....CDR....Okay. We're going to be busy for a minute. .
04 13 23 38....CDR....[After suiting up and exiting the Lunar Module (LM), Armstrong was ready to descend to the Moon's surface]. I'm at the foot of the ladder. The LM footpads are only depressed in the surface about 1 or 2 inches, although the surface appears to be very, very fine grained, as you get close to it. It's almost like a powder. Down there, it's very fine.
04 13 23 43....CDR....I'm going to step off the LM now.
04 13 24 48....CDR....\BTHAT'S ONE SMALL STEP FOR MAN, ONE GIANT LEAP FOR MANKIND. \b
04 13 24 48....CDR....And the -- the surface is fine and powdery. I can -- I can pick it up loosely with my toe. It does adhere in fine layers like powdered charcoal to the sole and sides of my boots. I only go in a small fraction of an inch, maybe an eighth of an inch, but I can see the footprints of my boots and the treads in the fine, sandy particles.
04 13 25 30....CC....Neil, this is Houston. We're copying.
04 13 25 45....CDR....There seems to be no difficulty in moving around as we suspected. It's even perhaps easier than the simulations at one-sixth g that we performed in the various simulations on the ground. It's actually no trouble to walk around. Okay. The descent engine did not leave a crater of any size. It has about 1 foot clearance on the ground. We're essentially on a very level place here. I can see some evidence of rays emanating from the descent engine, but a very insignificant amount. . . .
#
"Apollo 11, Twenty-Five Years On",104,0,0,0
\BApollo-11
Crew\b
Neil A. Armstrong (2), Commander
Edwin E. Aldrin (2), Jr., Lunar Module Pilot
Michael Collins (2), Command Module Pilot
\BBackup Crew\b
James Lovell (1), Backup Commander
Fred Haise (0), Backup Lunar Module Pilot
William A. Anders (1), Backup Command Module Pilot
\BMilestones\b
01/23/69 -- Hardware ondock at KSC
01/29/69 -- Command and Service Module Mated
04/14/69 -- Rollover of CSM from O&C to VAB
05/20/69 -- Rollout to Pad LC-39A
06/26/69 -- Countdown Demonstration Test
07/16/69 -- Launch
\BPayload\b
CSM-107 (Columbia) and LM-5 (Eagle)
\BMission Objective\b
Perform manned lunar landing and return mission safely. (Achieved.)
\BLaunch\b
July 16, 1969; 09:32:00 am EDT. Launch Complex 39-A Kennedy Space Center, FL. No launch delays.
\BOrbit\b
Altitude: 186km x 183km
Duration: 08 Days, 03 hours, 18 min, 35 seconds
\BLanding\b
July 24, 1969; 12:50 p.m. EDT. Splashdown area 13 deg 19 min North and 169 deg 9 min West; Splashdown at 195:18:35 MET. Crew on board U.S.S Hornet at 01:53 p.m. EDT; \Jspacecraft\j aboard ship at 03:50pm.
\BApollo 11 at Twenty-Five\b
July 1994 marks the twenty-fifth anniversary of the epochal lunar landing of Apollo 11 in the summer of 1969. Although President John F. Kennedy had made a public commitment on 25 May 1961 to land an American on the Moon by the end of the decade, up until this time Apollo had been all promise. Now the realization was about to begin. Kennedy's decision had involved much study and review prior to making it public, and his commitment had captured the American imagination, generating much positive support.
Project Apollo had originated as an effort to deal with an unsatisfactory situation (world perception of the Soviet Union's leadership in space and technology), and it addressed these problems very well. Even though Kennedy's political objectives were essentially achieved with the decision to go to the Moon, Project Apollo took on a life of its own over the years and left an important legacy to both the nation and the proponents of space exploration. Its success was enormously significant, coming at a time when American society was in crisis.
A unique confluence of political necessity, personal commitment and activism, scientific and technological ability, economic prosperity, and public mood made possible the 1961 decision to carry out an aggressive lunar landing program. It then fell to NASA, other organizations of the federal government, and the aerospace community to accomplish the task set out in a few short paragraphs by the president. By the time that the goal was accomplished in 1969, only a few of the key figures associated with the decision were still in leadership positions in the government.
Kennedy fell victim to an assassin's bullet in 1963, and science adviser Jerome B. Wiesner returned to MIT soon afterwards. Lyndon B. Johnson, of course, succeeded Kennedy as president but left office in January 1969 just a few months before the first landing. NASA Administrator James E. Webb resolutely guided NASA through most of the 1960s, but his image was tarnished by, among other things, a 1967 Apollo accident that killed three astronauts. He retired from office in October 1968. Several other early supporters of Apollo in Congress and elsewhere died during the 1960s and never saw the program successfully completed.
The first Apollo mission of public significance was the flight of Apollo 8. On 21 December 1968 it took off atop a Saturn V booster from the Kennedy Space Center. Three astronauts were aboard -- Frank Borman, James A. Lovell, Jr., and William A. Anders -- for an historic mission to orbit the Moon. At first that mission had been planned as a flight to test Apollo hardware in the relatively safe confines of low Earth orbit, but senior engineer George M. Low of the Manned \JSpacecraft\j Center at Houston, \JTexas\j, and Samuel C. Phillips, Apollo Program Manager at NASA headquarters, obtained approval to make it a circumlunar flight. The advantages of this could be important, they believed, both in technical and scientific knowledge gained as well as in a public demonstration of what the US could achieve.
After Apollo 8 made one and a half Earth orbits, its third stage began a burn to put the \Jspacecraft\j on a lunar trajectory. It orbited the Moon on 24-25 December and then fired the boosters for a return flight; it "splashed down" in the Pacific Ocean on 27 December. The public reaction to the Apollo 8 circumlunar mission was enthusiastic. It rekindled the excitement felt in the early 1960s during the first Mercury flights, and set the stage for the Apollo landing missions.
Perhaps most important, the flight was a significant accomplishment because it came at a time when American society was in crisis over Vietnam, race relations, urban problems, and a host of other difficulties. And if only for a few moments the nation united as one to focus on this epochal event. Two Apollo Earth-orbital missions occurred before the climax of the program, but they did little more than confirm that the time had come in mid-1969 for a lunar landing.
That landing came during the flight of Apollo 11, which lifted off on 16 July 1969 and, after confirmation that the hardware was working well, began the three day trip to the Moon. Then, at 4:18 p.m. EST on 20 July 1969 the Lunar Module -- with astronauts Neil A. Armstrong and Edwin E. Aldrin aboard -- landed on the surface of the Moon while Michael Collins orbited overhead in the Apollo Command Module. After checkout, Armstrong set foot on the surface, telling millions who saw and heard him on Earth that it was "one small step for man -- one giant leap for mankind."
Aldrin soon followed him out, and the two plodded around the landing site in the 1/6 lunar gravity, planted an American flag but omitted claiming the land for the US as had been routinely done during European exploration of the Americas, collected soil and rock samples, and set up scientific experiments. The next day they launched back to the Apollo capsule orbiting overhead and began the return trip to Earth, splashing down in the Pacific on 24 July.
The flight of Apollo 11 met with an ecstatic reaction around the globe, as everyone shared in the success of the mission. Ticker tape parades, speaking engagements, public relations events, and a world tour by the astronauts served to create good will both in the US and abroad.
Five more landing missions followed at approximately six month intervals through December 1972, each of them increasing the time spent on the Moon. Three of the latter Apollo missions used a lunar rover vehicle to travel in the vicinity of the landing site, but none of them equaled the excitement of Apollo 11.
The scientific experiments placed on the Moon and the lunar soil samples returned through Project Apollo have provided grist for scientists' investigations of the Solar System ever since. The scientific return was significant, but the Apollo program did not answer conclusively the age-old questions of lunar origins and evolution.
Project Apollo in general, and the flight of Apollo 11 in particular, should be viewed as a watershed in the nation's history. It was an endeavor that demonstrated both the technological and economic virtuosity of the United States and established national preeminence over rival nations -- the primary goal of the program when first envisioned by the Kennedy administration in 1961. It had been an enormous undertaking, costing $25.4 billion (about $95 billion in 1990 dollars) with only the building of the Panama Canal rivaling the Apollo program's size as the largest non-military technological endeavor ever undertaken by the United States and only the Manhattan Project to build the atomic bomb in World War II being comparable in a wartime setting.
There are several important legacies (or conclusions) about Project Apollo that need to be remembered at the twenty-fifth anniversary of the Apollo 11 landing. First, and probably most important, the Apollo program was successful in accomplishing the political goals for which it had been created. Kennedy had been dealing with a Cold War crisis in 1961 brought on by several separate factors -- the Soviet orbiting of Yuri Gagarin and the disastrous Bay of Pigs invasion only two of them -- that Apollo was designed to combat.
At the time of the Apollo 11 landing, Mission Control in Houston flashed the words of President Kennedy announcing the Apollo commitment on its big screen. Those phrases were followed with these: "TASK ACCOMPLISHED, July 1969." No greater understatement could probably have been made. Any assessment of Apollo that does not recognize the accomplishment of landing an American on the Moon and safely returning before the end of the 1960s is incomplete and inaccurate, for that was the primary goal of the undertaking.
Second, Project Apollo was a triumph of management in meeting enormously difficult systems \Jengineering\j and technological \Jintegration\j requirements. James E. Webb, the NASA Administrator at the height of the program between 1961 and 1968, always contended that Apollo was much more a management exercise than anything else, and that the technological challenge, while sophisticated and impressive, was largely within grasp at the time of the 1961 decision. More difficult was ensuring that those technological skills were properly managed and used.
Webb's contention was confirmed in spades by the success of Apollo. NASA leaders had to acquire and organize unprecedented resources to accomplish the task at hand. From both a political and technological perspective, management was critical. For seven years after Kennedy's Apollo decision, through October 1968, James Webb politicked, coaxed, cajoled, and maneuvered for NASA in Washington. In the process he acquired for the agency sufficient resources to meet its Apollo requirements.
More to the point, NASA personnel employed a "program management" concept that centralized authority over design, \Jengineering\j, procurement, testing, construction, manufacturing, spare parts, logistics, training, and operations. The systems management of the program was recognized as critical to Apollo's success in November 1968, when Science magazine, the publication of the American Association for the Advancement of Science, observed:
In terms of numbers of dollars or of men, NASA has not been our largest national undertaking, but in terms of complexity, rate of growth, and technological sophistication it has been unique. . . . It may turn out that [the space program's] most valuable spin-off of all will be human rather than technological: better knowledge of how to plan, coordinate, and monitor the multitudinous and varied activities of the organizations required to accomplish great social undertakings.
Understanding the management of complex structures for the successful completion of a multifarious task was a critical outgrowth of the Apollo effort.
Third, Project Apollo forced the people of the world to view the planet Earth in a new way. Apollo 8 was critical to this fundamental change, for on its outward voyage the crew focused a portable \Jtelevision\j camera on Earth and for the first time humanity saw its home from afar, a tiny, lovely, and fragile "blue marble" hanging in the blackness of space.
When the Apollo 8 \Jspacecraft\j arrived at the Moon on Christmas Eve of 1968 the image of Earth was even more strongly reinforced when the crew sent images of the planet back while reading the first part of the \JBible\j -- "And God created the heavens and the Earth, and the Earth was without form and void" -- before sending holiday greetings to humanity.
Writer Archibald MacLeish summed up the feelings of many people when he wrote at the time of Apollo, that "To see the Earth as it truly is, small and blue and beautiful in that eternal silence where it floats, is to see ourselves as riders on the Earth together, brothers on that bright loveliness in the eternal cold -- brothers who know now that they are truly brothers." The modern environmental movement was galvanized in part by this new perception of the planet and the need to protect it and the life that it supports.
Finally, the Apollo program, while an enormous achievement, left a divided legacy for NASA and the aerospace community. The perceived "golden age" of Apollo created for the agency an expectation that the direction of any major space goal from the president would always bring NASA a broad consensus of support and provide it with the resources and license to dispense them as it saw fit. Something most NASA officials did not understand at the time of the Moon landing in 1969, however, was that Apollo had not been conducted under normal political circumstances and that the exceptional circumstances surrounding Apollo would not be repeated.
The Apollo decision was, therefore, an anomaly in the national decision-making process. The dilemma of the "golden age" of Apollo has been difficult to overcome, but moving beyond the Apollo program to embrace future opportunities has been an important goal of the agency's leadership in the recent past. Exploration of the Solar System and the universe remains as enticing a goal and as important an objective for humanity as it ever has been. Project Apollo was an important early step in that ongoing process of exploration.
#
"Apollo 13 Mission",105,0,0,0
This mission was designed to land on the moon but after an on board explosion the only concern was how to get the astronauts back to earth safely. This aborted moon project held the world spellbound for days and finally resulted in a safe landing back on earth.
The following is a summary of the mission, the problems encountered, and the magnificent effort performed by the ground crew and the astronauts to bring everyone home safely.
\BCrew\b
James A Lovell, Jr; John L Swigert, Jr; Fred W Haise, Jr.
\BMission Objective\b
Apollo 13 was supposed to land in the Fra Mauro area. An explosion onboard forced Apollo 13 to circle the moon without landing.
The Fra Mauro site was reassigned to Apollo 14.
\BLaunch\b
Saturday, April 11, 1970 at 13:13 CST.
At five and a half minutes after liftoff, Swigert, Haise, and Lovell felt a little vibration. Then the center engine of the S-II stage shutdown two minutes early. This caused the remaining four engines to burn 34 seconds longer than planned, and the S-IVB third stage had to burn nine seconds longer to put Apollo 13 in orbit.
Ground tests before launch indicated the possibility of a poorly insulated supercritical \Jhelium\j tank in the LM's descent stage so the flight plan was modified to enter the LM three hours early in order to obtain an onboard readout of \Jhelium\j tank pressure.
\BMission Highlights:\b
Third lunar landing attempt. Mission was aborted after rupture of service module oxygen tank. Classed as "successful failure" because of experience in rescuing crew. Spent upper stage successfully impacted on the Moon.
The first two days the crew ran into a couple of minor surprises, but generally Apollo 13 was looking like the smoothest flight of the program. At 46 hours 43 minutes Joe Kerwin, the CapCom on duty, said, "The \Jspacecraft\j is in real good shape as far as we are concerned. We're bored to tears down here." It was the last time anyone would mention boredom for a long time.
At 55 hours 46 minutes, the crew finished a 49 minute TV broadcast showing how comfortably they lived and worked in weightlessness. Nine minutes later, Oxygen tank No 2 blew up, causing No 1 tank also to fail. The Apollo 13 command modules normal supply of electricity, light, and water was lost, and they were about 200,000 miles from Earth. The message came in the form of a sharp bang and vibration.
The time was 2108 hours on April 13. Next, the warning lights indicated the loss of two of Apollo 13's three fuel cells, which were the space craft's prime source of electricity. With warning lights blinking on, one Oxygen tank appeared to be completely empty, and there were indications that the oxygen in the second tank was rapidly being depleted.
The first thing the crew did, even before discovering the oxygen leak, was to try to close the hatch between the CM and the LM. They reacted spontaneously, like submarine crews, closing the hatches to limit the amount of flooding. First Jack and then Lovell tried to lock the reluctant hatch, but the stubborn lid wouldn't stay shut. Exasperated, and realizing that there wasn't a cabin leak, they strapped the hatch to the CM couch. The pressure in the No 1 oxygen tank continued to drift downward; passing 300 psi, now heading toward 200 psi.
Months later, after the accident investigation was complete, it was determined that, when No. 2 tank blew up, it either ruptured a line on the No. 1 tank, or caused one of the valves to leak. When the pressure reached 200 psi, the crew and ground controllers knew that they would lose all oxygen, which meant that the last fuel cell would also die.
At 1 hour and 29 seconds after the bang, Jack Lousma, then CapCom, said after instructions from Flight Director Glynn Lunney: "It is slowly going to zero, and we are starting to think about the LM lifeboat." Swigert replied, "That's what we have been thinking about too."
Ground controllers in Houston faced a formidable task. Completely new procedures had to be written and tested in the simulator before being passed up to the crew. The navigation problem had to be solved; essentially how, when, and in what attitude to burn the LM descent engine to provide a quick return home.
With only 15 minutes of power left in the CM, CapCom told the crew to make their way into the LM. Fred and Jim Lovell quickly floated through the tunnel, leaving Jack to perform the last chores in the Command Module. The first concern was to determine if there were enough consumables to get home? The LM was built for only a 45 hour lifetime, and it needed to be stretched to 90.
Oxygen wasn't a problem. The full LM descent tank alone would suffice, and in addition, there were two ascent engine oxygen tanks, and two backpacks whose oxygen supply would never be used on the lunar surface. Two emergency bottles on top of those packs had six or seven pounds each in them. (At LM jettison, just before re-entry, 285 pounds of oxygen remained, more than half of what was available after the explosion.)
Power was also a concern. There were 2181 \Jampere\j hours in the LM batteries. Ground controllers carefully worked out a procedure where the CM batteries were charged with LM power. All non-critical systems were turned off and energy consumption was reduced to a fifth of normal, which resulted in having 20 percent of LM electrical power left when \JAquarius\j was jettisoned. There was one electrical close call during the mission. One of the CM batteries vented with such force that it momentarily dropped off the line. Had the battery failed, there would be insufficient power to return the ship to Earth.
Water was the main consumable concern. It was estimated that the crew would run out of water about five hours before Earth reentry, which was calculated at around 151 hours. However, data from Apollo 11 (which had not sent its LM ascent stage crashing into the Moon as in subsequent missions) showed that its mechanisms could survive seven or eight hours in space without water cooling. The crew conserved water. They cut down to six ounces each per day, a fifth of normal intake, and used fruit juices; they ate hot dogs and other wet pack foods when they ate at all.
The crew became dehydrated throughout the flight and set a record that stood up throughout Apollo: Lovell lost fourteen pounds, and the crew lost a total of 315 pounds, nearly 50 percent more than any other crew. Those stringent measures resulted in the crew finishing with 282 pounds of water, about 9 percent of the total.
Removal of carbon dioxide was also a concern. There were enough \Jlithium\j \Jhydroxide\j canisters, which remove carbon dioxide from the \Jspacecraft\j, but the square canisters from the Command Module were not compatible with the round openings in the Lunar Module environmental system. There were four cartridges from the LM, and four from the backpacks, counting backups. However, the LM was designed to support two men for two days and was being asked to care for three men nearly four days. After a day and a half in the LM, a warning light showed that the carbon dioxide had built up to a dangerous level. Mission Control devised a way to attach the CM canisters to the LM system by using plastic bags, cardboard, and tape -- all materials carried on board.
One of the big questions was, "How to get back safely to Earth?". The LM navigation system wasn't designed to help in this situation. Before the explosion, at 30 hours and 40 minutes, Apollo 13 had made the normal midcourse correction, which would take it out of a free return to Earth trajectory and put it on a lunar landing course. Now the task was to get back on a free return course. The ground computed a 35 second burn and fired it 5 hours after the explosion. As they approached the Moon, another burn was computed; this time a long 5 minute burn to speed up the return home. It took place 2 hours after rounding the far side of the Moon.
The Command Module navigational platform alignment was transferred to the LM but verifying alignment was difficult. Ordinarily the alignment procedure uses an on board \Jsextant\j device, called the Alignment Optical \JTelescope\j (AOT), to find a suitable navigation star. Then with the help of the on board computer it verifies the guidance platform's alignment. However, due to the explosion, a swarm of debris from the ruptured service module made it impossible to sight real stars.
An alternate procedure was developed to use the sun as an alignment star. Lovell rotated the \Jspacecraft\j to the attitude Houston had requested and when he looked through the AOT, the Sun was just where it was expected. The alignment with the Sun proved to be less than a half a degree off. The ground and crew then knew they could do the 5 minute PC + 2 burn with assurance, and that would cut the total time of the voyage to about 142 hours.
The trip was marked by discomfort beyond the lack of food and water. Sleep was almost impossible because of the cold. When the electrical systems were turned off, the \Jspacecraft\j lost an important source of heat. The temperature dropped to 38 degrees F and condensation formed on all the walls.
A most remarkable achievement of Mission Control was quickly developing procedures for powering up the CM after its long cold sleep. Flight controllers wrote the documents for this innovation in three days, instead of the usual three months. The Command Module was cold and clammy at the start of power up. The walls, ceiling, floor, wire harnesses, and panels were all covered with droplets of water. It was suspected conditions were the same behind the panels. The chances of short circuits caused apprehension, but thanks to the safe guards built into the command module after the disastrous Apollo 1 fire in January 1967, no arcing took place.
Four hours before landing, the crew shed the service module; Mission Control had insisted on retaining it until then because everyone feared what the cold of space might do to the unsheltered CM heat shield. Photos of the Service Module showed one whole panel missing, and wreckage hanging out, it was a sorry mess as it drifted away. Three hours later the crew left the Lunar Module \JAquarius\j and then splashed down gently in the Pacific Ocean near \JSamoa\j.
\BCause of Explosion\b
The Apollo 13 malfunction was caused by an explosion and rupture of oxygen tank No. 2 in the service module. The explosion ruptured a line or damaged a valve in the No. 1 oxygen tank, causing it to lose oxygen rapidly. The service module bay No. 4 cover was blown off. All oxygen stores were lost within about 3 hours, along with loss of water, electrical power, and use of the propulsion system.
The No. 2 oxygen tank used in Apollo had originally been installed in Apollo 10. It was removed from Apollo 10 for modification and during the extraction was dropped 2 inches, slightly jarring an internal fill line. The tank was replaced with another for Apollo 10, and the exterior inspected. The internal fill line was not known to be damaged, and this tank was later installed in Apollo 13.
The oxygen tanks had originally been designed to run off the 28 volt DC power of the command and service modules. However, the tanks were redesigned to also run off the 65 volt DC ground power at Kennedy Space Center. All components were upgraded to accept 65 volts except the heater thermostatic switches, which were overlooked. These switches were designed to open and turn off the heater when the tank temperature reached 80 degrees F (Normal temperatures in the tank were -300 to -100 F).
During preflight testing, tank No. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65 volt DC power. This may have damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches may have welded shut, allowing the temperature within the tank to rise to over 1,000 degrees F. The high temperature would have resulted in damage to the Teflon \Jinsulation\j on the electrical wires to the power fans within the tank.
56 hours into the mission, at about 03:06 UT on 14 April 1970 (10:06 PM, April 13 EST), the power fans were turned on within the tank. The exposed fan wires shorted and the Teflon \Jinsulation\j caught fire. This fire spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the nominal 1,000 psi pressure within the tank, causing the No. 2 oxygen tank to explode. This damaged the No. 1 tank and parts of the interior of the service module and blew off the bay No. 4 cover.
#
"Remarks by the President at the 25th Anniversary of Apollo 11",106,0,0,0
\BTHE PRESIDENT:\b Thank you very much, Mr. Vice President. Members of Congress. Veterans of the Apollo program. The friends of the space program in America and, most of all, to those whom we honor here today.
Just a day before he died, President Kennedy compared our space program to a boy who comes upon a wall in an orchard. The wall is tall, it looks insurmountable, but the boy is curious about what lies on the other side. So, he throws his cap over the wall and then he has no choice but to go after it.
Twenty-five years ago today, our nation, represented by these three brave men, made that climb. And, so, today we are gathered to celebrate their voyage and I honestly hope to recommit ourselves to their spirit of discovery. Apollo 11, Neil Armstrong Buzz Aldrin and Michael Collins were our guides for the wondrous, the unimaginable at that time, the true handiwork of God. They realized the dreams of a nation. They fulfilled an American destiny. They taught us that nothing is impossible if we set our sights high enough.
Today, we're honored to have them and all the other Apollo astronauts who are here with us. For every American who followed your journey especially for those of us who were young on that fateful day 25 years ago, and for the young Americans who still dream dreams of a future in space, we thank you all.
Looking back on that mission, one thing is clear that we ought to remember today. It wasn't easy. The ship to the heavens measured just 13 feet in diameter. The destination was three days and a world away. On the third day as the tiny module descended to the Moon, it came dangerously close to a crash landing -- that happens around here all the time -- (laughter) -- but Neil Armstrong took over the controls from the computer and landed safely. Man had not been rendered obsolete by the mechanical and that hasn't happened yet.
Not long after that when he stepped on the Moon, Mr. Armstrong marked the outer limit of the human experiment with those simple words, "One small step for man. One giant leap for mankind."
These men and the other astronauts who came before and after have helped us to step into another world right here on Earth. They've shown us that we can harness the technology of space in areas from the economy to the environment to education to information and technology. The products and knowledge that grew out of our space missions has changed our way of life forever and for the better. And in our quest we have relearned a sense of confidence that has always been an essential ingredient of our American Dream.
Today that journey continues. Our commitment to the space program is strong and unwavering. The best way to honor these men and all the others who have helped it so much, is to continue that quest. Many have risked their lives and some have given their lives so that we could go forward.
Today I ask that we remember, especially, the crews of Apollo 1 and the Challenger. On this day of celebration we must never forget the deep debt we owe to those brave Americans. And our thoughts should also be with their families and their loved ones for the sacrifice they have given helped to bring us all to new horizons.
Our space explorations today are important models for cooperation in the new post-Cold War world. The Vice President described that eloquently a moment ago. Sergei's mission was an important first step toward full Russian partnership in what must be our next great mission, the international Space Station.
This permanent orbiting space laboratory, to be built with help from 14 nations, will hasten discoveries in fields from the environment to medicine, to computers. We should also remember that the space station holds great promise for us here at home, as it strengthens our largest export sector, aerospace technology.
All these reasons explain why the House has fully funded already the Space Station. I want to thank many people who are responsible for that bipartisan victory but let me mention especially George Brown, Lou Stokes, Bob Walker and Jerry Lewis. I know the Vice President and Dan Goldin and a lot of other people burned up the phone lines before the House vote.
Let me say that we've fought a lot of battles for the future around here in the last 18 months, and sometimes it seems that the most important ones are decided by the narrowest of margins. The economic plan passed by a vote, the assault weapons ban passed by two votes. Last year the Space Station survived by the vote of a single member of the House of Representatives who changed his mind on the way down the aisle. But this year, thanks to the common endeavors of all of us and thanks to the promise of cooperation with \JRussia\j and with other nations, the House of Representatives voted to fund the Space Station by 122 votes, a bipartisan commitment to America's future. (Applause.)
I thank the members of the Senate who are here today who are pushing for passage. I know they won't miss this great opportunity which is coming on them very soon. I thank you Senator Mikulski and all the other members of the Senate who are here for the work that will be done in the Senate.
As we work toward building a better world, we also have to preserve the one we've got here. William Anders of the Apollo 8 was the first to see the entire Earth at a glance. He said it looked like a fragile "little Christmas tree ornament against an infinite backdrop of space, the only color in the whole universe we could see. It seemed so very finite."
Well, because we are so very finite our responsibility to our planet must not be limited. That's why NASA's "Mission to \JPlanet\j Earth" is also a very important part of our future in space. We have to continue to monitor the global environment from space and to act on what we learn. Above all, let us never forget that all this work is about renewing our hopes and the hopes of generations to come. About the ability of Americans and the ability of human beings everywhere to conquer the seemingly impossible. I don't think anybody can look at the faces of these young people here with us today, and we ought to take a little while and look at them and welcome them here, without seeing again in their eyes dreams that those of us who are older could not have dreamed.
The explorations we continue in space are clear evidence to them that they will grow up in exciting times without limits. Times that demand their imagination, their vision, their courage. Times that will reward them, too, for believing in themselves and their possibilities.
One of our Young Astronauts, 13 year old Wayne Gusman from New Orleans, sees a future where being an \Jastronaut\j will be like, and I quote, "driving a car; everyone will do it."
That's a great dream. But that and our other dreams are clearly the natural extensions of the space program which began a generation ago, the direct descendants of the dreams of the three men we are here to honor today. We can get there.
No one who was alive then will ever forget where they were as Michael Collins traveled his solitary vigil around the Moon and Neil Armstrong and Buzz Aldrin landed that tiny craft on the surface. The world was captivated not only by the risk and the daring, although they were risking and daring, they were captivated because the landing meant again that the human experiment in conquering new and uncharted worlds was reborn.
In that sense it was not an end but a beginning. So, to you, gentlemen, we say for your valor, your courage, your pioneering spirit, and for being here today to remind us again that all things are possible, we are deeply in your debt. Thank you very much. (Applause.)
#
"Magellan Program",107,0,0,0
A US space mission, that between late 1990 and 1992 successfully mapped Venus in its entirety, using an orbiter with side-looking radar to provide sub-kilometer resolution images. Radar is used because of the global cloud cover, the created images being interpreted much as are \Jtelevision\j images. The single Magellan \Jspacecraft\j was managed by NASA's Jet Propulsion Laboratory and was launched using the NASA Shuttle in May 1989.
\BMission Details
Launch Date:\b May 4, 1989
\BDescription\b
The Magellan \Jspacecraft\j was launched on May 4, 1989, arrived at Venus on August 10, 1990 and was inserted into a near-polar elliptical orbit with a periapsis altitude of 294 km at 9.5 degrees. N. Radio contact with Magellan was lost on October 12, 1994. The primary objectives of the Magellan mission were to map the surface of Venus with a synthetic aperture radar (SAR) and to determine the topographic relief of the planet.
At the completion of radar mapping, 98% of the surface was imaged at resolutions better than 100 m, and many areas were imaged multiple times. The mission was divided up into "cycles", each cycle lasted 243 days (the time necessary for Venus to rotate once under the Magellan orbit -- i.e. the time necessary for Magellan to "see" the entire surface once.)
The mission proceeded as follows:
04 May 1989 -- Launch.
10 Aug 1990 -- Venus orbit insertion and \Jspacecraft\j checkout.
15 May 1991 -- Cycle 2: Radar mapping (right-looking).
15 Jan 1992 -- Cycle 3: Radar mapping (left-looking).
14 Sep 1992 -- Cycle 4: Gravity data acquisition.
24 May 1993 -- Aerobraking to circular orbit.
03 Aug 1993 -- Cycle 5: Gravity data acquisition.
30 Aug 1994 -- Windmill experiment.
12 Oct 1994 -- Loss of radio signal.
13 Oct 1994 -- Expected loss of \Jspacecraft\j.
A total of 4,225 usable SAR imaging orbits was obtained by Magellan. Each orbit typically covered an area 20 km wide by 17,000 km long, at a resolution of 75 m/pixel. This raw SAR data was processed into image strips called full-resolution basic image data records (F-BIDRs). Adjacent F-BIDRs were then assembled into full-resolution mosaicked image data records (F-MIDRs). These images were then compressed once (by a factor of 3), twice (9), or 3 times (27), to give C1-, C2-, and C3-MIDRs.
The Magellan mission scientific objectives were to study land forms and tectonics, impact processes, erosion, deposition, chemical processes, and model the interior of Venus. Magellan showed us an Earth-sized planet with no evidence of Earth-like plate tectonics. At least 85% of the surface is covered with volcanic flows, the remainder by highly deformed mountain belts.
Even with the high surface temperature (475â•‘ C) and high atmospheric pressure (92 bars), the complete lack of water makes erosion a negligibly slow process, and surface features can persist for hundreds of millions of years. Some surface modification in the form of wind streaks was observed. Over 80% of Venus lies within 1 km of the mean radius of 6,051.84 km.
The mean surface age is estimated to be about 500 My. A major unanswered question concerns whether the entire surface was covered in a series of large events 500 My ago, or if it has been covered slowly over time. The gravity field of Venus is highly correlated with the surface \Jtopography\j, which indicates the mechanism of topographic support is unlike the Earth, and may be controlled by processes deep in the interior. Details of the global tectonics on Venus are still unresolved.
#
"Ulysses Project Information",108,0,0,0
A joint \JESA\j (European Space Agency)-NASA space exploration mission to observe the Sun and solar wind from a high solar latitude perspective. It requires the \Jspacecraft\j to fly on a trajectory which passes over the poles of the Sun, using a boost from Jupiter's gravity field.
Originally named the 'International Solar Polar Mission', it was planned as a two-spacecraft mission, before the cancellation of the NASA \Jspacecraft\j. It is managed by \JESA\j's Space Research and Technology Center and by NASA's Jet Propulsion Laboratory.
\BUlysses
Launch Date/Time:\b 1990-10-06 at 11:47:16 UTC
\BDescription\b
The primary objectives of Ulysses are to investigate, as a function of solar latitude, the properties of the solar wind and the interplanetary magnetic field, of galactic cosmic rays and neutral interstellar gas, and to study energetic particle composition and acceleration.
The 55 kg payload includes two magnetometers, two solar wind plasma instruments, a unified radio/plasma wave instrument, three energetic charged particle instruments, an interstellar neutral gas sensor, a solar X-ray/cosmic gamma-ray burst detector, and a cosmic dust sensor. The communications systems is also used to study the solar corona and to search for gravitational waves.
Secondary objectives included interplanetary and planetary physics investigations during the initial Earth-Jupiter phase and investigations in the Jovian magnetosphere. The \Jspacecraft\j used a Jupiter swingby in Feb. 1992 to transfer to a heliospheric orbit with high heliocentric inclination, and will pass over the rotational south pole of the sun in mid-1994 at 2 AU, and over the north pole in mid-1995.
A second solar orbit will take Ulysses again over the south and north poles in years 2000 and 2001, respectively. The \Jspacecraft\j is powered by a single radio-isotope generator. It is spin stabilized at a rate of 5 rpm and its high-gain antenna points continuously to the earth. A nutation anomaly after launch was controlled by CONSCAN. The original mission planned for two \Jspacecraft\j, one built by \JESA\j and the other by NASA. NASA canceled its \Jspacecraft\j in 1981.
#
"Viking Mission to Mars",109,0,0,0
\BDescription\b
NASA's Viking Mission to Mars was composed of two \Jspacecraft\j, Viking 1 and Viking 2, each consisting of an orbiter and a lander. The primary mission objectives were to obtain high resolution images of the Martian surface, characterize the structure and composition of the atmosphere and surface, and search for evidence of life. Viking 1 was launched on August 20, 1975 and arrived at Mars on June 19, 1976. The first month of orbit was devoted to imaging the surface to find appropriate landing sites for the Viking Landers. On July 20, 1976, Viking Lander 1 separated from the Orbiter and touched down at Chryse Planitia (22.27 N, 49.97 W, 2 km below the datum elevation).
Viking 2 was launched September 9, 1975 and entered Mars orbit on August 7, 1976. Viking Lander 2 touched down at Utopia Planitia (47.57N, 225.74 W, 3 km below the datum elevation) on September 3, 1976. The Orbiters imaged the entire surface of Mars at a resolution of 150 to 300 meters, and selected areas at 8 meters. The lowest periapsis altitude for both Orbiters was 300 km. Viking Orbiter 2 was powered down on July 25,1978 after 706 orbits, and Viking Orbiter 1 on August 17, 1980, after over 1,400 orbits.
The Viking Landers transmitted images of the surface, took surface samples and analyzed them for composition and signs of life, studied atmospheric composition and \Jmeteorology\j, and deployed seismometers. Viking Lander 2 ended communications on April 11,1980, and Viking Lander 1 on November 13, 1982, after transmitting over 1,400 images of the two sites.
The results from the Viking experiments give our most complete view of Mars to date. Volcanoes, \Jlava\j plains, immense canyons, cratered areas, wind-formed features, and evidence of surface water are apparent in the Orbiter images. The planet appears to be divisible into two main regions, northern low plains and southern cratered highlands. Superimposed on these regions are the Tharsis and Elysium bulges, which are high-standing volcanic areas, and Valles Marineris, a system of giant canyons near the equator.
The surface material at both landing sites can best be characterized as iron-rich clay. Measured temperatures at the landing sites ranged from 150 to 250 K, with a variation over a given day of 35 to 50 K. Seasonal dust storms, pressure changes, and transport of atmospheric gases between the polar caps were observed. The \Jbiology\j experiment produced no evidence of life at either landing site.
\BThe "Face on Mars"
Background: The Viking Images\b
The Viking missions to Mars in the late 1970s produced more information about the Red \JPlanet\j than had been gathered in all the previous centuries of study by Earth-bound astronomers and observers. The primary mission of the Viking program was to search for signs of life on the surface of Mars. Two landers containing sophisticated biological laboratories studied soil samples in a variety of tests which, it was hoped, would prove or disprove the existence of life.
The results of these tests indicated that Mars contained no life, at least at these landing sites. However, Viking gathered volumes of data on the weather, soil chemistry and other surface properties and mapped the surface using low-to-moderate resolution cameras on the two orbiters.
Shortly after mapping began in 1976, an interesting image taken by the Viking 1 Orbiter was received at the Jet Propulsion Laboratory, Pasadena, Calif., which contained a surface feature resembling a human or ape-like face. The photo was immediately released to the public as an interesting geological feature and dubbed the "Face on Mars." Shortly afterwards, other photos of the same area were taken, and some scientists believed that the formation appeared to be a face due to the lighting angles as seen from the Orbiter.
\BOrigin of Features Examined\b
Over the years, some people began to raise questions about the origins of the features. A few ideas and theories arose speculating that the features may have been built by aliens in the distant past. These theories are based largely on the results of computer photo enhancements, and other analytical techniques performed on the Viking images, beginning in the early 1980s.
Most planetary geologists familiar with the set of photos, however, concluded that the natural processes known to occur on Mars -- such as wind erosion, Mars quakes, and erosion from running water in the distant past -- could account for the formation of the complicated fretted terrain of the Cydonia region, including the face.
Because the entire data set includes only nine low-to-moderate resolution photos, scientists say that there just is not enough data available to justify what would be an extraordinary conclusion that the features are not natural in origin (many scientists question whether images alone would be enough to settle the matter). Such a proven discovery of extraterrestrial life or artifacts would be one of the greatest discoveries in human history, and, as such, demand the most rigorous scientific investigation.
However, despite the phenomenal nature of such a potential discovery, no one in the scientific community -- either in the US or worldwide -- has ever proposed an investigation for a mission to study these features. Until more data is gathered, many scientists consider the probability that the features are anything other than natural in origin are just too low to justify the major expenditure of public funds which such an investigation would entail (more on this below).
What is agreed on is that a greater number of high resolution images of this area should be gathered. Following the failure of the Mars Observer mission in August, 1993, NASA proposed a decade-long program of Mars exploration, including orbiters and landers. The program, called Mars Surveyor, would take advantage of launch opportunities about every 2 years to launch an orbiter and a lander to the Red \JPlanet\j.
The first mission, consisting of an orbiter to be launched in 1996, will map the surface and take high- and medium-resolution images of particular features on the Martian surface that are of high interest. NASA intends to make observations of the Cydonia region making the best effort feasible, either with the first orbiter or on follow-on missions, to obtain images of the "face" and nearby landforms.
Quite aside from the interest generated by these curious features, Cydonia has long been regarded as an area of high scientific importance, ever since the first detailed images were returned by NASA's Viking \Jspacecraft\j in the late 1970s. The Cydonia region of Mars is part of the so-called fretted terrain, a belt of landforms that circles Mars at about 30-40 degrees North Latitude. In this region, the ancient crust of Mars has been intensely eroded by weathering processes, leaving high remnants of older crust surrounded by lower plains of eroded debris.
The landforms of Cydonia resemble in some respects those of terrestrial deserts, but they probably have been shaped by a unique range of peculiarly martian agencies: wind, frost and possibly running water in ancient times. Deciphering the geological age and origin of this terrain will yield important insights into the evolution of the martian surface, into the role of ice and water in its development, and into the nature of the martian climate in times past.
\BProposing Investigations\b
The selection of goals and scientific priorities for NASA to undertake on future space science missions starts in the scientific and academic communities, as well as within NASA. Scientific associations, such as the National Academy of Science, determine the research priorities in any given field of science. For instance, the most important questions remaining about Mars include gaining an understanding of the amount of water on the planet; mapping the surface in detail to gain a complete understanding of the geological processes, history and composition; and gaining a global understanding of the atmosphere, including climate and weather.
When NASA receives permission to proceed with a science mission, the Agency publishes an Announcement of Opportunity (AO). The AO solicits interest in providing high priority scientific investigations and instruments that will be part of the new mission. The AO receives the widest possible circulation throughout the university and research communities and industry.
Proposals are submitted and reviewed through a competitive peer review process. In this process, scientists from various institutions and organizations evaluate each proposal's scientific and technical merit, and then rank the relative merit of each. NASA receives the reports of the review panels and makes a final selection as to which instruments will be built and actually flown. This rational selection process ensures that only the most useful research, with a high probability of returning good science, is done at taxpayer expense.
After selection, each Mars Surveyor Principle Investigator (PI) team will develop its instrument, build it, test it and prepare it for launch and the 10-month journey to Mars. They are also charged with developing, testing, and using the software required to properly calibrate their instrument's data. Most of the scientists working on the various Mars Surveyor missions will have several years invested in their instrument before the \Jspacecraft\j arrives at Mars and they can actually receive the bulk of the data they have been waiting for.
\BObtaining Images of the "Face" and other Planetary Data\b
Since the release and subsequent widespread circulation of the "face" images, scientists and individual members of the public have freely drawn their own conclusions about the nature and origin of this feature. NASA encourages anyone seriously interested in this topic to obtain the photo(s) and decide for themselves, just as every day many hundreds of independent researchers and scientists make use of NASA-provided data on a variety of subjects.
The most noteworthy image of the "face" feature is available to the public, for a nominal fee, through Headquarters and JPL. A photo catalogue can be provided to select images.
All imaging data obtained by the Mars Surveyor program, as well as other types of data, will be deposited in open data \Jarchives\j. Two such \Jarchives\j widely used are the Planetary Data System (PDS), an open archive accessible to thousands of scientists and other individuals, and the National Space Science Data Center (NSSDC) where images and other data will be readily available to the general public (generally on CD-ROMs or as hard copy, as appropriate), for a nominal charge that covers the materials and time needed to produce the copies. For information about ordering copies of NASA science mission images, including on CD-ROM format, contact the NSSDC at:
National Space Science Data Center
Request Coordination Center
Goddard Space Flight Center
Greenbelt, MD 20771
Telephone (301) 286-6695
#
"Skylab Program",110,0,0,0
\JSkylab Introduction\j
\JSkylab I\j
\JSkylab II\j
\JSkylab III\j
#
"Skylab Introduction",111,0,0,0
\BLaunch Date/Time:\b 1973-05-14 at 00:00:00 UTC
\BDescription\b
The Skylab (SL) was a manned, orbiting \Jspacecraft\j composed of five parts, the Apollo \Jtelescope\j mount (ATM), the multiple docking adapter (MDA), the airlock module (AM), the instrument unit (IU), and the orbital workshop(OWS). The Skylab was in the form of a cylinder, with the ATM being positioned 90 degrees from the longitudinal axis after insertion into orbit. The ATM was a solar observatory, and it provided attitude control and experiment pointing for the rest of the cluster. It was attached to the MDA and AM at one end of the OWS. The retrieval and installation of film used in the ATM was accomplished by astronauts during extravehicular activity(EVA).
The MDA served as a dock for the command and service modules, which served as personnel taxis to the Skylab. The AM provided an airlock between the MDA and the OWS, and contained controls and instrumentation. The IU, which was used only during launch and the initial phases of operation, provided guidance and sequencing functions for the initial deployment of the ATM, solar arrays, etc. The OWS was a modified Saturn 4B stage suitable for long duration manned habitation in orbit. It contained provisions and crew quarters necessary to support three-person crews for periods of up to 84 days each.
All parts were also capable of unmanned, in-orbit storage, reactivation, and reuse. The Skylab itself was launched on May 14, 1973. It was first manned during the period May 25 to June 22, 1973, by the crew of the SL-2 mission (73-032A). Next, it was manned during the period July 28 to September 25, 1973, by the crew of the SL-3 mission (73-050A). The final manned period was from November 16, 1973, to February 8, 1974, when it was manned by the crew from the SL-4 mission.
#
"Skylab I",112,0,0,0
NASA had studied various concepts for a space station, including inflatable donuts, Chesley Bonestell's magnificent "Wheel," and various other designs since the earliest beginnings of the space program. When the Saturn rocket was developed in the mid-'60s, enabling some heavy lifting into space, the Skylab Program began to take shape. Following cancellation of Apollo 18, 19 and 20, we had a lot of hardware lying around gathering dust, so we put it to some remarkably good use.
At first there were two competing concepts for a space station. The first, called the "Wet Concept," called for launching a Saturn 1B, venting the S IV-B upper stage and refurbishing it, converting it to the space station, while in orbit. The second, or "Dry Concept," called for outfitting the S IV-B while still on the ground and launching it atop a Saturn V. While the Apollo 11 astronauts were actually on the moon in July, 1969, the decision was made to go with the Dry Concept.
Skylab weighed about 100 tons, and its launch marked the last launch of the wonderful Saturn V, the rocket that never failed. It had a volume of 283.17 cubic meters and was separated into two "floors;" the "upper" floor contained storage lockers and a large empty space for conducting experiments, and two airlocks, one pointed "down" toward the earth and the other "up" toward the sun; the "lower" floor was divided into rooms including a dining room with a table, three bedrooms, a work area, a bathroom and a shower. The floors consisted of an open gridwork that fit cleats on the bottom of the astronauts' shoes.
The station was also equipped with an airlock module for the many spacewalks that were required to change film in the external cameras and make repairs to the station. The Apollo Command and Service Modules remained attached to the station's docking mechanism for the duration of the astronauts' stays aboard the station.
The largest piece of scientific equipment was the "Apollo \JTelescope\j Mount" or ATM, which had its own solar panels for electricity generation and was used to make apectrographic analyses of the Sun without interference from Earth's atmosphere.
Launch of the unoccupied Skylab, designated Skylab 1 (the occupied missions were officially designated Skylabs 2, 3 and 4, but are generally referred to as Skylabs I, II and III, and are referred to in that manner here) took place on May 14, 1973, and problems set in early on.
One minute and three seconds into launch, the \Jmeteorite\j shield/sunshade was torn loose by the aerodynamic forces, destroying one of the solar arrays and damaging the other. The first crew, Skylab I, which was supposed to launch the next day, was delayed for ten days while mission personnel devised a method to repair the crippled station.
In all, three crews were launched, each in turn setting a record for longest human space flight; the all-time American record, which stood until Norm Thagard broke it aboard Mir in 1995 (and now held by Shannon Lucid), was set by Skylab III at over 2,017 hours (three months) and 1,214 orbits of the earth.
The Skylab program totalled 513 man-days in orbit, conducted thousands of experiments in many different disciplines, and even viewed the rather disappointing \Jcomet\j Kohoutek from Skylab III.
Skylab's orbit slowly deteriorated and it finally burned up in the atmosphere on July 11, 1979, more than five years after the last crew left for home.
#
"Skylab II",113,0,0,0
Launched: July 28, 1973
Splashed Down: September 25, 1973
Duration: 1,427 hours, 9 minutes, 4 seconds
Orbits: 858
Crew:
ò Alan L. Bean
ò Jack R. Lousma
ò Owen K. Garriott
This mission once again more than doubled the previous endurance record in space, just set by the astronauts of Skylab I just a month earlier. After an early bout with motion sickness, the crew settled down for their two-month mission, deploying a second sun shield on a space walk lasting six hours, 30 minutes.
They conducted many experiments, and brought with them live spiders to conduct a student-designed experiment to see what kinds of webs the spiders would spin in weightlessness.
Also on this mission the astronauts finally got to test the \JAstronaut\j Maneuvering Unit, or AMU, which had initially been carried into space aboard Gemini IX but could not be tested then because of problems with the old Gemini space suit.
The AMU experiments assisted engineers in designing the Manned Maneuvering Unit, which was first flown aboard the Shuttle flight STS 41B in February, 1984, and was still in use until quite recently.
#
"Skylab III",114,0,0,0
Launched: November 16, 1973
Splashed Down: February 8, 1974
Duration: 2,017 hours, 15 minutes, 31 seconds
Orbits: 1,214
Crew
ò Gerald P. Carr
ò William R. Pogue
ò Edward C. Gibson
At 84 days, 1 hour, 15 minutes and 31 seconds, Skylab III (actually officially designated "Skylab 4") remains by far the longest American space flight, a record that will certainly stand until the permanent human occupation of space begins with the international Space Station.
To help keep the crew in physical condition during their almost three months in orbit, they walked treadmills and rode an on-board stationary \Jbicycle\j, and came home in far better condition than had the previous Skylab crews.
Among the thousands of experiments they conducted during this long, long flight, the astronauts took four space walks, including one on Christmas Day to observe the \Jcomet\j Kohoutek.
#
"Galileo Project Information",115,0,0,0
\JGalileo Project Summary\j
\JCallisto Summary\j
\JGalilean Satellites (Jupiter)\j
\JEuropa, Natural and False Color Views\j
\JEuropa's Surface Reshaped\j
\JGalileo Returns To Europa For Another Close Look\j
\JGalileo Mission Status (1)\j
\JGalileo Mission Status (2)\j
\JGalileo Mission Status (3)\j
\JGalileo Images Hint At History For Europa\j
\JGalileo Mission Status (4)\j
\JGalileo Mission Status (5)\j
\JGalileo To Take One Last Close Look At Ganymede\j
\JGalileo Returns New Insights Into Callisto and Europa\j
\JJupiter's Dry Spots and Glowing Auroras To Be Unveiled\j
\JGalileo Finds New View of Jupiter's Light Show\j
\JGalileo Finds Europa Has An Atmosphere\j
\JGanymede G1 & G2 Encounters\j
\JGanymede, Fractures in Transitional Terrain\j
\JGalileo Calendar of Events\j
#
"Galileo Project Summary",116,0,0,0
\BOrbiter Launch Date:\b 12 October 1989
\BProbe Release Date:\b 13 July 1995
\BDescription\b
The \JGalileo\j \Jspacecraft\j flew by the Earth and Moon on Dec. 8, 1990 and Dec. 7, 1992. The primary mission of the \JGalileo\j orbiter and probe is to explore Jupiter and its satellites. Due to the great distance to Jupiter of over 600 million kilometers, and the on board fuel limitations, a series of planetary flybys has taken place in order to give \JGalileo\j a gravity assist to Jupiter:
Launch: 12 Oct 1989
Venus: 10 Feb 1990
Earth/Moon 1: 08 Dec 1990
Gaspra: 29 Oct 1991
Earth/Moon 2: 08 Dec 1992
Ida: 28 Aug 1993
Jupiter arrival: 07 Dec 1995
These flybys gave \JGalileo\j an opportunity to image the Moon at various wavelengths with the Solid State Imaging (SSI) camera. The camera uses a high-resolution, 800 x 800 charge-coupled device (CCD) array with a field of view of 0.46 degrees. Multi-spectral coverage is provided by an eight-position filter wheel on the camera, consisting of three broad-band filters: violet (404 nm), green (559 nm), and red (671 nm); four near-infrared filters: 727 nm, 756 nm, 889 nm, and 986 nm; and one clear filter (611 nm) with a very broad (440 nm) passband.
#
"Callisto Summary",117,0,0,0
With a diameter of over 4,800 km (2,985 miles), Callisto is the third largest satellite in the solar system (only Ganymede and Titan are bigger), and is almost the size of Mercury. Callisto is the outermost of the Galilean satellites, and orbits beyonds Jupiter's main radiation belts.
Callisto has the lowest density of the Galilean satellites (1.86 grams/cubic centimeter). Its interior is probably similar to Ganymede except the inner rocky core is smaller, and this core is surrounded by a large icy mantle. Callisto's surface is the darkest of the Galileans, but it is twice as bright our our own Moon.
Callisto is the most heavily cratered object in the solar system. It is thought to be a long dead world, with a nearly complete absence of any geologic activity on its surface. In fact, Callisto is the only body greater than 1000 km in diameter in the solar system that has shown no signs of undergoing any extensive resurfacing since impacts have molded its surface. With a surface age of about 4 billion years, Callisto has the oldest landscape in the solar system.
There are no large mountains on Callisto, which is probably due to the icy nature of the satellite's surface. The surface features are dominated by impact craters and rings, and the craters are quite shallow. There are two large "bulleye" structure on Callisto, thought to be the result of a massive impact. The largest structure, \JValhalla\j, has a bright patch 600 km across with rings extending out to almost 3000 km. The other ring structure, Asgard, is about 1600 km in diameter.
Seven impact crater chains have been mapped on Callisto. These chains probably formed when fragments of a \Jcomet\j were split apart by Jupiter's gravity and impacted on Callisto. In a similar scenario, \JComet\j Shoemaker-Levy 9 split into 21 fragments and impacted Jupiter in 1994.
No atmosphere has been detected on Callisto.
The \JGalileo\j \Jspacecraft\j will make three close passes by Callisto during its 2 year orbital tour around Jupiter. The first close encounter occurred November 4, 1996.
Callisto Quick-Look Statistics: (Refer to Table)
#
"Galilean Satellites (Jupiter)",118,0,0,0
This composite includes the four largest moons of Jupiter which are known as the Galilean satellites. From left to right, the moons shown are Ganymede, Callisto, Io, and Europa. The Galilean satellites were first seen by the Italian astronomer \JGalileo\j Galilei in 1610. In order of increasing distance from Jupiter, Io is closest, followed by Europa, Ganymede, and Callisto.
The order of these satellites from the planet Jupiter helps to explain some of the visible differences among the moons. Io is subject to the strongest tidal stresses from the massive planet. These stresses generate internal heating which is released at the surface and makes Io the most volcanically active body in our solar system.
Europa appears to be strongly differentiated with a rock/iron core, an ice layer at its surface, and the potential for local or global zones of water between these layers. Tectonic resurfacing brightens terrain on the less active and partially differentiated moon Ganymede. Callisto, furthest from Jupiter, appears heavily cratered at low resolutions and shows no evidence of internal activity.
North is to the top of this composite picture in which these satellites have all been scaled to a common factor of 10 kilometers (6 miles) per picture element.
The Solid State Imaging (CCD) system aboard NASA's \JGalileo\j \Jspacecraft\j obtained the Io and Ganymede images in June 1996, while the Europa images were obtained in September 1996. Because \JGalileo\j focusses on high resolution imaging of regional areas on Callisto rather than global coverage, the portrait of Callisto is from the 1979 flyby of NASA's Voyager \Jspacecraft\j.
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Europa, Natural and False Color Views",119,0,0,0
November 12, 1996
The images on this page show different views of the ice-covered satellite, Europa. Dark brown areas represent rocky material derived from the interior, implanted by impact, or from a combination of interior and exterior sources. Bright plains in the polar areas (top and bottom) are shown in tones of blue to distinguish possibly coarse-grained ice (dark blue) from fine-grained ice (light blue).
Long, dark lines are fractures in the crust, some of which are more than 3,000 kilometers (1,850 miles) long. The bright feature containing a central dark spot in the lower third of the image is a young impact crater some 50 kilometers (31 miles) in diameter. This crater has been provisionally named 'Pwyll' for the Celtic god of the underworld.
Europa is about 3,160 kilometers (1,950 miles) in diameter, or about the size of Earth's moon. This image was taken on September 7, 1996, at a range of 677,000 kilometers (417,900 miles) by the solid state imaging \Jtelevision\j camera onboard the \JGalileo\j \Jspacecraft\j during its second orbit around Jupiter. The image was processed by Deutsche Forschungsanstalt fuer Luft- und Raumfahrt e.V., Berlin, \JGermany\j.
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Europa's Surface Reshaped",120,0,0,0
January 17, 1997
Ice-spewing volcanoes and the grinding and tearing of tectonic plates have reshaped the chaotic surface of Jupiter's frozen moon Europa, images from NASA's \JGalileo\j \Jspacecraft\j reveal.
\BFlows on Europa\b
The images, captured when \JGalileo\j flew within just 430 miles (692 kilometers) of Europa on Dec. 19, were released at a news briefing today at NASA Headquarters, Washington, DC.
Although the images do not show currently active ice volcanoes or geysers, they do reveal flows of material on the surface that probably originated from them, said \JGalileo\j imaging team member Dr. Ronald Greeley of \JArizona\j State University, Tempe.
"This is the first time we've seen actual ice flows on any of the moons of Jupiter," said Greeley. "These flows, as well as dark scarring on some of Europa's cracks and ridges, appear to be remnants of ice volcanoes or geysers."
The new images appear to enhance Europa's prospects as one of the places in the Solar System that could have hosted the development of life, said Greeley.
"There are three main criteria to consider when you are looking for the possibility of life outside the Earth -- the presence of water, organic compounds and adequate heat," said Greeley. "Europa obviously has substantial water ice, and organic compounds are known to be prevalent in the Solar System. The big question mark has been how much heat is generated in the interior.
"These new images demonstrate that there was enough heat to drive the flows on the surface. Europa thus has a high potential to meet the criteria for exobiology," Greeley added.
"This doesn't prove that there is an ocean down there under the surface of Europa, but it does demonstrate that it is a scientifically exciting place," said \JGalileo\j imaging team member Dr. Robert Sullivan, also of \JArizona\j State University.
\BGalileo Image of Europa\b
The images also reveal a remarkable diversity in the geological age of various regions of Europa's surface. Some areas appear relatively young, with smooth, crater-free terrain, while others contain large craters and numerous pits, suggesting that they are much older.
The icy crust bears the signs of having been disrupted by the motion of tectonic plates. "There appear to be signs of different styles of tectonism," said Greeley. "In many areas we see that the crust was pulled apart in a spreading similar to the processes on the sea floor on Earth. This is different from the tectonic processes at work on, say, Jupiter's moon Ganymede. This suggests that Europa's interior may be different from Ganymede's."
Galileo scientists will have a better chance to understand Europa's interior when the \Jspacecraft\j gathers gravity data on another flyby next November. The gravity field is measured by tracking how the frequency of \JGalileo\j's radio signal changes as it flies past the moon. This was not possible during the recent flyby because radio conditions were degraded as Jupiter passed behind the Sun from Earth's point of view.
\BRidges on Europa\b
Europa is crisscrossed by an amazingly complex network of ridges, according to Sullivan. "Ridges are visible at all resolutions," he explained. "Closely paired ridges are most common. With higher resolution, ridges seen previously as singular features are revealed to be double."
Some of the ridges may have formed by tension in the icy crust: as two plates pull apart slightly, warmer material from below might push up and freeze to form a ridge. Other ridges may have been formed by compression: as two plates push together, the material where they meet might crumple to form the ridge.
In addition to ice flows and tectonics, Greeley and Sullivan noted that some areas on Europa seem to have been modified by unknown processes that scientists are still debating. Greeley said that some areas, for example, seem to have been modified by "sublimation erosion" -- the \Jevaporation\j of water and other volatiles such as ammonia and \Jmethane\j into the vacuum of space. "Something is destroying the topography," said Greeley, "and this sublimation erosion is a good candidate for what is at work."
During last month's encounter, \JGalileo\j flew more than 200 times closer to Europa than the Voyager 2 \Jspacecraft\j did in 1979. After a swing past Jupiter next week in what mission engineers call a "phasing orbit," \JGalileo\j's next targeted flyby will take it again past Europa as it passes within 364 miles (587 kilometers) on Feb. 20.
#
"Galileo Returns To Europa For Another Close Look",121,0,0,0
February 19, 1997
NASA's \JGalileo\j \Jspacecraft\j will make an encore appearance at Jupiter's icy moon, Europa, on Thursday, Feb. 20, marking the closest planned Europa flyby of the initial two-year mission.
The encounter will be \JGalileo\j's closest flyby yet of Europa. The craft will swoop past the Jovian moon at an altitude of 580 kilometers (360 miles) on Thursday, Feb. 20, at 9:06 a.m. Pacific time (12:06 p.m. Eastern time).
Galileo made its first pass of Europa in December 1996, revealing remarkable detail of that moon's terrain. This week's flyby will look at other areas of Europa's surface, which is covered by ice and a series of criss-crossed, dark lines. Europa holds great fascination for scientists because of the possibility that liquid oceans may be hidden underneath the icy surface. The presence of liquid water would boost the odds that Europa could host some form of life.
"I think this flyby may provide additional clues regarding the prospect of liquid water oceans on Europa," said \JGalileo\j Mission Director Bob Mitchell.
With its diameter of 3,138 kilometers (1,946 miles), Europa is just slightly smaller than Earth's moon. Because the \Jgeometry\j of the upcoming flyby will be somewhat different from the path taken by \JGalileo\j's previous Europa encounter, it will yield data and images of different portions of the moon.
"This position will allow for high resolution of different terrain," said Mitchell. "It will help us learn more about Europa's structure and surface and how the surface was formed."
The current Europa encounter phase began on Sunday, Feb. 16, and will continue through Saturday, Feb. 22. The \Jspacecraft\j has already begun returning real-time encounter data, with recorded data scheduled to be transmitted to Earth beginning on the evening of Saturday, Feb. 22 (Pacific time).
This encounter will include the return of magnetospheric measurements from Europa's vicinity. Other science highlights will include the study of surface features of Europa's lineated regions, images of two other, smaller Jovian moons, Thebe and Amalthea, and studies of such Jovian atmospheric features as the south equatorial belt-zone boundary and the aurora borealis.
This flyby provides a period of radio \Joccultation\j, when Europa crosses between Earth and \JGalileo\j, temporarily cutting off the \Jspacecraft\j's radio signal. This affords a prime opportunity for \JGalileo\j to study atmospheric data just before and after radio contact is lost, when the signal passes through the Europa's atmosphere.
"As the fifth encounter in \JGalileo\j's series of 10 flybys, this marks the approximate halfway point for this series, which began in June 1996," said \JGalileo\j Project Manager Bill O'Neil. "It's been eight months since then, and it will be another eight months before the series' final encounter."
A third Europa flyby is planned for Nov. 6, 1997, and JPL has asked NASA to extend the \JGalileo\j mission by two years to include eight more Europa flybys and ultimately a flyby of Io. The proposed extended mission might be shortened if the \Jspacecraft\j's operations were to deteriorate as a result of its continuous exposure to Jupiter's extreme radiation environment.
"NASA has assured us that the extended mission will be funded," said O'Neil. "The $30 million needed for the extension will come from within the existing NASA budget, enabled by cost savings due to improved efficiencies in JPL's \Jspacecraft\j tracking and mission operations."
The 2,223-kilogram (2-1/2 ton) \JGalileo\j orbiter \Jspacecraft\j was launched aboard Space Shuttle Atlantis on October 18, 1989.
#
"Galileo Mission Status (1)",122,0,0,0
February 27, 1997
NASA's \JGalileo\j \Jspacecraft\j has begun returning data from its latest flyby of Jupiter's icy moon, Europa, which took place last Thursday, Feb. 20. The \Jspacecraft\j's closest approach to Europa was at 9:06 a.m. Pacific time, with confirmation received on Earth at 9:56 a.m. \JGalileo\j's flyby was within one kilometer of the target altitude of 586 kilometers.
Galileo engineers will send commands to the \Jspacecraft\j tonight to correct a magnetometer software glitch that was discovered following the flyby. Although the problem caused the loss of flyby magnetometer data, similar data was collected during \JGalileo\j's previous Europa flyby and more will be gathered during the third encounter of that moon in November. Engineers expect the problem to be resolved within a few more days.
Recorded data playback began last Saturday evening and will continue through March 28, with initial data to include recent observations of Jupiter and its moon, Io. Transmission of the new Europa observations will begin later this week and continue through next week. New Europa images may be released within a few weeks.
This flyby, the closest of three Europa encounters for \JGalileo\j's primary mission, recorded data on a linea region which may represent a stress-controlled eruption or intrusion of material from beneath Europa's icy surface. This could be another sign of liquid oceans hidden underneath.
Scientists gathered data about Europa's atmosphere during the flyby when Europa crossed between Earth and \JGalileo\j, temporarily cutting off the \Jspacecraft\j's radio signal. This radio \Joccultation\j experiment uses the disturbances to the \Jspacecraft\j radio signal as it passes close to Europa to infer information about Europa's tenuous atmosphere.
Data to be returned this week include observations of Jupiter's "white ovals" -- huge storms fueled by an unknown energy source; surface changes and volcanic plume activity on Jupiter's moon, Io; and an image of one of Jupiter's smaller moons, Thebe.
Galileo has five more encounters of Jupiter's moons on its two-year primary mission journey through the Jovian system. The next destination is Jupiter's largest moon, Ganymede on Apr. 5, 1997. A planned follow-on mission would run from December 1997 through December 1999 and would include 15 encounters.
#
"Galileo Mission Status (2)",123,0,0,0
March 13, 1997
Playback from the latest Europa encounter on February 20 is proceeding on schedule, returning data from \JGalileo\j's most recent pass close to Jupiter and its satellites. The data include observations of white oval atmospheric features taken by the \Jspacecraft\j's near infrared mapping spectrometer and photopolarimeter-radiometer over a range of solar phase angles.
This week's data return will also include fields and particles instruments' high data rate recording as \JGalileo\j made its closest approach while flying through Jupiter's magnetic equator. Much of the plasma in Jupiter's inner magnetosphere is confined to the equatorial region and is whisked away to the outer boundaries of the magnetosphere through various processes. This data will help scientists understand more about those processes.
Other observations expected to be returned this week include chemical monitoring of volcanic hot spots on Io and a surface map from the Europa encounter. Ganymede observations to be returned are part of multi-orbit efforts to characterize the moon's surface and flesh out information obtained when the Voyager \Jspacecraft\j flew by the Jovian system. New observations of the small moon Amalthea will be used to determine the body's global shape and morphology.
Last week's playback focused on Europa observations, the prime target of last month's flyby. During the approach to Europa, most of the satellite was illuminated by the Sun. This vantage point enabled \JGalileo\j to gather information on the icy moon's surface composition and shape, as well as crater feature observations which may offer clues to what lies underneath Europa's surface. Near infrared mapping spectrometer data included observations of a lineated region and icy regions of varying ages.
A previous magnetometer glitch on \JGalileo\j has been corrected, and it's now believed the failure was caused by radiation effects. Although there are indications that the same radiation-induced faults have occurred with the magnetometer and the spectrometer, in each case, a reloading procedure has corrected the problem.
An orbital trim maneuver will begin today to put \JGalileo\j on track for its next destination, Ganymede. The flyby occurs on April 4 (PST). \JGalileo\j has five additional encounters of Jupiter's moons scheduled during its two-year primary journey through the Jovian system.
#
"Galileo Mission Status (3)",124,0,0,0
March 27, 1997
This is the final week for playback of data gathered during \JGalileo\j's February 20th Europa encounter. Included in the potpourri of Europa information is data on surface composition, an observation expected to help scientists distinguish between new and old ice, images of bright plains and craters that may help explain the formation of these features, and a fields and particles observation of Europa's interaction with Jupiter's magnetosphere.
The playback also features Io observations of surface chemistry, volcanic activity, and the surface while eclipsed from the Sun. Observations of two of Jupiter's white ovals, global observations of Ganymede and Callisto, and a single Amalthea observation have also be transmitted.
With the wrap-up of Europa 6 playback, the end of this week marks the start of preparations for \JGalileo\j's next encounter, a Ganymede flyby at 11:11 p.m. Pacific Standard Time on Friday, April 4. An orbital trim maneuver is planned for Monday, March 31, one day after the beginning of the Ganymede encounter period. Maintenance activities will also be performed, including the flushing of the thruster lines to prevent debris blockage, and conditioning of the tape recorder.
After this next Ganymede flyby, \JGalileo\j has four more encounters of Jupiter's moons scheduled during its two-year primary journey through the Jovian system. A planned two-year continuation of the mission, referred to as the \JGalileo\j Europa Mission (GEM), will include eight more Europa flybys and an Io flyby, as long as the \Jspacecraft\j remains healthy.
#
"Galileo Images Hint At History For Europa",125,0,0,0
April 9, 1997
Chunky ice rafts and relatively smooth, crater-free patches on the surface of Jupiter's frozen moon Europa suggest a younger, thinner icy surface than previously believed, according to new images from \JGalileo\j's \Jspacecraft\j released today.
The images were captured during \JGalileo\j's closest flyby of Europa on February 20, when the \Jspacecraft\j came within 586 kilometers (363 miles) of the Jovian moon. These features, which lend credence to the idea of hidden, subsurface oceans, are also stirring up controversy among scientists who disagree about the age of Europa's surface.
Dr. Ronald Greeley, an \JArizona\j State University geologist and \JGalileo\j imaging team member, said the ice rafts reveal that Europa had, and may still have, a very thin ice crust covering either liquid water or slush.
"We're intrigued by these blocks of ice, similar to those seen on Earth's polar seas during springtime thaws," Greeley said. "The size and \Jgeometry\j of these features lead us to believe there was a thin icy layer covering water or slushy ice, and that some motion caused these crustal plates to break up."
"These rafts appear to be floating and may, in fact, be comparable to icebergs here on Earth," said another \JGalileo\j imaging team member, Dr. Michael Carr, a geologist with the U.S. Geological Survey. "The puzzle is what causes the rafts to rotate. The implication is that they are being churned by convection."
The new images of Europa's surface have also sparked a lively debate among scientists. \JGalileo\j imaging team member Dr. Clark Chapman is among those who believe the smoother regions with few craters indicate Europa's surface is much younger than previously believed.
In essence, Chapman, a planetary scientist at Southwest Research Institute, Boulder, CO, believes the fewer the craters, the younger the region. Chapman based his estimate on current knowledge about cratering rates, or the rate at which astronomical bodies are bombarded and scarred by hits from comets and \Jasteroids\j.
"We're probably seeing areas a few million years old or less, which is about as young as we can measure on any planetary surface besides Earth," said Chapman. "Although we can't pinpoint exactly how many impacts occurred in a given period of time, these areas of Europa have so few craters that we have to think of its surface as young."
Chapman added, "Europa's extraordinary surface \Jgeology\j indicates an extreme youthfulness -- a very alive world in a state of flux."
However, Carr sees things differently. He puts Europa's surface age at closer to one billion years old.
"There are just too many unknowns," Carr said. "Europa's relatively smooth regions are most likely caused by a different cratering rate for Jupiter and Earth. For example, we believe that both Earth's moon and the Jovian moon, Ganymede, have huge craters that are 3.8 billion years old. But when we compare the number of smaller craters superimposed on these large ones, Ganymede has far fewer than Earth's moon. This means the cratering rate at Jupiter is less than the cratering rate in the Earth-moon system."
Scientists hope to find answers to some of the questions surrounding Europa and its possible oceans as the \JGalileo\j \Jspacecraft\j continues its journey through the Jovian system.
"We want to look for evidence of current activity on Europa, possibly some erupting geysers," Greeley said. "We also want to know whether Europa's surface has changed since the Voyager \Jspacecraft\j flyby in 1979, or even during the time of the \JGalileo\j flybys."
The craft will return for another Europa flyby on November 6, 1997, the final encounter of \JGalileo\j's primary mission. However, eight more Europa flybys are planned as part of \JGalileo\j's two-year extended mission, which will also include encounters with two other Jovian moons, Callisto and Io.
#
"Galileo Mission Status (4)",126,0,0,0
April 17, 1997
The \JGalileo\j \Jspacecraft\j is operating normally in its second week of "cruise" following the craft's latest encounter with Ganymede at 11:10pm Pacific Standard Time on April 4, with the signal received on the ground 46 minutes later. \JGalileo\j flew by the satellite at an altitude of 3,102 kilometers.
The fields and particles survey of Jupiter's magnetosphere continues. Other scheduled activities include one of the periodic \Jspacecraft\j turns to keep the antenna pointed near Earth, and transmission of commands to prepare for next week's orbit adjustment.
This week's playback includes observations taken by the \Jspacecraft\j during its non-targeted flyby of the increasingly- popular moon Europa. The playback includes a near-infrared mapping spectrometer (NIMS) observation at regional resolution, part of a plan to map all the Galilean satellites.
Another spectrometer observation was designed to look for differences in the mineral composition of the Tyre Macula region, a circular feature. Return from other instruments will include thermal observations and images of crater features near the terminator, the dividing line between day and night.
The playback schedule also features the return of observations of Jupiter's Great Red Spot, a smaller red spot, and a hot spot near the same latitude as the atmospheric probe entry site.
Galileo will return for another Ganymede flyby on May 7, with two Callisto encounters and another Europa flyby planned for the final orbits of \JGalileo\j's primary mission. A planned two- year continuation of the mission, known as the \JGalileo\j Europa Mission (GEM), will include eight more Europa flybys and one or two Io flybys, as long as the \Jspacecraft\j remains healthy.
#
"Galileo Mission Status (5)",127,0,0,0
April 28, 1997
As the \JGalileo\j \Jspacecraft\j prepares for another encounter with Jupiter's largest moon, Ganymede, playback of data from the craft's previous Ganymede flyby on April 4 Pacific Standard Time is nearing completion.
This final batch of data return from the April flyby will conclude around noon Pacific Daylight Time on Saturday, May 3. It includes observations of Ganymede's bright, dark and dark- rayed regions from the remote sensing instruments, and high resolution fields and particles data on the magnetospheres around Jupiter and Ganymede and the interaction between the two.
Playback also includes observations of Europa's lineated circular regions. This will be used to help determine how these features originated and to construct a global map of Europa at regional resolution. Regional observations of Jupiter to be returned this week will also be used for a global map of Jupiter.
A few Jupiter observations from the remote sensing instruments will be returned, including a study of a small red spot in the Jovian atmosphere, a hot spot on the planet and a single image of Adrastea, one of Jupiter's small inner moons. The fields and particles instruments' survey of Jupiter's magnetosphere will resume on Friday, marking the start of the second magnetospheric "mini-tour."
The \JGalileo\j flight team will transmit the first set of encounter sequence commands to the \Jspacecraft\j later this week, as it prepares for the next encounter with Ganymede on May 7. This will be \JGalileo\j's final close flyby of Ganymede.
#
"Galileo To Take One Last Close Look At Ganymede",128,0,0,0
May 6, 1997
NASA's \JGalileo\j \Jspacecraft\j will fly by Jupiter's largest moon, Ganymede, for the fourth and final time on Wednesday, May 7.
The closest approach will take place at 8:56 a.m. Pacific Daylight Time as the craft travels 1,600 kilometers (994 miles) above Ganymede at a speed of 8.6 kilometers per second (more than 19,000 miles per hour).
During the encounter, \JGalileo\j will collect data on the moon's surface shape and atmosphere. High resolution studies by the craft's remote-sensing instruments will include observations of \JOsiris\j, a dome structure; Tiamat Sulcus, a region of craters, grooves and furrows; a multi-ringed structure; and caldera-like features and dark floor craters.
Galileo will also begin its second "mini-tour" of the Jovian magnetosphere to learn more about the composition and dynamics of this tremendously vast region around Jupiter controlled by the Jovian magnetic field. This second "mini-tour" will continue until the end of \JGalileo\j's primary mission on Dec. 7. The tour, to take place this summer, will include a deep penetration into Jupiter's magnetotail, the region of the magnetosphere opposite the Sun's direction.
During this encounter, Ganymede will block the \Jspacecraft\j from the Earth and the Sun for about seven minutes. This will provide scientists with an opportunity to measure changes in the \Jspacecraft\j's radio signal as it passes very close to Ganymede, but just before it's blocked out by the Jovian moon. These measurements will allow for further study of Ganymede's tenuous atmosphere.
In addition to being the largest of the Jovian satellites, Ganymede is the largest moon in the solar system. Although this marks \JGalileo\j's last encounter with Ganymede, the craft will fly by two other Jovian moons, Callisto and Europa, before its primary mission ends in December. A two-year extension of the \JGalileo\j mission will enable further studies of Europa and Io, depending on the \Jspacecraft\j's health.
Galileo was launched in 1989 and entered orbit around Jupiter on Dec. 7, 1995.
#
"Galileo Returns New Insights Into Callisto and Europa",129,0,0,0
May 23, 1997
Jupiter's icy moon Europa has a metallic core and layered internal structure similar to the Earth's, while the heavily cratered moon Callisto is a mixture of metallic rock and ice with no identifiable central core, according to new results from NASA's \JGalileo\j mission.
In addition, recent plasma wave observations from \JGalileo\j show no evidence of a magnetic field or magnetosphere around Callisto, but do hint at the prospect of a tenuous atmosphere.
These peer-reviewed findings, reported in today's issue of Science magazine and the May 16 issue of Nature magazine, are based on data gathered during \JGalileo\j's Nov. 4, 1996, flyby of Callisto and its Europa encounters on Dec. 19, 1996, and Feb. 20, 1997.
"Before \JGalileo\j, we could only make educated guesses about the structure of the Jovian moons," said Dr. John Anderson, a planetary scientist at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "Now, with the help of the \Jspacecraft\j, we can measure the gravitational fields of the satellites and determine their interior structure and density. We can determine how the matter is distributed inside."
While scientists use seismic waves to study Earth's interior, \JGalileo\j performs remote studies of Jupiter's moons by measuring small changes in the \Jspacecraft\j's trajectory as it passes each body.
"These new results from the gravity data are very consistent with the idea of subsurface oceans on Europa," Anderson said. "We know that Europa has a very deep layer of water in some form, but we don't yet know whether that water is liquid or frozen."
In an article appearing in the May 23 edition of Science, Dr. Margaret Kivelson, principal investigator for \JGalileo\j's magnetometer, reports that during its December 1996 pass by Europa, the magnetometer detected what she described as "a substantial magnetic signature," and also found that Europa's north magnetic pole is pointed in an odd direction.
Based on these observations, Kivelson, a professor at the University of \JCalifornia\j at Los Angeles, said Europa may have a magnetic field about one-quarter the strength of Ganymede's magnetic field.
Although the magnetometer was malfunctioning during \JGalileo\j's Europa flyby in February 1997, Kivelson said the problem is corrected and the device is expected to return valuable data during its upcoming Europa flybys. The next Europa encounter is scheduled for November, with a series of flybys planned during a two-year \JGalileo\j extended mission.
Galileo's findings on the Jovian moon Callisto revealed a much different structure than Europa. Scientists believe that because Callisto is the Galilean moon located farthest from Jupiter, it was never subjected to the same gravitational pull as the inner moons and, therefore, never experienced enough heating to form different layers.
"Callisto had a much more sedate, predictable and peaceful history than the other Galilean moons," Anderson explained, "and, therefore, it is a more typical solar system object." The findings indicate Callisto has no core, but instead has a homogeneous structure, with 60 percent of its ingredients being rock, including iron and iron sulfide, and 40 percent made of compressed ice.
Dr. Donald Gurnett, principal investigator for the \JGalileo\j \Jspacecraft\j's plasma wave instrument, said the instrument displayed a very minor response from Callisto and, consequently, showed no evidence of a magnetic field or magnetosphere. The latest issue of Nature magazine contains these findings, as well as supportive data from magnetometer studies of Callisto, as reported by Dr. Krishan Khurana of UCLA.
However, Gurnett added, "There is some evidence of a plasma source on Callisto, which might indicate a very tenuous atmosphere." Gurnett is a professor at the University of \JIowa\j at \JIowa\j City.
The \JGalileo\j \Jspacecraft\j was launched in October 1989 and entered orbit around Jupiter on Dec. 7, 1995.
#
"Jupiter's Dry Spots and Glowing Auroras To Be Unveiled",130,0,0,0
June 2, 1997
New images from NASA's \JGalileo\j mission revealing dry spots and auroral light patterns on Jupiter will be presented at a press briefing on Thursday, June 5, at 11 a.m. Pacific Standard Time. The briefing will originate from NASA's Jet Propulsion Laboratory, Pasadena, CA, and will be carried live on NASA \JTelevision\j.
The latest images and data reveal the existence of areas where winds converge and cause clouds and moisture to evaporate, but also indicate that the giant, gaseous planet is not as dry as scientists had believed. This may clear up the controversy which arose after \JGalileo\j's probe entered the Jovian atmosphere on Dec. 7, 1995, and found no moisture.
Scientists will also discuss new images and data which show that Jupiter's glowing auroras stretch in a thin, patchy ribbon- like strand near the poles. Scientists believe that despite some similarities, auroras on Jupiter and on Earth are driven by different forces.
NASA \JTelevision\j is available through GE-2, transponder 9C at 85 degrees west longitude, vertical polarization, with a frequency of 3880 Mhz, and audio at 6.8 Mhz.
#
"Galileo Finds New View of Jupiter's Light Show",131,0,0,0
June 5, 1997
Jupiter has both wet and dry regions, just as Earth has tropics and deserts, according to new images and data from the \JGalileo\j \Jspacecraft\j released today. The data may explain why \JGalileo\j's atmospheric probe found much less water than scientists had anticipated when it dropped into the Jovian atmosphere in December 1995.
"We had suspected that the probe landed in the 'Sahara Desert of Jupiter,'" said Dr. Andrew Ingersoll, a professor at the \JCalifornia\j Institute of Technology, Pasadena, CA, and member of the \JGalileo\j science team. "But the new data show there is moisture in the surrounding areas. Jupiter is not as dry overall as we thought."
The area where the probe entered was a clearing in the clouds -- a dry spot through which deeper, warmer layers can be seen. By studying various areas, including those resembling the probe entry site, the \JGalileo\j orbiter has helped scientists understand the probe results.
In fact, the air around a dry spot has 100 times more water than the dry spot itself, according to Dr. Robert Carlson of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, principal investigator for the imaging spectrometer instrument onboard \JGalileo\j.
Such dry spots cover less than one percent of the Jovian atmosphere, and they appear to be regions where the winds converge and create a giant downdraft, according to \JCal\j Tech graduate student Ashwin Vasavada. In fact, the water content of the giant, gaseous planet varies at least as much as the moisture varies from place to place on Earth.
"Winds rise from the deep atmosphere and lose water and ammonia," explained Dr. Glenn Orton, a \JGalileo\j interdisciplinary scientist at JPL and Photopolarimeter-Radiometer co-investigator. "At the top, when they converge and drop back down, nothing is left to condense back into clouds, and a dry clearing is created. These dry spots may grow and diminish, but they recur in the same places, possibly because of the circulation patterns on Jupiter."
Ingersoll said the dry spots are found in a northern hemisphere band at five to seven degrees latitude. When the \JGalileo\j probe was released near the tops of the clouds, it found dry air underneath. But at other locations, the weather might be rather Earth-like.
In the months since the probe's descent, \JGalileo\j mission scientists have debated whether the dry conditions it encountered were due to the downdraft concept, or whether Jupiter's water had somehow been concentrated deep in the gas planet's interior as it formed and evolved four billion years ago. "There was a cosmo- chemical explanation and a meteorological explanation, and our latest analysis clearly favors the idea that the dry spots are a consequence of weather-related activity," Ingersoll said.
"Fifty miles below the cloud top."
#
"Galileo Finds Europa Has An Atmosphere",132,0,0,0
July 18, 1997
NASA's \JGalileo\j \Jspacecraft\j has found an \Jionosphere\j on Jupiter's moon Europa, an indication that the icy moon also has an atmosphere, \JGalileo\j scientists reported today.
"While this discovery does not relate to the question of possible life on Europa, it does show us there is a surface process occurring there, and Europa is not just some dead hunk of material," said lead investigator Dr. Arvydas Kliore of NASA's Jet Propulsion Laboratory, Pasadena, CA. Kliore reports his findings in the July 18 issue of Science magazine.
The \Jionosphere\j was detected through a series of six \Joccultation\j experiments performed during \JGalileo\j's encounters with Europa in December 1996 and February 1997. During \Joccultation\j, Europa was positioned between the \Jspacecraft\j and Earth, causing interruption in the radio signal.
Measurements of the \JGalileo\j radio signal received at the Deep Space Network stations in Goldstone, CA, and \JCanberra\j, \JAustralia\j, showed that the radio beam was refracted by a layer of electrons, or charged particles, in Europa's \Jionosphere\j.
An \Jionosphere\j is a layer of charged particles (ions and electrons) found in the upper levels of an atmosphere, created when gas molecules in the atmosphere are ionized. On Europa, this ionized layer can be caused either by the Sun's ultraviolet radiation or by energetic particles trapped in Jupiter's magnetic field, known as the magnetosphere.
Europa and the other Jovian satellites are immersed in this magnetosphere. "Most likely the charged particles in Jupiter's magnetosphere are hitting Europa's icy surface with great energy, knocking atoms of water molecules off the moon's surface," Kliore said,
Europa's \Jionosphere\j has a maximum density of 10,000 electrons per cubic centimeter, which is significantly lower than the average density of 20,000 to 250,000 electrons per cubic centimeter found in Jupiter's \Jionosphere\j. This indicates that Europa's \Jionosphere\j is tenuous; nonetheless, it is strong enough for scientists to infer the presence of an atmosphere.
The latest \JGalileo\j findings follow last year's observations by NASA's Hubble Space \JTelescope\j of oxygen emissions on Europa, a strong hint that an atmosphere might exist on that moon.
The existence of an \Jionosphere\j and, by inference, an atmosphere, on another Jovian moon, Io, was observed in 1973 during a radio \Joccultation\j conducted by NASA's Pioneer 10 \Jspacecraft\j and confirmed by recent \JGalileo\j occultations.
Io is believed to have an unusual atmosphere affected by sulfur dioxide spewing from the moon's volcanic vents. Kliore and his colleagues are currently studying two of Jupiter's other largest moons, Ganymede and Callisto, to determine whether they also have ionospheres and atmospheres.
"You could say an \Jionosphere\j or some kind of atmosphere has been found on most solar system bodies studied so far," said Kliore.
Participating with Kliore in the Europa radio \Joccultation\j experiments were Dr. David Hinson, professor at Stanford University, Stanford, CA; Dr. Michael Flasar of NASA's Goddard Space Flight Center, Greenbelt, MD; Dr. Andrew Nagy, professor at the University of \JMichigan\j, Ann Arbor, MI, and Dr. Thomas Cravens, professor at the University of Kansas, Lawrence, KS.
The \JGalileo\j mission is managed by JPL for NASA's Office of Space Science, Washington, DC. The \Jspacecraft\j entered the Jovian system on December 7, 1995, and its primary mission will end in November of this year. However, the mission has been extended for two more years so the craft can conduct an intensive study of Europa, with additional flybys of Callisto and Io, depending on \Jspacecraft\j health.
#
"Ganymede G1 & G2 Encounters",133,0,0,0
Voyager images are used to create a global view of Ganymede. The cut-out reveals the interior structure of this icy moon. This structure consists of four layers based on measurements of Ganymede's gravity field and theoretical analyses using Ganymede's known mass, size and density.
Ganymede's surface is rich in water ice and Voyager and \JGalileo\j images show features which are evidence of geological and tectonic disruption of the surface in the past. As with the Earth, these geological features reflect forces and processes deep within Ganymede's interior.
Based on geochemical and geophysical models, scientists expected Ganymede's interior to either consist of: a) an undifferentiated mixture of rock and ice or b) a differentiated structure with a large lunar sized 'core' of rock and possibly iron overlain by a deep layer of warm soft ice capped by a thin cold rigid ice crust.
Galileo's measurement of Ganymede's gravity field during its first and second encounters with the huge moon have basically confirmed the differentiated model and allowed scientists to estimate the size of these layers more accurately. In addition the data strongly suggest that a dense metallic core exists at the center of the rock core.
This metallic core suggests a greater degree of heating at sometime in Ganymede's past than had been proposed before and may be the source of Ganymede's magnetic field discovered by \JGalileo\j's space physics experiments.
Galileo's primary 24 month mission includes eleven orbits around Jupiter and will provide observations of Jupiter, its moons and its magnetosphere.
#
"Ganymede, Fractures in Transitional Terrain",134,0,0,0
This area of dark terrain on Jupiter's moon Ganymede lies near a transitional area between dark and bright terrain. The dark surface is cut by a pervasive network of fractures, which range in width from the limit of resolution up to 2.2 kilometers (1.4 miles).
Bright material is exposed in the walls of the chasms, and dark material fills the troughs. The impurities which darken the ice on the surface of dark terrain may be only a thin veneer over a brighter ice crust. Over time, these materials may be shed down steep slopes, where they collect in low areas.
The image is 68 by 54 kilometers (42 by 33 miles), and has a resolution of 190 meters (623 feet) per picture element (pixel). North is to the top. This image was obtained on September 6, 1996 by the Solid State Imaging (CCD) system aboard NASA's \JGalileo\j \Jspacecraft\j
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Galileo Calendar of Events",135,0,0,0
Public events, conferences, symposiums and workshops that members of the \JGalileo\j team will be participating in. For \JGalileo\j's upcoming encounters with Jupiter's satellites, see the Orbital Tour Highlights.
1997
òJan 07-10 - Conference on the Three Galileos: the Man, the \JSpacecraft\j, the \JTelescope\j, Padova, \JItaly\j
òJan 19-24 - Remote Sensing of Volcanoes on Earth and the Planets, Puerto Vallarta, Mexico
òJan 22 - \JGalileo\j Encounter with Europa Event, \JHonolulu\j, Hawaii
òFeb 17 - Carl Sagan Public Memorial, Pasadena, \JCalifornia\j
òSep 22-24 - Io During The \JGalileo\j Era Conference, Flagstaff, \JArizona\j
#
"Clementine Project",136,0,0,0
\BLaunch Date:\b 25 January 1994
\BDescription\b
The objective of the Clementine Mission was to test sensors and \Jspacecraft\j components under extended exposure to the space environment and to make scientific observations of the Moon and the near-Earth asteroid 1620 Geographos. The Clementine mission mapped most of the lunar surface at a number of resolutions and wavelengths from UV to IR. The \Jspacecraft\j was launched on January 25, 1994 at 16:34 and the nominal lunar mission lasted until the \Jspacecraft\j left lunar orbit on May 3. Clementine had five different imaging systems on-board. The UV/Visible camera had a filter wheel with six different filters, ranging from 415 nm to 1000 nm, and including a broad-band filter covering 400 to 950 nm.
The Near Infrared camera also had a six-filter wheel, ranging from 1100 nm to 2690 nm. The Longwave Infrared camera had a wavelength range of 8000 to 9500 nm. The Hi-Res imager had a broad-band filter from 400 to 800 nm and four other filters ranging from 415 to 750 nm. The Star Tracker camera was also used for imaging.
#
"Voyager Project Information",137,0,0,0
A multiple outer-planet flyby mission undertaken by NASA to make the first detailed exploration beyond Mars, and designed to take advantage of a rare (every 175 years) celestial alignment of Jupiter, Saturn, Uranus, and Neptune. Twin \Jspacecraft\j in the Mariner series were launched on Titan-Centaur vehicles in 1977. They flew through the Jovian system (March, July 1979) and, with a boost from Jupiter's gravity, flew on to Saturn (encounters November 1980, August 1981).
Following the Saturn flyby, Voyager 1's trajectory is taking it upward out of the \Jecliptic\j; Voyager 2 used Saturn's gravity to fly on to the historic encounters with Uranus (Jan 1986) and Neptune (August 1989). The \Jspacecraft\j, powered by \Jradioisotope\j thermo-electric generators, were built and are operated by NASA's Jet Propulsion Laboratory. They may send back data about the outermost reaches of the Solar System until well into the 21st-century.
\BMission Details
Launch Date:\b September 5, 1977 (Voyager 1); August 20, 1977 (Voyager 2)
\BDescription\b
The last two \Jspacecraft\j of NASA's Mariner series, Voyager 1 and 2 were the first in that series to be sent to explore the outer solar system. Preceded by the Pioneer 10 (1972 launch) and Pioneer 11 (1973 launch) missions, Voyager 1 and 2 were to make studies of Jupiter and Saturn, their satellites, and their magnetospheres as well as studies of the interplanetary medium. An option designed into the Voyager 2 trajectory, and ultimately exercised, would direct it toward Uranus and Neptune to perform similar studies.
Although launched sixteen days after Voyager 2, Voyager 1's trajectory was a faster path, arriving at Jupiter in March of 1979. Voyager 2 arrived about four months later in July 1979. Both \Jspacecraft\j were then directed on to Saturn with arrival times in November 1980 (Voyager 1) and August 1981 (Voyager 2). Voyager 2 was then diverted to the remaining gas giant, Uranus (January 1986) and Neptune (August 1989).
Data collected by Voyager 1 and 2 were not confined to the periods surrounding encounters with the outer gas giants, with the various fields and particles experiments and the ultraviolet spectrometer collecting data nearly continuously during the interplanetary cruise phases of the mission. Data collection continues as the recently renamed Voyager Interstellar Mission searches for the edge of the solar wind's influence (the heliopause) and exits the solar system.
\BSome Scientific Results of the Voyager Mission\b
A comprehensive list of the achievements of Voyager 1 and 2 would be so extensive that space doesn't permit. Here, then, are a (very) few results that would rank near the top of many such lists:
Discovery of the Uranian and Neptunian magnetospheres, both of them highly inclined and offset from the planets' rotational axes, suggesting their sources are significantly different from other magnetospheres.
The Voyagers found 22 new satellites: 3 at Jupiter, 3 at Saturn, 10 at Uranus, and 6 at Neptune.
Io was found to have active volcanism, the only solar system body other than the Earth to be so confirmed. Triton was found to have active geyser-like structures and an atmosphere.
Auroral zones were discovered at Jupiter, Saturn, and Neptune.
Jupiter was found to have rings. Saturn's rings were found to contain spokes in the B-ring and a braided structure in the F-ring. Two new rings were discovered at Uranus, and Neptune's rings, originally thought to be only ring arcs, were found to be complete, albeit composed of fine material.
At Neptune, originally thought to be too cold to support such atmospheric disturbances, large-scale storms (notably the Great Dark Spot) were discovered.
#
"Giotto Mission",138,0,0,0
\BLaunch Date/Time:\b 1985-07-02 at 11:23:13
\BDescription\b
The Giotto mission was designed to study \JComet\j P/Halley, and also to study \JComet\j P/Grigg-Skjellerup during its extended mission. The major objectives of the mission were to: (1) obtain color photographs of the nucleus; (2) determine the elemental and isotopic composition of volatile components in the cometary coma, particularly parent molecules; (3) characterize the physical and chemical processes that occur in the cometary atmosphere and \Jionosphere\j; (4) determine the elemental and isotopic composition of dust particles; (5) measure the total gas-production rate and dust flux and size/mass distribution and derive the dust-to-gas ratio; and (6) investigate the macroscopic systems of plasma flows resulting from the cometary-solar wind interaction.
The \Jspacecraft\j encountered the \Jcomet\j on March 13, 1986, at a distance of 0.89 AU from the sun and 0.98 AU from the Earth and an angle of 107 degrees from the comet-sun line. The \Jspacecraft\j was based as much as possible on the ESA-GEOS \Jspacecraft\j and was spin stabilized with a rate of 15 rpm. During the encounter with Halley's \Jcomet\j, the spin axis was aligned with the relative velocity vector. The 1.5m dish antenna, operating in the X-band, was inclined and despun in order to point at the Earth (44 degrees with respect to the velocity vector).
The goal was to come within 500 km of Halley's \Jcomet\j at closest encounter. The actual closest approach was measured at 596 km. The \Jspacecraft\j had a dust shield consisting of a front sheet of Aluminum (1 mm thick) and a 12 mm Kelvar near sheet separated by 25 cm, which could withstand impacts of particles up to 0.1 g. The scientific payload was comprised of ten hardware experiments: a narrow-angle camera, three mass spectrometers for neutrals, ions and dust, various dust detectors, a photo polarimeter and a set of plasma experiments. All experiments performed well and returned a wealth of new scientific results, of which perhaps the most important was the clear identification of the cometary nucleus.
Fourteen seconds before closest approach, Giotto was hit by a 'large' dust particle. The impact caused the \Jspacecraft\j angular momentum vector to shift by 0.9 degrees. Scientific data were received intermittently for the next 32 minutes. Some experiment sensors suffered damage during this 32-minute interval.
Other experiments (the camera baffle and deflecting mirror, the dust detector sensors on the front sheet of the bumper shield, and most experiment apertures) were exposed to dust particles regardless of the accident and also suffered damage. Many of the sensors survived the encounter with little or no damage. Questionable or partially damaged sensors included the camera (later proved to not be functional) and one of the plasma analyzers (RPA). Inoperable experiments included the neutral and ion mass spectrometers and one sensor each on the dust detector and the other plasma analyzer (JPA).
During the Giotto extended mission, the \Jspacecraft\j successfully encountered \JComet\j P/Grigg-Skjellerup on July 10, 1992. The closest approach was approximately 200 km. The heliocentric distance of the \Jspacecraft\j was 1.01 AU, and the geocentric distance, 1.43 AU at the time of the encounter. The payload was switched-on in the evening of July 9. Eight experiments were operated and provided a surprising wealth of exciting data.
The Johnstone Plasma Analyser detected the first presence of cometary ions 600,000 km from the nucleus at 12 hours before the closest approach. The Dust Impact Detectors reported the first impact of a fairly large particle at 15:30:56. Bow shocks/waves and acceleration regions were also detected. After the P/Grigg-Skjellerup encounter, the \Jspacecraft\j was retargeted for a possible 1999 encounter pending the existence of sufficient fuel and funding for ground operations support
#
"Hubble Space Telescope (HST)",139,0,0,0
\JHubble Space Telescope (HST) Summary\j
\JHubble's Look At Mars, Conditions For Pathfinder Landing\j
\JHubble's Sharpest View Of Mars\j
\JHubble Captures A Full Rotation Of Mars\j
\JHubble Monitors Weather On Neighboring Planets\j
\JHubble Finds Intergalactic Stars\j
\JBlack Holes Dwell In Most Galaxies, Hubble Census\j
\JSupernova Blast Begins Taking Shape\j
\JProto-Planetary Disks Destruction Explained\j
\JLagoon Nebula, Giant 'Twisters' and Star Wisps \j
\JHubble Space Telescope Check-Out Finds Successes, Concerns\j
\JHubble Images Of Comet Hale-Bopp\j
\JHubble Tracks Fading Optical Counterpart Of Gamma-Ray Burst\j
\JHubble's Upgrades Show Birth and Death Of Stars\j
\JHubble Captures Volcanic Eruption Plume From Io\j
#
"Hubble Space Telescope (HST) Summary",140,0,0,0
\BLaunch Date:\b 1990-04-25
\BDescription\b
The Hubble Space \JTelescope\j (HST) was the first and flagship mission of NASA's Great Observatories program. Designed to complement the wave length capabilities of the other \Jspacecraft\j in the program (CGRO, AXAF, and SIRTF), HST was a 2.4 m, f/24 Ritchey-Chretien \Jtelescope\j capable of performing observations in the visible, near-ultraviolet, and near-infrared(1150 A to 1 mm).
Placed into a low-earth orbit by the space shuttle, HST was designed to be modular so that on subsequent shuttle missions it could be recovered, have faulty or obsolete parts replaced with new and/or improved instruments, and be re-released. HST was roughly cylindrical in shape, 13.1 m end-to-end and 4.3 m in diameter at its widest point.
HST used an elaborate scheme for attitude control to improve the stability of the \Jspacecraft\j during observations. Maneuvering was performed by four of six gyros, or reaction wheels. Pointing could be maintained in this mode (coarse track) or the Fine Guidance Sensors (FGSs) could be used to lock onto guide stars (fine lock) to reduce the \Jspacecraft\j drift and increase the pointing accuracy.
Power to the two on-board computers and the scientific instruments was provided by two 2.4 x 12.1 m solar panels. The power generated by the arrays was also used to charge six nickel-hydrogen batteries which provided power to the \Jspacecraft\j during the roughly 25 minutes per orbit in which HST was within the Earth's shadow.
Communications with the satellite were maintained with the TDRS satellites. Observations taken during the time when neither TDRS was visible from the \Jspacecraft\j were recorded on tape recorder and dumped during periods of visibility. The \Jspacecraft\j also supported real-time interactions with the ground system during times of TDRS visibility, enabling observers to make small offsets in the \Jspacecraft\j pointing to perform their observations. HST was the first scientific \Jspacecraft\j designed to utilize the full capabilities of TDRSS, communicating over either multiple-access or single-access channels at any of the supported transmission rates.
HST was operated in three distinct phases. During the first phase of the mission (Orbital Verification or OV), responsibility for the \Jspacecraft\j was given to Marshall Space Center. OV consisted of an extended, eight-month checkout of the \Jspacecraft\j, including test of the on-board computers, pointing control system, solar arrays, etc. This phase was followed by the Science Verification (SV) phase, lasting nearly another year, during which each of the six science instruments was tested to verify their capabilities and set limits on their safe operations during the remainder of the mission.
Responsibility for the \Jspacecraft\j during SV was given to Goddard Space Flight Center. The last phase of the mission, known as the General Observer (GO) phase, was planned to last from the end of SV through the end of the mission and was the responsibility of the Space \JTelescope\j Science Institute. General observations were phased in gradually, however, during the SV phase because the OV and SV portions of the mission were considerably longer than expected prior to deployment.
The mission was troubled soon after launch by the discovery that the primary mirror was spherically aberrated. In addition, problems with the solar panels flexing as the \Jspacecraft\j passed from the Earth's shadow into sunlight caused problems with the pointing stability. Steps were taken to correct these problems, including replacement of the solar panels, replacement of the Wide Field and Planetary Camera with a second-generation version with built-in corrective \Joptics\j, and replacement of the High-Speed Photometer with COSTAR (Corrective \JOptics\j Space \JTelescope\j Axial Replacement) to correct the aberration for the remaining instruments.
#
"Hubble's Look At Mars, Conditions For Pathfinder Landing",141,0,0,0
July 1, 1997
Hubble Space \JTelescope\j pictures of Mars, taken on June 27 in preparation for the July 4 landing of the Pathfinder \Jspacecraft\j, show a dust storm churning through the deep canyons of Valles Marineris, just 600 miles (1000 km) south of the Pathfinder \Jspacecraft\j landing site.
"Unless the dust storm were to evolve into a massive, global event, its effects on the Pathfinder mission should be minimal, says Steve Lee of the University of \JColorado\j in Boulder, \JColorado\j. "This is something we did not expect to see".
The Hubble astronomers also report the presence of patchy cirrus clouds over the landing site and very thick clouds to the north. Because there are so many clouds (related to low temperatures in the atmosphere causing water vapor to freeze), the dust will probably stay confined to the canyons, they conclude.
If dust rises to the elevations where the water-ice clouds form, ice condenses on dust grains and the heavier ice/dust particles quickly fall back out of the atmosphere. Though the dust could extend at low altitudes over the landing site, researchers say current prevailing winds should not take the dust northward.
"If dust diffuses to the landing site, the sky could turn out to be pink like that seen by Viking," says Philip James of the University of Toledo. Otherwise, Pathfinder will likely show blue sky with bright clouds."
#
"Hubble's Sharpest View Of Mars",142,0,0,0
March 20, 1997
The sharpest view of Mars ever taken from Earth was obtained by the recently refurbished NASA Hubble Space \JTelescope\j (HST). This stunning portrait was taken with the HST Wide Field Planetary Camera-2 (WFPC2) on March 10, 1997, just before Mars opposition, when the red planet made one of its closest passes to the Earth (about 60 million miles or 100 million km).
At this distance, a single picture element (pixel) in WFPC2's Planetary Camera spans 13 miles (22 km) on the Martian surface.
The Martian north pole is at the top (near the center of the bright polar cap) and East is to the right. The center of the disk is at about 23 degrees north latitude, and the central longitude is near 305 degrees.
This view of Mars was taken on the last day of Martian spring in the northern hemisphere (just before summer solstice). It clearly shows familiar bright and dark markings known to astronomers for more than a century.
The annual north polar carbon dioxide frost (dry ice) cap is rapidly sublimating (evaporating from solid to gas), revealing the much smaller permanent water ice cap, along with a few nearby detached regions of surface frost. The receding polar cap also reveals the dark, circular sea' of sand dunes that surrounds the north pole (Olympia Planitia).
Other prominent features in this hemisphere include Syrtis Major Planitia, the large dark feature seen just below the center of the disk. The giant impact basin Hellas (near the bottom of the disk) is shrouded in bright water ice clouds.
Water ice clouds also cover several great volcanos in the Elysium region near the eastern edge of the planet (right). A diffuse water ice haze covers much of the Martian equatorial region as well.
The WFPC2 was used to monitor dust storm activity to support the Mars Pathfinder and Mars Global Surveyor Orbiter Missions, which are currently en route to Mars. Airborne dust is most easily seen in WFPC2's red and near-infrared images.
Hubble's "weather report" from these images in invaluable for Mars Pathfinder, which is scheduled for a July 4 landing. Fortunately, these images show no evidence for large-scale dust storm activity, which plagued a previous Mars mission in the early 1970s.
The WFPC2 was used to observe Mars in nine different colors spanning the ultraviolet to the near infrared. The specific colors were chosen to clearly discriminate between airborne dust, ice clouds, and prominent Martian surface features. This picture was created by combining images taken in blue (433 nm), green (554 nm), and red (763 nm) colored filters.
#
"Hubble Captures A Full Rotation Of Mars",143,0,0,0
Pictures of the planet Mars taken with the recently refurbished NASA Hubble Space \JTelescope\j (HST) will provide the most detailed global view of the red planet ever obtained from Earth.
The images were taken by HST's Wide Field Planetary Camera-2 on March 10, 1997, just before Mars opposition, when the red planet made one of its closest to the Earth (about 60 million miles or 100 million km).
These pictures were taken during three HST orbits that were separated by about six hours. This timing was chosen so that Mars, with its 24-hour 39-minute day, would rotate about 90 degrees between orbits. This imaging sequence therefore covers most of the Martian surface. These observations will be combined with others planned for March 30 to provide complete coverage.
During each orbit, Mars was observed in nine different colors spanning the ultraviolet to the near infrared. The specific colors were chosen to clearly discriminate between airborne dust, ice clouds, and prominent Martian surface features.
The color picture shown here was created by combining images taken in blue (433 nm), green (554 nm), and red (763 nm) colored filters. The Martian north pole is at the top (near the center of the bright polar cap) and East is to the right. The center of the disk is at about 23 degrees north latitude, and the central longitudes are near 160, 210, and 305 degrees.
These images show the planet on the last day of Martian spring in the northern hemisphere (just before summer solstice). The annual north polar carbon dioxide frost (dry ice) cap is rapidly sublimating, revealing the much smaller permanent water ice cap.
This polar cap remnant, along with a few nearby detached regions of surface frost are most obvious in pictures taken through ultraviolet, blue, and green filters. These filters also show numerous bright water ice clouds.
The brightest clouds are in the vicinity of the giant volcanos on the Tharsis Plateau (to right of center on left image), and in the giant impact basin, Hellas (near bottom of right-hand image), but a diffuse haze covers much of the Martian tropics as well.
The familiar bright and dark markings on the Martian surface are most obvious in images taken through red and near-infrared filters. These images clearly reveal the large, dark, circular "sea" of sand dunes (Olympia Planitia) that surrounds the north pole, as well a number of other familiar features, including the giant Tharsis volcanos.
The 16-mile (27 km) high Olympus Mons is near the center of the left-hand image, with Arsia, Povonis, and Ascraeus Mons forming a south-west to north-east line just to its right. The \Jvolcano\j, Elysium Mons is near the center of the middle image. The prominent dark feature just below the center on the disk on the rightmost image is Syrtis Major Planitia.
Hubble is being used to monitor dust storm activity to support the Mars Pathfinder and Mars Global Surveyor Orbiter Missions, which are currently en route to Mars. Airborne dust is most easily seen in WFPC2's red and near-infrared images.
Weather reports derived from these observations are particularly valuable for Mars Pathfinder, which is scheduled for a July 4, 1997 landing on the red planet. A preliminary analysis of these HST data reveals enhanced dust activity over the dark Vastitas Borealis region in the northern hemisphere, and over the Noachis Terra and Terra Tyrrhena regions just south of the Martian equator.
There is also evidence for airborne dust and ice clouds in the Hellas basin. However, these images show no evidence for large-scale dust storm activity.
#
"Hubble Monitors Weather On Neighboring Planets",144,0,0,0
MARS: A COOLER, CLEARER WORLD
Four years, (or two Mars years') worth of Hubble observations show that the Red \JPlanet\j's climate has changed since the mid-1970's. "The Hubble results show us that the Viking years are not the rule, and perhaps not typical. Our early assumptions about the Martian climate were wrong," said Philip James of the University of Toledo.
"There has been a global drop in temperature. The planet is cooler and the atmosphere clearer than seen before," said Steven Lee of the University of \JColorado\j in Boulder. "This shows the need for continuous monitoring of Mars. Space probes provided a close-up look, but it's difficult to extrapolate to long-term conditions based upon these brief encounters."
The researchers attribute the cooling of the Martian atmosphere to diminished dust storm activity, which was rampant when a pair of NASA Viking orbiter and lander \Jspacecraft\j arrived at Mars in 1976. Two major dust storms occurred during the first year of the Viking visits, which left fine dust particles suspended in the Martian atmosphere for longer than normal.
Warmed by the Sun, these dust particles (some only a \Jmicron\j in diameter, about the size of smoke particles) are the primary source of heat in the Martian atmosphere.
"Hubble is showing that our early understanding based on these visits is wrong. We just happened to visit Mars when it was dusty, and now the dust has settled out," Lee said. "We are going to have to look at Mars for many years to truly understand the workings of the climate," said Todd Clancy, of the Space Science Institute, Boulder, \JColorado\j.
Knowledge about the Martian climate has been limited by the fact that ground-based telescopes can only see weather details when Earth and Mars are closest -- an event called opposition -- that happens only once every two years. Though Hubble has observed Mars only for four years, the observations are equivalent to 15 years of ground- based observing because Hubble can follow seasonal changes through most of Mars' orbit.
Though the Mariner and Viking series of flyby, orbiter and lander \Jspacecraft\j that visited Mars in the late 60's and 70's provided a close-up look at Martian weather, these were snapshots of the planet's complex climate.
Hubble provides the advantage of a global view - much like the satellites that monitor Earth's weather, and can follow martian seasonal changes over many years. When Mars is closest to Earth, Hubble returns near-weather satellite resolution.
MARS -- NO LACK OF OZONE
Although there has been concern about a lack of ozone (a form of molecular oxygen created by the effects of sunlight on an atmosphere), dubbed the "ozone hole" over Earth's poles, there are no ozone holes on Mars.
By contrast, the planet has a surplus of ozone over its northern polar cap, as first identified by the Mariner 9 \Jspacecraft\j in 1971. (However the Martian atmosphere is different enough from Earth's that few parallels can be drawn about processes controlling the production and destruction of ozone.)
Hubble's ultraviolet sensitivity is ideal for monitoring ozone levels on a global scale. The Martian ozone is yet another indication the planet has grown drier, because the water in the atmosphere that normally destroys ozone has frozen-out to become ice-crystal clouds.
Spectroscopic observations made with the Faint Object Spectrograph (FOS) show that ozone now extends down from Mars' north pole to mid and lower latitudes. However, the Martian atmosphere is so thin, even this added ozone would offer future human explorers little protection from the Sun's harmful ultraviolet rays.
SEASONS ON MARS
The fourth planet from the Sun, Mars is one of the most intensely scrutinized worlds because of its Earth-like characteristics. Mars is tilted on its axis by about the same amount Earth is, hence Mars goes through seasonal changes.
However, because Mars' atmosphere is much thinner than Earth's, it is far more sensitive to minor changes in the amount of light and heat received from the Sun. This is intensified by Mars' orbit that is more elliptical than Earth's, so it's range of distance from the Sun is greater during the Martian year. Mars is now so distant, the sun is nearly 25% dimmer than average.
This chills Mars' average temperature by 36 degrees Fahrenheit (20 degrees Kelvin). At these cold temperatures, water vapor at low altitudes freezes out to form ice-crystal clouds now seen in abundance by Hubble.
"Clouds weren't considered to be very important to the Martian climate during the Viking visits because they were so scarce," says Clancy. "Now we can see where they may play a role in transporting water between the north and south poles during the Martian year."
Seasonal winds also play a major role is transporting dust across Mars' surface, and rapidly changing the appearance of a region. This gave early astronomers the misperception that Mars' shifting surface color was evidence of vegetation following a season cycle.
As clearly seen in the Hubble images, past dust storms in Mars' southern hemisphere have scoured the plains of fine light dust and transported the dust northward. This leaves behind a relatively coarser, less reflective sand in the southern hemisphere.
VENUS: NO EVIDENCE FOR NEW VOLCANIC ERUPTIONS
Hubble spectroscopic observations of Venus taken with the Goddard High Resolution Spectrograph provide a new opportunity to look for evidence of volcanic activity on the planet's surface. Though radar maps of the Venusian surface taken by the Magellan orbiter revealed numerous volcanoes, Magellan did not find clear cut evidence for active volcanoes.
Hubble can trace atmospheric changes that might be driven by volcanism. An abundance of sulfur dioxide in the atmosphere could be a tell-tale sign of an active volcanos. Sulfur dioxide was first detected by the Venus Pioneer probe in the late 1970s and has been declining ever since.
The Hubble observations show that sulfur dioxide levels continue to decline. This means there is no evidence for the recurrence of large scale volcanic eruptions in the last few years.
Ejected high into Venus' murky atmosphere, this sulfur dioxide is broken apart by sunlight to make an acid rain of concentrated sulfuric acid. This is similar to what happens on Earth above coal-burning power plants - but on a much larger and more intense scale.
FUTURE PLANS
More Hubble observations of Mars and Venus are critical to planning visits by future space probes. In particular, both robotic and human missions to Mars will need to be targeted for times during the Martian year when there is a minimal chance of getting caught in a dust storm.
Knowing whether the atmosphere is relatively hot or cold is crucial to planning aerobraking maneuvers, where \Jspacecraft\j use the aerodynamic drag of an atmosphere to slow down and enter an orbit around the planet.
This reduces the amount of propellant needed for the journey. "If the atmosphere is more extended than expected the added friction could burn up an aerobraking \Jspacecraft\j, just as Earth's atmosphere incinerates infalling meteors," says James.
Ultimately, knowing the Martian climate will be an fundamental prerequisite for any future plans to establish a permanent human outpost on the Red \JPlanet\j.
#
"Hubble Finds Intergalactic Stars",145,0,0,0
January 14, 1997
NASA's Hubble Space \JTelescope\j has found a long sought population of "stellar outcasts" -- stars tossed out of their home galaxy into the dark emptiness of intergalactic space. This is the first time stars have been found more than 300,000 light-years (three Milky Way diameters) from the nearest big galaxy.
The isolated stars dwell in the Virgo cluster of galaxies, about 60 million light-years away. The results suggest this population of "lone stars" accounts for 10 percent of the Virgo cluster's mass, or 1 trillion Sun-like stars adrift among the 2,500 galaxies in Virgo. "Our discovery provides a new tool for studying clusters of galaxies," says Harry Ferguson of Space \JTelescope\j Science Institute in Baltimore, Maryland.
The distribution of the stars in Virgo could help astronomers probe the distribution of dark matter in the cluster. (Dark matter is an unknown type of matter that accounts for most of the mass of the universe.) Another possible spinoff is that the stars detected, which are the brightest members of the red giant class, may serve as "standard candles" (stars that can be used for calibrating distances), providing an independent method for measuring cosmological distances to Virgo.
Such measurements are key to estimating the expansion rate and age of the universe. These results are being presented at the 189th Meeting of the American Astronomical Society in Toronto, Canada, by Ferguson and co-investigators Nial Tanvir (University of Cambridge, Cambridge, United Kingdom) and Ted von Hippel (University of Wisconsin).
Intergalactic stars have been predicted to exist as a result of galaxy interactions and mergers early in a galaxy cluster's history. These close encounters should have ripped stars out of their home galaxies and tossed them into intergalactic space, where they drift free of the gravitational influence of any single galaxy.
It was predicted that the stars should appear as a diffuse excess of light in Virgo, and there have previously been observations from ground-based telescopes that report just such an excess. "However, there are large uncertainties in the ground-based measurements, and it is not clear whether the diffuse light originates from galaxies too faint to detect individually or from a more uniform sea of stars," says Ferguson.
The accidental discovery in 1996 of planetary nebulae (stellar remnants) in Virgo, which are far removed from any galaxy, offered additional evidence that such an intergalactic population really exists.
The Hubble astronomers found the background stars by taking an exposure of a "blank" portion of sky in Virgo. The position is in the vicinity of the giant elliptical galaxy M87 in the center of Virgo, but far enough from the galaxy for the stars not to be members of M87's halo. The Virgo field was compared to the Hubble Deep Field (HDF) image which represents a region of sky devoid of any nearby galaxy cluster. With the HDF serving as the control, the astronomers counted approximately 600 sources down to 27.8 magnitude.
The stars are bright red giants -- stars late in their lives. Presumably there are many fainter stars -- perhaps as many as 10 million -- in the same field but are below Hubble's sensitivity.
"These stars are truly intergalactic because they are so isolated their motion is probably governed by the gravitational field of the cluster as a whole, rather than the pull of any one galaxy," says Ferguson. The Space \JTelescope\j Imaging Spectrograph (STIS) and Near Infrared Camera and Multi-Object Spectrometer (NICMOS) planned for installation on Hubble this February will be used to help understand the history of the stellar outcasts.
Comparison of heavy element abundances in the "loner" stars and in the Virgo galaxies should help to uncover whether the stars wandered off from the outskirts of galaxies that still exist, are the remnants of galaxies that were completely disrupted, or were somehow formed in the dark reaches of intergalactic space.
The Space \JTelescope\j Science Institute is operated by the Association of Universities for Research in \JAstronomy\j, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, Maryland. The Hubble Space \JTelescope\j is a project of international cooperation between NASA and the European Space Agency (ESA).
#
"Black Holes Dwell In Most Galaxies, Hubble Census",146,0,0,0
January 13, 1997
Announcing the discovery of three black holes in three normal galaxies, an international team of astronomers suggests nearly all galaxies may harbor supermassive black holes which once powered quasars (extremely luminous nuclei of galaxies), but are now quiescent.
This conclusion is based on a \Jcensus\j of 27 nearby galaxies carried out by NASA's Hubble Space \JTelescope\j and ground-based telescopes in Hawaii, which are being used to conduct a spectroscopic and photometric survey of galaxies to find black holes which have consumed the mass of millions of Sun-like stars.
The findings, being presented today at the 189th Meeting of the American Astronomical Society in Toronto, Canada, should provide insights into the origin and evolution of galaxies, as well as clarify the role of quasars in galaxy evolution.
The key results are:
Supermassive black holes are so common, nearly every large galaxy has one.
A black hole's mass is proportional to the mass of the host galaxy, so that, for example, a galaxy twice as massive as another would have a black hole that is also twice as massive. This discovery suggests that the growth of the black hole is linked to the formation of the galaxy in which it is located.
The number and masses of the black holes found are consistent with what would have been required to power the quasars.
"We believe we are looking at "fossil quasars" and that most galaxies at one time burned brightly as a quasar," says team leader Doug Richstone of the University of \JMichigan\j, Ann Arbor, \JMichigan\j. These conclusions are consistent with previous Hubble Space \JTelescope\j observations showing quasars dwelling in a variety of galaxies, from isolated normal-looking galaxies to colliding pairs.
Two of the black holes "weigh in" at 50 million and 100 million solar masses in the cores of galaxies NGC 3379 (also known as M105) and NGC 3377 respectively. These galaxies are in the "Leo Spur", a nearby group of galaxies about 32 million light-years away and roughly in the direction of the Virgo cluster.
Located 50 million light-years away in the Virgo cluster, NGC 4486B possesses a 500-million solar mass black hole. It is a small satellite of the galaxy M87, a very bright galaxy in the Virgo cluster. M87 has an active nucleus and is known to have a black hole of about 2 billion solar masses.
Though several groups have previously found massive black holes dwelling in galaxies the size of our Milky Way or larger, these new results suggest smaller galaxies have lower-mass black holes, below Hubble's detection limit. The survey shows the black hole's mass is proportional to the host galaxy's mass. Like shoe sizes on adults, the bigger the galaxy, the larger the black hole.
It remains a challenging puzzle as to why black holes are so abundant, or why they should be proportional to a galaxy's mass. One idea, supported by previous Hubble observations, is that galaxies formed out of smaller "building blocks" consisting of star clusters. A massive "seed" black hole may have been present in each of these protogalaxies. The larger number of building blocks needed to merge and form very luminous galaxies would naturally have provided more seed black holes to coalesce into a single, massive black hole residing in a galaxy's nucleus.
An alternative model is that galaxies start at some early epoch with a modest black hole (not necessarily approaching the masses discussed here), but that the black hole consumes some fixed fraction of the total gas shed by the stars in the galaxy during their normal evolution. If that fraction is around 1 percent, the black holes could easily weigh as much as they do now, and would naturally track the current luminosity of the galaxy.
Critical ground-based observations to identify candidates were obtained for all three of these objects by John Kormendy with the Canada-France-Hawaii \JTelescope\j (CFHT) on Mauna Kea, Hawaii. The NGC 4486b black hole detection was also based on CFHT spectra.
Hubble's high resolution then allowed the team to peer deep into the cores of the galaxies with extraordinary resolution unavailable from ground-based telescopes, and measure velocities of stars orbiting the black hole.
A sharp rise in velocity means that a great deal of matter is locked away in the galaxy's core, creating a powerful gravitational field that accelerates nearby stars. The team is confident their statistical search technique has allowed them to pinpoint all the black holes they expect to see, above a certain mass limit. "However, our result is complicated by the fact that the observational data for the galaxies are not of equal quality, and that the galaxies are at different distances," says Richstone.
One of the features of the February 1997 servicing mission to the Hubble will be the installation of the Space \JTelescope\j Imaging Spectrograph (STIS). This spectrograph will greatly increase the efficiency of projects, such as this black hole \Jcensus\j, that require spectra of several nearby positions in a single object.
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"Supernova Blast Begins Taking Shape",147,0,0,0
January 14, 1997
Though the brightest \Jsupernova\j in four centuries lit up the southern sky almost exactly 10 years ago on Feb. 23, 1987, astronomers have waited a decade for the ballooning fireball to become large enough -- about one-sixth of a light-year -- to be resolved from Earth's orbit with NASA's Hubble Space \JTelescope\j (HST).
Astronomers announced today that a close monitoring of the \Jsupernova\j (designated SN1987A) with HST's sharp view has resolved a one-tenth light-year long dumbbell-shaped structure consisting of two blobs of debris expanding apart at nearly 6 million miles per hour from each other.
"This structure is a bit of a surprise," says Jason Pun of Goddard Space Flight Center, Greenbelt, Maryland. "This is the first time we can see the \Jgeometry\j of the explosion and relate it to the \Jgeometry\j of the large glowing ring system around the \Jsupernova\j, which has an hourglass shape. The images may yield important clues to the dynamics of the \Jsupernova\j explosion and the structure of the progenitor star."
Pun says the dim area between the blobs may be related to the equatorial belt of material seen around the \Jsupernova\j that existed before the star exploded. The ring was illuminated by the \Jsupernova\j in 1987 during the explosion and has been slowly fading since then.
Ever since the star self-destructed in 1987 astronomers realized that it offered a once-in-a-lifetime possibility, because of its close proximity to Earth, to obtain images of the explosion at various stages and look for any changes in the shape and dynamics.
The latest findings are the result of the \JSupernova\j Intensive Study collaboration, headed by Professor Robert Kirshner of the Harvard-Smithsonian Center for Astrophysics in Cambridge, \JMassachusetts\j. Images of SN1987A were taken in September 1994, March 1995, and February 1996 with the Wide Field and Planetary Camera 2 (WFPC2). These results are being presented today at the 189th Meeting of the American Astronomical Society in Toronto, Canada, by co-investigator Pun.
The explosion of the \Jsupernova\j debris appears to be perpendicular to the plane of the inner ring. This suggests that whatever properties that the pre-supernova star has, such as rotation or the existence of a companion star, that is responsible for the formation of the inner ring, may also have influenced the dynamics of the explosion.
The explosion was triggered 10 years ago when the collapse of the star's core sent a blast wave of neutrinos which heated the star's inner layers to 10 billion degrees Fahrenheit. This triggered a shockwave which then ripped the star apart and sent the debris hurtling into space. The fireball has since cooled down (to a few hundred degrees Fahrenheit) and the debris is now heated by nuclear energy from the decay of radioactive nuclei produced in the explosion.
The Space \JTelescope\j Imaging Spectrograph (STIS) and Near Infrared Camera and Multi-Object Spectrometer (NICMOS), planned for installation on Hubble this February, will be used to obtain a spatially resolved velocity map of the debris, providing information on the physical conditions of the two blobs.
The debris is expected to collide with the inner ring as early as the year 2002. This will light up all of the dark nebulosity surrounding the \Jsupernova\j, providing new clues to the nature and evolution of the stellar explosion.
Theoretical models, coupled with NASA Hubble Space \JTelescope\j observations of the Trapezium cluster in the Orion nebula, suggest that disks around young cluster stars may not survive long enough for planets to form within them. This implies that there are certain hostile environments in star-forming regions that may inhibit planet formation.
The findings, presented by an international collaboration of astronomers today at the meeting of the American Astronomical Society in Toronto, Canada, explain the destruction of circumstellar disks in Orion's Trapezium, a star cluster at the very center of the nebula.
The report is being presented by Doug Johnstone, a Natural Sciences and \JEngineering\j Research Council (NSERC) Post-Doctoral Fellow at the Canadian Institute for Theoretical Astrophysics, University of Toronto. "For the first time we have a complete evolutionary picture for the stunning objects observed in the Trapezium," says Johnstone.
Their work provides an innovative technique for analyzing circumstellar disks, determining disk masses, and constraining the gestation period for planet embryos around stars in dense clusters, say researchers. "This may help produce a consensus within the star and planet formation community on standard disk properties," says Johnstone.
The team's results show the disks of dust and gas, which can be several billion kilometers across, are initially similar to the disk which is believed to have formed the planets in our own Solar system, but quickly evaporate in the glare of bright massive neighboring stars in the Trapezium. Radiation from these stars photoionizes, or heats and disperses, the cold gas. Within 1 million years the disk is eroded, a time scale shorter than the 1 to 10 million years it would take for planets to form according to current models.
"The theory of disk destruction predicts most efficient destruction at large distances from the embedded, central star. Near the center of the disk, perhaps even at the same distance as the Earth is from the Sun, the remnant disk might survive long enough to form planet embryos," says Johnstone. "Without a more detailed understanding of planet formation it is not possible to predict the future of these disks, but standard models based on our own Solar system suggest that giant planets like Jupiter and Saturn, at comparable distances from their central star, would be ruled out."
Using the Planetary Camera on the Hubble Space \JTelescope\j (HST), Johnstone's collaborators John Bally and Dave Devine of the University of \JColorado\j, and Ralph Sutherland of the Australian National University observed the Trapezium, a young, million-year-old star-forming region just below the belt of the \Jconstellation\j Orion.
Located nearby, only fifteen hundred light years away, the Trapezium region is the closest star formation site containing both Sun-like stars and stars much more massive than the Sun. While ground-based observations have hinted at extended structures surrounding the Sun-like stars, HST has produced images with incredible detail revealing that these stars are embedded in circumstellar disks and surrounded by diffuse hot ionized gas.
The idea that these young stars are embedded in evaporating disks was first proposed by Ed Churchwell of the University of \JWisconsin\j, based on ground based radio-wave observations. Recognizing that these disks might be the birth sites for planets, C. Robert O'Dell of Rice University in Houston, \JTexas\j confirmed the disk hypothesis using HST, and named the objects "proplyds" as an \Jacronym\j for proto-planetary disk.
"However, until our work there was no satisfactory model detailing the origin of the diffuse cloud of hot gas observed around each of these Sun-like stars," says Johnstone. Johnstone, along with David Hollenbach and Herbert Stoerzer of NASA Ames Research Center in \JCalifornia\j, have developed theoretical models describing the destructive effect of high energy radiation on disks, combining the results with Hubble observations by John Bally's team to produce a coherent evolutionary picture.
They report that the disk surface is initially heated to temperatures in excess of 1000 \JCelsius\j by the impinging radiation, evaporating the surface layer much like steam evaporates from the surface of boiling water.
As this material flows away from the central star and disk, higher energy photons ionize the gas, heating it to temperatures reaching 10000 degrees \JCelsius\j and in the process producing the nebulous glow seen in the images. "We are witnessing the destructive event through the illumination of the evaporated material" according to Johnstone.
Evaporation of the circumstellar disk erodes approximately three moon masses of material per year according to the theoretical model, a number which is verified by the HST observations. The exact \Jevaporation\j rate from the circumstellar disk is directly related to the size of the disk and thus, as the disk evaporates and shrinks, the erosion rate decreases.
By using this knowledge, and fitting the \Jevaporation\j model to the HST observations, the collaboration shows that the original circumstellar disks surrounding stars in the Trapezium were similar in appearance to disks around young stars in other systems, and more importantly to the hypothesized Solar proto-planetary disk from which our own nine planets formed.
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"Lagoon Nebula, Giant 'Twisters' and Star Wisps",149,0,0,0
January 22, 1997
This NASA Hubble Space \JTelescope\j (HST) image reveals a pair of one-half light-year long interstellar "twisters" -- eerie funnels and twisted-rope structures (upper left) -- in the heart of the Lagoon Nebula (Messier 8) which lies 5,000 light-years away in the direction of the \Jconstellation\j \JSagittarius\j.
The central hot star, O Herschel 36 (upper left), is the primary source of the ionizing radiation for the brightest region in the nebula, called the Hourglass. Other hot stars, also present in the nebula, are ionizing the extended optical nebulosity. The ionizing radiation induces photo-evaporation of the surfaces of the clouds (seen as a blue "mist" at the right of the image), and drives away violent stellar winds tearing into the cool clouds.
Analogous to the spectacular phenomena of Earth tornadoes, the large difference in temperature between the hot surface and cold interior of the clouds, combined with the pressure of starlight, may produce strong horizontal shear to twist the clouds into their tornado-like appearance. Though the spiral shapes suggest the clouds are "twisting", future observations will be needed, perhaps with Hubble's next generation instruments, with the spectroscopic capabilities of the Space \JTelescope\j Imaging Spectrograph (STIS) or the Near Infrared Camera and Multi-Object Spectrometer (NICMOS), to actually measure velocities.
This Hubble picture reveals a variety of small scale structures in the interstellar medium, small dark clouds called Bok globules, bow shocks around stars, ionized wisps, rings, knots and jets.
The Lagoon Nebula and nebulae in other galaxies are sites where new stars are being born from dusty molecular clouds. These regions are the "space laboratories" for the astronomers to study how stars form and the interactions between the winds from stars and the gas nearby. By studying the wealth of data revealed by HST, astronomers will understand better how stars form in the nebulae.
These color-coded images are the combination of individual exposures taken in July and September, 1995 with Hubble's Wide Field and Planetary Camera 2 (WFPC2) through three narrow-band filters (red light -- ionized sulphur atoms, blue light -- double ionized oxygen atoms, green light -- ionized hydrogen).
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"Hubble Space Telescope Check-Out Finds Successes, Concerns",150,0,0,0
March 25, 1997
The Servicing Mission Observatory Verification (SMOV) for NASA's Hubble Space \JTelescope\j (HST), currently about halfway through its detailed check-out prior to returning to scientific operations, has found Hubble in overall excellent health, with seven of the eight components replaced or installed during the servicing mission functioning very well to date. However, a few concerns with one of the science instruments are being evaluated.
"The Hubble Space \JTelescope\j is checking out extremely well overall, and the few anomalies we see give us no reason to believe we will not be able to meet all our scientific goals," said Dr. Ed Weiler, HST Program Scientist, NASA Headquarters, Washington, DC.
"I'm very impressed that in just the few weeks since the servicing mission, we've already seen Hubble take the best images of Mars ever obtained from Earth's distance. Every observatory commissioning encounters some problems, but we're on track to clear up all our remaining concerns. That's good news for the many, many astronomers who are lined up for observing time on Hubble."
Earlier this month science observations with the Wide Field and Planetary Camera-2 resumed, and on March 10 the science team obtained images of Mars. Also, further optimization and alignment of the mirrors in the new Fine Guidance Sensor (FGS), installed during the servicing mission, were completed with excellent results following its first star observation. Project management officials say it's clearly the best FGS aboard HST.
Commissioning of the new Space \JTelescope\j Imaging Spectrograph (STIS) has proceeded very well, according to project officials. In the coming two weeks team members will test the instrument's ability to acquire targets in the narrow slits. Once this is demonstrated, the instrument will be ready to begin science operations.
Checkout of the Near Infrared Camera and Multi- Object Spectrometer (NICMOS), installed during the second servicing mission, has provided both excellent results and some areas of concern.
The NICMOS, designed to observe the universe in near-infrared light, contains three cameras and a set of highly advanced light sensors which must be maintained at a very cold temperature -- nominally 58 degrees Kelvin (-355 degrees Fahrenheit). These sensors, along with filters and other components, are housed in a large cryogenic dewar (a high-technology insulated bottle filled with about 225 lbs of solid \Jnitrogen\j embedded in aluminum foam).
The NICMOS Principal Investigator, Dr. Rodger Thompson, University of \JArizona\j, said NICMOS high resolution cameras 1 and 2 have shown excellent images in preliminary focus tests. However, these tests also show that camera 3 focus is currently beyond the range of the NICMOS internal mechanical adjustment capability. Analysis indicates the situation may be due to unexpected thermal contact in the dewar, which results in a slightly warmer cryogen temperature and a subsequent reduction of dewar lifetime.
The most likely explanation is that as the solid \Jnitrogen\j warms up it expands, and exerts pressure on the internal structure of the dewar. This expansion resulted in an unwanted physical contact between two internal structural components of the dewar, providing a pathway for excess heat to travel from the warmer outer structure of the dewar to its colder internal parts, warming the solid \Jnitrogen\j to a higher than desired operating temperature. This expansion also is affecting the performance of Camera 3.
The analysis team expects that the thermal contact might release in the future, returning NICMOS to its nominal state. Under these conditions, analysts predict that camera 3 should move back into the instrument's range of focus. Rearrangement of the NICMOS observing schedule could allow the full implementation of the NICMOS science program.
It will take several weeks or months for team engineers to be able to determine for certain the amount of reduction in the lifetime of the cryogen; however, the reduction can be compensated for by rearrangement of observing schedules.
Current plans call for SMOV activities to continue for the next few weeks with results of the Early Release Observation program available in early May.
During the STS-82 HST Second Servicing Mission in February, astronauts aboard the Space Shuttle Discovery replaced two older science instruments aboard Hubble with STIS and NICMOS, and also replaced a Fine Guidance Sensor, a Reaction Wheel Assembly, a Data Interface Unit, a Solar Array Drive \JElectronics\j package, an Engineering/Science Tape Recorder, and a Solid State Recorder. In addition, the astronauts performed other maintenance on the observatory, including patching of some \Jinsulation\j and installing covers on the Magnetic Sensing System.
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"Hubble Images Of Comet Hale-Bopp",151,0,0,0
March 27, 1997
This is a series of Hubble Space \JTelescope\j observations of the region around the nucleus of Hale-Bopp, taken on eight different dates since September 1995. They chronicle changes in the evolution of the nucleus as it moves ever closer to, and is warmed by, the sun.
The first picture shows a strong dust outburst on the \Jcomet\j that occurred when it was beyond the orbit of Jupiter. Images in the Fall of 1996 show multiple jets that are presumably connected to the activation of multiple vents on the surface of the nucleus.
In these false color images, taken with the Wide Field and Planetary Camera 2, the faintest regions are black, the brightest regions are white, and intermediate intensities are represented by different levels of red. All images are processed at the same spatial scale of 280 miles per pixel (470 kilometers), so the solid nucleus, no larger than 25 miles across, is far below Hubble's resolution.
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"Hubble Tracks Fading Optical Counterpart Of Gamma-Ray Burst",152,0,0,0
April 1, 1997
NASA's refurbished Hubble Space \JTelescope\j has made an important contribution toward solving one of \Jastronomy\j's greatest enigmas by allowing astronomers to continue watching the fading visible-light counterpart of a gamma-ray burst (GRB), one of the most energetic and mysterious events in the universe.
The so-called optical counterpart is presumably a cooling fireball from the catastrophic event that triggered the massive burst of invisible gamma rays -- the highest-energy radiation in the universe. This event may have unleashed as much energy in a few seconds as the Sun does in ten billion years!
The burst was discovered on February 28 by the Gamma-Ray Burst Monitor aboard the Italian-Dutch BeppoSAX satellite. The burst was also within the field of view of one of the SAX Wide Field Cameras. Followup observations were conducted by several other space-based astronomical observatories. The visible GRB counterpart, the first ever detected, was then discovered in a pair of ground-based telescopic images of the region where the burst occurred. Taken a week apart, the later picture showed that an object that could be seen in the first image had disappeared in the field, suggesting it was the decaying fireball from the event. A week after that discovery, astronomers at the New Technology \JTelescope\j and the Keck \Jtelescope\j identified an extended source at the location of the suspected GRB.
Hubble's high resolution and sensitivity were brought in to hunt down the rapidly dimming fireball -- plunging from 21st to below 23rd magnitude in eight days -- after it had grown so faint that it could not be resolved by ground-based telescopes by March 13. On March 26, Hubble allowed astronomers to reacquire the lost remnant, and continue following the behavior of the fading source. The Hubble observation clearly shows that the visible GRB source has two components: a point-like object and an extended feature.
This observation demonstrates Hubble's unique capability for monitoring the aftermath of gamma-ray bursts, long after they have faded from the view of Earth-based telescopes. And there will be no shortage of targets: once a day, a gamma-ray burst occurs somewhere in the universe.
"Now we know that, at least in some cases, we can follow the aftermath of GRBs for several weeks, using a coordinated effort between ground-based telescopes, Hubble and other spacecraft," said Kailash Sahu, leader of a team of scientists at The Space \JTelescope\j Science Institute, Baltimore
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"Hubble's Upgrades Show Birth and Death Of Stars",153,0,0,0
May 12, 1997
Three months after an orbital house call by astronauts, new instruments aboard NASA's Hubble Space \JTelescope\j are helping astronomers probe the universe in greater detail than ever before.
New data released by NASA today include direct evidence of a supermassive black hole and remarkable new details on the explosive life cycle of stars. NASA also reported that all new Hubble instruments and upgrades are generally performing well.
"We're extremely excited about the quality and precision of the images from Hubble," said Wes Huntress, NASA Associate Administrator for Space Science. "Following check-out of the instruments, Hubble will return to full science operations, and we can expect a continuing flow of new and exciting discoveries."
These initial results clearly demonstrate the ability of the new instruments to fulfill their science goals with the Hubble \JTelescope\j, say project astronomers. Project officials are pleased to report that other instruments and \Jelectronics\j installed during the second servicing mission are performing well.
Among Hubble's recent observations:
Jets and Gaseous Disk Around the Egg Nebula -- A new infrared instrument peered deep into the dust-obscured central region around a dying star embedded in the Egg nebula. A nebula is a cloud of dust and gas 3,000 light years from Earth.
The new images provide a clear view of a twin pair of narrow bullet-shaped "jets" of gas and dust blasted into space. The instrument, called the Near Infrared Camera and Multi-Object Spectrometer, also revealed an unusual scalloped edge along a doughnut-shaped molecular \Jhydrogen\j cloud in the nebula.
"Because we can now see these 'missing pieces' in infrared and visible light, we have a more complete view of the dynamic and complicated structure of the star," said Rodger Thompson of the University of \JArizona\j, Tucson, the principal investigator for the infrared instrument. "It also allows us to see a 'fossil record' of the star's late evolutionary stages."
Unveiling Violent Starbirth in the Orion Nebula - The new infrared instrument penetrated the shroud of dust along the back wall of the Orion nebula, located in the "sword" of the \Jconstellation\j Orion. Data revealed what can happen to a stellar neighborhood when massive young stars begin to violently eject material into the surrounding molecular cloud.
Although ground-based infrared cameras have previously observed this hidden region known as OMC-1, the Hubble's new instrument provides the most detailed look yet at the heart of this giant molecular cloud. Hubble reveals a surprising array of complex structures, including clumps, bubbles, and knots of material. Most remarkable are "bullets" composed of molecular \Jhydrogen\j -- the fastest of which travels at more than one million mph (500 km/s). These bullets are colliding with slower-moving material, creating bow shocks, like a speedboat racing across water.
Monster Black Hole in Galaxy M84 - In a single exposure, a new powerful instrument called the Space \JTelescope\j Imaging Spectrograph discovered a black hole at least 300 million times the mass of the Sun. The spectrograph made a precise observation along a narrow slit across the center of galaxy M84, located 50 million light-years away.
This allowed the instrument to measure the increasing velocity of a disk of gas orbiting the black hole. To scientists, this represents the signature of a black hole, among the most direct evidence obtained to date. Due to their nature, it's impossible to directly photograph black holes. Scientists must instead look for clues to show the effects of black holes on surrounding dust, gas and stars.
"Hubble proved the existence of supermassive black holes three years ago," said Bruce Woodgate of the Goddard Space Flight Center, Greenbelt, MD, and principal investigator for the new spectrograph. "With this new instrument, we can do it 40 times faster than we used to."
Composition and Structure of the Ring Around \JSupernova\j 1987A - The new spectrograph also provides an unprecedented look at a unique and complex structure in the universe -- a light-year-wide ring of glowing gas around \JSupernova\j 1987A, the closest \Jsupernova\j explosion in 400 years.
The spectrograph dissects the ring's light to tell scientists which elements are in the ring and helps paint a picture of the physics and stellar processes which created the ring. This gives astronomers better insight into how stars evolve and become a \Jsupernova\j, and into the origin of the chemical elements created in these massive explosions.
Hubble Status -- NASA project officials are encouraged that a problem detected earlier with one of the cameras on the infrared instrument has shown some improvement. The problem stems from the unexpected movement of the dewar -- an insulated vessel containing solid \Jnitrogen\j at extremely cold temperatures.
After launch, the \Jnitrogen\j expanded more than expected as it warmed, moving the dewar into contact with another surface in the mechanism and pushing one of the cameras out of its range of focus. The camera has moved back about one-third of the distance required to be within reach of the instrument's internal focusing mechanism.
This is because the dewar is "relaxing" toward its normal state, as pressure caused by the expansion of the \Jnitrogen\j is reduced. The ice keeps the sensitive infrared detector cooled. Project officials also are considering how to deal with unexpected, excessive coolant loss.
"We are anticipating a shorter lifetime for the instrument, but we don't know how much shorter," said Goddard Hubble Project Scientist David Leckrone. "We are taking steps to work around the problem, and will increase the percentage of time this instrument will be used."
NASA officials also report that other upgrades to Hubble are performing well, including the newly installed solid state recorder, fine guidance sensor and solar array drive \Jelectronics\j. The solid state recorder has significantly improved data storage and playback, and the new fine guidance sensor is by far the best of the three on Hubble.
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"Black Hole Signature Recorded By STIS",154,0,0,0
May 12, 1997
The colorful "zigzag" found is not the work of a flamboyant artist, but the signature of a supermassive black hole in the center of galaxy M84, discovered by Hubble Space \JTelescope\j's Space \JTelescope\j Imaging Spectrograph (STIS).
The image, taken with Hubble's Wide Field Planetary and Camera 2 shows the core of the galaxy where the suspected black hole dwells. Astronomers mapped the motions of gas in the grip of the black hole's powerful gravitational pull by aligning the STIS's spectroscopic slit across the nucleus in a single exposure.
The STIS data shows the rotational motion of stars and gas along the slit. The change in wavelength records whether an object is moving toward or away from the observer. The larger the excursion from the centerline, the greater the rotational velocity. If no black hole were present, the line would be nearly vertical across the scan.
Instead, STIS's detector found the S-shape at the center of this scan, indicating a rapidly swirling disk of trapped material encircling the black hole. Along the S-shape from top to bottom, velocities skyrocket as seen in the rapid, dramatic swing to the left (blueshifted or approaching gas), then the region in the center simultaneously records the enormous speeds of the gas both approaching and receding for orbits in the immediate vicinity of the black hole, and then an equivalent swing from the right, back to the center line.
STIS measures a velocity of 880,000 miles per hour (400 kilometers per second) within 26 light-years of the galaxy's center, where the black hole dwells. This motion allowed astronomers to calculate that the black hole contains at least 300 million solar masses. (Just as the mass of our Sun can be calculated from the orbital radii and speeds of the planets.)
This observation demonstrates a direct connection between a supermassive black hole and activity (such as radio emission) in the nucleus of an active galaxy. It also shows that STIS is ideally suited for efficiently conducting a survey of galaxies to determine the distribution of the black holes and their masses.
Each point on STIS's solid-state CCD (Charge Coupled Device) detector samples a square patch at the galaxy that is 12 light-years on a side. The detection of black holes at the centers of galaxies is about 40 times faster than the earlier Faint Object Spectrograph. STIS was configured to record five spectral features in red light from glowing \Jhydrogen\j atoms as well as \Jnitrogen\j and sulfur ions in orbit around the center of M84. At each sampled patch the velocity of the entrapped gas was measured. Because the patches are contiguous, the astronomers could map the change of velocity in detail.
M84 is located in the Virgo Cluster of galaxies, 50 million light-years from Earth.
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"NICMOS Captures The Heart Of OMC-1",155,0,0,0
May 12, 1997
The infrared vision of the Hubble Space \JTelescope\j's Near Infrared Camera and Multi-Object Spectrometer (NICMOS) is providing a dramatic new look at the beautiful Orion Nebula which contains the nearest nursery for massive stars.
For comparison, Hubble's Wide Field and Planetary Camera 2 (WFPC2) image shows a large part of the nebula as it appears in visible light. The heart of the giant Orion molecular cloud, OMC-1, is included in the relatively dim and featureless area inside the blue outline near the top of the image. Light from a few foreground stars seen in the WFPC2 image provides only a hint of the many other stars embedded in this dense cloud.
NICMOS's infrared vision reveals a chaotic, active star birth region. Here, stars and glowing interstellar dust, heated by and scattering the intense starlight, appear yellow-orange. Emission by excited \Jhydrogen\j molecules appears blue. The image is oriented with north up and east to the left. The diagonal extent of the image is about 0.4 light-years. Some details are as small as the size of our solar system.
The brightest object in the image is a massive young star called BN (Becklin-Neugebauer). Blue "fingers" of molecular \Jhydrogen\j emission indicate the presence of violent outflows, probably produced by a young star or stars still embedded in dust (located to the lower left, southeast, of BN).
The outflowing material may also produce the crescent-shaped "bow shock" on the edge of a dark feature north of BN and the two bright "arcs" south of BN. The detection of several sets of closely spaced double stars in these observations further demonstrates NICMOS's ability to see fine details not possible from ground-based telescopes.
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"Seyfert Galaxy NGC 4151, Fireworks in",156,0,0,0
June 9, 1997
The Space \JTelescope\j Imaging Spectrograph (STIS) simultaneously records, in unprecedented detail, the velocities of hundreds of gas knots streaming at hundreds of thousands of miles per hour from the nucleus of NGC 4151, thought to house a supermassive black hole.
This is the first time the velocity structure in the heart of this object, or similar objects, has been mapped so vividly this close to its central black hole.
The twin cones of gas emission are powered by the energy released from the supermassive black hole believed to reside at the heart of this Seyfert galaxy. The STIS data clearly show that the gas knots illuminated by one of these cones is rapidly moving towards us, while the gas knots illuminated by the other cone are rapidly receding.
The highest velocity material expelled in a cataclysmic stellar explosion ten years ago has been detected for the first time by the Hubble Space \JTelescope\j's Space \JTelescope\j Imaging Spectrograph (STIS).
An image taken with Hubble's Wide Field and Planetary Camera 2 in 1995, shows orange-red rings surrounding \JSupernova\j 1987A in the Large Magellanic Cloud. The glowing debris of the \Jsupernova\j explosion, which occurred in February 1987, is at the center of the inner ring.
The STIS spectrograph viewed the entire inner ring in far-ultraviolet light, spreading it into a spectrum. The Earth's atmosphere completely blocks ultraviolet radiation from reaching the Earth's surface, hence astronomers can study the ultraviolet universe only from orbiting telescopes.
The STIS data shows the presence of glowing \Jhydrogen\j expanding at a speed of 33 million miles per hour (15,000 kilometers per second) coming from an extended area inside the inner ring. In addition to \Jhydrogen\j emission STIS also detected emission from high-velocity ionized \Jnitrogen\j.
This is the first time that astronomers have measured the very fast moving gas ejected by the \Jsupernova\j explosion, which was invisible until observed by Hubble with the STIS ultraviolet detectors. This gas is glowing in the ultraviolet because it is slamming into the remains of the gas lost by the \Jsupernova\j star about 20,000 years before it exploded.
The STIS spectrum also reveals the presence of emissions from hot gasses (oxygen, \Jnitrogen\j, and helium) coming from the inner ring itself. The ring is about 1.2 light-years in diameter.
Supernova 1987A is located 167,000 light-years from Earth in the Large Magellanic Cloud.
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"Gamma-Ray Bursts Common To Normal Galaxies?",158,0,0,0
June 10, 1997
Nature's most powerful explosions, gamma-ray bursts, occur among the normal stellar population inside galaxies scattered across the universe. This means that, on average, a gamma-ray burst goes off once every few million years inside our Milky Way galaxy.
The energy released in such a \Jtitanic\j explosion, which can last from a fraction of a second to a few hundred seconds, is equal to all of the Sun's energy generated over its 10 billion year lifetime.
A team of astronomers, led by Kailash Sahu of the Space \JTelescope\j Science Institute (STScI) in Baltimore, MD, is reporting this conclusion at the 190th Meeting of the American Astronomical Society in \JWinston-Salem\j, NC. They used the Hubble Space \JTelescope\j to study the fading optical counterpart to a burst that happened on February 28, 1997.
The team's results are based on Hubble images taken on March 26 and April 7, 1997 that show the gamma-ray burst is offset from the center of an extended "fuzzy" object that looks like a galaxy. The researchers argue, statistically, that this must be more than a chance alignment. The gamma-ray burst was embedded inside the galaxy.
"Hubble's unmatched resolution was crucial in pinpointing the fact that the gamma-ray burst is away from the center," Sahu says. "This would rule out massive black holes, thought to dwell in the cores of most galaxies, as the source of these incredible explosions."
"These observations definitely represent a huge step forward towards the full understanding of these enigmatic objects," says co-investigator Mario Livio, of STScI. However, he points out that this conclusion assumes the mechanism for creating a gamma-ray burst is basically the same.
Keck \Jtelescope\j spectroscopic measurements of the optical counterpart to another gamma-ray burst, which exploded on May 8, found that its distance from Earth is several billion light-years. The Keck observation establishes that gamma-ray bursts are truly extragalactic in location and origin.
However, a June 2nd Hubble observation of this newer burst source, made with both of the newly installed science instruments, failed to reveal a galaxy adjacent to the optical counterpart.
"If the gamma-ray burster is at the distance indicated by the Keck spectrum, then its host galaxy is far less luminous than is the Milky Way," says Andrew Fruchter, one of the leaders of another STScI team in making the observation. "What is clear is that further observations of new bursters, and of these two bursters at later times, will be required to better understand the nature and location of these astonishing objects."
One possible mechanism for unleashing such a \Jtitanic\j fireball of energy is the collision of a neutron star with another neutron star or a black hole. The Hubble observations support this model because it appears gamma-ray bursts occur in the disk of a galaxy, where there is ongoing stellar formation, and so there should be an abundance of neutron stars from recently exploded supernovae.
The astronomers estimate a gamma-ray burst may explode within a few thousand light-years from Earth every few hundred million years.
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"Fireball From A Cataclysmic Explosion",159,0,0,0
June 10, 1997
The CCD camera (Charge Coupled Device) on the Space \JTelescope\j Imaging Spectrograph, a new instrument on Hubble Space \JTelescope\j, captured a visible fireball from a \Jtitanic\j explosion in deep space, called a gamma-ray burst.
The burst occurred on May 8, and Hubble observations to acquire the fading fireball were made on June 2. No accompanying object, such as a host galaxy, can be found near the burst. This result adds to the puzzlement over of the source of these enigmatic explosions, because a previous Hubble image of another gamma-ray burst counterpart identified a potential host galaxy. If a galaxy is present, and at the distance suggested by Keck \Jspectroscopy\j, it is much fainter than our Milky Way. A few faint galaxies are, however, seen several arcseconds from the source. If one of these is the host, then the gamma-ray burst is very far out in the galaxy's halo, well outside the galaxy's stellar disk.
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"Hubble Captures Volcanic Eruption Plume From Io",160,0,0,0
The Hubble Space \JTelescope\j has snapped a picture of a 400-km-high (250-mile-high) plume of gas and dust from a volcanic eruption on Io, Jupiter's large innermost moon.
Io was passing in front of Jupiter when this image was taken by the Wide Field and Planetary Camera 2 in July 1996. The plume appears as an orange patch just off the edge of Io in the eight o'clock position, against the blue background of Jupiter's clouds. Io's volcanic eruptions blasts material hundreds of kilometers into space in giant plumes of gas and dust. In this image, material must have been blown out of the \Jvolcano\j at more than 2,000 mph to form a plume of this size, which is the largest yet seen on Io.
Until now, these plumes have only been seen by \Jspacecraft\j near Jupiter, and their detection from the Earth-orbiting Hubble Space \JTelescope\j opens up new opportunities for long-term studies of these remarkable phenomena.
Io's volcanic plumes are much taller than those produced by terrestrial volcanos because of a combination of factors. The moon's thin atmosphere offers no resistance to the expanding volcanic gases; its weak gravity (one-sixth that of Earth) allows material to climb higher before falling; and its biggest volcanos are more powerful than most of Earth's volcanos.
This image is a contrast-enhanced composite of an ultraviolet image (2600 Angstrom wavelength), shown in blue, and a violet image (4100 Angstrom wavelength), shown in orange. The orange color probably occurs because of the absorption and/or scattering of ultraviolet light in the plume. This light from Jupiter passes through the plume and is absorbed by sulfur dioxide gas or is scattered by fine dust, or both, while violet light passes through unimpeded. Future HST observations may be able to distinguish between the gas and dust explanations.
#
"Mariner Program",161,0,0,0
A series of \Jspacecraft\j launched by NASA to begin the exploration of the inner and outer Solar System. The program included the first planetary flyby (Mariner 2, Venus, December 1962), the first planetary orbiter (Mariner 9, Mars, November 1971), and the first mission to Mercury (Mariner 10, Mar 1974, September 1974, and Mar 1975). A brief summary of some mission highlights follows.
\BMariner 2
Launch Date/Time:\b 1962-08-27 at 06:53:13 UTC
\BDescription\b
The Mariner 2 \Jspacecraft\j was the second of a series of \Jspacecraft\j used for planetary exploration in the flyby, or nonlanding, mode. Mariner 2 was a backup for the Mariner 1 mission which failed shortly after launch to Venus. The \Jspacecraft\j was attitude-stabilized using the Sun and Earth as references. It was solar powered and capable of continuous telemetry operation.
The \Jspacecraft\j obtained data on the interplanetary medium during the flight to Venus and beyond, and it obtained planetary data during the encounter of Venus. The \Jspacecraft\j passed Venus at a distance of 41,000 km on December 14, 1962.
\BMariner 5
Launch Date/Time:\b 1967-06-14 at 06:01:00 UTC
\BDescription\b
The Mariner 5 \Jspacecraft\j was the fifth in a series of \Jspacecraft\j used for planetary exploration in the flyby mode. Mariner 5 was a refurbished backup \Jspacecraft\j for the Mariner 4 mission and was converted from a Mars mission to a Venus mission. The \Jspacecraft\j was fully attitude stabilized, using the sun and Canopus as references. A central computer and sequencer subsystem supplied timing sequences and computing services for other \Jspacecraft\j subsystems.
The \Jspacecraft\j passed 4,000 km from Venus on October 19, 1967. The \Jspacecraft\j instruments measured both interplanetary and Venusian magnetic fields, charged particles, and plasmas, as well as the radio refractivity and UV emissions of the Venusian atmosphere. The mission was termed a success.
\BMariner 6
Launch Date/Time:\b 1969-02-24 at 01:29:00 UTC
\BDescription\b
Mariner 6 was the sixth in a series of \Jspacecraft\j used for planetary exploration in the flyby mode. Mariner 6 was attitude stabilized in three axes (referenced to the sun and the star, Canopus). The \Jspacecraft\j was solar powered and capable of continuous telemetry transmission. It was fully automatic in operation, although it could be reprogrammed from earth during the mission.
The \Jspacecraft\j was oriented entirely to planetary data acquisition, and no data were obtained during the trip to Mars or beyond Mars. Mariner 6 passed 3,431 km from Mars on July 31, 1969. The \Jspacecraft\j instruments took TV images of Mars and measured the radio refractivity and UV and IR emissions of the Martian atmosphere. The mission was a success, and data from it were used to program Mariner 7.
\BMariner 9
Launch Date:\b 1971-05-30
\BDescription\b
The Mariner 9 Mars mission was planned to consist of two \Jspacecraft\j on complementary missions, but due to the failure of Mariner 8 to launch properly, only one \Jspacecraft\j was available. Mariner 9 combined mission objectives of both Mariner 8 (mapping 70 % of the Martian surface) and Mariner 9 (a study of temporal changes in the Martian atmosphere and on the Martian surface). For the survey portion of the mission, the planetary surface was to be mapped with the same resolution as planned for the original mission, although the resolution of pictures of the polar regions would be decreased due to the increased slant range.
The variable features experiments were changed from studies of six given areas every 5 days to studies of smaller regions every 17 days. Mariner 9 arrived at Mars on November 14, 1971. The \Jspacecraft\j gathered data on the atmospheric composition, density, pressure, and temperature and also the surface composition, temperature, and \Jtopography\j of Mars. After depleting its supply of attitude control gas, the \Jspacecraft\j was turned off October 27,1972.
\BMariner 10
Launch Date/Time:\b 1973-11-03 at 05:45:00 UTC
\BDescription\b
Mariner 10 was the seventh successful launch in the Mariner series and the first \Jspacecraft\j to use the gravitational pull of one planet (Venus) to reach another (Mercury). The \Jspacecraft\j structure was an eight-sided framework with eight \Jelectronics\j compartments. It measured 1.39 m diagonally and 0.457 m in depth. Two solar panels, each 2.7 m long and 0.97 m wide, were attached at the top, supporting 5.1 sq m of solar cell area. The rocket engine was liquid-fueled, with two sets of reaction jets used to stabilize the \Jspacecraft\j on three axes.
It carried a low-gain omnidirectional antenna, composed of a honeycomb-disk parabolic reflector, 1.37 m in diameter, with focal length 55 cm. Feeds enabled the \Jspacecraft\j to transmit at S- and X-band frequencies. The \Jspacecraft\j carried a Canopus star tracker, located on the upper ring structure of the octagonal satellite, and acquisition sun sensors on the tips of the solar panels. The interior of the \Jspacecraft\j was insulated with multilayer thermal blankets at top and bottom. A sunshade was deployed after launch to protect the \Jspacecraft\j on the solar-oriented side.
Instruments on-board the \Jspacecraft\j measured the atmospheric, surface, and physical characteristics of Mercury and Venus. Experiments included \Jtelevision\j photography, magnetic field, plasma, infrared \Jradiometry\j, ultraviolet \Jspectroscopy\j, and radio science detectors. An experimental X-band, high-frequency transmitter was flown for the first time on this \Jspacecraft\j.
Mariner 10 was placed in a parking orbit after launch for approximately 25 minutes, then placed in orbit around the Sun en route to Venus. The orbit direction was opposite to the motion of the Earth around the Sun. Mid-course corrections were made. The \Jspacecraft\j passed Venus on February 5, 1974, at a distance of 4,200 km. It crossed the orbit of Mercury on March 29, 1974, at 2046 UT, at a distance of about 704 km from the surface. The TV and UV experiments were turned on the \Jcomet\j Kohoutek while the \Jspacecraft\j was on the way to Venus.
A second encounter with Mercury, when more photographs were taken, occurred on September 21, 1974, at an altitude of about 47,000 km. A third and last Mercury encounter at an altitude of 327 km, with additional photography of about 300 photographs and magnetic field measurements occurred on March 16, 1975. \JEngineering\j tests were continued until March 24, 1975, when the supply of attitude-control gas was depleted and the mission was terminated.
#
"Mars Observer Project (detailed)",162,0,0,0
US Mars orbiter \Jspacecraft\j launched in September 1992 and designed to acquire global maps of the surface chemistry, \Jmineralogy\j, and elevations, as well as to provide very high resolution (3 m / 7 ft) surface images and measurements of magnetic field and atmosphere dynamics. The \Jspacecraft\j was equipped with a French radio receiver to communicate with balloons and landed Russian \Jspacecraft\j to be launched in 1994 and 1996. Contact with the \Jspacecraft\j was lost on 21 August 1993, three days before the scheduled Mars orbit insertion.
\BMars Observer
Launch Date/Time:\b 1992-09-25 at 17:05:01 UTC
\BDescription \b
Mars Observer, the first of the Observer series of planetary missions, was designed to study the geoscience and climate of Mars. The primary science objectives for the mission were to:
(1) determine the global elemental and mineralogical character of the surface material;
(2) define globally the \Jtopography\j and gravitational field;
(3) establish the nature of the Martian magnetic field;
(4) determine the temporal and spatial distribution, abundance, sources, and sinks of volatiles and dust over a seasonal cycle; and,
(5) explore the structure and circulation of the atmosphere.
The bus and \Jelectronics\j of the Observer series of \Jspacecraft\j, used to study the terrestrial planets and near-Earth \Jasteroids\j, were derived from the Satcom-K and DMSP/TIROS \Jspacecraft\j. The rectangular bus section was 2.1 x 1.5 x 1.1 m. During the cruise phase of the mission, the high-gain antenna and the booms for the magnetometer (MAG/ER) and gamma-ray spectrometer (GRS) were partially deployed.
When fully deployed, the two booms were each 6 m long. The 1.5 m diameter high-gain antenna was, when fully deployed, on a 5.5 m boom to allow for clearance over the solar array when the antenna was pointed toward Earth. Pointing control for the \Jspacecraft\j was maintained through the use of four reaction wheels. Attitude information was provided by a horizon sensor (which defined the direction of the nadir), a star mapper (for inertial attitude), gyros and accelerometers (for measuring angular rates and linear accelerations), and multiple Sun sensors.
Power was provided through a six-panel solar array which, when fully deployed, measured 7.0 x 3.7 m. During the cruise phase, however, only four panels were deployed (due to the proximity of the \Jspacecraft\j to the sun) to reduce the amount of power generated. During periods when the \Jspacecraft\j was in Mars' shadow, energy was provided by two Ni-Cd batteries, each with a capacity of 43 amp-hours.
The interplanetary cruise phase of the mission was intended primarily for \Jspacecraft\j and instrument checkout and \Jcalibration\j. Two periods of data collection for the MAG/ER and GRS and one for the gravity wave experiment were planned for this phase as well. During the four month period from Mars orbital insertion until the \Jspacecraft\j achieved its final mapping orbit, only data collection for the MAG/ER, GRS, and thermal emission spectrometer (TES) were scheduled.
The mapping phase of the mission was scheduled to nominally last one Martian year. Mars Observer also supported the acquisition of data from the Russian Mars 1994 mission through the use of the joint French-Russian-American Mars \JBalloon\j Relay instrument. Contact with Mars Observer was lost on August 21, 1993, three days before scheduled orbit insertion, for unknown reasons and has not been re-established.
It is not known whether the \Jspacecraft\j was able to follow its automatic programming and go into Mars orbit or if it flew by Mars and is now in a heliocentric orbit. Although none of the primary objectives of the mission were achieved, cruise mode data were collected up to loss of contact.
Sponsoring Agencies/Countries included:
ò NASA-Office of Space Science Applications/United States
ò Center National d'Etudes Spatiales/France
ò Russian Space Agency
#
"Pioneer Venus Program",163,0,0,0
The Pioneer Venus mission consisted of two components, launched separately: an Orbiter and a Multiprobe.
\BThe Pioneer Venus Orbiter
Launch Date:\b 20 May 1978
\BLaunch Vehicle:\b Atlas-Centaur
\BDescription\b
The Pioneer Venus Orbiter was inserted into an elliptical orbit around Venus on December 4, 1978. The Orbiter was a flat cylinder 2.5 m in diameter and 1.2 m high. All instruments and \Jspacecraft\j subsystems were mounted on the forward end of the cylinder, except the magnetometer, which was at the end of a 4.7 m boom. A solar array extended around the \Jcircumference\j of the cylinder. A 1.09 m despun dish antenna provided S and X band communication with Earth.
The Pioneer Venus Orbiter carried 17 experiments (with a total mass of 45 kg):
ò a cloud photopolarimeter to measure the vertical distribution of the clouds.
ò a surface radar mapper to determine \Jtopography\j and surface characteristics.
ò an infrared radiometer to measure IR emissions from the Venus atmosphere.
ò an airglow ultraviolet spectrometer to measure scattered and emitted UV light.
ò a neutral mass spectrometer to determine the composition of the upper atmosphere.
ò a solar wind plasma analyzer to measure properties of the solar wind.
ò a magnetometer to characterize the magnetic field at Venus.
ò an electric field detector to study the solar wind and its interactions.
ò an \Jelectron\j temperature probe to study the thermal properties of the \Jionosphere\j.
ò an ion mass spectrometer to characterize the ionospheric ion population.
ò a charged particle retarding potential analyzer to study ionospheric particles.
ò two radio science experiments to determine the gravity field of Venus.
ò a radio \Joccultation\j experiment to characterize the atmosphere.
ò an atmospheric drag experiment to study the upper atmosphere.
ò a radio science atmospheric and solar wind turbulence experiment.
ò a gamma ray burst detector to record gamma ray burst events.
From Venus orbit insertion to July 1980, periapsis was held between 142 and 253 km (at 17 degrees north latitude) to facilitate radar and ionospheric measurements. The \Jspacecraft\j was in a 24 hour orbit with an apoapsis of 66,900 km. Thereafter, the periapsis was allowed to rise (to 2,290 km at maximum) and then fall, to conserve fuel. In 1991, the Radar Mapper was reactivated to investigate previously inaccessible southern portions of the planet.
In May, 1992, Pioneer Venus began the final phase of its mission, in which the periapsis was held between 150 and 250 km until the fuel ran out and atmospheric entry destroyed the \Jspacecraft\j the following August.
\BPioneer Venus Multiprobe
Launch Date:\b 08 August 1978
\BLaunch Vehicle:\b Atlas-Centaur
\BDescription\b
The Pioneer Venus Multiprobe consisted of a bus which carried one large and three small atmospheric probes. The large probe was released on November 16, 1978 and the three small probes on November 20. All four probes entered the Venus atmosphere on December 9, followed by the bus.
The Pioneer Venus large probe was equipped with 7 science experiments, contained within a sealed spherical pressure vessel. This pressure vessel was encased in a nose cone and aft protective cover. After deceleration from initial atmospheric entry at about 11.5 km/s near the equator on the Venus night side, a parachute was deployed at 47 km altitude. The large probe was about 1.5 m in diameter and the pressure vessel itself was 73.2 cm in diameter. The science experiments were:
ò a neutral mass spectrometer to measure the atmospheric composition.
ò a gas chromatograph to measure the atmospheric composition.
ò a solar flux radiometer to measure solar flux penetration in the atmosphere.
ò an infrared radiometer to measure distribution of infrared radiation.
ò a cloud particle size spectrometer to measure particle size and shape.
ò a nephelometer to search for cloud particles.
ò temperature, pressure, and acceleration sensors.
The three small probes were identical to each other, 0.8 m in diameter. These probes also consisted of spherical pressure vessels surrounded by an aeroshell, but unlike the large probe, they had no parachutes and the aeroshells did not separate from the probe. Each small probe carried a nephelometer and temperature, pressure, and acceleration sensors, as well as a net flux radiometer experiment to map the distribution of sources and sinks of radiative energy in the atmosphere.
The radio signals from all four probes were also used to characterize the winds, turbulence, and propagation in the atmosphere. The small probes were each targeted at different parts of the planet and were named accordingly. The North probe entered the atmosphere at about 60 degrees north latitude on the day side. The night probe entered on the night side. The day probe entered well into the day side, and was the only one of the four probes which continued to send radio signals back after impact, for over an hour.
The Pioneer Venus bus also carried two experiments, a neutral mass spectrometer and an ion mass spectrometer to study the composition of the atmosphere. With no heat shield or parachute, the bus survived and made measurements only to about 110 km altitude before burning up. The bus was a 2.5 m diameter cylinder weighing 290 kg, and afforded us our only direct view of the upper Venus atmosphere, as the probes did not begin making direct measurements until they had decelerated lower in the atmosphere.
#
"Sakigake and Suisei Projects",164,0,0,0
The first Japanese interplanetary \Jspacecraft\j, launched to intercept Halley's \Jcomet\j. Sakigake ('forerunner') was launched (January 1985) as a test \Jspacecraft\j for Suisei ('comet', launched August 1985). They were instrumented to measure solar wind interaction and the \Jhydrogen\j cloud of the \Jcomet\j. They were built and operated by the Institute for Space and Aeronautical Science at the University of \JTokyo\j.
\BSakigake
Launch Date/Time:\b 1985-01-08 at 19:26:00 UTC
\BDescription\b
Sakigake is a test \Jspacecraft\j similar to Suisei. It flew by \JComet\j P/Halley on its sunward side at a distance of about 7 million kilometers on March 11, 1986. It carries three instruments to measure plasma wave spectra, solar wind ions, and interplanetary magnetic fields, all of which worked normally. The \Jspacecraft\j is spin-stabilized at two different rates (5 and 0.2 rpm). It is equipped with hydrazine thrusters for attitude and velocity control, star and sun sensors for attitude determination, and a mechanically despun off-set parabolic dish for long-range communication.
Sakigake made an Earth swingby on January 8, 1992. The closest approach was at 23h 08m 47s (JST = UTC+9h) with a geocentric distance of 88,997 km. This was the first planet-swingby for a Japanese \Jspacecraft\j. During the approach, Sakigake observed the geotail. Some geotail passages are scheduled during ISTP's multi-spacecraft investigation of that region.
The second Earth swingby was on June 14, 1993 at 40 Re, and the third on October 28, 1994 at 86 Re. Further mission planning targets a 23.6 km/s, 10,000 km flyby of \JComet\j P/Honda-Mrhos-Pajdusakova on Feb 3, 1996 at about 21.00 GMT (approaching the nucleus along the tail) some 0.17 AU from the Sun, and a 14 million km passage of \JComet\j P/Giacobini-Zinner on Nov 29, 1998.
\BSuisei
Launch Date/Time:\b 1985-08-18 at 23:33:00 UTC
\BDescription \b
Suisei (the Japanese name meaning 'Comet') was launched on March 18, 1985 into heliocentric orbit to fly by \JComet\j P/Halley. It is identical to Sakigake apart from its payload: a CCD UV imaging system and a solar wind instrument. The main objective of the mission was to take UV images of the \Jhydrogen\j corona for about 30 days before and after \JComet\j Halley's descending crossing of the \Jecliptic\j plane.
Solar wind parameters were measured for a much longer time period. The \Jspacecraft\j is spin-stabilized at two different rates (5 and 0.2 rpm). Hydrazine thrusters are used for attitude and velocity control; star and sun sensors are for attitude control; and a mechanically despun off-set parabolic dish is used for long range communication.
#
"Soviet/Russian Space Programs continued",165,0,0,0
This section summarizes some of the major achievements of the Soviet Space Program. This program was replaced by the Russian Space Agency after the demise of the Soviet Union in 1992.
The programs discussed in this section include the following:
\BEarly Soviet Satellites\b - Including \JSputnik\j, Vostok and Soyuz
\BSpace Stations\b - Salyut and Mir
\BSoviet Lunar Missions\b - Luna Series and Zond Series
\BVenera Program\b - Covering Venera missions to Venus
\BVega Project\b - Covering missions to Venus and Halley's \JComet\j
\BPhobos\b - Mission to Mars
A summary of each of the above projects follows.
\B Early Soviet Satellites\b
\B\ISputnik Satellites\b\i
The first Soviet satellite program consisted of four \JSputnik\j satellites. However, the \JSputnik\j launched between \JSputnik\j 2 and 3 failed to reach orbit.
Sputnik 1, launched on October 4, 1957, was designed to send radio signals to Earth and determine the density of the upper atmosphere. However, it only transmitted signals to Earth for a short time after launch. Its orbit decayed and it fell to Earth on January 4, 1958. \JSputnik\j 1 had an orbital period of 98.6 minutes and an orbital altitude of 228-947 km.
Sputnik 2 was launched on November 3, 1957, and carried aboard it a dog, Laika. Biological data was returned for approximately a week (the first data of its kind). However, there was no safe re-entry possible at the time, and Laika was put to sleep after a week in orbit. The satellite itself remained in orbit 162 days. \JSputnik\j 2 had an orbital period of 103.75 minutes and an orbital altitude of 225-1,671 km.
Sputnik 3 was launched on May 15, 1958. It may originally have been intended as the first launch in the \JSputnik\j program, however it was apparently decided to be more cautious in the launch schedule. It was designed to be a geophysical laboratory, performing experiments on the Earth's magnetic field, radiation belt, and \Jionosphere\j. It orbited Earth and transmitted data until April 6, 1960, when its orbit decayed.
All Sputniks were launched using the SS-6, or Sapwood rocket. The SS-6 was originally designed as a ballistic missile, and had its upper stage modified slightly to hold the \JSputnik\j payload. It had two stages, four strap-on booster rockets for the first stage, connected to the second stage rocket. Total mass at launch for \JSputnik\j 3 was 267 tonnes, with a length of 29.17 meters. The primary stage used RD-107 engines, which provided 100,000 kg of thrust. Both stages were powered by LOX/Kerosene.
\B\IVostok Spacecraft\b\i
The first generation of Soviet crewed \Jspacecraft\j, carrying a single member. Vostok 1 carried the first human into space (12 April 1961) -- Yuri Gagarin, who orbited Earth once on a flight of 118 min. Crew were recovered over land after ejection from the capsule at 7,000 m / 23,000 ft altitude after re-entry. The last Vostok flight carried Valentina Tereshkova, the first woman to fly in space (Vostok 6, 16 June 1963).
Voskhod ('Sunrise') was an intermediate-generation Soviet-crewed \Jspacecraft\j following Vostok and preceding Soyuz; it made only two flights, in 1964 and 1965.
\B\ISoyuz\b\i
A Soviet basic space capsule, consisting of three modules (orbiter, descent, and instrumentation), and carrying a crew of one to three. It has been flown on several dozen missions, and is capable of precision targeting to soft-land in Central Asia (contrasting with the US ocean-recovery technique). It has been used to ferry crew to and from Salyut and, later, Mir space stations with which docking takes place. The first flight was in 1967, when its pilot, V Komarov, was killed in a landing accident.
Soyuz 19 docked with the Apollo \Jspacecraft\j in Earth orbit after its rendezvous (July 1975). There was a notable launch accident (September 1983) in which the crew of Soyuz T-10A ejected to safety as the launch vehicle exploded.
\BSpace Stations\b
\B\ISalyut\b\i
The first-generation Soviet space station, capable of docking with the Soyuz crew ferry and Progress resupply vehicle; it provides 100 cubic meters / 3,500 cu ft of living space for up to five cosmonauts. Two versions have been flown in a program spanning 1971 to the present, aimed at accumulating data on long-duration space-flight experience and biomedical experiments, Earth remote sensing, and \Jmicrogravity\j science. The first station was flown in 1971, the last (Salyut 7) in 1982. The station's orbit eventually decays, with the vehicle re-entering the atmosphere and burning up. Crews have accumulated many hundred days of flight experience. A notably dangerous repair mission was undertaken to Salyut 7 in 1985, by V Dzhanibekov and V Savinykh, after the station seriously malfunctioned between crew occupancies.
\B\IMir\b\i
A Soviet space station (launched February 1986) which evolved from Salyut, having more power (solar panels) and more docking ports (five) than previous \Jspacecraft\j, allowing for the build-up of a modular station. It is used for long-duration spaceflight experience, and for biomedical, science, and applications experiments. Yuri Romanenko occupied Mir for 326 days in 1987. The current space endurance record for men is held by Russian \Jcosmonaut\j Valeri Poliakov, who lived in space for 438 days (Jan 1994-Mar 1995), and for women by Yelena Kondakova.
\BSoviet Lunar Missions\b
The Soviet Lunar program had 20 successful missions to the Moon and achieved a number of notable lunar "firsts": first probe to impact the Moon, first flyby and image of the lunar farside, first soft landing, first lunar orbiter, and the first circumlunar probe to return to Earth. The two successful series of Soviet probes were the Luna (15 missions) and the Zond (5 missions).
Lunar flyby missions (Luna 3, Zond 3, 6, 7, 8) obtained photographs of the lunar surface, particularly the limb and farside regions. The Zond 6, 7, and 8 missions circled the Moon and returned to Earth where they were recovered, Zond 6 and 7 in \JSiberia\j and Zond 8 in the Indian Ocean. The purpose of the photography experiments on the lunar landers (Luna 9, 13, 22) was to obtain close-up images of the surface of the Moon for use in lunar studies and determination of the feasibility of manned lunar landings.
\B\IThe Luna Series
Luna 2
Launched 12 September 1959\b\i
Impacted Moon 13 September 1959 at 22:02:04 UT
Latitude 29.10 N, Longitude 0.00 -- Palus Putredinis
\B\ILuna 3
Launched 04 October 1959\b\i
Lunar Flyby
\B\ILuna 9
Launched 31 January 1966\b\i
Landed on Moon 03 February 1966 at 18:44:52 UT
Latitude 7.08 N, Longitude 295.63 E -- \JOceanus\j Procellarum
\B\ILuna 10
Launched 31 March 1966\b\i
Lunar Orbiter
\B\ILuna 11
Launched 24 August 1966\b\i
Lunar Orbiter
\B\ILuna 12
Launched 22 October 1966\b\i
Lunar Orbiter
\B\ILuna 13
Launched 21 December 1966\b\i
Landed on Moon 24 December 1966 at 18:01:00 UT
Latitude 18.87 N, 297.95 E -- \JOceanus\j Procellarum
\B\ILuna 14
Launched 7 April 1968\b\i
Lunar Orbiter
\B\ILuna 16
Launched 12 September 1970\b\i
Landed on Moon 20 September 1970 at 05:18:00 UT
Latitude 0.68 S, Longitude 56.30 E -- Mare Fecunditatis
Lunar Sample Return
\B\ILuna 17
Launched 10 November 1970\b\i
Landed on Moon 17 November 1970 at 03:47:00 UT
Latitude 38.28 N, Longitude 325.00 E -- Mare Imbrium
Lunar Rover -- Lunokhod 1
\B\ILuna 19
Launched 28 September 1971\b\i
Lunar Orbiter
\B\ILuna 20
Launched 14 February 1972\b\i
Landed on Moon 21 February 1972 at 19:19:00 UT
Latitude 3.57 N, Longitude 56.50 E -- Mare Fecunditatis
Lunar Sample Return to Earth 25 February 1972
\B\ILuna 21
Launched 08 January 1973\b\i
Landed on Moon 15 January 1973 at 23:35:00 UT
Latitude 25.51 N, Longitude 30.38 E -- Mare Serenitatis
Lunar Rover -- Lunokhod 2
\B\ILuna 22
Launched 02 June 1974\b\i
Lunar Orbiter
\B\ILuna 24
Launched 14 August 1976\b\i
Landed on Moon 18 August 1976 at 02:00:00 UT
Latitude 12.25 N, Longitude 62.20 E -- Mare Crisium
Lunar Sample Return
\B\IThe Zond Series
Zond 3
Launched 18 July 1965\b\i
Lunar Flyby
\B\IZond 5
Launched 15 September 1968\b\i
Circumlunar
Returned to Earth 21 September 1968
\B\IZond 6
Launched 10 November 1968\b\i
Circumlunar
Returned to Earth 17 November 1968
\B\IZond 7
Launched 07 August 1969\b\i
Circumlunar
Returned to Earth 14 August 1969
\B\IZond 8
Launched 20 October 1970\b\i
Circumlunar
Returned to Earth 27 October 1970
\BVenera Program\b
A highly successful evolutionary series of Soviet space missions to Venus 1961-83 (plus Vega landers and balloons of 1985). Its highlights include: the first successful atmospheric entry probe (Venera 4, 1967); the first complete descent to the surface (Venera 5, 1969); the first science measurements on the surface (Venera 7, 1970); the first TV pictures from the surface (Venera 9, 1975); the first chemical analysis of the soil (Venera 13, 1981); the first high resolution images of the surface from orbit using radar to penetrate clouds (Venera 15/16, 1983); and the first balloons deployed and tracked in the atmosphere (Vega, 1985).
\B\IVenera 4
Launch Date:\b\i 1967-06-12
\B\IOn-orbit dry mass:\b\i 1106.00 kg
\B\IDescription\b\i
Venera 4 was launched from a Tyazheliy \JSputnik\j (67-058B) towards the planet Venus with the announced mission of direct atmospheric studies. On October 18, 1967, the \Jspacecraft\j entered the Venusian atmosphere and released two thermometers, a \Jbarometer\j, a radio \Jaltimeter\j, and atmospheric density gauge, 11 gas analyzers, and two radio transmitters operating in the DM waveband.
The main bus, which had carried the capsule to Venus, carried a magnetometer, cosmic ray detectors, \Jhydrogen\j and oxygen indicators, and charged particle traps. Signals were returned by the \Jspacecraft\j, which braked and then deployed a parachute system after entering the Venusian atmosphere, until it reached an altitude of 24.96 km.
\B\IVenera 5
Launch Date:\b\i 1969-01-05
\B\IOn-orbit dry mass:\b\i 1130.00 kg
\B\IDescription\b\i
Venera 5 was launched from a Tyazheliy \JSputnik\j (69-001C) towards Venus to obtain atmospheric data. The \Jspacecraft\j was very similar to Venera 4 although it was of a stronger design. When the atmosphere of Venus was approached, a capsule weighing 405 kg and containing scientific instruments was jettisoned from the main \Jspacecraft\j. During satellite descent towards the surface of Venus, a parachute opened to slow the rate of descent.
For 53 minutes on May 16, 1969, while the capsule was suspended from the parachute, data from the Venusian atmosphere were returned. The \Jspacecraft\j also carried a medallion bearing the coat of arms of the USSR and a bas-relief of V. I. Lenin to the night side of Venus.
\B\IVenera 6
Launch Date:\b\i 1969-01-10
\B\IOn-orbit dry mass:\b\i 1130.00 kg
\B\IDescription\b\i
Venera 6 was launched from a Tyazheliy \JSputnik\j (69-002C) towards Venus to obtain atmospheric data. The \Jspacecraft\j was very similar to Venera 4 although it was of a stronger design. When the atmosphere of Venus was approached, a capsule weighing 405 kg was jettisoned from the main \Jspacecraft\j. This capsule contained scientific instruments. During descent towards the surface of Venus, a parachute opened to slow the rate of descent.
For 51 minutes on May 17, 1969, while the capsule was suspended from the parachute, data from the Venusian atmosphere were returned. The \Jspacecraft\j also carried a medallion bearing the coat of arms of the USSR and a bas-relief of V. I. Lenin to the night side of Venus.
\B\IVenera 7
Launch Date:\b\i 1970-08-17
\B\IOn-orbit dry mass:\b\i 1180.00 kg
\B\IDescription\b\i
Venera 7 was launched from a Tyazheliy \JSputnik\j in an earth parking orbit towards Venus to study the Venusian atmosphere and other phenomena of the planet. Venera 7 entered the atmosphere of Venus on December 15, 1970, and a landing capsule was jettisoned. After aerodynamic braking, a parachute system was deployed. The capsule antenna was extended, and signals were returned for 35 min.
Another 23 minutes of very weak signals were received after the \Jspacecraft\j landed on Venus. The capsule was the first man-made object to return data after landing on another planet.
\B\IVenera 9 Descent Craft
Description\b\i
On October 20, 1975, this \Jspacecraft\j was separated from the Orbiter, and landing was made with the sun near zenith at 0513 UT on October 22. A system of circulating fluid was used to distribute the heat load. This system, plus precooling prior to entry, permitted operation of the \Jspacecraft\j for 53 minutes after landing.
During descent, heat dissipation and deceleration were accomplished sequentially by protective hemispheric shells, three parachutes, a disk-shaped drag brake, and a compressible, metal, doughnut-shaped, landing cushion. The landing was about 2,200 km from the Venera 10 landing site.
Preliminary results indicated: (A) clouds 30-40 km thick with bases at 30-35 km altitude, (B) atmospheric constituents including HCl, HF, Br, and I, (C) surface pressure about 90 (earth) atmospheres, (D) surface temperature 485 deg C, (E) light levels comparable to those at earth mid-latitudes on a cloudy summer day, and (F) successful TV photography showing shadows, no apparent dust in the air, and a variety of 30-40 cm rocks which were not eroded.
\B\IVenera 11 Descent Craft
Launch Date:\b\i 1978-09-09
\B\IDescription\b\i
The Venera 11 descent craft carried instruments designed to study the detailed chemical composition of the atmosphere, the nature of the clouds, and the thermal balance of the atmosphere. Separating from its flight platform on December 25, 1978, it made a soft landing on the surface after a descent time of approximately 1 hour. During this time, it employed aerodynamic braking followed by parachute braking and ending with atmospheric braking.
The touchdown speed was 7-8 m/s. Information was transmitted to the flight platform for retransmittal to earth. It is unknown whether the Lander Probe carried an imaging system. No mention of it occurs in the Soviet literature examined by the author. Two other experiments on the Lander did fail, and their failure was acknowledged by the Soviets. Some US literature on the subject notes that the imaging system "failed" but did return some data.
\B\IVenera 12 Descent Craft
Launch Date:\b\i 1978-09-14
\B\IDescription \b\i
The Venera 12 descent craft carried instruments designed to study the detailed chemical composition of the atmosphere, the nature of the clouds, and the thermal balance of the atmosphere. Separating from its flight platform on December 21, 1978, it made a soft landing on the surface after a descent time of approximately 1 hour. During this time, it employed aerodynamic braking followed by parachute braking and ending with atmospheric braking.
The touchdown speed was 7-8 m/s. Information was transmitted to the flight platform for retransmittal to earth. It is unknown whether the Lander Probe carried an imaging system. No mention of it occurs in the Soviet literature examined by the author. Two other experiments on the Lander did fail, and their failure was acknowledged by the Soviets. Some US literature on the subject notes that the imaging system "failed" but did return some data.
\B\ITo continue please click\b\i \JSoviet/Russian Space Programs continued 2\j
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"Soviet/Russian Space Programs continued 2",166,0,0,0
\B\IVenera 13 Descent Craft
Launch Date:\b\i 1981-10-30
\B\IDescription\b\i
Venera 13 landed at 7 deg 30 minutes S by 303 deg, just east of the eastern extension of an elevated region known as Phoebe Regio. It survived for 2 h 7 minutes in an environment with a temperature of 457 deg C and a pressure of 89 earth atmospheres.
Venera 13 carried instruments to take chemical and isotopic measurements, monitored the spectrum of scattered sunlight, and recorded electric discharges during its descent phase through the Venusian atmosphere. The \Jspacecraft\j utilized a camera system, an X-ray \Jfluorescence\j spectrometer, and a seismometer to conduct investigations on the surface.
\B\IVenera 14 Descent Craft
Launch Date:\b\i 1981-11-04
\B\IDescription\b\i
Venera 14 landed at 13 deg 15 minutes S by 310 deg, about 950 km southwest of Venera 13. Surface temperature was 465 deg C and pressure was 94 earth atmospheres. Venera 14 carried instruments to take chemical and isotopic measurements, monitored the spectrum of scattered sunlight, and recorded electric discharges during its descent phase through the Venusian atmosphere. The \Jspacecraft\j utilized a camera system, an X-ray \Jfluorescence\j spectrometer, and a seismometer to conduct investigations on the surface.
\B\IVenera 15
Launch Date:\b\i 1983-06-02
\B\IOn-orbit dry mass:\b\i 4000.00 kg
\B\IDescription\b\i
Venera 15 was part of a two \Jspacecraft\j mission (along with Venera 16) designed to use 8 cm band side-looking radar maps to study the surface properties of Venus. The two \Jspacecraft\j were inserted into Venus orbit a day apart with their orbital planes shifted by an angle of approximately 4 degrees relative to one another. This made it possible to reimage an area if necessary. Each \Jspacecraft\j was in a nearly polar orbit with a periapsis at 62 N latitude. Together, the two \Jspacecraft\j imaged the area from the north pole down to about 30 degrees N latitude over the 8 months of mapping operations.
The Venera 15 and 16 \Jspacecraft\j were identical and were based on modifications to the orbiter portions of the Venera 9 and 14 probes. Each \Jspacecraft\j consisted of a 5 m long cylinder with a 6 m diameter, 1.4 m tall parabolic dish antenna for the synthetic aperture radar (SAR) at one end. A 1 meter diameter parabolic dish antenna for the radio \Jaltimeter\j was also located at this end. The electrical axis of the radio \Jaltimeter\j antenna was lined up with the axis of the cylinder. The electrical axis of the SAR deviated from the \Jspacecraft\j axis by 10 degrees.
During imaging, the radio \Jaltimeter\j would be lined up with the center of the planet (local vertical) and the SAR would be looking off to the side at 10 degrees. A bulge at the opposite end of the cylinder held fuel tanks and propulsion units. Two square solar arrays extended like wings from the sides of the cylinder. A 2.6 m radio dish antenna for communications was also attached to the side of the cylinder.
\B\IVenera 16
Launch Date:\b\i 1983-06-07
\B\IOn-orbit dry mass:\b\i 4000.00 kg
\B\IDescription\b\i
Venera 16 was part of a two \Jspacecraft\j mission (along with Venera 15) designed to use 8 cm band side-looking radar maps to study the surface properties of Venus. The two \Jspacecraft\j were inserted into Venus orbit a day apart with their orbital planes shifted by an angle of approximately 4 degrees relative to one another. This made it possible to reimage an area if necessary. Each \Jspacecraft\j was in a nearly polar orbit with a periapsis at 62 N latitude.
Together, the two \Jspacecraft\j imaged the area from the north pole down to about 30 degrees N latitude over the 8 months of mapping operations. In June 1984, Venus was at superior conjunction and passed behind the Sun as seen from Earth. No transmissions were possible, so the orbit of Venera 16 was rotated back 20 degrees at this time to map the areas missed during this period.
The Venera 15 and 16 \Jspacecraft\j were identical and were based on modifications to the orbiter portions of the Venera 9 and 14 probes. Each \Jspacecraft\j consisted of a 5 m long cylinder with a 6 m diameter, 1.4 m tall parabolic dish antenna for the synthetic aperture radar (SAR) at one end. A 1 meter diameter parabolic dish antenna for the radio \Jaltimeter\j was also located at this end. The electrical axis of the radio \Jaltimeter\j antenna was lined up with the axis of the cylinder. The electrical axis of the SAR deviated from the \Jspacecraft\j axis by 10 degrees.
During imaging, the radio \Jaltimeter\j would be lined up with the center of the planet (local vertical) and the SAR would be looking off to the side at 10 degrees. A bulge at the opposite end of the cylinder held fuel tanks and propulsion units. Two square solar arrays extended like wings from the sides of the cylinder. A 2.6 m radio dish antenna for communications was also attached to the side of the cylinder.
\BVega Project\b
Vega was a highly successful Soviet mission to Venus and Halley's \JComet\j undertaken in 1984-86. Two \Jspacecraft\j each deployed a lander and a \Jballoon\j at Venus, then used a Venus gravity assist to fly on to intercept Halley's \JComet\j, passing within an estimated 3,000 and 10,000 km of its nucleus. The mission carried an ambitious science payload, highly international in scope; all elements of the mission were boldly undertaken in public view and proved to be highly successful.
\B\IVega 1 and Vega 2
Launch Date:\b\i 1984-12-15 (Vega 1) and 1984-12-21 (Vega 2)
\B\IOn-orbit dry mass:\b\i 2500.00 kg
\B\IDescription\b\i
This \Jspacecraft\j mission combined a Venus swingby and a \JComet\j Halley flyby. Two identical \Jspacecraft\j, Vega 1 and Vega 2, were launched December 15 and 21, 1984, respectively. After carrying Venus entry probes to the vicinity of Venus (arrival and deployment of probes were scheduled for June 11-15, 1985), the two \Jspacecraft\j were to be retargetted using Venus gravity field assistance to intercept \JComet\j Halley in March 1986.
The first \Jspacecraft\j was to encounter \JComet\j Halley on March 6, 1986, and the second about three days later. The flyby velocity was to be 77.7 km/s. Although the \Jspacecraft\j could be targetted with a precision of 100 km, the position of the \Jspacecraft\j relative to the \Jcomet\j nucleus was estimated to be known only to within a few thousand kilometers. This, together with the problem of dust protection, led to estimated flyby distances of 10,000 km for the first \Jspacecraft\j and 3,000 km for the second.
The \Jspacecraft\j was three-axis stabilized. Its main features were large solar panels, a high-gain antenna dish, and an automatic pointing platform carrying those experiments that required pointing at the \Jcomet\j nucleus. The automatic platform could rotate through + or -110 degrees and + or -40 degrees in two perpendicular directions with a pointing accuracy of 5 arc-min and a stability of 1 arc-min/s. It carried the narrow- and the wide-angle camera, the three-channel spectrometer, and the infrared sounder.
All other experiments were body-mounted, with the exception of two magnetometer sensors and various plasma probes and plasma wave analyzers which were mounted on a 5-m boom. The total scientific payload weighed 125 kg and had a data rate of 65 kbs in fast telemetry mode for encounter. There was also a slow telemetry mode for the cruise mode. The comet-encounter science data-take was from 2.5 h before until 0.5 h after the closest approach, with several periods of data-take before and after, each lasting about 2 h.
Continuous coverage for plasma and dust instruments was provided by an onboard memory (5-megabit tape recorder). The \Jspacecraft\j was shielded from hypervelocity dust impacts by a shield consisting of a 100-micrometer multilayer sheet 20 to 30 cm from the \Jspacecraft\j, and a 1-mm Al sheet 5 to 10 cm from the \Jspacecraft\j.
Approximately half of the VEGA \Jspacecraft\j was devoted to the Halley module, and half to the Venus lander package. The total scientific payload weight was 144.3 kg. The Venus package consisted of a sphere 240 cm in diameter, which was to be separated two days before arrival at Venus and enter the planet's atmosphere on an inclined path, without active maneuvers, as was done on previous Venera missions. The lander probe was identical to those of Venera 9 through 14 and similarly had two objectives, the study of the atmosphere and the study of the superficial crust.
In addition to temperature and pressure measuring instruments, the descent probe carried a UV spectrometer for measurement of minor atmospheric constituents, an instrument dedicated to measurement of the concentration of water, and other instruments for determination of the chemical composition of the condensed phase: a gas-phase chromatograph; an X-ray spectrometer observing the \Jfluorescence\j of grains or drops; and a mass spectrograph measuring the chemical composition of the grains or drops.
The X-ray spectrometer separated the grains according to their sizes using a laser imaging device, while the mass spectrograph separated them according to their sizes using an aerodynamical inertial separator. After landing, a small surface sample near the probe was to be analyzed by gamma \Jspectroscopy\j and X-ray \Jfluorescence\j.
The UV spectrometer, the mass spectrograph, and the pressure- and temperature-measuring instruments were developed in cooperation between French and Soviet investigators. In addition to the lander probe, a constant-pressure instrumented \Jballoon\j was to be deployed immediately after entry into the atmosphere. The \Jballoon\j, with a 5-kg payload and 25-kg total mass, was to float at approximately 50 km altitude in the middle, most active layer of the Venus three-tiered cloud system.
Data from the \Jballoon\j instruments were to be transmitted directly to Earth for the 60-h lifetime of the batteries. Onboard instruments were to measure temperature, pressure, vertical wind velocity, and visibility (density of local aerosols). Very long baseline interferometry was to be used to track the motion of the \Jballoon\j to provide the wind velocity in the clouds.
The balloons were deployed at 54 km altitude and tracked for two days by an international network of antennas, including NASA's Deep Space Network (DSN) - a notable example of international space cooperation. Likewise for the Halley flyby, the Vega project supplied optical navigation data, and DSN supplied radio tracking inputs to the European Space Agency's Giotto Project. This provided the first close-up view of the \JComet\j's nucleus, and the first measurement of gas and dust properties. The \Jspacecraft\j were severely battered by the 75 km impact of Halley's dust, becoming non-operational.
\BPhobos Project\b
\B\ILaunch Date:\b\i July 7, 1988 (Phobos 1) and July 21, 1988 (Phobos 2)
\B\ILaunch Vehicle:\b\i Proton
\B\IMass:\b\i 2,600 Kg (6,220 Kg with orbital insertion hardware attached)
\B\IPower System:\b\i Solar panels
\B\IDescription\b\i
Phobos 1, and its companion \Jspacecraft\j \JPhobos\j 2, were the next-generation in the Venera-type planetary missions, succeeding those last used during the Vega 1 and 2 missions to \Jcomet\j P/Halley. The objectives of the \JPhobos\j missions were to:
(1) conduct studies of the interplanetary environment;
(2) perform observations of the Sun;
(3) characterize the plasma environment in the Martian vicinity;
(4) conduct surface and atmospheric studies of Mars; and,
(5) study the surface composition of the Martian satellite \JPhobos\j.
The main section of the \Jspacecraft\j consisted of a pressurized toroidal \Jelectronics\j section surrounding a modular cylindrical experiment section. Below these were mounted four spherical tanks containing hydrazine for attitude control and, after the main propulsion module was to be jettisoned, orbit adjustment. A total of 28 thrusters (twenty-four 50 N thrusters and four 10 N thrusters) were mounted on the spherical tanks with additional thrusters mounted on the \Jspacecraft\j body and solar panels. Attitude was maintained through the use of a three-axis control system with pointing maintained with sun and star sensors.
Phobos 1 operated normally until an expected communications session on 2 September 1988 failed to occur. The failure of controllers to regain contact with the \Jspacecraft\j was traced to an error in the software uploaded on 29/30 August which had deactivated the attitude thrusters. This resulted in a loss of lock on the Sun, resulting in the \Jspacecraft\j orienting the solar arrays away from the Sun, thus depleting the batteries.
Phobos 2 operated normally throughout its cruise and Mars orbital insertion phases, gathering data on the Sun, interplanetary medium, Mars, and \JPhobos\j. Shortly before the final phase of the mission, during which the \Jspacecraft\j was to approach within 50 m of \JPhobos\j' surface and release two landers, one a mobile 'hopper', the other a stationary platform, contact with \JPhobos\j 2 was lost. The mission ended when the \Jspacecraft\j signal failed to be successfully reacquired on 27 March 1989. The cause of the failure was determined to be a malfunction of the on-board computer.
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"Mercury Photo Gallery",167,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Pluto Photo Gallery",168,0,0,0
Click on the \BCaption\b button to read a description about this image.
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"Uranus Photo Gallery",169,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Neptune Photo Gallery",170,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Venus Photo Gallery",171,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Asteroid Photo Gallery",172,0,0,0
This page contains a series of nine images of \Jasteroids\j taken from the \JGalileo\j \Jspacecraft\j.
The \JPhobos\j and Deimos images were obtained by the Viking Orbiter \Jspacecraft\j in 1977.
The \JGalileo\j project, whose primary mission is the exploration of the Jupiter system in 1995-97, is managed for NASA's Office of Space Science and Applications by the Jet Propulsion Laboratory.
\BPictures:\b Courtesy of NASA
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"Comets",173,0,0,0
This page contains several images. Click on the \BCaption\b button to read a description for each image.
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"Earth and Moon Photo Gallery",174,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Earth Photo Gallery",175,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Jupiter Photo Gallery",176,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Mars Photo Gallery",177,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Moon Photo Gallery",178,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Saturn Photo Gallery",179,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
Planned objectives were deployment of Tracking Data Relay Satellite-2 (TDRS-2) and flying of Shuttle-Pointed Tool for \JAstronomy\j (SPARTAN-203)/Halley's \JComet\j Experiment Deployable, a free-flying module designed to observe tail and coma of Halleys \Jcomet\j with two ultraviolet spectrometers and two cameras.
Other payloads were Fluid Dynamics Experiment (FDE); \JComet\j Halley Active Monitoring Program CHAMP); Phase Partitioning Experiment (PPE); three Shuttle Student Involvement Program (SSIP) experiments; and set of lessons for Teacher in Space Project (TISP).
\BLaunch:\b
January 28, 1986,11:38:00 a.m. EST. First Shuttle liftoff scheduled from Pad B. Launch set for 3:43 p.m. EST, Jan. 22, slipped to Jan. 23, then Jan. 24, due to delays in mission 61-C. Launch reset for Jan. 25 because of bad weather at transoceanic abort landing (TAL) site in \JDakar\j, \JSenegal\j.
To utilize \JCasablanca\j (not equipped for night landings) as alternate TAL site, T-zero moved to morning liftoff time. Launch postponed a day when launch processing unable to meet new morning liftoff time. Prediction of unacceptable weather at KSC led to launch rescheduled for 9:37 a.m. EST, Jan. 27. Launch delayed 24 hours again when ground servicing equipment hatch closing fixture could not be removed from orbiter hatch.
Fixture sawed off and attaching bolt drilled out before closeout completed. During delay, cross winds exceeded return-to-launch-site limits at KSC's Shuttle Landing Facility. Launch Jan. 28 delayed two hours when hardware interface module in launch processing system, which monitors fire detection system, failed during liquid \Jhydrogen\j tanking procedures.
Just after liftoff at .678 seconds into the flight, photographic data show a strong puff of gray smoke was spurting from the vicinity of the aft field joint on the right Solid Rocket Booster. Computer graphic analysis of film from pad cameras indicated the initial smoke came from the 270 to 310-degree sector of the \Jcircumference\j of the aft field joint of the right Solid Rocket Booster.
This area of the solid booster faces the External Tank. The vaporized material streaming from the joint indicated there was not complete sealing action within the joint.
Eight more distinctive puffs of increasingly blacker smoke were recorded between .836 and 2.500 seconds. The smoke appeared to puff upwards from the joint. While each smoke puff was being left behind by the upward flight of the Shuttle, the next fresh puff could be seen near the level of the joint.
The multiple smoke puffs in this sequence occurred at about four times per second, approximating the frequency of the structural load dynamics and resultant joint flexing. As the Shuttle increased its upward velocity, it flew past the emerging and expanding smoke puffs. The last smoke was seen above the field joint at 2.733 seconds.
The black color and dense composition of the smoke puffs suggest that the grease, joint \Jinsulation\j and rubber O-rings in the joint seal were being burned and eroded by the hot propellant gases.
At approximately 37 seconds, Challenger encountered the first of several high-altitude wind shear conditions, which lasted until about 64 seconds. The wind shear created forces on the vehicle with relatively large fluctuations. These were immediately sensed and countered by the guidance, navigation and control system.
The steering system (thrust vector control) of the Solid Rocket Booster responded to all commands and wind shear effects. The wind shear caused the steering system to be more active than on any previous flight.
Both the Shuttle main engines and the solid rockets operated at reduced thrust approaching and passing through the area of maximum dynamic pressure of 720 pounds per square foot. Main engines had been throttled up to 104 percent thrust and the Solid Rocket Boosters were increasing their thrust when the first flickering flame appeared on the right Solid Rocket Booster in the area of the aft field joint.
This first very small flame was detected on image enhanced film at 58.788 seconds into the flight. It appeared to originate at about 305 degrees around the booster \Jcircumference\j at or near the aft field joint.
One film frame later from the same camera, the flame was visible without image enhancement. It grew into a continuous, well-defined plume at 59.262 seconds. At about the same time (60 seconds), telemetry showed a pressure differential between the chamber pressures in the right and left boosters. The right booster chamber pressure was lower, confirming the growing leak in the area of the field joint.
As the flame plume increased in size, it was deflected rearward by the aerodynamic slipstream and circumferentially by the protruding structure of the upper ring attaching the booster to the External Tank. These deflections directed the flame plume onto the surface of the External Tank. This sequence of flame spreading is confirmed by analysis of the recovered wreckage. The growing flame also impinged on the strut attaching the Solid Rocket Booster to the External Tank.
The first visual indication that swirling flame from the right Solid Rocket Booster breached the External Tank was at 64.660 seconds when there was an abrupt change in the shape and color of the plume. This indicated that it was mixing with leaking \Jhydrogen\j from the External Tank.
Telemetered changes in the \Jhydrogen\j tank pressurization confirmed the leak. Within 45 milliseconds of the breach of the External Tank, a bright sustained glow developed on the black-tiled underside of the Challenger between it and the External Tank.
Beginning at about 72 seconds, a series of events occurred extremely rapidly that terminated the flight. Telemetered data indicate a wide variety of flight system actions that support the visual evidence of the photos as the Shuttle struggled futilely against the forces that were destroying it.
At about 72.20 seconds the lower strut linking the Solid Rocket Booster and the External Tank was severed or pulled away from the weakened \Jhydrogen\j tank permitting the right Solid Rocket Booster to rotate around the upper attachment strut. This rotation is indicated by divergent yaw and pitch rates between the left and right Solid Rocket Boosters.
At 73.124 seconds,. a circumferential white vapor pattern was observed blooming from the side of the External Tank bottom dome. This was the beginning of the structural failure of \Jhydrogen\j tank that culminated in the entire aft dome dropping away.
This released massive amounts of liquid \Jhydrogen\j from the tank and created a sudden forward thrust of about 2.8 million pounds, pushing the \Jhydrogen\j tank upward into the intertank structure. At about the same time, the rotating right Solid Rocket Booster impacted the intertank structure and the lower part of the liquid oxygen tank. These structures failed at 73.137 seconds as evidenced by the white vapors appearing in the intertank region.
Within milliseconds there was massive, almost explosive, burning of the \Jhydrogen\j streaming from the failed tank bottom and liquid oxygen breach in the area of the intertank.
At this point in its trajectory, while traveling at a Mach number of 1.92 at an altitude of 46,000 feet, the Challenger was totally enveloped in the explosive burn. The Challenger's reaction control system ruptured and a hypergolic burn of its propellants occurred as it exited the oxygen-hydrogen flames. The reddish brown colors of the hypergolic fuel burn are visible on the edge of the main fireball.
The Orbiter, under severe aerodynamic loads, broke into several large sections which emerged from the fireball. Separate sections that can be identified on film include the main engine/tail section with the engines still burning, one wing of the Orbiter, and the forward fuselage trailing a mass of umbilical lines pulled loose from the payload bay.
The Explosion 73 seconds after liftoff claimed crew and vehicle. Cause of explosion was determined to be an O-ring failure in right SRB. Cold weather was a contributing factor. Launch Weight: 268,829 lbs.
Orbit:
Altitude: 150nm (planned)
Inclination: 28.5 degrees (planned)
Orbits: 0
Duration: 01 min 13 seconds
Distance: 18 miles
Hardware:
SRB: BI-026
SRM: L025(HPM)
ET : 26/LWT-19
MLP : 2
SSME-1: SN-2023
SSME-2: SN-2020
SSME-3: SN-2021
Landing:
None. KSC Landing planned after a 6 day, 34 minute mission.
Mission Highlights:
The planned orbital activities of the Challenger 51-L mission were as follows:
On Flight Day 1, after arriving into orbit, the crew was to have two periods of scheduled high activity. First they were to check the readiness of the TDRS-B satellite prior to planned deployment. After lunch they were to deploy the satellite and its Inertial Upper Stage (IUS) booster and to perform a series of separation maneuvers. The first sleep period was scheduled to be eight hours long starting about 18 hours after crew wake up the morning of launch.
On Flight Day 2, the \JComet\j Halley Active Monitoring Program (CHAMP) experiment was scheduled to begin. Also scheduled were the initial "teacher in space" (TISP) video taping and a firing of the orbital maneuvering engines (OMS) to place Challenger at the 152-mile orbital altitude from which the Spartan would be deployed.
On Flight Day 3, the crew was to begin pre-deployment preparations on the Spartan and then the satellite was to be deployed using the remote manipulator system (RMS) robot arm. Then the flight crew was to slowly separate from Spartan by 90 miles.
On Flight Day 4, the Challenger was to begin closing on Spartan while Gregory B. Jarvis continued fluid dynamics experiments started on day two and day 3. Live telecasts were also planned to be conducted by Christa McAuliffe.
On Flight Day 5, the crew was to rendezvous with Spartan and use the robot arm to capture the satellite and re-stow it in the payload bay.
On Flight Day 6, re-entry preparations were scheduled. This included flight control checks, test firing of maneuvering jets needed for reentry, and cabin stowage. A crew news conferences was also scheduled following the lunch period.
On Flight Day 7, the day would have been spent preparing the Space Shuttle for deorbit and entry into the atmosphere. The Challenger was scheduled to land at the Kennedy Space Center 144 hours and 34 minutes after launch.
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"Space Shuttle, Living in a",181,0,0,0
\JSpace Shuttle Living\j
\JShuttle, The Air Inside\j
\JShuttle Meals\j
\JShuttle Sanitation\j
\JSpace Suit, Unisex\j
\JShuttle Recreation and Sleep\j
\JShuttle Weightlessness\j
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"Space Shuttle Living",182,0,0,0
The idea that ordinary people would someday live and work in space has fascinated science fiction fans as well as serious scientists and engineers. NASA's Space Shuttle is the first step in turning this dream into reality.
The Space Shuttle is a reusable aerospace vehicle that takes off like a rocket, can be maneuvered in space, and lands like an airplane.
The \Jspacecraft\j, called the orbiter, is about the size of a DC-9 commercial jetliner. The orbiter carries people and cargo between the ground and Earth orbit. It can also be used as an observation post in space and as a space platform for a fully equipped laboratory for medical, scientific, \Jengineering\j, and industrial experiments.
One of the key attributes of the Shuttle and its operation is the relatively low g-force exerted on crew and passengers during launch and reentry. Launch and reentry forces are less than 4 g's -- well within the limits which can be tolerated by healthy people.
Orbiter living accommodations are relatively comfortable. They incorporate advances made through nearly 2 decades of experimental manned space missions and an even longer period of ground studies.
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"Shuttle, The Air Inside",183,0,0,0
The orbiter's air is cleaner than Earth's, and hay fever sufferers will welcome its pollen-free atmosphere.
Orbiter air pressure is the same as Earth's at sea level: 1,033 grams per square centimeter (14.7 pounds per square foot). Its air is made up of 80 percent \Jnitrogen\j and 20 percent other gases such as \Jargon\j and neon. The orbiter's environmental control system circulates air through filters to remove carbon dioxide and other impurities. Excess moisture is also removed, keeping \Jhumidity\j at comfortable levels. Temperature in the orbiter can be regulated between 16 and 32 degrees \JCelsius\j (61 and 90 degrees Fahrenheit). The orbiter crew requires only ordinary clothing. People can move about, work, and relax unencumbered by bulky space suits.
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"Shuttle Meals",184,0,0,0
Shuttle meals are tasty and nutritious. They can be eaten anywhere, although crew members normally congregate in the middeck area for their meals. Trays holding the food can be attached to a crew member's legs or to any orbiter surface with adhesive straps, removing the need for a table and chairs at mealtime. Meals are served in a special tray which separates the different food containers and keeps them from lifting off and soaring around in the weightless cabin.
Packages of food that have to be warmed are placed in the galley oven before going into the tray. Hot and cold water are available for preparation of foods or beverages.
Studies have shown that despite zero gravity, most foods can be eaten with ordinary spoons and forks as long as there are no sudden starts, stops, or spinning. As a result, dining in space is almost like dining on Earth.
The orbiter menu includes more than 70 food items and 20 beverages. With so many different items, Shuttle travelers can have varied menus every day for 6 days.
Earth-bound chefs might envy orbiter meal preparation -- one crewmember can ready meals for four people in about 5 minutes.
What are orbiter meals like? A typical day's menus include orange drink, peaches, scrambled eggs, sausage, \Jcocoa\j, and a sweet roll for breakfast; cream of mushroom soup, ham and cheese sandwich, stewed tomatoes, banana, and cookies for lunch; and shrimp cocktail, beefsteak, broccoli au gratin, strawberries, pudding, and \Jcocoa\j for dinner.
Menus provide about 2,700 calories daily. Previous space missions demonstrated that astronauts need at least as many calories in space as they do on Earth.
The orbiter does not have a refrigerator. Most of the Shuttle foods are preserved by \Jdehydration\j, which saves weight and storage space. Water for rehydration is ample since it is a byproduct of the fuel cells which generate electricity. Some foods are thermostabilized, that is, they are heat sterilized and then sealed in conventional cans or plastic pouches. A few, such as cookies and nuts, are available in ready-to-eat form.
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"Shuttle Sanitation",185,0,0,0
Eating utensils are cleaned with wet wipes. The difference between orbiter wet wipes and those used on Earth is that the orbiter's contain a strong \Jdisinfectant\j.
Sanitation is more important in the confines of the orbiter than on Earth. Space studies have shown the population of some microbes can increase extraordinarily in a confined weightless area such as a \Jspacecraft\j cabin. This could potentially spread illness to everyone on board. As a result, not only eating components but also the dining area, the toilet, and sleeping areas are regularly cleaned. Since there are no washing machines in space, trousers (changed weekly), socks, shirts, and underwear (changed every 2 days) are sealed in airtight plastic bags after being worn. Garbage and trash also are sealed in plastic bags.
A favorite question of people interested in space is how the astronauts took care of digestive elimination. The orbiter travelers use a toilet very much the same as one on Earth. Air flow directs waste to the bottom of the toilet, substituting for gravity. Waste goes directly into a sealed container where it is processed and stored.
Some of the waste may be used for post-flight laboratory analyses. Such analyses have told doctors which minerals are lost excessively in space and have helped to increase their understanding of body functions.
Orbiter travelers have facilities and supplies available for sponge baths while in space. They can obtain water from the water dispensing system. Water temperature can be set at any comfortable level from 18 to 35 degrees \JCelsius\j (65 to 95 degrees Fahrenheit).
Because of weightlessness, water droplets would float about in the cabin. This could be not only a nuisance but also potentially hazardous to equipment and crew. To prevent this from happening, an airflow system directs waste water into the orbiter's waste collection system, where the waste water is sealed in plastic watertight bags.
Whiskers cut off in shaving and floating about weightlessly in a cabin could be a nuisance and foul up equipment. This problem is avoided by using conventional shaving cream and a safety razor and cleaning off the face with a disposable towel. Also available is a wind-up shaver that works like an electric razor and contains a vacuum device to prevent the escape of cut whiskers.
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"Space Suit, Unisex",186,0,0,0
In the past, space suits were tailor-made for each \Jastronaut\j, a time-consuming and costly process. The Shuttle space suit is manufactured in small, medium, and large sizes and can be worn by men or women. The suit comes with an upper and lower torso equivalent to a shirt and trousers. Each piece snaps together with sealing rings. A life-support system is built into the upper torso. Previous pressure suits had separate life support systems which had to be connected to the suits.
The Shuttle space suit is lighter, more durable, and easier to move about in than previous space suits. When an \Jastronaut\j has to work outside the space- craft, the Shuttle suit is used for extravehicular activity.
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"Shuttle Recreation and Sleep",187,0,0,0
Just as on Earth, recreation and sleep are important to good health in space. A scientifically planned exercise program is provided, largely as a countermeasure for cardiovascular deconditioning and atrophy of muscles in a weightless environment. Cards and other games, books, writing material, and tape recorders and tapes to chronicle personal observations or to listen to music, are available.
#
"Shuttle Weightlessness",188,0,0,0
Many of the problems of going into space have been resolved. However, the physiological effects of weightlessness are still not completely understood. Among them are leaching of minerals from bones, reduction in rate of bone formation, atrophy of muscles when not exercised, and motion sickness.
All of the effects of zero gravity have so far been reversed after return to the normal gravity on Earth. In addition, some of the effects have been countered by exercise and food supplements.
However, even vigorous exercise in space does not appear to stop bone loss or decrease in the rate of bone formation. As a result, NASA is engaged in an intense and sustained effort aimed at understanding the causes underlying these changes and developing ways to prevent them. The increased information about body functions derived from this effort will pave the way for prolonged missions in space and contribute to our understanding of the \Jphysiology\j of living things on Earth.
#
"Shuttle Location",189,0,0,0
\JShuttle Location Over Earth\j
\JShuttle Position in the Universe\j
#
"Shuttle Location Over Earth",190,0,0,0
When a space shuttle crew wants to know where it is over the Earth, the easiest way to find out is to look out a window. Sometimes, when the shuttle is orbiting over the middle of the Pacific Ocean, or when its windows are facing away from the Earth, this can be a problem.
To help the astronauts with their situational awareness and to provide them with information for documenting their Earth observations with cameras, the flight control team provides them with an application called WORLDMAP which runs on a Payload and General Support Computer (PGSC). These commercially available laptops (at this time, they're using an \JIBM\j Thinkpad 755) are used to run various applications on-board. WORLDMAP is used to display the orbiter's location and to display the location of various Earth observation sites. WORLDMAP uses a state vector to determine its position over the Earth.
The image shown is a snapshot of the same program--developed by the SpOC (Space Operations Computing) team programmers at the Johnson Space Center and running in Microsoft Windows--that the crew on board is using. The image may not be an exact copy of the one used by the astronauts, since they can change the program's preferences, but they could duplicate this view if they so desired.
Both the flight version and the ground version of the program access the orbiter state vector information via the shuttle's pulse code master \Jmodulation\j unit (PCMMU, pronounced "puckamoo"). The PCMMU data stream is read into another commercially available program called PCDeCOM, which extracts specified data parameters and feeds them to the WORLDMAP program.
The small window in the lower left portion of the image shows the overall world view. The largest part of the image shows the Earth below from the same vantage, but magnified eight times. The application's title bar includes the current Greenwich Mean Time (GMT), Mission Elapsed Time (MET), latitude, longitude and altitude.
The map title bar displays the name of the country the shuttle is flying over. The timer window provides a variety of information, including the next S-band and Ku-band acquisition and loss of signal through the Tracking and Data Relay Satellite System, and sunrise and sunset times (remember that the sun rises or sets every 45 minutes on orbit).
Sometimes, areas will be outlined in red and red numerals will be displayed in the large map. These are upcoming Earth observations photography opportunities. The crew on board can click on these red areas and get additional information about the site and the best lenses and film speeds to use when photographing the area.
Flight controllers in the Mission Control Center use a different display called the Distributed Earth Modeling and Orbiter System (DEMOS) to help them visualize the location and attitude of the shuttle.
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"Shuttle Position in the Universe",191,0,0,0
The shuttle's ground track over the Earth's surface is shown on the Distributed Earth Model and Orbiter System (DEMOS).
Flight controllers in NASA's Mission Control Center use this DEMOS computer modeling software to help them visualize exactly where the space shuttle is over the Earth and what attitude it is in (how is it facing in three-dimensional perspective to the Earth). Crewmembers aboard the shuttle can access the same information using a laptop computer and WORLDMAP software.
These computer-generated images are translated into \Jtelevision\j signals that are displayed in the Flight Control Room, and frequently are broadcast on NASA \JTelevision\j. The NASA Shuttle Web uses a frame-grabber to produce still images of the DEMOS display every few minutes. If you are using Netscape as your browser, the image will be updated automatically. If not, you must reload to update the image. Image quality will vary depending on the quality of the \Jtelevision\j signal at the the frame was grabbed.
DEMOS uses the same state vectors used by NASA's Mission Control Center to keep track of the Shuttle. The state vector defines very exactly the position of the \Jspacecraft\j. It contains orbital elements, also called Keplerian elements, such as the tilt of the orbit plane, altitude, and how circular (round) the orbit is.
The Mission Control Center in Houston regularly updates the state vector on the shuttle's General Purpose Computers so that the shuttle's navigation systems will know exactly where the \Jspacecraft\j is and where it will be in the future.
You can use these same orbital elements to track \Jspacecraft\j on your home computer system. There are a number of freeware and shareware tracking programs that will allow you to input the orbital elements and follow a \Jspacecraft\j (or a planet or a satellite for that matter) from home. To use these programs, you will need both the software and the latest orbital elements.
You must update these elements regularly (usually about once every 12 hours, but more often when the \Jspacecraft\j is maneuvering). Use these software programs at your own risk; NASA takes no responsibility for their effect on the stability of your computer system.
The Launch Control Center (LCC) is a four-story building that is the electronic "brain" of Launch Complex 39. Attached to the southeast corner of the Vehicle Assembly Building (VAB) , it is 5,535 meters (18,159 ft) from Pad 39A . At the time it was constructed, advances in \Jelectronics\j had made it unnecessary to continue locating blockhouses adjacent to launch pads.
The first floor contains offices and computer operations. The second floor houses telemetry, RF and tracking, instrumentation, and data reduction and evaluation equipment. The computers of the Central Data Subsystem (CDS), one of the two major components of the Launch Processing System (LPS) that automatically performs most checkout and launch functions, is also on the second floor.
The third floor contains the four firing rooms, and each room contains its own copy of the second major component of the Launch Processing System - the Checkout, Control and Monitor Subsystem (CCMS).
The system consoles of the CCMS system are manned by the team which oversees all aspects of a checkout and launch operation. Firing rooms 1 and 3 are configured for full control of launch and orbiter operations while Firing room 2 is usually used for software development and testing. Firing room 4 is only a partial firing room and is primarily used as an \Jengineering\j analysis and support area for launch and checkout operations.
The fourth floor of the LCC contains conference rooms, offices and mechanical equipment. The LCC is 23.5 meters (77ft) high, 115.2 meters (378ft) long and 55.1 meters (181 ft) wide.
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"Mobile Launch Platforms (MLP)",194,0,0,0
The three Mobile Launchers used in Apollo/Saturn operations were modified for use in Shuttle operations. With cranes, umbilical towers, and swing arms removed, the Mobile Launchers were redesignated Mobile Launcher Platforms (MLP). In place of one large opening in the platform, three smaller openings accommodate flames and hot exhaust gases from the solid rocket boosters and the orbiter engines.
Segments of the dismantled umbilical towers are part of the permanent installation at the launch pad, where they serve as sections of the Fixed Service Structure (FSS). A third Apollo umbilical tower, removed from MLP-3, has been cut into 20 ft sections and placed in a field in the KSC industrial area. It may someday become reconstructed as part of the KSC tour route.
Tail Service Masts (TSM's), one on each side of the main engines exhausts hole provide umbilical connections for fuel and oxidizer, gases, ground electrical power and communications links. These Masts are 4.6m (15ft) long, 2.7m (9ft) wide and 9.4m (31 ft) high.
MLP Statistics
Launch Platform: Two-Story steel structure, 7.6 meters (25ft) high, 49 Meters (160ft) long and 41 meters (135ft)wide.
ò Empty Weight: 4.19 million kg (9.25 million lb)
ò With unfueled Shuttle: 5.45 million kg (12.02 million lb)
ò With fueled Shuttle: 6.22 million kg (13.72 million lb)
Positioned on 6 steel pedestals 7m (22ft) high when in the VAB or at the launch pad. At the pad, 4 extensible columns were used during the Apollo program were used to stiffen the MLP against rebound loads, should engine cutoff occur. They are no longer used for the Shuttle program.
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"Operations and Checkout Building",195,0,0,0
Horizontally integrated payloads are received, assembled and integrated in the Operations and Checkout Building before they are mated with the orbiter at the Orbiter Processing Facility. The Spacelab and its payloads constitute a large majority of the horizontal payloads.
The Operations and Checkout Building is a five-story structure containing 600,000 square feet of offices, laboratories, \Jastronaut\j crew quarters and \Jspacecraft\j assembly areas. It is located in the industrial area immediately east of the KSC headquarters building.
The \JSpacecraft\j Assembly and Encapsulation Facility -2 (SAEF-2) is located at F Avenue and 7th Street in the Hypergol Maintenance Facility Area, KSC Industrial Area. The facility is used for the assembly, test, encapsulation, ordnance work, propellant loading, and pressurization of \Jspacecraft\j. The facility contains approximately 1556.08 meters (m)2 (16,750 feet (ft)2) of usable floor space. Construction is of reinforced concrete and concrete block. The high bay is a steel frame structure with insulated aluminum siding.
Functionally, the building is divided into the following areas: a clean work area (CWA) complex consisting of an airlock, a high bay, and two low bays; a test cell; a sterilization oven (non- operational); support office areas; and mechanical equipment rooms (figure 2-3).
Floors in the airlock, high bay, and test cell are designed for 3175.20 \Jkilogram\j (kg) (7000 pounds (lb)) per wheel plus 20 percent impact loading. The airlock, located at the north end of the building, measures 12.5 m by 17.7 m (41 ft wide by 58 ft long), providing a usable floor area of 221 m\U2\u (2378 ft\U2\u) and is rated as a Class 300,000 CWA. The airlock has a clear ceiling height of 15.9 m (52 ft).
Access is by means of personnel doors, vestibule, and a 6.4 m by 12.2 m (21.5 ft wide by 40 ft high) vertical lift equipment door. A 6.5 m by 12.1 m (21 ft wide by 39.5 ft high) horizontal sliding lift door separates the airlock from the adjacent high bay.
The high bay measures 14.9 m by 30.2 m (49 ft wide by 99 ft long), providing a usable floor area of 450.7 m2 (4851 ft\U2\u); clear ceiling height is 22.6 m (74 ft) and is rated as a Class 100,000 CWA. Personnel and small equipment can enter the high bay through the equipment airlock, equipped with air showers. Clear access is 1.4 m by 2.1 m (4 ft 5 in wide by 6 ft 11 in high).
Two large bays are located along the west side of the high bay. One of the bays measures 5.8 m by 21.9 m (19 ft wide by 72 ft long) with a clear ceiling height of 7.62 m (25 ft); the other bay is 5.8 m by 8.2 m (19 ft wide by 27 ft long) and has a clear ceiling height of 13.3 m (43.5 ft). The combined bay areas provide a usable floor space of 174.8 m2 (1881 ft2). The low bays are also rated as Class 100,000 CWA's.
The test cell is located at the northeast corner of the facility. The cell measures 11.3 m by 11.3 m (37 ft by 37 ft), providing a usable floor area of 127.2 m2 (1369 ft2). The clear ceiling height is 15.9 m (52 ft). Access to the test cell is by means of personnel doors and three 6.7 m by 12.2 m (22 ft wide by 40 ft high) vertical lift doors. The test cell can be used for \Jspacecraft\j and payload support activities not requiring Class 100,000 CWA conditions.
The remainder of the building consists of such support rooms as a mechanical equipment room, communication equipment room, entry and observation room, change room, storage room, miscellaneous equipment rooms, and limited office space.
The sterilization oven is located at the south end of the facility in a 13.1 m by 16.2 m (43 ft wide by 53 ft long) open- sided shed. Ceiling height of the shed in the oven area is 4.9 m (16 ft). (The oven is not operational.)
There is a 13.7 m by 8.5 m (45 ft long by 28 ft wide) conference area located on the second floor above rooms 117 and 119. Room 109 is used as the clean room garment storage and issue room. There are five 3.7 m by 18.3 m (12 ft by 60 ft) office trailers, figure 2-3, available for payload personnel use when their payload is in SAEF-2. Two are on the north side of the building; two, on the west side; and one, on the east side. One of the trailers located on the north side is used by the NASA and PGOC facility managers.
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"Space Station Processing Facility (SSPF)",197,0,0,0
The SSPF is located in the KSC industrial area, just east of the Operations and Checkout Building. It was built for the processing of the International Space Station flight hardware. The three-story SSPF, a 457,000 square foot building, includes two processing bays, an airlock, operational control rooms, laboratories, logistics areas, office space, and a cafeteria. The processing areas, airlock, and laboratories were designed to support non-hazardous Station and Shuttle payloads in 100,000 class clean work areas.
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"Vehicle Assembly Building",198,0,0,0
The Vehicle Assembly Building (VAB) is one of the largest buildings in the world. It was originally built for assembly of Apollo/Saturn vehicles and was later modified to support Space Shuttle operations. High Bays 1 and 3 are used for \Jintegration\j and stacking of the complete Space Shuttle vehicle.
High Bay 2 is used for external tank (ET) checkout and storage and as a contingency storage area for orbiters. High Bay 4 is also used for ET checkout and storage, as well as for payload canister operations and solid rocket booster (SRB) contingency handling.
The Low Bay area contains Space Shuttle main engine maintenance and overhaul shops, and serves as a holding area for SRB forward assemblies and aft skirts.
During Space shuttle build-up operations inside the VAB, integrated SRB segments are transfered from nearby SRB assembly and checkout facilities, hoisted onto a Mobile Launcher Platform in High Bays 1 or 3 and mated together to form two complete SRBs. The ET, after arrival by barge, is inspected and checked out in High Bays 2 or 4 and then transfered to High Bay's 1 or 3 to be attached to the SRBs already in place.
The orbiter is then towed over from the Orbiter Processing Facility to the VAB transfer aisle, raised to a vertical position, lowered onto the Mobile Launcher Platform and then mated to the rest of the stack. When assembly and checkout is complete, the crawler-transporter enters the High Bay, picks up the platform and assembled shuttle vehicle and carries them to the launch pad.
The VAB covers 3.25 hectares (8 acres). It is 160 meters (525 ft) tall, 218 meters (716 ft) long and 158 meters (518 ft) wide. It encloses 3,664,883 cubic meters (129,428,000 cubic feet) of space.
ò Flag & Bicentennial Emblem: Added in 1976, required 6,000 gallons of paint. The flag is 64 x 33.5 meters (209 x 110 ft) in size. Each strip on the flag is as big as the tour buses used to transport visitors around KSC
ò Doors: There are 4 High Bay doors. Each opening is 139 meters (456 ft) high. The north entry to the transfer aisle was widened 12.2 meters (40ft) to permit entry of the Orbiter, and slotted at the center to accommodate its vertical stabilizer.
Comparisons:
ò Height: VAB - 160meters (525 ft) Statue of Liberty - 93 meters (305 ft)
ò Volume: VAB - 3,665,013 cu meters (129,428,000 cub ft) Pentagon 2,181,117 cu meters (77,025,000 cu ft).
ò VAB equals 3.75 Empire State Buildings
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"Shuttle Flights To Date",199,0,0,0
(Refer to Table)
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"Space Shuttle Launch Team",200,0,0,0
\JShuttle Launch Team\j
\JShuttle Launch Team Structure\j
\JCountdown To Launch\j
\JShuttle Operation Communication\j
\JShuttle Firing Room Protocol\j
\JLaunch Pad Activities During Countdown\j
\JLaunch: Final Go\j
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"Shuttle Launch Team",201,0,0,0
Launch Complex 39 at the Kennedy Space Center has served as the springboard for U.S. human spaceflight since the Apollo lunar landing program in the 1960s. The first launch conducted from one of the two LC 39 pads was Apollo 4 on Nov. 9, 1967. Since then, teams staffing the consoles in one of the Launch Control Center firing rooms have sent into space astronauts who walked on the moon, interplanetary explorer \Jspacecraft\j destined for the far reaches of the solar system and Space Shuttle crews to conduct research and to service, deploy and retrieve \Jspacecraft\j.
The first Space Shuttle launch was conducted from Pad 39A on April 12, 1981. Today, Shuttles lift off regularly from either Pad 39A or B under the management of the launch team in the Launch Control Center.
The Space Shuttle launch team is a highly organized and disciplined group of approximately 500 professionals. Membership of the launch team reflects the complexity involved in preparing for and conducting a human space launch. Civil service and contractor personnel from Kennedy occupy central roles, but other NASA centers, contractor personnel and agencies also contribute.
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"Shuttle Launch Team Structure",202,0,0,0
The Shuttle launch team is organized into three groups, according to major functional responsibility. All three groups report to the Shuttle Launch Director, who has overall technical and safety responsibility for the countdown.
1. The prime launch team is responsible for test, checkout and monitoring of the flight hardware and ground support equipment to ensure that all system parameters meet the criteria to commit the vehicle to launch. This team includes the group stationed in the prime firing room in the Launch Control Center, as well as offsite personnel with critical launch support responsibilities. The prime launch team is headed by the NASA Test Director (NTD), who reports to the Shuttle Launch Director.
Approximately 200 of the 300 members of this team are stationed in the prime firing room for the mission at hand. The NTD and other managers overseeing the countdown process are stationed at consoles facing into the firing room. Reporting to the NTD are 11 test conductors, each responsible for a specific subset of requirements to be met prior to launch.
For example, the Orbiter Test Conductor (OTC) is the individual in charge of the system-level engineers monitoring the hardware and software on board the orbiter itself. Other test conductors are responsible for the external tank and solid rocket boosters, the payloads, support operations, landing operations, safety, communications and other functions.
Also stationed in the prime firing room is the Shuttle Project Engineer, who is the lead engineer and reviews all technical issues for the NTD.
The \Jastronaut\j assigned to command the Shuttle mission represents his flight crew as he communicates with the NTD from inside the vehicle at the pad.
Seated at seven console sets facing the windows of the prime firing room and looking toward the two launch pads are the teams of system-level experts who report to the test conductors. They represent all orbiter, external tank and solid rocket booster components. Mission-specific payload experts are assigned to this team also.
Also part of this configuration is an eighth console called the \Jintegration\j console, which includes the Ground Launch Sequencer that automatically controls the final nine minutes of the countdown.
At the rear of the room is the Master Console, which monitors the Launch Processing System computer network with which all Shuttle processing, checkout and launch are performed.
Each console is arranged in its own semicircle and includes up to a dozen individual operator stations. The number of stations dedicated to a major function, such as payloads, varies. Besides the software giving the operator control over a particular system, each console has radio hookups and video displays monitoring Shuttle and pad hardware. Click here for photo of a typical console.
A console chief is assigned to each console. This typically is a senior system engineer. One of the console chief's functions is to communicate with personnel in an adjacent backup firing room so the expertise of both rooms is combined.
The system engineers at the seven console sets have their own method of assigning the various responsibilities that encompass their particular set of requirements. For example, the lead engineer for the liquid \Jhydrogen\j system has approximately eight engineers reporting to him or her, each of whom are responsible for a portion of the entire system. This includes the storage tank at the pad, pressurization systems, the cross-country pipelines used to load the external tank, the \Jhydrogen\j portion of the tank itself, and so on. In turn, each of these elements has its own set of responsibilities and requirements to monitor.
Seated behind the console operators are additional engineers and technicians. They are available to provide technical expertise or to help troubleshoot any glitches that may occur during the countdown.
Approximately 100 members of the primary launch team are located outside the confines of the Launch Control Center, yet oversee systems and hardware critical to the launch process. These include representatives from Johnson Space Center, which manages the Space Shuttle program and is responsible for on-orbit operations. U.S. and overseas contingency landing site readiness is the JSC team's job also. The Eastern Range, managed by the U.S. Air Force, is responsible for ensuring range safety as well as providing weather information. Located nearby is the Merritt Island Launch Area (MILA) Station, which provides voice and data transmission paths to and from the vehicle. Goddard Space Flight Center manages the MILA station.
2. The \Jengineering\j support team has a similar composition and organization to that of the launch team, but is not directly responsible for system management. Acting in more of an oversight capacity, they provide technical support should problems arise. This extra set of eyes is composed of extremely experienced engineers who can assist in solving problems in real time. These engineers are stationed primarily in the backup firing room, Firing Room 2, located between Firing Rooms 1 and 3 on the LCC third floor. They work at computer hardware and software similar to that of the prime team, but perform systems-monitoring only, with no command capability or responsibility.
Because of space limitations, an additional room on the third floor also is used, called the \JEngineering\j Support Area (ESA). The \Jengineering\j support team also includes personnel outside KSC, such as other centers and contractor facilities. All represent a pool of expertise from which the prime team can draw at a moment's notice if need be.
3. The senior government and contractor managers that comprise the Mission Management Team (MMT) are charged with reporting any issues that may affect the safety or success of the countdown or mission. Reportable issues can originate during any phase of the preflight hardware component processing as well as during the countdown itself. All issues raised must be resolved prior to clearing the launch vehicle for flight. During the countdown, the MMT is located in the Operations Support Room area of the prime firing room.
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"Countdown To Launch",203,0,0,0
The launch of the Space Shuttle marks the finale of many thousands of individual tasks performed by highly trained and motivated workers at the Kennedy Space Center and elsewhere. Approximately four months of system tests, refurbishment and unique configurations for the upcoming mission are required to prepare the Shuttle for its next flight.
The countdown formally begins with the call to stations, issued by the NTD from the Launch Control Center. The countdown clock begins ticking at T-43 hours about three days before liftoff. With built-in hold time included, it takes roughly 72 hours to conduct a Shuttle launch countdown.
The reference tool for conducting a Shuttle launch countdown is a five-volume manual encompassing some 5,000 pages of instructions. More commonly known by its identification number, S0007, the Shuttle countdown manual is the lengthiest Operations and Maintenance Instructions (OMIs) used at KSC to document procedures for assembly, processing, testing and launching of the Shuttle. S0007 is reviewed and updated prior to each mission.
The first volume of S0007 contains all the preparations necessary to lead up to the beginning of the three-day countdown. Volume 2 is the actual set of countdown instructions that logically and sequentially configure the vehicle for launch. It is this integrated set of instructions that contains all the requirements necessary to launch.
Volume 3 contains all the instructions to follow in the event a launch is scrubbed. In this case, the vehicle and facilities are recycled for a subsequent launch attempt. This next attempt could be in as little as 24 hours or could be several days later. Volume 4 contains all the specific individual system instructions that are initiated from Volume 2. Volume 5 is a set of preplanned contingency procedures and emergency instructions available in the unlikely event they are required.
This entire set of instructions is performed under the direction of the NTD.
The beginning of a countdown is not unlike a routine power-up of the orbiter. System checks are conducted and the vehicle configured for later operations. Once preparations get under way to load the orbiter's fuel cell power reaction and storage distribution system at around T-28 hours, the spaceship's configuration is getting more oriented toward flight. The remaining countdown milestones focus on pad and vehicle closeouts: closure of the payload bay doors; retraction of the Rotating Service Structure at the launch pad; installation of the crew escape pole; activation of the onboard fuel cells; loading of the external tank, and boarding of the flight crew.
After the countdown clock starts ticking, the prime firing room is staffed around the clock. Personnel are typically selected for this prime launch team at about the time an orbiter is being transferred from the Orbiter Processing Facility to the Vehicle Assembly Building for \Jintegration\j with the other Shuttle flight elements.
As subsequent countdown milestones are met, the composition of the prime firing room team will change. Assigned shifts of teams will report on station at varying times, depending on which system they oversee. For example, a loading crew comes on station at T-9 hours to begin preparing for cryogenic loading of the Shuttle external tank at the T-6 hour mark. Once its task is completed by the T-3 hour mark, this group of individuals will be succeeded by a fresh crew to carry on through launch.
Another example of a specialized countdown crew are the personnel who oversee the ground launch sequencer (GLS). GLS operators may be on hand earlier in the countdown, but they do not have a formal role until the day of launch. The GLS crew -- a primary operator, two backups and a fourth to retrieve data -- are on station by the T-2 hour mark to call up the GLS software and prepare for the final count. At T-20 minutes, the GLS will begin issuing active commands, and at T-9 minutes, it assumes automatic control of the count.
As they work through S0007, the launch team members at the consoles in the prime firing room are monitoring vehicle and support system performance according to acceptable measurements and parameters. The full set of measurements that must be checked totals about 25,000. About two thirds of these originate from the flight vehicle and one third from the ground support equipment and facilities. Measurements generally fall into one of two categories, Launch Commit Criteria or supporting data.
Launch Commit Criteria are those parameters that have safety-related or mission success implications; they define what constitutes a vehicle that is ready to fly as well as the conditions under which it is permissible to launch. Launch Commit Criteria is implemented at T-6 hours (prior to ET load). There are Launch Commit Criteria that guard against flight hardware damage and those designed specifically for \Jastronaut\j safety. For instance, there is a permissible concentration of gaseous \Jhydrogen\j within the orbiter's aft fuselage that deals with \Jastronaut\j safety and a slightly different permissible concentration that may affect shuttle hardware issues. Even the weather must meet specific Launch Commit Criteria requirements in order for a liftoff to proceed.
The other set of data represents supporting information available for the engineers to help maintain a specific hardware configuration or to aid in troubleshooting problems. An example would be the temperature limits for the \Jhydrogen\j vent line that is used to safely transport the gaseous \Jhydrogen\j from the external tank to its flarestack at the pad for disposal by burning.
This vast data base on system performance is part of the Launch Processing System and is available for the engineers to access at any point in the launch countdown. About 2,300 Launch Commit Criteria measurements are monitored. As the countdown progresses toward liftoff, all attention focuses on the Launch Commit Criteria. All the parameters must be met prior to passing a "Go" for launch.
#
"Shuttle Operation Communication",204,0,0,0
Communication in the prime firing room is carefully routed through the Operational Intercommunications System (OIS). It is a closed-loop digital voice system utilizing fiber optic cable. During countdown, the NTD uses one frequency as the command channel for overall countdown \Jintegration\j. The Test Conductors use separate channels to individually lead their specific subset of the countdown. The Test Conductors and the NASA Test Director communicate with each other on an as-needed basis on issues such as status checks, safety and command responses. The use of different channels separates communication traffic and keeps it at a manageable level.
Acronyms are used liberally by the entire team to keep verbiage to a minimum. All Firing Room console positions are assigned unique 'call signs' that are used by the team for quick and positive identification of who is talking. The protocol has the initiator of the dialogue calling the person he wishes to talk to followed by his own call sign. This allows the person called to know who is calling. For instance,
"OTC, NTD, Begin crew module closeout" means the NTD is directing the OTC to begin preparing the crew module for closing the hatch and launch. Other acronyms are used for system descriptions such as PRSD for the Power Reactant Storage and Distribution System which is the liquid oxygen and liquid \Jhydrogen\j fuels that power the orbiter's on-board fuel cells that create electric power. When the flight crew enters the orbiter, the astronauts will hook up to the same communication channel as the OTC. At the T-20 minute mark, the NTD switches over to this channel and it now becomes the Command channel. The \Jintegration\j console in the prime and backup firing rooms maintain communication channels with the other centers, such as the Mission Evaluation Room at Johnson Space Center in Houston, and the \JHuntsville\j Operations Support Center at the Marshall Space Flight Center in \JAlabama\j. The Mission Evaluation Room plans and implements flight data retrieval, processing, exchange analysis, evaluation and reporting, and post-mission evaluation. The \JHuntsville\j center provides technical support on the Shuttle main engines, external tank and solid rocket boosters.
Hookups also are established with the contractors' home offices. Orbiter manufacturer Rockwell International maintains a support room at its Downey, Calif., plant that remains open around the clock throughout a Shuttle mission.
#
"Shuttle Firing Room Protocol",205,0,0,0
A special type of discipline is exercised in the prime firing room, commensurate with its importance. Launch team members undergo training in the rules and regulations governing their conduct. These include limiting conversation to the business at hand, no personal \Jtelephone\j calls except in emergency, and no reading of non-work related materials. During time-critical operations, personnel remain at their assigned stations.
From T-3 hours on, entrance into the prime firing room is restricted. Only personnel with firing room badges are allowed in and movement is minimized. A prime firing room badge is issued only to personnel having a direct console position related to the terminal portion of the count. At T-20 minutes, the door to the prime firing room is locked. The intent is to eliminate distractions and allow the team to focus its attention on the countdown.
#
"Launch Pad Activities During Countdown",206,0,0,0
While countdown activities are controlled from the LCC prime firing room, personnel at the pad perform different tasks required for launch preparations. From T-11 hours to T-6 hours, a great deal of final preparation work occurs at the pad: rollback of the Rotating Service Structure; installation of time-critical flight crew equipment; performance of the pre-ingress switch list; sampling of crew seat oxygen; and installation of the crew escape pole in the orbiter. Overseeing these activities and keeping the NTD informed of their progress is the pad leader.
After the External Tank is loaded, only critical and highly specialized teams will travel to the pad again before liftoff. One of these is the Final Inspection Team, also referred to as the ice team, which conducts a preflight walkdown of the vehicle and pad during the two-hour hold at T-3 hours. Another ice team is stationed in the backup firing room. Its job is to monitor the external tank's \Jinsulation\j and attachment struts for excessive ice formation before, during and after loading of the supercold liquid \Jhydrogen\j and liquid oxygen. Click here for a picture of the ice team on the Mobile Launch Platform.
Another specialized team is the white room closeout crew, which also proceeds to the pad during the two-hour-hold at the T-3 hour mark. Their task is to insure that the orbiter cockpit is properly configured for flight and to assist the astronauts with entry into the orbiter. They also ensure the side hatch is properly closed and that the white room is configured for launch.
Handling of any anomaly at the pad that should occur during or after external tank loading is the responsibility of a red crew. This is not a pre-existing unit, but a team assembled from a pool of specially trained workers with experience in the particular problem area. Members have been specially trained in fire and rescue techniques and must have undergone special certification. Their activity at the pad would be conducted by the system engineer responsible for the anomalous system and under the strict direction of the NASA Test Director..
#
"Launch: Final Go",207,0,0,0
Beginning approximately 15 minutes before launch, readiness polls are conducted by the three teams that together comprise the Shuttle Launch Team. The NTD verifies that the prime launch team is reporting no violation of the Launch Commit Criteria. The \JEngineering\j Director who heads up the \JEngineering\j Support Team verifies no constraints to continuing with the final count. And the Mission Management Team Chairman verifies that there are no open issues with any of the senior element managers.
These three verifications are passed on to the Shuttle Launch Director, who conducts a KSC management poll. Assuming all responsible personnel are in agreement, the Launch Director gives his permission to proceed with the countdown to the NTD. The NTD in turn sets in motion the final nine minutes of the countdown, automatically controlled by the Ground Launch Sequencer.
Once the Shuttle's twin solid rocket boosters ignite at T-0, responsibility for the mission switches from KSC to the Mission Control Center at Johnson Space Center in Houston. KSC once again assumes responsibility after the orbiter has landed and the flight crew has exited the vehicle.
Jan. 5 President Nixon proposes development of a reusable space transportation system, the Space Shuttle.
March 15 NASA selects the three-part configuration for the Space Shuttle -- reusable orbiter, partly reusable SRB and an expendable external tank.
Aug. 9 Rockwell receives NASA contract for construction of the Space Shuttle orbiter.
1975
Oct. 17 First Space Shuttle main engine tested at the National Space Technology Laboratories, Miss.
Sept. 17 Rollout of orbiter Enterprise (OV-101).
1976
July 18 Thiokol conducts 2-minute firing of an SRB at Brigham City, \JUtah\j.
Aug. 12 First free flight Approach and Landing Test (ALT) of orbiter Enterprise from Shuttle carrier \Jaircraft\j at Dryden Flight Research Center, Calif. Flight duration: 5 minutes, 21 seconds. Landing occurred on Runway 17.
Sept. 13 Second Enterprise ALT flight of 5 minutes, 28 seconds; landing on Runway 15. (Three more ALT flights were flown by Enterprise on Sept. 23 Oct. 12 and Oct. 25.)
1978
Jan. 18 Thiokol conducts second test firing of an SRB.
1979
March 8 Orbiter Columbia (OV-102) transported 38 miles overland from Palmdale to Dryden Flight Research Center.
March 20-24 Columbia flown on Shuttle carrier \Jaircraft\j to Kennedy Space Center with overnight stops at El Paso and San Antonio, \JTexas\j, and Eglin AFB, Fla.
June 15 First SRB qualification test firing; 122 seconds.
1980
Feb. 20 Flight readiness firing of Columbia's main engines; 20 seconds.
April 20-21 Columbia returned to KSC by Shuttle carrier \Jaircraft\j via Tinker AFB, Okla.
Aug. 4 Columbia mated with SRBs and external tank for STS-2 mission.
Aug. 26 Space Shuttle vehicle moved to Launch Complex 39A for STS-2 mission.
Nov. 12-14 STS-2, first flight of an orbiter previously flown in space
Nov. 24-25 Columbia transported back to KSC via Bergstrom AFB, \JTexas\j.
Nov. 26 Columbia mated to SRBs and external tank at Vehicle Assembly Building (VAB) for STS-l mission.
Dec. ll Spacelab l arrives at KSC.
Dec. 29 Space Shuttle vehicle moved from VAB to Launch Complex 39A for STS-l mission.1981
1982
Feb. 3 Columbia moved to VAB for mating in preparation for STS-3 mission.
Feb. 16 Assembled Space Shuttle vehicle moved from VAB to launch pad for STS-3 mission.
March 22-30 STS-3 mission; landing at White Sands, N.M.
April 6 Columbia returned to KSC from White Sands.
May 16 Columbia moved to VAB for mating in preparation for STS-4.
May 25 STS-4 vehicle moved to launch pad.
June 27-July 4 STS-4 mission flown; first concrete runway landing at Edwards AFB.
June 30 Orbiter Challenger (OV-099) rolled out at Palmdale.
July l Challenger moved overland to Dryden.
July 4-5 Challenger flown to KSC via Ellington AFB, \JTexas\j.
July 14-15 Columbia flown to KSC via Dyess AFB, \JTexas\j.
Sept. 9 Columbia mated with SRBs and external tank in preparation for STS-5.
Sept. 21 STS-5 vehicle moved to launch pad.
Nov. ll-16 STS-5 mission; landing at Edwards AFB.
Nov. 21-22 Columbia returned to KSC via Kelly AFB, \JTexas\j
Nov. 23 Challenger moved to VAB and mated for STS-6.
Nov. 30 STS-6 vehicle moved to launch pad.
Dec. 18 Flight readiness firing of Challenger's main engines; 20 seconds.
1983
Jan. 22 Second flight readiness firing of Challenger's main engines; 22 seconds.
April 4-9 STS-6 mission, first flight of Challenger.
May 21 Challenger moved to VAB for mating in preparation for STS-7 mission.
May 26 Challenger moved to launch pad for STS-7.
June 18-24 STS-7 mission flown with landing at Edwards AFB.
July 26 Challenger moved to VAB for mating in preparation for STS-8.
June 28-29 Challenger flown back to KSC via Kelly AFB.
Aug. 2 STS-8 vehicle moved to launch pad.
Aug. 30-Sept. 5 STS-8 mission; first night launch and landing at Edwards AFB.
Sep. 9 Challenger returned to KSC via Sheppard AFB, \JTexas\j.
Sept. 23 Columbia moved to VAB for mating in preparation for STS-9.
Sept. 28 STS-9vehicle moved to launch pad.
Oct. 17 STS-9launch vehicle moved back to VAB from pad because of SRB nozzle problem.
Oct. 19 Columbia moved to Orbiter Processing Facility.
Nov. 5 Orbiter Discovery (OV-103) moved overland to Dryden.
Nov. 6 Discovery transported to Vandenberg AFB, Calif.
Nov. 8 STS-9vehicle again moved to launch pad.
Nov. 8-9 Discovery flown from Vandenberg AFB to KSC via Carswell AFB, \JTexas\j.
Nov. 28-Dec. 8 STS-9mission; landing at Edwards AFB.
Dec. 14-15 Columbia flown to KSC via El Paso, Kelly AFB and Eglin AFB.
l984
Jan. 6 Challenger moved to VAB for mating in preparation of STS 41 B mission.
Jan. ll STS 41-B vehicle moved to launch pad.
Feb. 3-ll STS 41-B mission; first landing at KSC.
March 14 Challenger moved to VAB for mating in preparation for STS 41-C mission.
March 19 STS 41-C vehicle moved to launch pad.
April 6-13 STS 41-C mission; landing at Edwards AFB.
April 17-18 Challenger flown back to KSC via Kelly AFB.
May 12 Discovery moved to VAB for mating in preparation for STS 41-D.
May 19 STS 41-D vehicle moved to launch pad.
June 2 Flight readiness firing of Discovery's main engines.
June 25 STS 41-D launch attempt scrubbed because of computer problem.
June 26 STS 41-D launch attempt scrubbed following main engine shutdown at T minus 4 seconds.
July 14 STS 41-D vehicle moved back to VAB for remanifest of payloads.
Aug. 9 STS 41-D vehicle again moved out to the launch pad.
Aug. 30-Sept. 5 STS 41-D mission; first flight of Discovery; landing at Edwards AFB.
Sept. 8 Challenger moved to VAB for mating in preparation for STS 41-G mission.
Sept. 9-10 Discovery returned to KSC via Altus AFB, Okla.
Sept. 13 STS 41-G launch vehicle moved to launch pad.
Oct. 5-13 STS 41-G mission; landing at KSC.
Oct. 18 Discovery moved to VAB for mating in preparation for STS 51-A mission.
Oct. 23 STS 51-A launch vehicle moved to launch pad.
Nov. 7 STS 51-A launch scrubbed because of high shear winds.
Nov. 8-16 STS 51-A mission; landing at KSC.
1985
Jan. 5 Discovery moved to launch pad for STS 51-C mission.
Jan. 24-27 STS 51-C mission landing at KSC.
Feb. 10 Challenger moved to VAB for mating in preparation for STS 51-E mission.
Feb. 15 STS 51-E vehicle moved to launch pad.
March 4 STS 51-E vehicle rolled back to VAB; mission cancelled; payloads combined with STS 51-B.
March 23 Discovery moved to VAB for mating in preparation for STS 51-D mission.
March 28 STS 51-D vehicle moved to launch pad.
April 6 Atlantis (OV-104) rollout at Palmdale.
April 10 Challenger moved to VAB for mating in preparation for STS 51-B mission.
April 12-19 STS 51-D mission; landing at KSC.
April 13 Atlantis ferried to KSC via Ellington AFB, \JTexas\j.
April 15 Challenger moved to launch pad for 51-B missing.
April 29-May 6 STS 51-B mission; landing at Edwards AFB.
May 10 Challenger transported back to KSC via Kelly AFB.
May 28 Discovery moved to VAB for mating in preparation for STS 51-G.
June 4 STS 51-G vehicle moved to the launch pad.
June 17-24 STS 51-G mission; landing Edwards AFB.
June 24 Challenger moved to VAB for mating in preparation for STS 51-F.
June 28 Discovery ferried back to KSC via Bergstrom AFB, \JTexas\j.
June 29 STS 51-F vehicle moved to the launch pad.
July ll Refurbished Columbia moved overland from Palmdale to Dryden.
July 12 STS 51-F launch scrubbed at T-minus 3 seconds because of main engine shutdown.
July 14 Columbia returned to KSC via Offutt AFB, Neb.
July 29-Aug. 6 STS 51-F mission landing at Edwards AFB.
July 30 Discovery moved to VAB for mating in preparation for STS 51-I mission.
Aug. 6 STS 51-I vehicle moved to the launch pad.
Aug. 10-ll Challenger flown to KSC via Davis-Monthan AFB, Ariz.; Kelly AFB; and Eglin AFB.
Aug. 24 STS 51-I mission scrubbed at T minus 5 minutes because of bad weather.
Aug. 25 STS 51-I mission scrubbed at T-minus 9 minutes because of an onboard computer problem.
Aug. 27-Sept. 3 STS 51-I mission; landing at Edwards AFB.
August 29 Atlantis moved to launch pad for the 51-J mission.
Sept. 7-8 Discovery flown back to KSC via Kelly AFB.
Sept. 12 Flight readiness firing of Atlantis' main engines; 20 seconds.
Oct. 3-7 STS 51-J mission; landing at Edwards AFB.
Oct. ll Atlantis returned to KSC via Kelly AFB.
Oct. 12 Challenger moved to VAB for mating in preparation for the STS 61-A mission.
Oct. 16 Challenger vehicle moved to the launch pad for STS 61-A mission.
Oct. 30-Nov. 6 STS 61-A mission; landing at Edwards AFB.
Nov. 8 Atlantis moved to VAB for mating in preparation for the STS 61-B.
Nov. 10-ll Challenger flown back to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
Nov. 12 STS 61-B vehicle moved to the launch pad.
Nov. 18 Enterprise (OV-101) flown from KSC to Dulles Airport, Washington, D.C., and turned over to the Smithsonian Institution.
Nov. 22 Columbia moved to the VAB for mating in preparation STS 61-C.
Nov. 26-Dec. 3 STS 61-B mission landing at Edwards AFB.
Dec. l STS 61-C vehicle moved to launch pad.
Dec. 7 Atlantis returned to KSC via Kelly AFB.
Dec. 16 Challenger moved to VAB for mating in preparation for the STS 51-L mission.
Dec. 19 STS 61-C mission scrubbed at T minus 13 seconds because of SRB auxiliary power unit problem.
Dec. 22 STS 51-L vehicle moved to Launch Pad 39B.
1986
Jan. 6 STS 61-C mission scrubbed at T minus 31 seconds because of liquid oxygen valve problem on pad.
Jan. 7 STS 61-C mission scrubbed at T minus 9 minutes because of weather problems at contingency landing sites.
Jan. 10 STS 61-C mission scrubbed T minus 9 minutes because of bad weather at KSC.
Jan. 12-18 STS 61-C mission; landing at Edwards AFB.
Jan. 22-23 Columbia returned to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
Jan. 27-28 STS 51-L launched from Pad B. Vehicle exploded 1 minute, 13 seconds after liftoff resulting loss of seven crew members.
Feb. 3 President Reagan announced the formation of the Presidential Commission on the Space Shuttle Challenger Accident, headed by William P. Rogers, former Secretary of State.
March 24 NASA publishes "Strategy for Safely Returning the Space Shuttle to Flight Status."
May 12 President Reagan appoints Dr. James C. Fletcher NASA Administrator.
July 8 NASA establishes Safety, Reliability Maintainability, and Quality Assurance Office.
July 14 NASA's plan to implement the recommendations of the Rogers commission was submitted to President Reagan.
Aug. 15 President Reagan announced his decision to support a replacement for the Challenger. At the same time, it was announced that NASA no longer would launch commercial satellites, except for those which are Shuttle-unique or have national security or foreign policy implications.
Aug. 22 NASA announced the beginning of a series of tests designed to verify the ignition pressure dynamics of the Space Shuttle solid rocket motor field joint.
Sept. 5 Study contracts were awarded to five aerospace firms for conceptual designs of an alternative or Block II Space Shuttle solid rocket motor.
Sept. 10 \JAstronaut\j Bryan O'Connor was named chairman of Space Flight Safety Panel. This panel, with oversight responsibility for all NASA manned space program activities, reports to the Associate Administrator for Safety, Reliability, Maintainability and Quality Assurance.
Oct. 2 After an intensive study, NASA announced the decision to test fire the redesigned solid rocket motor in a horizontal attitude to best simulate the critical conditions on the field joint which failed during the 51-L mission.
Oct. 30 Discovery moved to OPF where more than 200 modifications are accomplished for STS-26 mission.
Nov. 6 Office of the Director, National Space Transportation System, established in the NASA Headquarters Office of Space Flight.
1987
July 31 Rockwell International awarded contract to build a fifth orbiter to replace the Challenger.
Aug. 3 Discovery in the Orbital Processing Facility is powered up for STS-26 mission.
1988
Mid-Jan. Main engines are installed in Discovery.
March 28 Stacking of Discovery's SRBs gets underway.
May 28 Stacking of Discovery's SRBs completed.
June 10 SRBs and External Tank are mated.
June 14 The fourth full-duration test firing of the redesigned SRB motor is carried out.
June 21 Discovery rolls over from OPF to the VAB.
July 4 Discovery moved to Launch Pad 39B for STS-26 mission.
Aug. 10 Flight Readiness Firing of Discovery's main engines is conducted successfully.
#
"Space Shuttle System",209,0,0,0
\JSpace Shuttle Program\j
\JSpace Shuttle Requirements\j
\JShuttle Launch Sites\j
\JShuttle Background and Status\j
\JShuttle Mission Profile\j
\JShuttle Abort Modes\j
\JOrbiter Ground Turnaround\j
\JOrbiter Operational Improvements\j
\JSpace Shuttle Main Engine Margin Improvement Program\j
\JSSME Flight Program\j
\JShuttle Solid Rocket Motor (SRM)\j
\JShuttle Solid Rocket Boosters\j
\JShuttle External Tank\j
\JShuttle Liquid Oxygen Tank\j
\JShuttle Intertank\j
\JShuttle Liquid Hydrogen Tank\j
\JShuttle ET Thermal Protection System\j
\JShuttle ET Hardware\j
\JShuttle ET Range Safety System\j
\JShuttle Orbiter Structures\j
\JOrbiter Passive Thermal Control\j
\JOrbiter Purge, Vent and Drain System\j
\JOrbiter In-Flight Crew Escape System\j
\JOrbiter Emergency Egress Slide\j
\JOrbiter Secondary Emergency Egress\j
\JOrbiter Side Hatch Jettison\j
\JShuttle Crew Equipment\j
\JSpace Shuttle Orbiter Systems\j
\JShuttle Launch and Flight Operations\j
#
"Shuttle Abort Modes",210,0,0,0
\JShuttle Aborts\j
\JShuttle, Return To Launch Site\j
\JShuttle, Transatlantic Landing Abort\j
\JShuttle, Abort To Orbit\j
\JShuttle, Abort Once Around\j
\JShuttle, Contingency Abort\j
#
"Orbiter Operational Improvements",211,0,0,0
\JOrbital Maneuvering System and Reaction Control System (Modifications)\j
\JOrbiter, Fuel Cell Power Plants (Modifications)\j
\JOrbiter, Auxiliary Power Units (Modifications)\j
\JOrbiter, Main Landing Gear (Modifications)\j
\JOrbiter, Nose Wheel Steering (Modifications)\j
\JOrbiter, Thermal Protection System (Modifications)\j
\JOrbiter, Wing Modification\j
\JOrbiter, Mid-Fuselage Modifications\j
\JOrbiter, General Purpose Computers (Modifications)\j
\JOrbiter, Inertial Measurement Units (Modifications)\j
\JShuttle Payload Deployment and Retrieval System\j
\JShuttle Payload Retention Mechanisms\j
\JSpace Flight Tracking and Data Network\j
\JSatellite System, Tracking and Data Relay\j
\JShuttle Avionics Systems\j
\JShuttle Data Processing System\j
\JShuttle Guidance, Navigation and Control\j
\JShuttle Flight Control System Hardware\j
\JShuttle Navigation Aids\j
\JShuttle Inertial Measurement Units\j
\JShuttle Star Trackers\j
\JShuttle Crewman Optical Alignment Sight\j
\JShuttle Air Data System\j
\JShuttle Microwave Scan Beam Landing System\j
\JShuttle Radar Altimeter\j
\JShuttle Accelerometer Assemblies\j
\JOrbiter Rate Gyro Assemblies\j
\JSolid Rocket Booster Rate Gyro Assemblies\j
\JShuttle Rotational Hand Controller\j
\JShuttle Translational Hand Controller\j
\JShuttle Control Stick Steering Push Button Light Indicators\j
\JShuttle Rudder Pedals\j
\JShuttle Speed Brake/Thrust Controller\j
\JShuttle Body Flap Switch\j
\JShuttle Aerosurface Servoamplifiers\j
\JShuttle Digital Autopilot\j
\JShuttle Rendezvous Thrusting Maneuvers\j
\JShuttle Component Locations\j
\JShuttle, Dedicated Display Systems\j
\JShuttle, Attitude Director Indicator\j
\JShuttle, Horizontal Situation Indicator\j
\JShuttle, Alpha Mach Indicator\j
\JShuttle, Altitude/Vertical Velocity Indicator\j
\JShuttle, Surface Position Indicator\j
\JShuttle, Flight Control System Push Button Light Indicators\j
\JShuttle, Reaction Control System Command Lights\j
\JShuttle, G-Meter\j
\JShuttle, Head-Up Display\j
#
"Shuttle Launch and Flight Operations",217,0,0,0
\JShuttle, Pre-Launch Operations\j
\JShuttle, Pre-Launch Propellant-Loading\j
\JShuttle, Final Pre-Launch Activities\j
\JShuttle, Launch Control Center\j
\JShuttle, Launch Countdown\j
\JShuttle, Mission Control Center\j
\JShuttle, Marshall Payload Operations Control Center\j
\JSpace Tracking and Data Acquisition\j
\JShuttle Support Ground Network\j
\JSpace Network\j
\JShuttle Flight Operations\j
\JShuttle Launch Abort Modes\j
\JShuttle, On-Orbit Operations\j
\JShuttle, Maneuvering In Orbit\j
#
"Space Shuttle Program",218,0,0,0
The Space Shuttle is developed by the National Aeronautics and Space Administration. NASA coordinates and manages the Space Transportation System (NASA's name for the overall Shuttle program), including intergovernmental agency requirements and international and joint projects. NASA also oversees the launch and space flight requirements for civilian and commercial use.
The Space Shuttle system consists of four primary elements: an orbiter \Jspacecraft\j, two Solid Rocket Boosters (SRB), an external tank to house fuel and oxidizer and three Space Shuttle main engines.
The orbiter is built by Rockwell International's Space Transportation Systems Division, Downey, Calif., which also has responsibility for the \Jintegration\j of the overall space transportation system. Both orbiter and \Jintegration\j contracts are under the direction of NASA's Johnson Space Center in Houston, \JTexas\j.
The SRB motors are built by the Wasatch Division of Morton Thiokol Corp., Brigham City, \JUtah\j, and are assembled, checked out and refurbished by United Space Boosters Inc., Booster Production Co., Kennedy Space Center. Cape Canaveral, Fla.
The external tank is built by Martin Marietta Corp. at its Michoud facility, New Orleans, La., and the Space Shuttle main engines are built by Rockwell's Rocketdyne Division, Canoga Park, Calif. These contracts are under the direction of NASA's George C. Marshall Space Flight Center, \JHuntsville\j, Ala.
#
"Space Shuttle Requirements",219,0,0,0
The Shuttle will transport cargo into near Earth orbit 100 to 217 nautical miles (115 to 250 statute miles) above the Earth. This cargo -- or payload -- is carried in a bay 15 feet in diameter and 60 ft long.
Major system requirements are that the orbiter and the two solid rocket boosters be reusable.
Other features of the Shuttle:
The orbiter has carried a flight crew of up to eight persons. A total of 10 persons could be carried under emergency conditions. The basic mission is 7 days in space. The crew compartment has a shirtsleeve environment, and the acceleration load is never greater than 3 Gs. In its return to Earth, the orbiter has a cross-range maneuvering capability of 1,100 nautical miles (1,265 statute miles).
The Space Shuttle is launched in an upright position, with thrust provided by the three Space Shuttle engines and the two SRB. After about 2 minutes, the two boosters are spent and are separated from the external tank. They fall into the ocean at predetermined points and are recovered for reuse.
The Space Shuttle main engines continue firing for about 8 minutes. They shut down just before the craft is inserted into orbit. The external tank is then separated from the orbiter. It follows a ballistic trajectory into a remote area of the ocean but is not recovered.
There are 38 primary Reaction Control System (RCS) engines and six vernier RCS engines located on the orbiter. The first use of selected primary reaction control system engines occurs at orbiter/external tank separation. The selected primary reaction control system engines are used in the separation sequence to provide an attitude hold for separation. Then they move the orbiter away from the external tank to ensure orbiter clearance from the arc of the rotating external tank. Finally, they return to an attitude hold prior to the initiation of the firing of the Orbital Maneuvering System (OMS) engines to place the orbiter into orbit.
The primary and/or vernier RCS engines are used normally on orbit to provide attitude pitch, roll and yaw maneuvers as well as translation maneuvers.
The two OMS engines are used to place the orbiter on orbit, for major velocity maneuvers on orbit and to slow the orbiter for reentry, called the deorbit maneuver. Normally, two OMS engine thrusting sequences are used to place the orbiter on orbit, and only one thrusting sequence is used for deorbit.
The orbiter's velocity on orbit is approximately 25,405 feet per second (17,322 statute miles per hour). The deorbit maneuver decreases this velocity approximately 300 fps (205 mph) for reentry.
In some missions, only one OMS thrusting sequence is used to place the orbiter on orbit. This is referred to as direct insertion. Direct insertion is a technique used in some missions where there are high-performance requirements, such as a heavy payload or a high orbital altitude. This technique uses the Space Shuttle main engines to achieve the desired apogee (high point in an orbit) altitude, thus conserving orbital maneuvering system propellants. Following jettison of the external tank, only one OMS thrusting sequence is required to establish the desired orbit altitude.
For deorbit, the orbiter is rotated tail first in the direction of the velocity by the primary reaction control system engines. Then the OMS engines are used to decrease the orbiter's velocity.
During the initial entry sequence, selected primary RCS engines are used to control the orbiter's attitude (pitch, roll and yaw). As aerodynamic pressure builds up, the orbiter flight control surfaces become active and the primary reaction control system engines are inhibited.
During entry, the thermal protection system covering the entire orbiter provides the protection for the orbiter to survive the extremely high temperatures encountered during entry. The thermal protection system is reusable (it does not burn off or ablate during entry).
The unpowered orbiter glides to Earth and lands on a runway like an airplane. Nominal touchdown speed varies from 184 to 196 knots (213 to 225 miles per hour).
The main landing gear wheels have a braking system for stopping the orbiter on the runway, and the nose wheel is steerable, again similar to a conventional airplane.
There are two launch sites for the Space Shuttle. Kennedy Space Center (KSC) in \JFlorida\j is used for launches to place the orbiter in equatorial orbits (around the equator), and Vandenberg Air Force Base launch site in \JCalifornia\j will be used for launches that place the orbiter in polar orbit missions.
Landing sites are located at the KSC and Vandenberg. Additional landing sites are provided at Edwards Air Force Base in \JCalifornia\j and White Sands, N.M. Contingency landing sites are also provided in the event the orbiter must return to Earth in an emergency.
#
"Shuttle Launch Sites",220,0,0,0
Space Shuttles destined for equatorial orbits are launched from the KSC, and those requiring polar orbital planes will be launched from Vandenberg.
Orbital mechanics and the complexities of mission requirements, plus safety and the possibility of infringement on foreign air and land space, prohibit polar orbit launches from the KSC.
Kennedy Space Center launches have an allowable path no less than 35 degrees northeast and no greater than 120 degrees southeast. These are \Jazimuth\j degree readings based on due east from KSC as 90 degrees.
A 35-degree \Jazimuth\j launch places the \Jspacecraft\j in an orbital inclination of 57 degrees. This means the \Jspacecraft\j in its orbital trajectories around the Earth will never exceed an Earth latitude higher or lower than 57 degrees north or south of the equator.
A launch path from KSC at an \Jazimuth\j of 120 degrees will place the \Jspacecraft\j in an orbital inclination of 39 degrees (it will be above or below 39 degrees north or south of the equator).
These two azimuths - 35 and 120 degrees - represent the launch limits from the KSC. Any \Jazimuth\j angles further north or south would launch a \Jspacecraft\j over a habitable land mass, adversely affect safety provisions for abort or vehicle separation conditions, or present the undesirable possibility that the SRB or external tank could land on foreign land or sea space.
Launches from Vandenberg have an allowable launch path suitable for polar insertions south, southwest and southeast.
The launch limits at Vandenberg are 201 and 158 degrees. At a 201-degree launch \Jazimuth\j, the \Jspacecraft\j would be orbiting at a 104-degree inclination. Zero degrees would be due north of the launch site, and the orbital trajectory would be within 14 degrees east or west of the north-south pole meridian.
At a launch \Jazimuth\j of 158 degrees, the \Jspacecraft\j would be orbiting at a 70-degree inclination, and the trajectory would be within 20 degrees east or west of the polar meridian. Like KSC, Vandenberg has allowable launch azimuths that do not pass over habitable areas or involve safety, abort, separation and political considerations.
Mission requirements and payload weight penalties also are major factors in selecting a launch site.
The Earth rotates from west to east at a speed of approximately 900 nautical miles per hour (1,035 mph). A launch to the east uses the Earth's rotation somewhat as a springboard. The Earth's rotational rate also is the reason the orbiter has a cross-range capability of 1,100 nautical miles (1,265 statute miles) to provide the abort-once-around capability in polar orbit launches.
Attempting to launch and place a \Jspacecraft\j in polar orbit from KSC to avoid habitable landmass would be uneconomical because the Shuttle's payload would be reduced severely-down to approximately 17,000 pounds. A northerly launch into polar orbit of 8 to 20 degrees \Jazimuth\j would necessitate a path over a landmass; and most safety, abort, and political constraints would have to be waived. This prohibits polar orbit launches from the KSC.
NASA's latest assessment of orbiter ascent and landing weights incorporates currently approved modifications to all vehicle elements, including crew escape provisions, and assumes a maximum Space Shuttle main engine throttle setting of 104 percent. It is noted that the resumption of Space Shuttle flights initially requires more conservative flight design criteria and additional instrumentation, which reduces the following basic capabilities by approximately 1,600 pounds:
Kennedy Space Center Eastern Space and Missile Center (ESMC) satellite deploy missions. The basic cargo-lift capability for a due east (28.5 degrees) launch is 55,000 pounds to a 110-nautical-mile (126-statute-mile) orbit using OV-103 (Discovery) or OV-104 (Atlantis) to support a 4-day satellite deploy mission. This capability will be reduced approximately 100 pounds for each additional nautical mile of altitude desired by the customer.
The payload capability for the same satellite deploy mission with a 57-degree inclination is 41,000 pounds.
The performance for intermediate inclinations can be estimated by allowing 500 pounds per degree of plane change between 28.5 and 57 degrees.
If OV-102 (Columbia) is used, the cargo-lift weight capability must be decreased by approximately 8,400 pounds. This weight difference is attributed to an approximately 7,150-pound difference in inert weight, 850 pounds of orbiter experiments, 300 pounds of additional thermal protection system and 100 pounds to accommodate a fifth cryogenic liquid oxygen and liquid \Jhydrogen\j tank set for the power reactant storage and distribution system.
Vandenberg Air Force Base Western Space and Missile Center (WSMC) satellite deploy missions. Using OV-103 (Discovery) or OV-104 (Atlantis), the cargo-lift weight capability is 29,600 pounds for a 98-degree launch inclination and 110-nautical-mile (126-statute-mile) polar orbit.
Again, an increase in altitude costs approximately 100 pounds per nautical mile. NASA assumes also that the advanced solid rocket motor will replace the filament-wound solid rocket motor case previously used for western test range assessments.
The same mission at 68 degrees inclination (minimum western test range inclination based on range safety limitations) is 49,600 pounds. Performance for intermediate inclinations can be estimated by allowing 660 pounds for each degree of plane change between inclinations of 68 and 98 degrees.
Landing weight limits. All the Space Shuttle orbiters are currently limited to a total vehicle landing weight of 240,000 pounds for abort landings and 230,000 pounds for nominal end-of-mission landings. It is noted that each additional crew person beyond the five-person standard is chargeable to the cargo weight allocation and reduces the payload capability by approximately 500 pounds. (This is an increase of 450 pounds to account for the crew escape equipment.)
#
"Shuttle Background and Status",221,0,0,0
On July 26, 1972, NASA selected Rockwell's Space Transportation Systems Division in Downey, Calif., as the industrial contractor for the design, development, test and evaluation of the orbiter. The contract called for fabrication and testing of two orbiters, a full-scale structural test article, and a main propulsion test article. The award followed years of NASA and Air Force studies to define and assess the feasibility of a reusable space transportation system.
NASA previously (March 31, 1972) had selected Rockwell's Rocketdyne Division to design and develop the Space Shuttle main engines. Contracts followed to Martin Marietta for the external tank (Aug. 16, 1973) and Morton Thiokol's Wasatch Division for the solid rocket boosters (June 27, 1974).
In addition to the orbiter DDT&E contract, Rockwell's Space Transportation Systems Division was given contractual responsibility as system integrator for the overall Shuttle system.
Rockwell's Launch Operations, part of the Space Transportation Systems Division, was under contract to NASA's Kennedy Space Center for turnaround, processing, prelaunch testing, and launch and recovery operations from STS-1 through the STS-11 mission.
On Oct. 1, 1983, the Lockheed Space Operations Co. was awarded the Space Shuttle processing contract at KSC for turnaround processing, prelaunch testing, and launch and recovery operations.
The first orbiter \Jspacecraft\j, Enterprise (OV-101), was rolled out on Sept. 17, 1976. On Jan. 31, 1977, it was transported 38 miles overland from Rockwell's assembly facility at Palmdale, Calif., to NASA's Dryden Flight Research Facility at Edwards Air Force Base for the Approach and Landing Test (ALT) program.
The 9-month-long ALT program was conducted from February through November 1977 at Dryden and demonstrated the orbiter could fly in the atmosphere and land like an airplane except without power, a gliding flight.
The ALT program involved ground tests and flight tests.
The ground tests included taxi tests of the 747 shuttle carrier \Jaircraft\j (SCA) with the Enterprise mated atop the SCA to determine structural loads and responses and assess the mated capability in ground handling and control characteristics up to flight takeoff speed.
The taxi tests also validated 747 steering and braking with the orbiter attached. A ground test of orbiter systems followed the unmanned captive tests. All orbiter systems were activated as they would be in atmospheric flight. This was the final preparation for the manned captive-flight phase.
Five captive flights of the Enterprise mounted atop the SCA with the Enterprise unmanned and Enterprise systems inert were conducted to assess the structural integrity and performance-handling qualities of the mated craft.
Three manned captive flights that followed the five unmanned captive flights included an \Jastronaut\j crew aboard the orbiter operating its flight control systems while the orbiter remained perched atop the SCA. These flights were designed to exercise and evaluate all systems in the flight environment in preparation for the orbiter release (free) flights. They included flutter tests of the mated craft at low and high speed, a separation trajectory test and a dress rehearsal for the first orbiter free flight.
In the five free flights the \Jastronaut\j crew separated the \Jspacecraft\j from the SCA and maneuvered to a landing at Edwards Air Force Base. In the first four such flights the landings were on a dry lake bed; in the fifth, the landing was on Edwards' main concrete runway under conditions simulating a return from space.
The last two free flights were made without the tail cone, which is the \Jspacecraft\j's configuration during an actual landing from Earth orbit. These flights verified the orbiter's pilot-guided approach and landing capability; demonstrated the orbiter's subsonic terminal area energy management autoland approach capability; and verified the orbiter's subsonic airworthiness, integrated system operations and selected subsystems in preparation for the first manned orbital flight.
The flights demonstrated the orbiter's ability to approach and land safely with a minimum gross weight and using several center-of-gravity configurations.
For all of the captive flights and the first three free flights, the orbiter was outfitted with a tail cone covering its aft section to reduce aerodynamic drag and turbulence. The final two free flights were without the tail cone, and the three simulated Space Shuttle main engines and two orbital maneuvering system engines were exposed aerodynamically.
The final phase of the ALT program prepared the \Jspacecraft\j for four ferry flights. Fluid systems were drained and purged, the tail cone was reinstalled and elevon locks were installed.
The forward attachment strut was replaced to lower the orbiter's cant from 6 to 3 degrees. This reduces drag to the mated vehicles during the ferry flights.
After the ferry flight tests, OV-101 was returned to the NASA hangar at Dryden and modified for vertical ground vibration tests at NASA's Marshall Space Flight Center, \JHuntsville\j, Ala.
On March 13, 1978, the Enterprise was ferried atop the SCA to MSFC. At Marshall, Enterprise was mated with the external tank and SRB and subjected to a series of vertical ground vibration tests. These tested the mated configuration's critical structural dynamic response modes, which were assessed against analytical math models used to design the various element interfaces.
These were completed in March 1979. On April 10, 1979 the Enterprise was ferried to Kennedy Space Center, mated with the external tank and SRB and transported via the mobile launcher platform to Launch Complex 39-A. At Launch Complex 39-A, the Enterprise served as a practice and launch complex fit-check verification tool representing the flight vehicles.
It was ferried back to Dryden at Edwards AFB in \JCalifornia\j on Aug. 16, 1979, and then returned overland to Rockwell's Palmdale final assembly facility on Oct. 30, 1979. Certain components were refurbished for use on flight vehicles being assembled at Palmdale. The Enterprise was then returned overland to Dryden on Sept. 6, 1981.
During exhibition at the Paris, May and June 1983, Enterprise was ferried to \JFrance\j for the Air Show as well as to \JGermany\j, \JItaly\j, England and Canada before returning to Dryden.
From April to October 1984, Enterprise was ferried to Vandenberg AFB and to Mobile, Ala., where it was taken by barge to New Orleans, La., for the United States 1984 World's Fair.
In November 1984 it was transported to Vandenberg and used as a practice and fit-check verification tool. On May 24, 1985, Enterprise was ferried from Vandenberg to Dryden.
On Sept. 20, 1985, Enterprise was ferried from Dryden Flight Research Facility to KSC. On Nov. 18, 1985, Enterprise was ferried from KSC to Dulles Airport, Washington, D.C., and became the property of the Smithsonian Institution. The Enterprise was built as a test vehicle and is not equipped for space flight.
The second orbiter, Columbia (OV-102), was the first to fly into space. it was transported overland on March 8, 1979, from Palmdale to Dryden for mating atop the SCA and ferried to KSC. It arrived on March 25, 1979, to begin preparations for the first flight into space.
The structural test article, after 11 months of extensive testing at Lockheed's facility in Palmdale, was returned to Rockwell's Palmdale facility for modification to become the second orbiter available for operational missions. It was redesignated OV-099, the Challenger.
The main propulsion test article (MPTS-098) consisted of an orbiter aft fuselage, a truss arrangement that simulated the orbiter's mid-fuselage and the Shuttle main propulsion system (three Space Shuttle main engines and the external tank). This test structure is at the Stennis Space Center in Mississippi. A series of static firings was conducted from 1978 through 1981 in support of the first flight into space.
On Jan. 29, 1979, NASA contracted with Rockwell to manufacture two additional orbiters, OV-103 and OV-104 (Discovery and Atlantis), convert the structural test article to space flight configuration (Challenger) and modify Columbia from its development configuration to that required for operational flights.
NASA named the first four orbiter spacecrafts after famous exploration sailing ships. In the order they became operational, they are: Columbia (OV-102), after a sailing \Jfrigate\j launched in 1836, one of the first Navy ships to circumnavigate the globe.
Columbia also was the name of the Apollo 11 command module that carried Neil Armstrong, Michael Collins and Edward (Buzz) Aldrin on the first lunar landing mission, July 20, 1969. Columbia was delivered to Rockwell's Palmdale assembly facility for modifications on Jan. 30, 1984, and was returned to KSC on July 14, 1985, for return to flight.
Challenger (OV-099), also a Navy ship, which from 1872 to 1876 made a prolonged exploration of the Atlantic and Pacific oceans. It also was used in the Apollo program for the Apollo 17 lunar module. Challenger was delivered to DSC on July 5, 1982.
Discovery (OV-103), after two ships, the vessel in which Henry Hudson in 1610-11 attempted to search for a northwest passage between the Atlantic and Pacific oceans and instead discovered Hudson Bay and the ship in which Capt. Cook discovered the Hawaiian Islands and explored southern \JAlaska\j and western Canada.
Discovery was delivered to KSC on Nov. 9, 1983. Atlantis (OV-104), after a two-masted ketch operated for the Woods Hole Oceanographic Institute from 1930 to 1966, which traveled more than half a million miles in ocean research. Atlantis was delivered to KSC on April 3, 1985.
In April 1983, under contract to NASA, Rockwell's Space Transportation Systems Division, Downey, Calif., began the construction of structural spares for completion in 1987. The structural spares program consisted of an aft fuselage, crew compartment, forward reaction control system, lower and upper forward fuselage, mid-fuselage, wings (elevons), payload bay doors, vertical stabilizer (rudder/speed brake), body flap and one set of orbital maneuvering system/reaction control system pods.
On Sept. 12, 1985, Rockwell International's Shuttle Operations Co., Houston, \JTexas\j, was awarded the Space Transportation System operation contract at NASA's Johnson Space Center, consolidating work previously performed under 22 contracts by 16 different contractors.
On July 31, 1987, NASA awarded Rockwell's Space Transportation Systems Division, Downey, Calif., a contract to build a replacement Space Shuttle orbiter using the structural spares. The replacement orbiter will be assembled at Rockwell's Palmdale, Calif., assembly facility and is scheduled for completion in 1991. This orbiter is designated OV-105.
#
"Shuttle Mission Profile",222,0,0,0
In the launch configuration, the orbiter and two SRBs are attached to the external tank in a vertical (nose-up) position on the launch pad. Each SRB is attached at its aft skirt to the mobile launcher platform by four bolts.
Emergency exit for the flight crew on the launch pad up to 30 seconds before liftoff is by slidewire. There are seven 1,200-foot-long slidewires, each with one basket. Each basket is designed to carry three persons.
The baskets, 5 feet in diameter and 42 inches deep, are suspended beneath the slide mechanism by four cables. The slidewires carry the baskets to ground level. Upon departing the basket at ground level, the flight crew progresses to a bunker that is designed to protect it from an explosion on the launch pad.
At launch, the three Space Shuttle main engines - fed liquid \Jhydrogen\j fuel and liquid oxygen oxidizer from the external tank - are ignited first. When it has been verified that the engines are operating at the proper thrust level, a signal is sent to ignite the SRB. At the proper thrust-to-weight ratio, initiators (small explosives) at eight hold-down bolts on the SRB are fired to release the Space Shuttle for liftoff. All this takes only a few seconds.
Maximum dynamic pressure is reached early in the ascent, nominally approximately 60 seconds after liftoff. Approximately 1 minute later (2 minutes into the ascent phase), the two SRB have consumed their propellant and are jettisoned from the external tank. This is triggered by a separation signal from the orbiter.
The boosters briefly continue to ascend, while small motors fire to carry them away from the Space Shuttle. The boosters then turn and descend, and at a predetermined altitude, parachutes are deployed to decelerate them for a safe splashdown in the ocean. Splashdown occurs approximately 141 nautical miles (162 statute miles) from the launch site. The boosters are recovered and reused.
Meanwhile, the orbiter and external tank continue to ascend, using the thrust of the three Space Shuttle main engines. Approximately 8 minutes after launch and just short of orbital velocity, the three Space Shuttle engines are shut down (main engine cutoff), and the external tank is jettisoned on command from the orbiter.
The forward and aft reaction control system engines provide attitude (pitch, yaw and roll) and the translation of the orbiter away from the external tank at separation and return to attitude hold prior to the orbital maneuvering system thrusting maneuver.
The external tank continues on a ballistic trajectory and enters the atmosphere, where it disintegrates. Its projected impact is in the Indian Ocean (except for 57-degree inclinations) in the case of equatorial orbits KSC launch) and in the extreme southern Pacific Ocean in the case of a Vandenberg launch.
Normally, two thrusting maneuvers using the two OMS engines at the aft end of the orbiter are used in a two-step thrusting sequence: to complete insertion into Earth orbit and to circularize the \Jspacecraft\j's orbit. The OMS engines are also used on orbit for any major velocity changes.
In the event of a direct-insertion mission, only one OMS thrusting sequence is used.
The orbital altitude of a mission is dependent upon that mission. The nominal altitude can vary between 100 to 217 nautical miles (115 to 250 statute miles).
The forward and aft RCS thrusters (engines) provide attitude control of the orbiter as well as any minor translation maneuvers along a given axis on orbit.
At the completion of orbital operations, the orbiter is oriented in a tail first attitude by the reaction control system. The two OMS engines are commanded to slow the orbiter for deorbit.
The reaction control system turns the orbiter's nose forward for entry. The reaction control system controls the orbiter until atmospheric density is sufficient for the pitch and roll aerodynamic control surfaces to become effective.
Entry interface is considered to occur at 400,000 feet altitude approximately 4,400 nautical miles (5,063 statute miles) from the landing site and at approximately 25,000 feet per second velocity.
At 400,000 feet altitude, the orbiter is maneuvered to zero degrees roll and yaw (wings level) and at a predetermined angle of attack for entry. The angle of attack is 40 degrees. The flight control system issues the commands to roll, pitch and yaw reaction control system jets for rate \Jdamping\j.
The forward RCS engines are inhibited prior to entry interface, and the aft reaction control system engines maneuver the \Jspacecraft\j until a dynamic pressure of 10 pounds per square foot is sensed, which is when the orbiter's ailerons become effective.
The aft RCS roll engines are then deactivated. At a dynamic pressure of 20 pounds per square foot, the orbiter's elevators become active, and the aft RCS pitch engines are deactivated. The orbiter's speed brake is used below Mach 10 to induce a more positive downward elevator trim deflection. At approximately Mach 3.5, the rudder becomes activated, and the aft reaction control system yaw engines are deactivated at 45,000 feet.
Entry guidance must dissipate the tremendous amount of energy the orbiter possesses when it enters the Earth's atmosphere to assure that the orbiter does not either burn up (entry angle too steep) or skip out of the atmosphere (entry angle too shallow) and that the orbiter is properly positioned to reach the desired touchdown point.
During entry, energy is dissipated by the atmospheric drag on the orbiter's surface. Higher atmospheric drag levels enable faster energy dissipation with a steeper trajectory. Normally, the angle of attack and roll angle enable the atmospheric drag of any flight vehicle to be controlled. However, for the orbiter, angle of attack was rejected because it creates surface temperatures above the design specification. The angle of attack scheduled during entry is loaded into the orbiter computers as a function of relative velocity, leaving roll angle for energy control. Increasing the roll angle decreases the vertical component of lift, causing a higher sink rate and energy dissipation rate. Increasing the roll rate does raise the surface temperature of the orbiter, but not nearly as drastically as an equal angle of attack command.
If the orbiter is low on energy (current range-to-go much greater than nominal at current velocity), entry guidance will command lower than nominal drag levels. If the orbiter has too much energy (current range-to-go much less than nominal at the current velocity), entry guidance will command higher-than-nominal drag levels to dissipate the extra energy.
Roll angle is used to control cross range. \JAzimuth\j error is the angle between the plane containing the orbiter's position vector and the heading alignment cylinder tangency point and the plane containing the orbiter's position vector and velocity vector. When the \Jazimuth\j error exceeds a computer-loaded number, the orbiter's roll angle is reversed.
Thus, descent rate and down ranging are controlled by bank angle. The steeper the bank angle, the greater the descent rate and the greater the drag. Conversely, the minimum drag attitude is wings level. Cross range is controlled by bank reversals.
The entry thermal control phase is designed to keep the backface temperatures within the design limits. A constant heating rate is established until below 19,000 feet per second.
The equilibrium glide phase shifts the orbiter from the rapidly increasing drag levels of the temperature control phase to the constant drag level of the constant drag phase. The equilibrium glide flight is defined as flight in which the flight path angle, the angle between the local horizontal and the local velocity vector, remains constant. Equilibrium glide flight provides the maximum downrange capability. It lasts until the drag acceleration reaches 33 feet per second squared.
The constant drag phase begins at that point. The angle of attack is initially 40 degrees, but it begins to ramp down in this phase to approximately 36 degrees by the end of this phase.
In the transition phase, the angle of attack continues to ramp down, reaching the approximately 14-degree angle of attack at the entry Terminal Area Energy Management (TAEM) interface, at approximately 83,000 feet altitude, 2,500 feet per second, Mach 2.5 and 52 nautical miles (59 statute miles) from the landing runway. Control is then transferred to TAEM guidance.
During the entry phases described, the orbiter's roll commands keep the orbiter on the drag profile and control cross range.
TAEM guidance steers the orbiter to the nearest of two heading alignment cylinders, whose radii are approximately 18,000 feet and which are located \Jtangent\j to and on either side of the runway centerline on the approach end. In TAEM guidance, excess energy is dissipated with an S-turn; and the speed brake can be used to modify drag, lift-to-drag ratio and flight path angle in high-energy conditions.
This increases the ground track range as the orbiter turns away from the nearest Heading Alignment Circle (HAC) until sufficient energy is dissipated to allow a normal approach and landing guidance phase capture, which begins at 10,000 feet altitude. The orbiter also can be flown near the velocity for maximum lift over drag or wings level for the range stretch case. The \Jspacecraft\j slows to subsonic velocity at approximately 49,000 feet altitude, about 22 nautical miles (25.3 statute miles) from the landing site.
At TAEM acquisition, the orbiter is turned until it is aimed at a point \Jtangent\j to the nearest HAC and continues until it reaches way point 1. At WP-1, the TAEM heading alignment phase begins. The HAC is followed until landing runway alignment, plus or minus 20 degrees, has been achieved.
In the TAEM pre-final phase, the orbiter leaves the HAC; pitches down to acquire the steep glide slope, increases airspeed; banks to acquire the runway centerline and continues until on the runway centerline, on the outer glide slope and on airspeed. The approach and landing guidance phase begins with the completion of the TAEM pre-final phase and ends when the \Jspacecraft\j comes to a complete stop on the runway.
The approach and landing trajectory capture phase begins at the TAEM interface and continues to guidance lock-on to the steep outer glide slope. The approach and landing phase begins at about 10,000 feet altitude at an equivalent airspeed of 290, plus or minus 12, knots 6.9 nautical miles (7.9 statute miles) from touchdown.
Autoland guidance is initiated at this point to guide the orbiter to the minus 19- to 17-degree glide slope (which is over seven times that of a commercial airliner's approach) aimed at a target 0.86 nautical mile (1 statute mile) in front of the runway.
The \Jspacecraft\j's speed brake is positioned to hold the proper velocity. The descent rate in the later portion of TAEM and approach and landing is greater than 10,000 feet per minute (a rate of descent approximately 20 times higher than a commercial airliner's standard 3-degree instrument approach angle).
At 1,750 feet above ground level, a pre-flare maneuver is started to position the \Jspacecraft\j for a 1.5-degree glide slope in preparation for landing with the speed brake positioned as required. The flight crew deploys the landing gear at this point.
The final phase reduces the sink rate of the \Jspacecraft\j to less than 9 feet per second. Touchdown occurs approximately 2,500 feet past the runway threshold at a speed of 184 to 196 knots (213 to 226 mph).
#
"Shuttle Aborts",223,0,0,0
Selection of an ascent abort mode may become necessary if there is a failure that affects vehicle performance, such as the failure of a Space Shuttle main engine or an orbital maneuvering system. Other failures requiring early termination of a flight, such as a cabin leak, might require the selection of an abort mode.
There are two basic types of ascent abort modes for Space Shuttle missions: intact aborts and contingency aborts. Intact aborts are designed to provide a safe return of the orbiter to a planned landing site. Contingency aborts are designed to permit flight crew survival following more sever failures when an intact abort is not possible. A contingency abort would generally result in a ditch operation.
There are four types of intact aborts: Abort to Orbit (ATO), Abort Once Around (AOA), Transatlantic Landing (TAL) and Return to Launch Site (RTLS).
The ATO mode is designed to allow the vehicle to achieve a temporary orbit that is lower than the nominal orbit. This mode requires less performance and allows time to evaluate problems and then choose either an early deorbit maneuver or an orbital maneuvering system thrusting maneuver to raise the orbit and continue the mission.
The AOA is designed to allow the vehicle to fly once around the Earth and make a normal entry and landing. This mode generally involves two orbital maneuvering system thrusting sequences, with the second sequence being a deorbit maneuver. The entry sequence would be similar to a normal entry.
The TAL mode is designed to permit an intact landing on the other side of the Atlantic Ocean. This mode results in a ballistic trajectory, which does not require an orbital maneuvering system maneuver.
The RTLS mode involves flying downrange to dissipate propellant and then turning around under power to return directly to a landing at or near the launch site.
There is a definite order of preference for the various abort modes. The type of failure and the time of the failure determine which type of abort is selected. In cases where performance loss is the only factor, the preferred modes would be ATO, AOA, TAL and RTLS, in that order.
The mode chosen is the highest one that can be completed with the remaining vehicle performance. In the case of some support system failures, such as cabin leaks or vehicle cooling problems, the preferred mode might be the one that will end the mission most quickly. In these cases, TAL or RTLS might be preferable to AOA or ATO. A contingency abort is never chosen if another abort option exists.
The Mission Control Center-Houston is prime for calling these aborts because it has a more precise knowledge of the orbiter's position than the crew can obtain from onboard systems. Before main engine cutoff, Mission Control makes periodic calls to the crew to tell them which abort mode is (or is not) available. If ground communications are lost, the flight crew has onboard methods, such as cue cards, dedicated displays and display information, to determine the current abort region.
Which abort mode is selected depends on the cause and timing of the failure causing the abort and which mode is safest or improves mission success. If the problem is a Space Shuttle main engine failure, the flight crew and Mission Control Center select the best option available at the time a space shuttle main engine fails.
If the problem is a system failure that jeopardizes the vehicle, the fastest abort mode that results in the earliest vehicle landing is chosen. RTLS and TAL are the quickest options (35 minutes), whereas an AOA requires approximately 90 minutes. Which of these is elected depends on the time of the failure with three good Space Shuttle main engines.
The flight crew selects the abort mode by positioning an abort mode switch and depressing an abort push button.
#
"Shuttle, Return To Launch Site",224,0,0,0
The RTLS abort mode is designed to allow the return of the orbiter, crew, and payload to the launch site, Kennedy Space Center. Approximately 25 minutes after lift-off. The RTLS profile is designed to accommodate the loss of thrust from one space shuttle main engine between liftoff and approximately four minutes 20 seconds, at which time not enough main propulsion system propellant remains to return to the launch site.
An RTLS can be considered to consist of three stages-a powered stage, during which the main engines are still thrusting; an ET separation phase; and the glide phase, during which the orbiter glides to a landing at the KSC. The powered RTLS phase begins with the crew selection of the RTLS abort, which is done after SRB separation.
The crew selects the abort mode by positioning the abort rotary switch to RTLS and depressing the abort push button. The time at which the RTLS is selected depends on the reason for the abort. For example, a three-engine RTLS is selected at the last moment, approximately 3 minutes, 34 seconds into the mission; whereas an RTLS chosen due to an engine out at liftoff is selected at the earliest time, approximately two minutes 20 seconds into the mission (after SOR separation).
After RTLS is selected, the vehicle continues downrange to dissipate excess main propulsion system propellant. The goal is to leave only enough main propulsion system propellant to be able to turn the vehicle around, fly back towards KSC and achieve the proper main engine cutoff conditions so the vehicle can glide to the KSC after external tank separation.
During the downrange phase, a pitch-around maneuver is initiated (the time depends in part on the time of a main engine failure) to orient the orbiter/ external tank configuration to a heads up attitude, pointing toward the launch site. At this time, the vehicle is still moving away from the launch site, but the main engines are now thrusting to null the downrange velocity. In addition, excess orbital maneuvering system and reaction control system propellants are dumped by continuous orbital maneuvering system and reaction control system engine thrustings to improve the orbiter weight and center of gravity for the glide phase and landing.
The vehicle will reach the desired main engine cutoff point with less than 2 percent excess propellant remaining in the external tank. At main engine cutoff minus 20 seconds, a pitch-down maneuver (called powered pitch-down) takes the mated vehicle to the required external tank separation attitude and pitch rate. After main engine cutoff has been commanded, the external tank separation sequence begins, including a reaction control system translation that ensures that the orbiter does not recontact the external tank and that the orbiter has achieved the necessary pitch attitude to begin the glide phase of the RTLS.
After the reaction control system translation maneuver has been completed, the glide phase of the RTLS begins. From then on, the RTLS is handled similarly to a normal entry.
#
"Shuttle, Transatlantic Landing Abort",225,0,0,0
The TAL abort mode was developed to improve the options available when a main engine fails after the last RTLS opportunity but before the first time that an AOA can be accomplished with only two main engines or when a major orbiter system failure, for example, a large cabin pressure leak or cooling system failure, occurs after the last RTLS opportunity, making it imperative to land as quickly as possible.
In a TAL abort, the vehicle continues on a ballistic trajectory across the Atlantic Ocean to land at a predetermined runway. Landing occurs approximately 45 minutes after launch. The landing site is selected near the nominal ascent ground track of the orbiter in order to make the most efficient use of space shuttle main engine propellant.
The landing site also must have the necessary runway length, weather conditions and U.S. State Department approval. Currently, the three landing sites that have been identified for a due east launch are Moron, \JSpain\j; Banjul, The Gambia; and Ben Guerir, Morocco.
To select the TAL abort mode, the crew must place the abort rotary switch in the TAL/AOA position and depress the abort push button before main engine cutoff. (Depressing it after main engine cutoff selects the AOA abort mode.)
The TAL abort mode begins sending commands to steer the vehicle toward the plane of the landing site. It also rolls the vehicle heads up before main engine cutoff and sends commands to begin an orbital maneuvering system propellant dump (by burning the propellants through the orbital maneuvering system engines and the reaction control system engines).
This dump is necessary to increase vehicle performance (by decreasing weight), to place the center of gravity in the proper place for vehicle control, and to decrease the vehicle's landing weight. TAL is handled like a nominal entry.
#
"Shuttle, Abort To Orbit",226,0,0,0
An ATO is an abort mode used to boost the orbiter to a safe orbital altitude when performance has been lost and it is impossible to reach the planned orbital altitude. If a Space Shuttle main engine fails in a region that results in a main engine cutoff under speed, the Mission Control Center will determine that an abort mode is necessary and will inform the crew. The orbital maneuvering system engines would be used to place the orbiter in a circular orbit.
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"Shuttle, Abort Once Around",227,0,0,0
The AOA abort mode is used in cases in which vehicle performance has been lost to such an extent that either it is impossible to achieve a viable orbit or not enough Orbital Maneuvering System (OMS) propellant is available to accomplish the OMS thrusting maneuver to place the orbiter on orbit and the deorbit thrusting maneuver.
In addition, an AOA is used in cases in which a major systems problem (cabin leak, loss of cooling) makes it necessary to land quickly. In the AOA abort mode, one OMS thrusting sequence is made to adjust the post-main engine cutoff orbit so a second orbital maneuvering system thrusting sequence will result in the vehicle deorbiting and landing at the AOA landing site (White Sands, N.M.; Edwards AFB; or KSC). Thus, an AOA results in the orbiter circling the Earth once and landing approximately 90 minutes after liftoff.
After the deorbit thrusting sequence has been executed, the flight crew flies to a landing at the planned site much as it would for a nominal entry.
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"Shuttle, Contingency Abort",228,0,0,0
Contingency aborts are caused by loss of more than one main engine or failures in other systems. Loss of one main engine while another is stuck at a low thrust setting may also necessitate a contingency abort. Such an abort would maintain orbiter integrity for in-flight crew escape if a landing cannot be achieved at a suitable landing field.
Contingency aborts due to system failures other than those involving the main engines would normally result in an intact recovery of vehicle and crew. Loss of more than one main engine may, depending on engine failure times, result in a safe runway landing. However, in most three-engine-out cases during ascent, the orbiter would have to be ditched. The in-flight crew escape system would be used before ditching the orbiter.
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"Orbiter Ground Turnaround",229,0,0,0
Spacecraft recovery operations at the nominal end-of-mission landing site are supported by approximately 160 Space Shuttle launch operations team members. Ground team members wearing self-contained atmospheric protective ensemble suits that protect them from toxic chemicals approach the \Jspacecraft\j as soon as it stops rolling.
The ground team members take sensor measurements to ensure the atmosphere in the vicinity of the \Jspacecraft\j is not explosive. In the event of propellant leaks, a wind machine truck carrying a large fan will be moved into the area to create a turbulent airflow that will break up gas concentrations and reduce the potential for an explosion.
A ground support equipment air-conditioning purge unit is attached to the right-hand orbiter T-0 umbilical so cool air can be directed through the orbiter's aft fuselage, payload bay, forward fuselage, wings, vertical stabilizer, and orbital maneuvering system/reaction control system pods to dissipate the heat of entry.
A second ground support equipment ground cooling unit is connected to the left-hand orbiter T-0 umbilical \Jspacecraft\j Freon Coolant loops to provide cooling for the flight crew and avionics during the postlanding and system checks. The \Jspacecraft\j fuel cells remain powered up at this time. The flight crew will then exit the \Jspacecraft\j, and a ground crew will power down the \Jspacecraft\j.
AT KSC, the orbiter and ground support equipment convoy move from the runway to the Orbiter Processing Facility.
If the \Jspacecraft\j lands at Edwards, the same procedures and ground support equipment are used as at the KSC after the orbiter has stopped on the runway. The orbiter and ground support equipment convoy move from the runway to the orbiter mate and demate facility at Edwards. After detailed inspection, the \Jspacecraft\j is prepared to be ferried atop the Shuttle carrier \Jaircraft\j from Edwards to KSC. For ferrying, a tail cone is installed over the aft section of the orbiter.
In the event of a landing at an alternate site, a crew of about eight team members will move to the landing site to assist the \Jastronaut\j crew in preparing the orbiter for loading aboard the Shuttle carrier \Jaircraft\j for transport back to the KSC. For landings outside the United States, personnel at the contingency landing sites will be provided minimum training on safe handling of the orbiter with emphasis on crash rescue training, how to tow the orbiter to a safe area, and prevention of propellant conflagration.
Upon its return to the Orbiter Processing Facility (OPF) at KSC, the orbiter is safed (ordnance devices safed), the payload (if any) is removed, and the orbiter payload bay is reconfigured from the previous mission for the next mission. Any required maintenance and inspections are also performed while the orbiter is in the OPF. A payload for the orbiter's next mission may be installed in the orbiter's payload bay in the OPF or may be installed in the payload bay when the orbiter is at the launch pad.
The \Jspacecraft\j is then towed to the Vehicle Assembly Building and mated to the external tank. The external tank and solid rocket boosters are stacked and mated on the mobile launcher platform while the orbiter is being refurbished. Space Shuttle orbiter connections are made and the integrated vehicle is checked and ordnance is installed.
The mobile launcher platform moves the entire space shuttle system on four crawlers to the launch pad, where connections are made and servicing and checkout activities begin. If the payload was not installed in the OPF, it will be installed at the launch pad followed by prelaunch activities.
Space Shuttle launches from Vandenberg will use the Vandenberg Launch Facility (SL6), which was built but never used for the manned orbital laboratory program. This facility was modified for Space Transportation System use.
The runway at Vandenberg was strengthened and lengthened from 8,000 feet to 12,000 feet to accommodate the orbiter returning from space.
When the orbiter lands at Vandenberg, the same procedures and ground support equipment and convoy are used as at KSC after the orbiter stops on the runway. The orbiter and ground support equipment are moved from the runway to the Orbiter Maintenance and Checkout Facility at Vandenberg. The orbiter processing procedures used at this facility are similar to those used at the OPF at the KSC.
Space Shuttle buildup at Vandenberg differs from that of the KSC in that the vehicle is integrated on the launch pad. The orbiter is towed overland from the Orbiter Maintenance and Checkout Facility at Vandenberg to launch facility SL6.
SL6 includes the launch mount, access tower, mobile service tower, launch control tower, payload preparation room, payload changeout room, solid rocket booster refurbishment facility, solid rocket booster disassembly facility, and liquid \Jhydrogen\j and liquid oxygen storage tank facilities.
The SRB start the on-the-launch-pad buildup followed by the external tank. The orbiter is then mated to the external tank on the launch pad.
The launch processing system at the launch pad is similar to the one used at KSC.
Kennedy Space Center Launch Operations has responsibility for all mating, prelaunch testing and launch control ground activities until the Space Shuttle vehicle clears the launch pad tower. Responsibility is then turned over to Mission Control Center-Houston. The Mission Control Center's responsibility includes ascent, on-orbit operations, entry, approach and landing until landing runout completion, at which time the orbiter is handed over to the postlanding operations at the landing site for turnaround and re-launch. At the launch site the SRBs and external tank are processed for launch and the SRBs are recycled for reuse.
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"Orbital Maneuvering System and Reaction Control System (Modifications)",230,0,0,0
The 64 valves operated by AC-motors in the OMS and RCS were modified to incorporate a "sniff" line for each valve to permit monitoring of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of the valves during ground operations. This new line reduces the probability of floating particles in the electrical microswitch portion of each valve, which could affect the operation of the microswitch position indicators for onboard displays and telemetry. It also reduces the probability of \Jnitrogen\j tetroxide or monomethyl hydrazine leakage into the bellows of each ac-motor-operated valve.
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"Orbiter, Fuel Cell Power Plants (Modifications)",231,0,0,0
End-cell heaters on each fuel cell power plant were deleted because of potential electrical failures and replaced with Freon coolant loop passages to maintain uniform temperature throughout the power plants. In addition, the \Jhydrogen\j pump and water separator of each fuel cell power plant were improved to minimize excessive \Jhydrogen\j gas entrained in the power plant product water. A current measurement detector was added to monitor the \Jhydrogen\j pump of each fuel cell power plant and provide an early indication of \Jhydrogen\j pump overload.
The starting and sustaining heater system for each fuel cell power plant was modified to prevent overheating and loss of heater elements. A stack inlet temperature measurement was added to each fuel cell power plant for full visibility of thermal conditions.
The product water from all three fuel cell power plants flows to a single water relief control panel. The water can be directed from the single panel to the Environmental Control and Life Support System (ECLSS) potable water tank A or to the fuel cell power plant water relief nozzle. Normally, the water is directed to water tank A. In the event of a line rupture in the vicinity of the single water relief panel, water could spray on all three water relief panel lines causing them to freeze and preventing water discharge.
The product water lines from all three fuel cell power plants were modified to incorporate a parallel (redundant) path of product water to ECLSS potable water tank B in the event of a freeze-up in the single water relief panel. If the single water relief panel freezes up, pressure would build up and discharge through the redundant paths to water tank B.
A water purity sensor (pH) was added at the common product water outlet of the water relief panel to provide a redundant measurement of water purity (a single measurement of water purity in each fuel cell power plant was provided previously). If the fuel cell power plant \JpH\j sensor failed in the past, the flight crew had to sample the potable water.
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"Orbiter, Auxiliary Power Units (Modifications)",232,0,0,0
The APUs that have been in use to date have a limited life. Each unit was refurbished after 25 hours of operation because of cracks in the \Jturbine\j housing, degradation of the gas generator catalyst (which varied up to approximately 30 hours of operation) and operation of the gas generator valve module (which also varied up to approximately 30 hours of operation). The remaining parts of the APU were qualified for 40 hours of operation.
Improved APUs are scheduled for delivery in late 1988. A new \Jturbine\j housing increases the life of the housing to 75 hours of operation (50 missions); a new gas generator increases its life to 75 hours; a new standoff design of the gas generator valve module and fuel pump deletes the requirement for a water spray system that was required previously for each APU upon shutdown after the first OMS thrusting period or orbital checkout; and the addition of a third seal in the middle of the two existing seals for the shaft of the fuel pump/lube oil system (previously only two seals were located on the shaft, one on the fuel pump side and one on the gearbox lube oil side) reduces the probability of hydrazine leaking into the lube oil system.
The deletion of the water spray system for the gas generator valve module and fuel pump for each APU results in a weight reduction of approximately 150 pounds for each orbiter. Upon the delivery of the improved units, the life-limited APUs will be refurbished to the upgraded design.
In the even that a fuel tank valve switch in an auxiliary power unit is inadvertently left on or an electrical short occurs within the valve electrical coil, additional protection is provided to prevent overheating of the fuel isolation valves.
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"Orbiter, Main Landing Gear (Modifications)",233,0,0,0
The following modifications were made to improve the performance of the main landing gear elements:
The thickness of the main landing gear axle was increased to provide a stiffer configuration that reduces brake-to-axle deflections and precludes brake damage experienced in previous landings. The thicker axle should also minimize tire wear.
Orifices were added to hydraulic passages in the brake's piston housing to prevent pressure surges and brake damage caused by a wobble/pump effect.
The electronic brake control boxes were modified to balance hydraulic pressure between adjacent brakes and equalize energy applications. The anti-skid circuitry previously used to reduce brake pressure to the opposite wheel if a flat tire was detected has now been removed.
The carbon-lined \Jberyllium\j stator discs in each main landing gear brake were replaced with thicker discs to increase braking energy significantly.
A long-term structural carbon brake program is in progress to replace the carbon-lined \Jberyllium\j stator discs with a carbon configuration that provides higher braking capacity by increasing maximum energy absorption.
Strain gauges were added to each nose and main landing gear wheel to monitor tire pressure before launch, deorbit and landing.
Other studies involve arresting barriers at the end of landing site runways (except lakebed runways), installing a skid on the landing gear that could preclude the potential for a second blown tire on the same gear after the first tire has blown, providing "roll on rim" for a predictable roll if both tires are lost on a single or multiple gear and adding a drag chute.
Studies of landing gear tire improvements are being conducted to determine how best to decrease tire wear observed after previous KSC landings and how to improve crosswind landing capability.
Modifications were made to the KSC Shuttle Landing Facility runway. The full 300-foot width of 3,500-foot sections at both ends of the runway were ground to smooth the runway surface texture and remove cross grooves. The modified corduroy ridges are smaller than those they replaced and run the length of the runway rather than across its width. The existing landing zone light fixtures were also modified, and the markings of the entire runway and overruns were repainted. The primary purpose of the modifications is to enhance safety by reducing tire wear during landing.
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"Orbiter, Nose Wheel Steering (Modifications)",234,0,0,0
The nose wheel steering system was modified on Columbia (OV-102) for the 61-C mission, and Discovery (OV-103) and Atlantis (OV-104) are being similarly modified before their return to flight. The modification allows a safe high-speed engagement of the nose wheel steering system and provides positive lateral directional control of the orbiter during rollout in the presence of high crosswinds and blown tires.
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"Orbiter, Thermal Protection System (Modifications)",235,0,0,0
The area aft of the reinforced carbon-carbon nose cap to the nose landing gear doors has sustained damage (tile slumping) during flight operations from impact during ascent and overheating during reentry. This area, which previously was covered with high-temperature reusable surface \Jinsulation\j tiles, will now be covered with reinforced carbon-carbon.
The low-temperature thermal protection system tiles on Columbia's midbody, payload bay doors and vertical tail were replaced with advanced Flexible Reusable Surface \JInsulation\j (FRSI) blankets.
Because of evidence of plasma flow on the lower wing trailing edge and elevon landing edge tiles (wing/elevon cove) at the outboard elevon tip and inboard elevon, the low-temperature tiles are being replaced with Fibrous Refractory Composite \JInsulation\j (FRC1-12) and High-Temperature (HRSI-22) tiles along with gap fillers on Discovery and Atlantis. On Columbia only gap fillers are installed in this area.
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"Orbiter, Wing Modification",236,0,0,0
Before the wings for Discovery and Atlantis were manufactured, a weight reduction program was instituted that resulted in a redesign of certain areas of the wing structure. An assessment of wing air loads from actual flight data indicated greater loads on the wing structure than predicted. To maintain positive margins of safety during ascent, structural modifications were incorporated into certain areas of the wings.
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"Orbiter, Mid-Fuselage Modifications",237,0,0,0
Because of additional detailed analysis of actual flight data concerning descent-stress thermal-gradient loads, torsional straps were added to tie all the lower mid-fuselage stringers in bays 1 through 11 together in a manner similar to a box section. This eliminates rotational (torsional) capabilities to provide positive margins of safety.
Also, because of the detailed analysis of actual descent flight data, room-temperature vulcanizing silicone rubber material was bonded to the lower mid-fuselage from bays 4 through 11 to act as a heat sink, distributing temperatures evenly across the bottom of the mid-fuselage, reducing thermal gradients and ensuring positive margins of safety.
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"Orbiter, General Purpose Computers (Modifications)",238,0,0,0
New upgraded General Purpose Computers (GPC), \JIBM\j AP-101S, will replace the existing GPCs aboard the Space Shuttle orbiters in late 1988 or early 1989. The upgraded computers allow NASA to incorporate more capabilities into the orbiters and apply advanced computer technologies that were not available when the orbiter was first designed.
The new computer design began in January 1984, whereas the older design began in January 1972. The upgraded GPCs provide two-and-a-half times the existing memory capacity and up to three times the existing processor speed with minimum impact on flight software. They are half the size, weigh approximately half as much, and require less power to operate.
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"Orbiter, Inertial Measurement Units (Modifications)",239,0,0,0
The new High-Accuracy Inertial Navigation System (HAINS) will be phased in in 1988-89 to augment the present KT-70 inertial measurement units . These new Inertial Measurement Units (IMUs) will result in lower program costs over the next decade, ongoing production support, improved performance, lower failure rates and reduced size and weight.
The HAINS IMUs also contain an internal dedicated \Jmicroprocessor\j with memory for processing and storing compensation and scale factor data from the IMU manufacturer's \Jcalibration\j, thereby reducing the need for extensive initial load data for the orbiter's computers. The HAINS is both physically and functionally interchangeable with the KT-70 IMU.
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"Orbiter, Crew Escape System (Modifications)",240,0,0,0
The in-flight crew escape system is provided for use only when the orbiter is in controlled gliding flight and unable to reach a runway. This would normally lead to ditching. The crew escape system provides the flight crew with an alternative to water ditching or to landing on terrain other than a landing site. The probability of the flight crew surviving a ditching is very small.
The hardware changes required to the orbiters would enable the flight crew to equalize the pressurized crew compartment with the outside pressure via a depressurization valve opened by pyrotechnics in the crew compartment aft bulkhead that would be manually activated by a flight crew member in the middeck of the crew compartment; pyrotechnically jettison the crew ingress/ egress side hatch in the middeck of the crew compartment; and bail out from the middeck of the orbiter through the ingress/ egress side hatch opening after manually deploying the escape pole through, outside and down from the side hatch opening.
One by one, each crewmember attaches a lanyard hook assembly, which surrounds the deployed escape pole, to his parachute harness and egresses through the side hatch opening. Attached to the escape pole, the crewmember slides down the pole and off the end. The escape pole provides a trajectory that takes the crew members below the orbiter's left wing.
Changes were also made in the software of the orbiter's general purpose computers. The software changes were required for the primary avionics software system and the backup flight system for transatlantic-landing and glide-return-to-launch-site aborts. The changes provide the orbiter with an automatic-mode input by the flight crew through keyboards on the commander's and/or pilot's panel C3, which provides the orbiter with an automatic stable flight for crew bailout.
The side hatch jettison feature also could be used in a landing emergency.
The emergency egress slide provides orbiter flight crew members with a means for rapid and safe exit through the orbiter middeck ingress/egress side hatch after a normal opening of the side hatch or after jettisoning the side hatch at the nominal end-of-mission landing site or at a remote or emergency landing site.
The emergency egress slide replaces the emergency egress side hatch bar, which required the flight crewmembers to drop approximately 10.5 feet to the ground. The previous arrangement could have injured crewmembers or prevented an already-injured crewmember from evacuating and moving a safe distance from the orbiter.
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"Orbiter, 17-Inch Orbiter/External Tank Disconnects (Modifications)",242,0,0,0
Each mated pair of 17-inch disconnects contains two flapper valves: one on the orbiter side and one on the external tank side. Both valves in each disconnect pair are opened to permit propellant flow between the orbiter and the external tank. Prior to separation from the external tank, both valves in each mated pair of disconnects are commanded closed by pneumatic (helium) pressure from the main propulsion system.
The closure of both valves in each disconnect pair prevents propellant discharge from the external tank or orbiter at external tank separation. Valve closure on the orbiter side of each disconnect also prevents contamination of the orbiter main propulsion system during landing and ground operations.
Inadvertent closure of either valve in a 17-inch disconnect during main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.
To prevent inadvertent closure of the 17-inch disconnect valves during the Space Shuttle main engine thrusting period, a latch mechanism was added in each orbiter half of the disconnect. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so that it overlaps both flappers and obstructs closure for any reason.
In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic pressure from the main propulsion system causes the latch actuator to rotate the shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.
A backup mechanical separation capability is provided in case a latch pneumatic actuator malfunctions. When the orbiter umbilical initially moves away from the ET umbilical, the mechanical latch disengages from the ET flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.
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"Space Shuttle Main Engine Margin Improvement Program",243,0,0,0
Improvements to the Space Shuttle Main Engines (SSMEs) for increased margin and durability began with a formal Phase II program in 1983.
Phase II focused on turbo-machinery to extend the time between high-pressure turbopump overhauls by reducing the operating temperature in the high-pressure fuel turbopump and by incorporating margin improvements to the High Pressure Fuel Turbopump (HPFT) rotor dynamics (whirl), \Jturbine\j blade and HPFT bearings. Phase II certification was completed in 1985, and all the changes have been incorporated into the SSMEs for the STS-26 mission.
In addition to the Phase II improvements, additional changes in the SSME have been incorporated to further extend the engines' margin and durability. The main changes were to the high-pressure turbo-machinery, main combustion chamber, hydraulic actuators and high-pressure \Jturbine\j discharge temperature sensors. Changes were also made in the controller software to improve engine control.
Minor high-pressure turbo-machinery design changes resulted in margin improvements to the \Jturbine\j blades, thereby extending the operating life of the turbopumps. These changes included applying surface texture to important parts of the fuel \Jturbine\j blades to improve the material properties in the pressure of \Jhydrogen\j and incorporating a damper into the high-pressure oxidizer \Jturbine\j blades to reduce vibration.
Main combustion chamber life has been increased by plating a welded outlet manifold with nickel. Margin improvements have also been made to five hydraulic actuators to preclude a loss in redundancy on the launch pad. Improvements in quality have been incorporated into the servo-component coil design along with modifications to increase margin. To address a temperature sensor in-flight anomaly, the sensor has been redesigned and extensively tested without problems.
To certify the improvements to the SSMEs and demonstrate their reliability through margin (or limit testing), an aggressive ground test program was initiated in December 1986. From December 1986 to December 1987, 151 tests and 52.363 seconds of operation (equivalent to 100 Shuttle missions) were performed. The SSMEs have exceeded 300,000 seconds total test time, the equivalent of 615 Space Shuttle missions. These hot-fire ground tests are performed at the single-engine test stands NASA's Stennis Space Center in Mississippi and at Rockwell International's Rocketdyne Division's Santa Susana Field Laboratory in \JCalifornia\j.
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"SSME Flight Program",244,0,0,0
By January 1986, there have been 25 flights (75 engine launches with three SSMEs per flight) of the SSMEs. A total of 13 engines were flown, and SSME reusability was demonstrated. One engine (serial number 2012) has been flown 10 times; 10 other engines have flown between five and nine times. Two off-nominal conditions were experienced on the launch pad and one during flight.
Two fail-safe shutdowns occurred on the launch pad during engine start but before SRB ignition. In each case, the controller detected a loss of redundancy in the hydraulic actuator system and commanded engine shutdown in keeping with the launch commit criteria.
Another loss of redundancy occurred in flight with a loss of a red-line temperature sensor and its backup. The engine was commanded to shut down, but the other two engines safely delivered the Space Shuttle to orbit. A major upgrade of these components was implemented to prevent a recurrence of these conditions and will be incorporated for STS-26.
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"Solid Rocket Motor Redesign",245,0,0,0
On June 13, 1986, President Reagan directed NASA to implement, as soon as possible, the recommendations of the "Presidential Commission on the Space Shuttle Challenger Accident." NASA developed a plan to provide a Redesigned Solid Rocket Motor (RSRM).
The primary objective of the redesign effort was to provide an SRM that is safe to fly. A secondary objective was to minimize impact on the schedule by using existing hardware, to the extent practical, without compromising safety. A joint redesign team was established that included participation from Marshall Space Flight Center, Morton Thiokol and other NASA centers as well as individuals from outside NASA.
An "SRM Redesign Project Plan" was developed to formalize the methodology for SRM redesign and requalification. The plan provided an overview of the organizational responsibilities and relationships, the design objectives, criteria and process; the verification approach and process; and a master schedule. The companion "Development and Verification Plan" defined the test program and analyses required to verify the redesign and the unchanged components of the SRM.
All aspects of the existing SRM were assessed, and design changes were required in the field joint, case-to-nozzle joint, nozzle, factory joint, propellant grain shape, ignition system and ground support equipment. No changes were made in the propellant, liner or castable inhibitor formulations. Design criteria were established for each component to ensure a safe design with an adequate margin of safety. These criteria focused on loads, environments, performance, redundancy, margins of safety and verification philosophy.
The criteria were converted into specific design requirements during the Preliminary Requirements Reviews held in July and August 1986. The design developed from these requirements was assessed at the Preliminary Design Review held in September 1986 and baselined in October 1986.
The final design was approved at the Critical Design Review held in October 1987. Manufacture of the RSRM test hardware and the first flight hardware began prior to the Preliminary Design Review (PDR) and continued in parallel with the hardware certification program. The Design Certification Review will review the analyses and test results versus the program and design requirements to certify the redesigned SRM is ready to fly.
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"Shuttle SRM Field Joint Redesigned",246,0,0,0
The SRM (Solid Rocket Motor) field-joint metal parts, internal case \Jinsulation\j and seals were redesigned and a weather protection system was added.
In the STS 51-L design, the application of actuating pressure to the upstream face of the O-ring was essential for proper joint sealing performance because large sealing gaps were created by pressure-induced deflections, compounded by significantly reduced O-ring sealing performance at low temperature.
The major change in the motor case is the new tang capture feature to provide a positive metal-to-metal interference fit around the \Jcircumference\j of the tang and clevis ends of the mating segments. The interference fit limits the deflection between the tang and clevis O-ring sealing surfaces caused by motor pressure and structural loads. The joints are designed so that the seals will not leak under twice the expected structural deflection and rate.
The new design, with the tang capture feature, the interference fit and the use of custom shims between the outer surface of the tang and inner surface of the outer clevis leg, controls the O-ring sealing gap dimension.
The sealing gap and the O-ring seals are designed so that a positive compression (squeeze) is always on the O-rings. The minimum and maximum squeeze requirements include the effects of temperature, O-ring resiliency and compression set, and pressure. The clevis O-ring groove dimension has been increased so that the O-ring never fills more than 90 percent of the O-ring groove and pressure actuation is enhanced.
The new field joint design also includes a new O-ring in the capture feature and an additional leak check port to ensure that the primary O-ring is positioned in the proper sealing direction at ignition. This new or third O-ring also serves as a thermal barrier in case the sealed \Jinsulation\j is breached.
The field joint internal case \Jinsulation\j was modified to be sealed with a pressure-actuated flap called a J-seal, rather than with putty as in the STS 51-L configuration.
Longer field-joint-case mating pins, with a reconfigured retainer band, were added to improve the shear strength of the pins and increase the metal parts' joint margin of safety. The joint safety margins, both thermal and structural, are being demonstrated over the full ranges of ambient temperature, storage compression, grease effect, assembly stresses and other environments. External heaters with integral weather seals were incorporated to maintain the joint and O-ring temperature at a minimum of 75 F. The weather seal also prevents water intrusion into the joint.
The SRM case-to nozzle joint, which experienced several instances of O-ring erosion in flight, has been redesigned to satisfy the same requirements imposed upon the case field joint. Similar to the field joint, cast-to-nozzle joint modifications have been made in the metal parts, internal \Jinsulation\j and O-rings. Radial bolts with Stato-O-Seals were added to minimize the joint sealing gap opening.
The internal \Jinsulation\j was modified to be sealed adhesively, and third O-ring was included. The third O-ring serves as a dam or wiper in front of the primary O-ring to prevent the polysulfide adhesive from being extruded into the primary O-ring groove. It also serves as a thermal barrier in case the polysulfide adhesive is breached. The polysulfide adhesive replaces the putty used in the 51-L joint. Also, an additional leak check port was added to reduce the amount of trapped air in the joint during the nozzle installation process and to aid in the leak check procedure.
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"Shuttle Nozzle Redesigned",248,0,0,0
The internal joints of the nozzle metal parts have been redesigned to incorporate redundant and verifiable O-rings at each joint. The nozzle steel fixed housing part has been redesigned to permit the incorporation of the 100 radial bolts that attach the fixed housing to the case's aft dome. Improved bonding techniques are being used for the nozzle nose inlet, cowl/boot and aft exit cone assemblies.
The distortion of the nose inlet assembly's metal-part-to-ablative-parts bond line has been eliminated by increasing the thickness of the aluminum nose inlet housing and improving the bonding process. The tape-wrap angle of the carbon cloth fabric in the areas of the nose inlet and throat assembly parts was changed to improve the ablative \Jinsulation\j erosion tolerance. Some of these ply-angle changes were in progress prior to STS 51-L. The cowl and outer boot ring has additional structural support with increased thickness and contour changes to increase their margins of safety. Additionally, the outer boot ring ply configuration was altered.
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"Shuttle Factory Joint Modified",249,0,0,0
Minor modifications were made in the case factory joints by increasing the \Jinsulation\j thickness and lay-up to increase the margin of safety on the internal \Jinsulation\j. Longer pins were also added, along wit a reconfigured retainer band and new weather seal to improve factory joint performance and increase the margin of safety. Additionally, the O-ring and O-ring groove size was changed to be consistent with the field joint.
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"Shuttle Solid Rocket Motor Ignition System Modified",250,0,0,0
Several minor modifications were incorporated into the ignition system. The aft end of the igniter steel case, which contains the igniter nozzle insert, was thickened to eliminate a localized weakness. The igniter internal case \Jinsulation\j was tapered to improve the manufacturing process. Finally, although vacuum putty is still being used at the joint of the igniter and case forward dome, it was changed to eliminate \Jasbestos\j as one of its constituents.
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"Shuttle Ground Support Equipment Redesigned",251,0,0,0
The ground support equipment has been redesigned to (1) minimize the case distortion during handling at the launch site; (2) improve the segment tang and clevis joint measurement system for more accurate reading of case diameters to facilitate stacking; (3) minimize the risk of O-ring damage during joint mating; and (4) improve leak testing of the igniter, case and nozzle field joints.
A Ground Support Equipment (GSE) assembly aid guides the segment tang into the clevis and rounds the two parts with each other. Other GSE modifications include transportation monitoring equipment and lifting beam.
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"Shuttle Design Analysis Summary",252,0,0,0
Improved, state-of-the-art, analyses related to structural strength, loads, stress, dynamics, fracture mechanics, gas and thermal dynamics, and material characterization and behavior were performed to aid the field joint, nozzle-to-case joint and other designs. Continuing these analyses will ensure that the design integrity and system compatibility adhere to design requirements and operational use. These analyses will be verified by tests, whose results will be correlated with pre-test predictions.
The verification program demonstrates that the RSRM meets all design and performance requirements, and that failure modes and hazards have been eliminated or controlled. The verification program encompasses the following program phases: development, certification, acceptance, preflight checkout, flight and postflight.
Redesigned SRM (Solid Rocket Motor) certification is based on formally documented results of development motor tests; qualification motor tests and other tests and analyses. The certification tests are conducted under strict control of environments, including thermal and structural loads; assembly, inspection and test procedures; and safety, reliability, maintainability and quality assurance surveillance to verify that flight hardware meets the specified performance and design requirements. The "Development and Verification Plan" stipulates the test program, which follows a rigorous sequence wherein successive tests build on the results of previous tests leading to formal certification.
The test activities include laboratory and component tests, subscale tests, full-scale simulation and full-scale motor static test firings. Laboratory and component tests are used to determine component properties and characteristics. Subscale motor firings are used to simulate gas dynamics and thermal conditions for components and subsystem design.
Full-scale hardware simulators are used to verify analytical models; determine hardware assembly characteristics; determine joint deflection characteristics; determine joint performance under short-duration hot-gas tests, including joint flaws and flight loads; and determine redesigned hardware structural characteristics.
Fourteen full-scale joint assembly demonstration vertical mate/demate tests, with eight interspersed hydro tests to simulate flight hardware refurbishment procedures, were completed early for the redesigned capture-feature hardware. Assembly loads were as expected, and the case growth was as predicted with no measurable increase after three hydro-proof tests.
Flight-configuration aft and center segments were fabricated, loaded with live propellant, and used for assembly test article stacking demonstration tests at Kennedy Space Center. These tests were pathfinder demonstrations for the assembly of flight hardware using newly developed ground support equipment.
In a long-term stack test, a full-scale casting segment, with live propellant, has been mated vertically with a J-seal \Jinsulation\j segment and is undergoing temperature \Jcycling\j. This will determine the compression set of the J-seal, aging effects and long-term propellant slumping effects.
The Structural Test Article (STA-3), consisting of flight-type forward and aft motor segments and forward and aft skirts, was subjected to extensive static and dynamic structural testing, including maximum prelaunch, liftoff and flight (maximum dynamic pressure) structural loads.
Redesigned SRM certification includes testing the actual flight configuration over the full range of operating environments and conditions. The joint environment simulator, transient pressure test article, and the nozzle joint environment simulator test programs all utilize full-scale flight design hardware and subject the RSRM design features to the maximum expected operating pressure, maximum pressure rise rate and temperature extremes during ignition tests. Additionally, the Transient Pressure Test Article (TPTA) is subjected to ignition and liftoff loads as well as maximum dynamic pressure structural loads.
Four TPTA tests have been completed to subject the redesigned case field and case-to-nozzle joints to the above-described conditions. The field and case-to-nozzle joints were temperature-conditioned to 75 F. and contained various types of flaws in the joints so that the primary and secondary O-rings could be pressure-actuated, joint rotation and O-ring performance could be evaluated and the redesigned joints could be demonstrated as fail safe.
Six of the seven Joint Environment Simulators (JES) tests have been completed. The JES test program initially used the STS 51-L configuration hardware to evaluate the joint performance with prefabricated blowholes through the putty. The JES-1 test series, which consisted of two tests, established a structural and performance \Jdatabase\j for the STS 51-L configuration with and without a replicated joint failure.
The JES-2 series, two tests, also used the STS 51-L case metal-part joint but with a bonded labyrinth and U-seal \Jinsulation\j that was an early design variation of the J-seal. Tests were conducted with and without flaws built into the U-seal joint \Jinsulation\j; neither joint showed O-ring erosion or blow-by. The JES-3 series, three tests, uses almost exact flight configuration hardware, case field-joint capture feature with interference fit and J-seal \Jinsulation\j.
Four of five nozzle JES tests have been successfully conducted. The STS 51-L hardware configuration hydro test confirmed predicted case-to-nozzle-joint deflection. The other three tests used the radially bolted RSRM configuration.
Seven full-scale, full-duration motor static tests are being conducted to verify the integrated RSRM performance. These include one \Jengineering\j test motor used to (1) provide a data base for STS 51-L-type field joints; (2) evaluate new seal material; (3) evaluate the ply-angle change in the nozzle parts,; (4) evaluate the effectiveness of \Jgraphite\j composite stiffener rings to reduce joint rotation; and (5) evaluate field-joint heaters.
There were two development motor tests and three qualification motor tests for final flight configuration and performance certification. There will be one flight Production Verification Motor that contains intentionally induced defects in the joints to demonstrate joint performance under extreme worse case conditions. The QM-7 and QM-8 motors were subjected to liftoff and maximum dynamic pressure structural loads, QM-7 was temperature-conditioned to 90 F., and QM-8 was temperature-conditioned to 40 F.
An assessment was conducted to determine the full-duration static firing test attitude necessary to certify the design changes completely. The assessment included establishing test objectives, defining and quantifying attitude-sensitive parameters, and evaluating attitude options. Both horizontal and vertical (nozzle up and down) test attitudes were assessed.
In all three options, consideration was given to testing with and without externally applied loads. This assessment determined that the conditions influencing the joint and \Jinsulation\j behavior could best be tested to design extremes in the horizontal attitude. In conjunction with the horizontal attitude for the RSRM full-scale testing, it was decided to incorporate externally applied loads.
A second horizontal test stand for certification of the RSRM was constructed at Morton Thiokol. This new stand, designated as the T-97 Large Motor Static Test Facility, is being used to simulate environmental stresses, loads and temperatures experienced during an actual Shuttle launch and ascent. The new test stand also provides redundancy for the existing stand.
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"Shuttle, Non-Destructive Evaluation",254,0,0,0
The Shuttle 51-L and Titan 34D-9 vehicle failures, both of which occurred in 1986, resulted in major reassessments of each vehicle's design, processing, inspection and operations. While the Shuttle SRM insulation/ propellant integrity was not implicated in the 51-L failure, the intent is to preclude a failure similar to that experienced by Titan.
The RSRM field joint is quite tolerant of unbonded \Jinsulation\j. It has sealed \Jinsulation\j to prevent hot combustion products from reaching the insulation-to-case bond line. The bonding processes have been improved to reduce contamination potential, and the new \Jgeometry\j of the tang capture feature inherently provides more isolation of the edge \Jinsulation\j area from contaminating agents.
A greatly enhanced Non-Destructive Evaluation program for the RSRM has been incorporated. The enhanced non-destructive testing includes ultrasonic inspection and mechanical testing of propellant and \Jinsulation\j bonded surfaces. All segments will again be X-rayed for the first flight and near-term subsequent flights.
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"Shuttle, Contingency Planning",255,0,0,0
To provide additional program confidence, both near- and long-term contingency planning was implemented. Alternative designs, which might be incorporated into the flight program at discrete decision points, include field-joint graphite-composite overwrap bands and alternative seals for the field joint and case-to-nozzle joint. Alternative designs for the nozzle include a different composite lay-up technique and a steel nose inlet housing.
Alternative designs with long-lead-time implications were also developed. These designs focus on the field joint and cast-to-nozzle joint. Since fabrication of the large steel components dictates the schedule, long-lead procurement of maximum-size steel ingots was initiated. This allowed machining of case joints to either the new baseline or to an alternative design configuration. Ingot processing continued through forging and heat treating. At that time, the final design was selected. A principal consideration in this configuration decision was the result of verification testing on the baseline configuration.
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"NASA, Independent Oversight Panel",256,0,0,0
As recommended in the "Presidential Commission Report" and at the request of the NASA administrator, the National Research Council established an Independent Oversight Panel chaired by Dr. H. Guyford Stever, who reports directly to the NASA Administrator.
Initially, the panel was given introductory briefings on the Shuttle system requirements, implementation and control, the original design and manufacturing of the SRM, Mission 51-L accident analyses and preliminary plans for the redesign. The panel has met with major SRM manufacturers and vendors, and has visited some of their facilities.
The panel frequently reviewed the RSRM design criteria, \Jengineering\j analyses and design, and certification program planning. Panel members continuously review the design and testing for safe operation, selection and specifications for material, and quality assurance and control. The panel has continued to review the design as it progresses through certification and review the manufacturing and assembly of the first flight RSRM.
Panel members have participated in major program milestones, project requirements review, and preliminary design review; they also will participate in future review. Six written reports have been provided by the panel to the NASA administrator.
In addition to the NRC, the redesign team has a design review group of 12 expert senior engineers from NASA and the aerospace industry. They have advised on major program decisions and serve as a "sounding board" for the program.
Additionally, NASA requested the four other major SRM companies -- Aerojet Strategic Propulsion Co., Atlantic Research Corp., Hercules Inc. and United Technologies Corp.'s Chemical Systems Division -- to participate in the redesign efforts by critiquing the design approach and providing experience on alternative design approaches.
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"Solid Rocket Boosters (SRBs)",257,0,0,0
The two SRBs provide the main thrust to lift the space shuttle off the pad and up to an altitude of about 150,000 feet, or 24 nautical miles (28 statute miles). In addition, the two SRBs carry the entire weight of the external tank and orbiter and transmit the weight load through their structure to the mobile launcher platform. Each booster has a thrust (sea level) of approximately 3,300,000 pounds at launch.
They are ignited after the three space shuttle main engines' thrust level is verified. The two SRBs provide 71.4 percent of the thrust at lift- off and during first-stage ascent. Seventy- five seconds after SRB separation, SRB apogee occurs at an altitude of approximately 220,000 feet, or 35 nautical miles (41 statute miles). SRB impact occurs in the ocean approximately 122 nautical miles (141 statute miles) downrange.
The SRBs are the largest solid- propellant motors ever flown and the first designed for reuse. Each is 149.16 feet long and 12.17 feet in diameter.
Each SRB weighs approximately 1,300,000 pounds at launch. The propellant for each solid rocket motor weighs approximately 1,100,000 pounds. The inert weight of each SRB is approximately 192,000 pounds.
Primary elements of each booster are the motor (including case, propellant, igniter and nozzle), structure, separation systems, operational flight instrumentation, recovery avionics, pyrotechnics, deceleration system, thrust vector control system and range safety destruct system.
Each booster is attached to the external tank at the SRB's aft frame by two lateral sway braces and a diagonal attachment. The forward end of each SRB is attached to the external tank at the forward end of the SRB's forward skirt. On the launch pad, each booster also is attached to the mobile launcher platform at the aft skirt by four bolts and nuts that are severed by small explosives at lift-off.
During the downtime following the Challenger accident, detailed structural analyses were performed on critical structural elements of the SRB. Analyses were primarily focused in areas where anomalies had been noted during postflight inspection of recovered hardware.
One of the areas was the attach ring where the SRBs are connected to the external tank. Areas of distress were noted in some of the fasteners where the ring attaches to the SRB motor case. This situation was attributed to the high loads encountered during water impact. To correct the situation and ensure higher strength margins during ascent, the attach ring was redesigned to encircle the motor case completely (360 degrees). Previously, the attach ring formed a C and encircled the motor case 270 degrees.
Additionally, special structural tests were performed on the aft skirt. During this test program, an anomaly occurred in a critical weld between the hold-down post and skin of the skirt. A redesign was implemented to add reinforcement brackets and fittings in the aft ring of the skirt.
These two modifications added approximately 450 pounds to the weight of each SRB.
The propellant mixture in each SRB motor consists of an ammonium perchlorate (oxidizer, 69.6 percent by weight), aluminum (fuel, 16 percent), iron oxide (a catalyst, 0.4 percent), a polymer (a binder that holds the mixture together, 12.04 percent), and an epoxy curing agent (1.96 percent). The propellant is an 11-point star- shaped perforation in the forward motor segment and a double- truncated- cone perforation in each of the aft segments and aft closure. This configuration provides high thrust at ignition and then reduces the thrust by approximately a third 50 seconds after lift-off to prevent overstressing the vehicle during maximum dynamic pressure.
The SRBs are used as matched pairs and each is made up of four solid rocket motor segments. The pairs are matched by loading each of the four motor segments in pairs from the same batches of propellant ingredients to minimize any thrust imbalance. The segmented-casing design assures maximum flexibility in fabrication and ease of transportation and handling. Each segment is shipped to the launch site on a heavy- duty rail car with a specially built cover.
The nozzle expansion ratio of each booster beginning with the STS-8 mission is 7-to-79. The nozzle is gimbaled for thrust vector (direction) control. Each SRB has its own redundant auxiliary power units and hydraulic pumps. The all-axis gimbaling capability is 8 degrees. Each nozzle has a carbon cloth liner that erodes and chars during firing. The nozzle is a convergent- divergent, movable design in which an aft pivot- point flexible bearing is the gimbal mechanism.
The cone- shaped aft skirt reacts the aft loads between the SRB and the mobile launcher platform. The four aft separation motors are mounted on the skirt. The aft section contains avionics, a thrust vector control system that consists of two auxiliary power units and hydraulic pumps, hydraulic systems and a nozzle extension jettison system.
The forward section of each booster contains avionics, a sequencer, forward separation motors, a nose cone separation system, drogue and main parachutes, a recovery beacon, a recovery light, a parachute camera on selected flights and a range safety system.
Each SRB has two integrated electronic assemblies, one forward and one aft. After burnout, the forward assembly initiates the release of the nose cap and frustum and turns on the recovery aids. The aft assembly, mounted in the external tank/SRB attach ring, connects with the forward assembly and the orbiter avionics systems for SRB ignition commands and nozzle thrust vector control. Each integrated electronic assembly has a multiplexer/ demultiplexer, which sends or receives more than one message, signal or unit of information on a single communication channel.
Eight booster separation motors (four in the nose frustum and four in the aft skirt) of each SRB thrust for 1.02 seconds at SRB separation from the external tank. Each solid rocket separation motor is 31.1 inches long and 12.8 inches in diameter.
Location aids are provided for each SRB, frustum/ drogue chutes and main parachutes. These include a transmitter, antenna, strobe/ converter, battery and salt water switch \Jelectronics\j. The location aids are designed for a minimum operating life of 72 hours and when refurbished are considered usable up to 20 times. The flashing light is an exception. It has an operating life of 280 hours. The battery is used only once.
The SRB nose caps and nozzle extensions are not recovered.
The recovery crew retrieves the SRBs, frustum/ drogue chutes, and main parachutes. The nozzles are plugged, the solid rocket motors are dewatered, and the SRBs are towed back to the launch site. Each booster is removed from the water, and its components are disassembled and washed with fresh and deionized water to limit salt water \Jcorrosion\j. The motor segments, igniter and nozzle are shipped back to Thiokol for refurbishment.
Each SRB incorporates a range safety system that includes a battery power source, receiver/ decoder, antennas and ordnance.
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"SRB Hold-Down Posts",258,0,0,0
Each solid rocket booster has four hold- down posts that fit into corresponding support posts on the mobile launcher platform. Hold- down bolts hold the SRB and launcher platform posts together. Each bolt has a nut at each end, but only the top nut is frangible. The top nut contains two NASA standard detonators, which are ignited at solid rocket motor ignition commands.
When the two NSDs are ignited at each hold- down, the hold- down bolt travels downward because of the release of tension in the bolt (pretensioned before launch), NSD gas pressure and gravity. The bolt is stopped by the stud deceleration stand, which contains sand. The SRB bolt is 28 inches long and is 3.5 inches in diameter. The frangible nut is captured in a blast container.
The solid rocket motor ignition commands are issued by the orbiter's computers through the master events controllers to the hold- down pyrotechnic initiator controllers on the mobile launcher platform. They provide the ignition to the hold- down NSDs. The launch processing system monitors the SRB hold- down PICs for low voltage during the last 16 seconds before launch. PIC low voltage will initiate a launch hold.
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"SRB Ignition",259,0,0,0
SRB (Solid Rocket Booster) ignition can occur only when a manual lock pin from each SRB safe and arm device has been removed. The ground crew removes the pin during prelaunch activities. At T minus five minutes, the SRB safe and arm device is rotated to the arm position. The solid rocket motor ignition commands are issued when the three SSMEs are at or above 90-percent rated thrust, no SSME fail and/or SRB ignition PIC low voltage is indicated and there are no holds from the LPS.
The solid rocket motor ignition commands are sent by the orbiter computers through the MECs to the safe and arm device NSDs in each SRB. A PIC single-channel capacitor discharge device controls the firing of each pyrotechnic device.
Three signals must be present simultaneously for the PIC to generate the pyro firing output. These signals- arm, fire 1 and fire 2-originate in the orbiter general- purpose computers and are transmitted to the MECs. The MECs reformat them to 28-volt dc signals for the PICs. The arm signal charges the PIC capacitor to 40 volts dc (minimum of 20 volts dc).
The fire 2 commands cause the redundant NSDs to fire through a thin barrier seal down a flame tunnel. This ignites a pyro booster charge, which is retained in the safe and arm device behind a perforated plate. The booster charge ignites the propellant in the igniter initiator; and combustion products of this propellant ignite the solid rocket motor initiator, which fires down the length of the solid rocket motor igniting the solid rocket motor propellant.
The GPC launch sequence also controls certain critical main propulsion system valves and monitors the engine- ready indications from the SSMEs. The MPS start commands are issued by the onboard computers at T minus 6.6 seconds (staggered start- engine three, engine two, engine one- all approximately within 0.25 of a second), and the sequence monitors the thrust buildup of each engine. All three SSMEs must reach the required 90-percent thrust within three seconds; otherwise, an orderly shutdown is commanded and safing functions are initiated.
Normal thrust buildup to the required 90-percent thrust level will result in the SSMEs being commanded to the lift- off position at T minus three seconds as well as the fire 1 command being issued to arm the SRBs. At T minus three seconds, the vehicle base bending load modes are allowed to initialize (movement of approximately 25.5 inches measured at the tip of the external tank, with movement towards the external tank).
At T minus zero, the two SRBs are ignited, under command of the four onboard computers; separation of the four explosive bolts on each SRB is initiated (each bolt is 28 inches long and 3.5 inches in diameter); the two T-0 umbilicals (one on each side of the spacecraft) are retracted; the onboard master timing unit, event timer and mission event timers are started; the three SSMEs are at 100 percent; and the ground launch sequence is terminated.
The solid rocket motor thrust profile is tailored to reduce thrust during the maximum dynamic pressure region.
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"SRB Electrical Power Distribution",260,0,0,0
Electrical power distribution in each SRB (Solid Rocket Booster) consists of orbiter- supplied main dc bus power to each SRB via SRB buses A, B and C. Orbiter main dc buses A, B and C supply main dc bus power to corresponding SRB buses A, B and C. In addition, orbiter main dc bus C supplies backup power to SRB buses A and B, and orbiter bus B supplies backup power to SRB bus C. This electrical power distribution arrangement allows all SRB buses to remain powered in the event one orbiter main bus fails.
The nominal dc voltage is 28 volts dc, with an upper limit of 32 volts dc and a lower limit of 24 volts dc.
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"SRB Hydraulic Power Units",261,0,0,0
There are two self- contained, independent HPUs on each SRB . Each HPU consists of an auxiliary power unit, fuel supply module, hydraulic pump, hydraulic reservoir and hydraulic fluid manifold assembly. The APUs are fueled by hydrazine and generate mechanical shaft power to a hydraulic pump that produces hydraulic pressure for the SRB hydraulic system.
The two separate HPUs and two hydraulic systems are located on the aft end of each SRB between the SRB nozzle and aft skirt. The HPU components are mounted on the aft skirt between the rock and tilt actuators. The two systems operate from T minus 28 seconds until SRB separation from the orbiter and external tank. The two independent hydraulic systems are connected to the rock and tilt servoactuators.
The APU controller \Jelectronics\j are located in the SRB aft integrated electronic assemblies on the aft external tank attach rings.
The APUs and their fuel systems are isolated from each other. Each fuel supply module (tank) contains 22 pounds of hydrazine. The fuel tank is pressurized with gaseous \Jnitrogen\j at 400 psi, which provides the force to expel (positive expulsion) the fuel from the tank to the fuel distribution line, maintaining a positive fuel supply to the APU throughout its operation.
The fuel isolation valve is opened at APU startup to allow fuel to flow to the APU fuel pump and control valves and then to the gas generator. The gas generator's catalytic action decomposes the fuel and creates a hot gas. It feeds the hot gas exhaust product to the APU two- stage gas \Jturbine\j. Fuel flows primarily through the startup bypass line until the APU speed is such that the fuel pump outlet pressure is greater than the bypass line's. Then all the fuel is supplied to the fuel pump.
The APU \Jturbine\j assembly provides mechanical power to the APU gearbox. The gearbox drives the APU fuel pump, hydraulic pump and lube oil pump. The APU lube oil pump lubricates the gearbox. The \Jturbine\j exhaust of each APU flows over the exterior of the gas generator, cooling it, and is then directed overboard through an exhaust duct.
When the APU speed reaches 100 percent, the APU primary control valve closes, and the APU speed is controlled by the APU controller \Jelectronics\j. If the primary control valve logic fails to the open state, the secondary control valve assumes control of the APU at 112-percent speed.
Each HPU on an SRB is connected to both servoactuators on that SRB. One HPU serves as the primary hydraulic source for the servoactuator, and the other HPU serves as the secondary \Jhydraulics\j for the servoactuator. Each servoactuator has a switching valve that allows the secondary \Jhydraulics\j to power the actuator if the primary hydraulic pressure drops below 2,050 psi. A switch contact on the switching valve will close when the valve is in the secondary position. When the valve is closed, a signal is sent to the APU controller that inhibits the 100-percent APU speed control logic and enables the 112-percent APU speed control logic. The 100-percent APU speed enables one APU/HPU to supply sufficient operating hydraulic pressure to both servoactuators of that SRB.
The APU 100-percent speed corresponds to 72,000 rpm, 110-percent to 79,200 rpm, and 112-percent to 80,640 rpm.
The hydraulic pump speed is 3,600 rpm and supplies hydraulic pressure of 3,050, plus or minus 50, psi. A high- pressure relief valve provides overpressure protection to the hydraulic system and relieves at 3,750 psi.
The APUs/HPUs and hydraulic systems are reusable for 20 missions.
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"SRB Thrust Vector Control",262,0,0,0
Each SRB has two hydraulic gimbal servoactuators: one for rock and one for tilt. The servoactuators provide the force and control to gimbal the nozzle for thrust vector control.
The space shuttle ascent thrust vector control portion of the flight control system directs the thrust of the three shuttle main engines and the two SRB nozzles to control shuttle attitude and trajectory during lift- off and ascent. Commands from the guidance system are transmitted to the ATVC drivers, which transmit signals proportional to the commands to each servoactuator of the main engines and SRBs.
Four independent flight control system channels and four ATVC channels control six main engine and four SRB ATVC drivers, with each driver controlling one hydraulic port on each main and SRB servoactuator.
Each SRB servoactuator consists of four independent, two- stage servovalves that receive signals from the drivers. Each servovalve controls one power spool in each actuator, which positions an actuator ram and the nozzle to control the direction of thrust.
The four servovalves in each actuator provide a force- summed majority voting arrangement to position the power spool. With four identical commands to the four servovalves, the actuator force-sum action prevents a single erroneous command from affecting power ram motion. If the erroneous command persists for more than a predetermined time, differential pressure sensing activates a selector valve to isolate and remove the defective servovalve hydraulic pressure, permitting the remaining channels and servovalves to control the actuator ram spool.
Failure monitors are provided for each channel to indicate which channel has been bypassed. An isolation valve on each channel provides the capability of resetting a failed or bypassed channel.
Each actuator ram is equipped with transducers for position feedback to the thrust vector control system. Within each servoactuator ram is a splashdown load relief assembly to cushion the nozzle at water splashdown and prevent damage to the nozzle flexible bearing.
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"SRB Rate Gyro Assemblies",263,0,0,0
Each SRB contains two RGAs, with each RGA containing one pitch and one yaw gyro. These provide an output proportional to angular rates about the pitch and yaw axes to the orbiter computers and guidance, navigation and control system during first- stage ascent flight in conjunction with the orbiter roll rate gyros until SRB separation. At SRB separation, a switchover is made from the SRB RGAs to the orbiter RGAs.
The SRB RGA rates pass through the orbiter flight aft multiplexers/ demultiplexers to the orbiter GPCs. The RGA rates are then mid-value- selected in redundancy management to provide SRB pitch and yaw rates to the user software. The RGAs are designed for 20 missions.
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"SRB Separation",264,0,0,0
SRB (Solid Rocket Booster) separation is initiated when the three solid rocket motor chamber pressure transducers are processed in the redundancy management middle value select and the head- end chamber pressure of both SRBs is less than or equal to 50 psi. A backup cue is the time elapsed from booster ignition.
The separation sequence is initiated, commanding the thrust vector control actuators to the null position and putting the main propulsion system into a second-stage configuration (0.8 second from sequence initialization), which ensures the thrust of each SRB is less than 100,000 pounds. Orbiter yaw attitude is held for four seconds, and SRB thrust drops to less than 60,000 pounds.
The SRBs separate from the external tank within 30 milliseconds of the ordnance firing command.
The forward attachment point consists of a ball (SRB) and socket (ET) held together by one bolt. The bolt contains one NSD pressure cartridge at each end. The forward attachment point also carries the range safety system cross-strap wiring connecting each SRB RSS and the ET RSS with each other.
The aft attachment points consist of three separate struts: upper, diagonal and lower. Each strut contains one bolt with an NSD pressure cartridge at each end. The upper strut also carries the umbilical interface between its SRB and the external tank and on to the orbiter.
There are four booster separation motors on each end of each SRB. The BSMs separate the SRBs from the external tank. The solid rocket motors in each cluster of four are ignited by firing redundant NSD pressure cartridges into redundant confined detonating fuse manifolds.
The separation commands issued from the orbiter by the SRB separation sequence initiate the redundant NSD pressure cartridge in each bolt and ignite the BSMs to effect a clean separation.
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"SRB Range Safety System",265,0,0,0
The shuttle vehicle has three RSSs. One is located in each SRB and one in the external tank. Any one or all three are capable of receiving two command messages (arm and fire) transmitted from the ground station. The RSS is used only when the shuttle vehicle violates a launch trajectory red line.
An RSS consists of two antenna couplers, command receivers/ decoders, a dual distributor, a safe and arm device with two NSDs, two confined detonating fuse manifolds, seven CDF assemblies and one linear-shaped charge.
The antenna couplers provide the proper \Jimpedance\j for radio frequency and ground support equipment commands. The command receivers are tuned to RSS command frequencies and provide the input signal to the distributors when an RSS command is sent. The command decoders use a code plug to prevent any command signal other than the proper command signal from getting into the distributors. The distributors contain the logic to supply valid destruct commands to the RSS pyrotechnics.
The NSDs provide the spark to ignite the CDF, which in turn ignites the LSC for shuttle vehicle destruction. The safe and arm device provides mechanical isolation between the NSDs and the CDF before launch and during the SRB separation sequence.
The first message, called arm, allows the onboard logic to enable a destruct and illuminates a light on the flight deck display and control panel at the commander and pilot station. The second message transmitted is the fire command.
The SRB distributors in the SRBs and the ET are cross- strapped together. Thus, if one SRB received an arm or destruct signal, the signal would also be sent to the other SRB and the ET.
Electrical power from the RSS battery in each SRB is routed to RSS system A. The recovery battery in each SRB is used to power RSS system B as well as the recovery system in the SRB. The SRB RSS is powered down during the separation sequence, and the SRB recovery system is powered up. Electrical power for the ET RSS system A and system B is independently supplied by two RSS batteries on the ET.
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"SRB Descent and Recovery",266,0,0,0
The recovery sequence begins with the operation of the high-altitude baroswitch, which triggers the functioning of the pyrotechnic nose cap thrusters. This ejects the nose cap, which deploys the pilot parachute. This occurs at 15,704 feet altitude 225 seconds after separation. The 11.5-foot-diameter conical ribbon pilot parachute provides the force to pull the lanyard activating the zero-second cutter, which cuts the loop securing the drogue retention straps.
This allows the pilot chute to pull the drogue pack from the SRB, causing the drogue suspension lines to deploy from their stored position. At full extension of the 12 95-foot suspension lines, the drogue deployment bag is stripped away from the canopy, and the 54-foot-diameter conical ribbon drogue parachute inflates to its initial reefed condition. The drogue disreefs twice after specified time delays, and it reorients/stabilizes the SRB for main chute deployment. The drogue parachute can withstand a load of 270,000 pounds and weighs approximately 1,200 pounds.
After the drogue chute has stabilized the vehicle in a tailfirst attitude, the frustum is separated from the forward skirt by a charge triggered by the low-altitude baroswitch at an altitude of 5,975 feet 248 seconds after separation. It is then pulled away from the SRB by the drogue chute. The main chutes' suspension lines are pulled out from deployment bags that remain in the frustum.
At full extension of the lines, which are 204 feet long, the three main chutes are pulled from the deployment bags and inflate to their first reefed condition. The frustum and drogue parachute continue on a separate trajectory to splashdown. After specified time delays, the main chutes' reefing lines are cut and the chutes inflate to their second reefed and full open configurations.
The main chute cluster decelerates the SRB to terminal conditions. Each of the 136-foot-diameter, 20-degree conical ribbon parachutes can withstand a load of 180,000 pounds and weighs 2,180 pounds. The nozzle extension is severed by pyrotechnic charge either at apogee or 20 seconds after low baroswitch operation.
Water impact occurs 295 seconds after separation at a velocity of 81 feet per second. The water impact range is approximately 140 miles off the eastern coast of \JFlorida\j. Because the parachutes provide for a nozzlefirst impact, air is trapped in the empty (burned out) motor casing, causing the booster to float with the forward end approximately 30 feet out of the water.
The main chutes are released from the SRB at impact using the parachute release nut ordnance system. Residual loads in the main chutes deploy the parachute attach fittings with the redundant flotation tethered to each fitting. The drogue and frustum; each main chute, with its flotation; and the SRB are buoyant.
The SRB recovery aids are the radio beacon and flashing lights, which become operable at frustum separation. The radio transponder in each SRB has a range of 8.9 nautical miles (10.35 statute miles), and the flashing light has a nighttime range of 4.9 nautical miles (5.75 statute miles).
Various parameters of SRB operation are monitored and displayed on the orbiter flight deck control and display panel and are transmitted to ground telemetry.
#
"Shuttle External Tank",267,0,0,0
The external tank contains the liquid \Jhydrogen\j fuel and liquid oxygen oxidizer and supplies them under pressure to the three space shuttle main engines in the orbiter during lift-off and ascent. When the SSMEs are shut down, the ET is jettisoned, enters the Earth's atmosphere, breaks up, and impacts in a remote ocean area. It is not recovered.
The largest and heaviest (when loaded) element of the space shuttle, the ET has three major components: the forward liquid oxygen tank, an unpressurized intertank that contains most of the electrical components, and the aft liquid \Jhydrogen\j tank. The ET is 153.8 feet long and has a diameter of 27.6 feet.
Beginning with the STS-6 mission, a lightweight ET was introduced. Although future tanks may vary slightly, each will weigh approximately 66,000 pounds inert. The last heavyweight tank, flown on STS-7, weighed approximately 77,000 pounds inert. For each pound of weight reduced from the ET, the cargo-carrying capability of the space shuttle \Jspacecraft\j is increased almost one pound. The weight reduction was accomplished by eliminating portions of stringers (structural stiffeners running the length of the \Jhydrogen\j tank), using fewer stiffener rings and by modifying major frames in the \Jhydrogen\j tank.
Also, significant portions of the tank are milled differently to reduce thickness, and the weight of the ET's aft solid rocket booster attachments were reduced by using a stronger, yet lighter and less expensive \Jtitanium\j alloy. Earlier several hundred pounds were eliminated by deleting the anti-geyser line. The line paralleled the oxygen feed line and provided a circulation path for liquid oxygen to reduce accumulation of gaseous oxygen in the feed line while the oxygen tank was being filled before launch.
After propellant-loading data from ground tests and the first few space shuttle missions was assessed, the anti- \Jgeyser\j line was removed for STS-5 and subsequent missions. The total length and diameter of the ET remain unchanged.
The ET is attached to the orbiter at one forward attachment point and two aft points. In the aft attachment area, there are also umbilicals that carry fluids, gases, electrical signals and electrical power between the tank and the orbiter. Electrical signals and controls between the orbiter and the two solid rocket boosters also are routed through those umbilicals.
#
"Shuttle Liquid Oxygen Tank",268,0,0,0
The liquid oxygen tank is an aluminum monocoque structure composed of a fusion-welded assembly of preformed, chem-milled gores, panels, machined fittings and ring chords. It operates in a pressure range of 20 to 22 psig. The tank contains anti-slosh and anti-vortex provisions to minimize liquid residuals and damp fluid motion.
The tank feeds into a 17-inch- diameter feed line that conveys the liquid oxygen through the intertank, then outside the ET to the aft right-hand ET / orbiter disconnect umbilical. The 17-inch-diameter feed line permits liquid oxygen to flow at approximately 2,787 pounds per second with the SSMEs operating at 104 percent or permits a maximum flow of 17,592 gallons per minute.
The liquid oxygen tank's double-wedge nose cone reduces drag and heating, contains the vehicle's ascent air data system (for nine tanks only) and serves as a \Jlightning\j rod. The liquid oxygen tank's volume is 19,563 cubic feet. It is 331 inches in diameter, 592 inches long and weighs 12,000 pounds empty.
#
"Shuttle Intertank",269,0,0,0
The intertank is a steel / aluminum semimonocoque cylindrical structure with flanges on each end for joining the liquid oxygen and liquid \Jhydrogen\j tanks. The intertank houses ET instrumentation components and provides an umbilical plate that interfaces with the ground facility arm for purge gas supply, hazardous gas detection and \Jhydrogen\j gas boiloff during ground operations.
It consists of mechanically joined skin, stringers and machined panels of aluminum alloy. The intertank is vented during flight. The intertank contains the forward SRB-ET attach thrust beam and fittings that distribute the SRB loads to the liquid oxygen and liquid \Jhydrogen\j tanks. The intertank is 270 inches long, 331 inches in diameter and weighs 12,100 pounds.
#
"Shuttle Liquid Hydrogen Tank",270,0,0,0
The liquid \Jhydrogen\j tank is an aluminum semimonocoque structure of fusion-welded barrel sections, five major ring frames, and forward and aft ellipsoidal domes. Its operating pressure range is 32 to 34 psia. The tank contains an anti-vortex baffle and siphon outlet to transmit the liquid \Jhydrogen\j from the tank through a 17-inch line to the left aft umbilical.
The liquid \Jhydrogen\j feed line flow rate is 465 pounds per second with the SSMEs at 104 percent or a maximum flow of 47,365 gallons per minute. At the forward end of the liquid \Jhydrogen\j tank is the ET / orbiter forward attachment pod strut, and at its aft end are the two ET / orbiter aft attachment ball fittings as well as the aft SRB-ET stabilizing strut attachments. The liquid \Jhydrogen\j tank is 331 inches in diameter, 1,160 inches long, and has a volume of 53,518 cubic feet and a dry weight of 29,000 pounds.
#
"Shuttle ET Thermal Protection System",271,0,0,0
The ET thermal protection system consists of sprayed-on \Jfoam\j \Jinsulation\j and premolded ablator materials. The system also includes the use of phenolic thermal insulators to preclude air liquefaction. Thermal isolators are required for liquid \Jhydrogen\j tank attachments to preclude the liquefaction of air-exposed metallic attachments and to reduce heat flow into the liquid \Jhydrogen\j. The thermal protection system weighs 4,823 pounds.
#
"Shuttle ET Hardware",272,0,0,0
The external hardware, ET / orbiter attachment fittings, umbilical fittings, electrical and range safety system weigh 9,100 pounds.
Each propellant tank has a vent and relief valve at its forward end. This dual-function valve can be opened by ground support equipment for the vent function during prelaunch and can open during flight when the ullage (empty space) pressure of the liquid \Jhydrogen\j tank reaches 38 psig or the ullage pressure of the liquid oxygen tank reaches 25 psig.
The liquid oxygen tank contains a separate, pyrotechnically operated, propulsive tumble vent valve at its forward end. At separation, the liquid oxygen tumble vent valve is opened, providing impulse to assist in the separation maneuver and more positive control of the entry aerodynamics of the ET.
There are eight propellant-depletion sensors, four each for fuel and oxidizer. The fuel-depletion sensors are located in the bottom of the fuel tank. The oxidizer sensors are mounted in the orbiter liquid oxygen feed line manifold downstream of the feed line disconnect. During SSME thrusting, the orbiter general-purpose computers constantly compute the instantaneous mass of the vehicle due to the usage of the propellants. Normally, main engine cutoff is based on a predetermined velocity; however, if any two of the fuel or oxidizer sensors sense a dry condition, the engines will be shut down.
The locations of the liquid oxygen sensors allow the maximum amount of oxidizer to be consumed in the engines, while allowing sufficient time to shut down the engines before the oxidizer pumps cavitate (run dry). In addition, 1,100 pounds of liquid \Jhydrogen\j are loaded over and above that required by the 6-1 oxidizer / fuel engine mixture ratio. This assures that MECO from the depletion sensors is fuel-rich; oxidizer-rich engine shutdowns can cause burning and severe erosion of engine components.
Four pressure transducers located at the top of the liquid oxygen and liquid \Jhydrogen\j tanks monitor the ullage pressures.
Each of the two aft external tank umbilical plates mate with a corresponding plate on the orbiter. The plates help maintain alignment among the umbilicals. Physical strength at the umbilical plates is provided by bolting corresponding umbilical plates together. When the orbiter GPCs command external tank separation, the bolts are severed by pyrotechnic devices.
The ET has five propellant umbilical valves that interface with orbiter umbilicals: two for the liquid oxygen tank and three for the liquid \Jhydrogen\j tank. One of the liquid oxygen tank umbilical valves is for liquid oxygen, the other for gaseous oxygen. The liquid \Jhydrogen\j tank umbilical has two valves for liquid and one for gas. The intermediate-diameter liquid \Jhydrogen\j umbilical is a recirculation umbilical used only during the liquid \Jhydrogen\j chill-down sequence during prelaunch.
The ET also has two electrical umbilicals that carry electrical power from the orbiter to the tank and the two SRBs and provide information from the SRBs and ET to the orbiter.
A swing-arm-mounted cap to the fixed service structure covers the oxygen tank vent on top of the ET during the countdown and is retracted about two minutes before lift- off. The cap siphons off oxygen vapor that threatens to form large ice on the ET, thus protecting the orbiter's thermal protection system during launch.
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"Shuttle ET Range Safety System",273,0,0,0
A range safety system provides for dispersing tank propellants if necessary. It includes a battery power source, a receiver / decoder, antennas and ordnance.
Various parameters are monitored and displayed on the flight deck display and control panel and are transmitted to the ground.
The contractor for the external tank is Martin Marietta Aero space, New Orleans, La. The tank is manufactured at Michoud, La. Motorola, Inc., Scottsdale, Ariz., is the contractor for range safety receivers.
#
"Orbiter Structure",274,0,0,0
The orbiter structure is divided into nine major sections: the forward fuselage, which consists of upper and lower sections that fit clam-like around a pressurized crew compartment; wings; midfuselage; payload bay doors; aft fuselage; forward reaction control system; vertical tail; orbital maneuvering system/reaction control system pods; and body flap. The majority of the sections are constructed of conventional aluminum and protected by reusable surface \Jinsulation\j.
The forward fuselage structure is composed of 2024 aluminum alloy skin-stringer panels, frames and bulkheads.
The crew compartment is supported within the forward fuselage at four attachment points and is welded to create a pressure-tight vessel. The three-level compartment has a side hatch for normal passage and hatches in the airlock to permit extravehicular and intravehicular activities. The side hatch can be jettisoned.
The midfuselage is a 60-foot section of primary load-carrying structure between the forward and aft fuselages. It includes the wing carry-through structure and the payload bay doors. The skins consist of integral-machined aluminum panels and aluminum honeycomb sandwich panels.
The frames are constructed from a combination of aluminum panels with riveted or machined integral stiffeners and a truss structure center section. The upper half of the midfuselage consists of structural payload bay doors hinged along the side and split at the top centerline. The doors are \Jgraphite\j epoxy frames and honeycomb panel construction.
The aft fuselage includes a truss-type internal structure of diffusion-bonded elements that transfer the main engine thrust loads to the midfuselage and external tank. (In OV-105 , the truss-type internal structure is of a forging construction.) The aft fuselage's external surface is of standard construction except for the removable OMS/RCS pods, which are constructed of \Jgraphite\j epoxy skins and frames. An aluminum bulkhead shield with reusable \Jinsulation\j at the rear of the orbiter protects the rear portion of the aft fuselage.
The wing is constructed of a conventional aluminum alloy, using a corrugated spar web, truss-type ribs and riveted skin-stringer and honeycomb covers. The elevons are constructed of aluminum honeycomb and are split into two segments to minimize hinge binding and interaction with the wing.
The vertical tail, a conventional aluminum alloy structure, is a two-spar, multirib, integrally machined skin assembly. The tail is attached to the aft fuselage by bolted fittings at the two main spars. The rudder/speed brake assembly is divided into upper and lower sections, which are split longitudinally and actuated individually to serve as both rudder and speed brake.
These major structural assemblies are mated and held together by rivets and bolts. The midfuselage is joined to the forward and aft fuselage primarily by shear ties, with the midfuselage overlapping the bulkhead caps at stations Xo 582 and Xo 1307. The wing is attached to the midfuselage and aft fuselage primarily by shear ties, except in the area of the wing carry-through, where the upper panels are attached with tension bolts. The vertical tail is attached to the aft fuselage with bolts that work in both shear and tension. The body flap, which has aluminum honeycomb covers, is attached to the lower aft fuselage by four rotary actuators.
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"Orbiter Forward Fuselage",275,0,0,0
The forward fuselage consists of the upper and lower fuselages. It houses the crew compartment and supports the forward reaction control system module, nose cap, nose gear wheel well, nose gear and nose gear doors.
The forward fuselage is constructed of conventional 2024 aluminum alloy skin-stringer panels, frames and bulkheads. The panels are single curvature and stretch-formed skins with riveted stringers spaced 3 to 5 inches apart. The frames are riveted to the skin-stringer panels. The major frames are spaced 30 to 36 inches apart. The Y o 378 upper forward bulkhead is constructed of flat aluminum and formed sections riveted and bolted together; the lower is a machined section. The bulkhead provides the interface fitting for the nose section.
The nose section contains large machined beams and struts. The structure for the nose landing gear wheel well consists of two support beams, two upper closeout webs, drag-link support struts, nose landing gear strut and actuator attachment fittings, and the nose landing gear door fittings. The left and right landing gear doors are attached by hinge fittings in the nose section.
The doors are constructed of aluminum alloy honeycomb, and although the doors are the same length, the left door is wider than the right. Each door has an up-latch fitting at the forward and aft ends to lock the door closed when the gear is retracted, and each has a pressure seal in addition to a thermal barrier.
Lead ballast in the nose wheel well and on the X o 378 bulkhead provides weight and center-of-gravity control. The nose wheel well will accommodate 1,350 pounds of ballast, and the X o 378 bulkhead will accommodate a maximum of 2,660 pounds.
The forward fuselage carries the basic body-bending loads (a tendency to change the radius of a curvature of the body) and reacts nose landing gear loads.
The forward fuselage is covered with reusable \Jinsulation\j, except for the six windshields, two overhead windows and side hatch window areas around the forward RCS engines. The nose cap is also a reusable thermal protection system. It is constructed of reinforced carbon-carbon and has thermal barriers at the nose cap-structure interface.
The forward fuselage skin has structural provisions for installing antennas, deployable air data probes and the door eyelet openings for the two star trackers. Two openings are required in the upper forward fuselage for star tracker viewing. Each opening has a door for environmental control.
The forward orbiter/external tank attach fitting is at the Xo 378 bulkhead and the skin panel structure aft of the nose gear wheel well. Purge and vent control is provided by flexible boots between the forward fuselage and crew compartment around the windshield windows, overhead observation window, crew hatch window and star tracker openings. The forward fuselage is isolated from the payload bay by a flexible membrane between the forward fuselage and crew compartment at Xo 582.
Six forward outer pane windshields are installed on the forward fuselage. They are described in the section on windows. The window structural frames in the forward fuselage are five-axis machined parts.
The forward RCS module is constructed of conventional 2024 aluminum alloy skin-stringer panels and frames. The panels are composed of single-curvature and stretch-formed skins with riveted stringers. The frames are riveted to the skin-stringer panels. The forward RCS module is secured to the forward fuselage nose section and forward bulkhead of the forward fuselage with 16 fasteners, which permit the installation and removal of the module.
The components of the forward RCS are mounted and attached to the module, which will have a reusable thermal protection cover, in addition to thermal barriers installed around it and the RCS engine interfaces and the interface-attachment area to the forward fuselage.
The forward fuselage and forward RCS module are built by Rockwell's Space Transportation Systems Division, Downey, Calif.
#
"Orbiter Crew Compartment",276,0,0,0
The three-level crew compartment is constructed of 2219 aluminum alloy plate with integral stiffening stringers and internal framing welded together to create a pressure-tight vessel. The compartment has a side hatch for normal ingress and egress, a hatch into the airlock from the middeck, and a hatch from the airlock through the aft bulkhead into the payload bay for extravehicular activity and payload bay access.
Redundant pressure window panes are provided in the six forward windshields, the two overhead viewing windows, the two aft viewing windows and the side hatch windows; they are described in the window section. Approximately 300 penetrations in the pressure shell are sealed with plates and fittings.
A large removable panel in the aft bulkhead provides access to the interior of the crew compartment during initial fabrication and assembly and provides for airlock installation and removal. The compartment supports the environmental control and life support system; avionics; guidance, navigation and control equipment; inertial measurement units; displays and controls; star trackers; and crew accommodations for sleeping, waste management, seats and an optional galley.
The crew compartment is supported within the forward fuselage at only four attach points to minimize the thermal \Jconductivity\j between them. The two major attach points are located at the aft end of the crew compartment at the flight deck floor level. The vertical load reaction link is on the centerline of the forward bulkhead. The lateral load reaction is on the lower segment of the aft bulkhead.
The compartment is configured to accommodate a crew of four on the flight deck and three in the middeck. In OV-102, four can be accommodated in the middeck. The crew cabin arrangement consists of a flight deck, middeck and lower level equipment bay.
The crew compartment is pressurized to 14.7 psia, plus or minus 0.2 psia, and is maintained at an 80-percent \Jnitrogen\j and 20-percent oxygen composition by the ECLSS, which provides a shirt-sleeve environment for the flight crew. The crew compartment is designed for 16 psia.
The crew compartment's volume with the airlock in the middeck is 2,325 cubic feet. If the airlock is in the payload bay, the crew compartment's cabin volume is 2,625 cubic feet.
The flight deck is the uppermost compartment of the cabin. The commander's and pilot's work stations are positioned side by side in the forward portion of the flight deck. These stations have controls and displays for maintaining autonomous control of the vehicle throughout all mission phases. Directly behind and to the sides of the commander and pilot centerline are the mission specialist seats.
The commander's and pilot's seats have two shoulder harnesses and a lap belt for restraints. The shoulder harnesses have an \Jinertia\j reel lock/unlock feature. The unlocked position allows the shoulder harness to move. The commander and pilot can move their seats along the orbiter's Z (vertical) and X (longitudinal) axes so they can reach and see controls better during the ascent and entry phases of flight.
Seat movement for each axis is provided by a single ac motor. The total travel distance for the Z and X axes is 10 and 5 inches, respectively. Seat adjustment controls are located on the left side of the seat pan and consist of a three-position toggle switch for power bus selection and one spring-loaded, three-position toggle switch each to control horizontal and vertical seat movement.
Each mission and payload specialist's seat has two shoulder harnesses and a lap belt for restraints. The specialists' seats have controls to manually lock and unlock the tilt of the seat back. Each seat has removable seat cushions and mounting provisions for oxygen and communications connections to the CAPS. The specialists' seats are removed and stowed in the middeck on orbit. No tools are required since the legs of each seat have quick-disconnect fittings. Each seat is 25.5 inches long, 15.5 inches wide and 11 inches high when folded for stowage.
The aft flight deck has two overhead and aft viewing windows for viewing orbital operations. The aft flight deck station also contains displays and controls for executing attitude or translational maneuvers for rendezvous, stationkeeping, docking, payload deployment and retrieval, payload monitoring, remote manipulator system controls and displays, payload bay door operations and closed-circuit \Jtelevision\j operations.
The forward flight deck, which includes the center console and seats, is approximately 24 square feet. However, the side console controls and displays add approximately 3.5 square feet more. If the center console is subtracted from the 24 square feet, this would amount to approximately 5.2 square feet.
The aft flight deck is approximately 40 square feet.
Directly beneath the flight deck is the middeck. Access to the middeck is through two interdeck openings, which measure 26 by 28 inches. Normally, the right interdeck opening is closed and the left is open. A ladder attached to the left interdeck access allows easy passage in 1-g conditions. The middeck provides crew accommodations and contains three avionics equipment bays.
The two forward avionics bays utilize the complete width of the cabin and extend into the middeck 39 inches from the forward bulkhead. The aft bay extends into the middeck 39 inches from the aft bulkhead on the right side of the airlock. Just forward of the waste management system is the side hatch. The completely stripped middeck is approximately 160 square feet; the gross mobility area is approximately 100 square feet.
The side hatch in the middeck is used for normal crew entrance/exit and may be operated from within the crew cabin middeck or externally. It can be jettisoned for emergencies, as discussed in the escape system section. It is attached to the crew cabin tunnel by hinges, a \Jtorque\j tube and support fittings.
The hatch opens outwardly 90 degrees down with the orbiter horizontal or 90 degrees sideways with the orbiter vertical. It is 40 inches in diameter and has a 10-inch clear-view window in the center of the hatch. The window consists of three panes of glass. The side hatch has a pressure seal that is compressed by the side hatch latch mechanisms when the hatch is locked closed.
A thermal barrier of Inconel wire mesh spring with a ceramic fiber braided sleeve is installed between the reusable surface \Jinsulation\j tiles on the forward fuselage and the side hatch. The total weight of the side hatch is 294 pounds.
Depending on the mission requirements, bunk sleep stations and a galley can be installed in the middeck. In addition, three or four seats of the same type as the mission specialists' seats on the flight deck can be installed in the middeck. Three seats over the normal three could be installed in the middeck for rescue missions if the bunk sleep stations were removed.
The waste management system, located in the middeck, can also accommodate payloads in the pressurized crew compartment environment.
The middeck also provides a stowage volume of 140 cubic feet. Accommodations are included for dining, sleeping, maintenance, exercising and data management. On the orbiter centerline, just aft of the forward avionics equipment bay, an opening in the ceiling provides access to the inertial measurement units.
The middeck floor contains removable panels that provide access to the ECLSS equipment. The middeck equipment bay below the middeck floor houses the major components of the waste management and air revitalization systems, such as pumps, fans, \Jlithium\j \Jhydroxide\j, absorbers, heat exchangers and ducting. This compartment has space for stowing \Jlithium\j \Jhydroxide\j canisters and five separate spaces for crew equipment stowage with a volume of 29.92 cubic feet.
Modular stowage lockers are used to store the flight crew's personal gear, mission-necessary equipment, personal hygiene equipment and experiments. The modular lockers are made of sandwich panels of Kevlar/epoxy and a non-metallic core. This reduced the lockers' weight by 83 percent compared to all-aluminum lockers, a reduction of approximately 150 pounds. There are 42 identical boxes, which are 11 by 18 by 21 inches.
An airlock, located in the middeck, is composed of machined aluminum sections welded together to form a cylinder with hatch mounting flanges. The upper cylindrical section and bulkheads are constructed of aluminum honeycomb. Two semicylindrical aluminum sections are welded to the airlock's primary structure to house the ECLSS and electrical support equipment. Each semicylindrical section has three feedthrough plates for plumbing and cable routings from the orbiter to the airlock.
Normally, two extravehicular mobility units are stowed in the airlock. The EMU is an integrated space suit assembly and life support system that enables flight crew members to leave the pressurized orbiter crew cabin and work outside the cabin in space.
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"Orbiter Airlock",277,0,0,0
The airlock is normally located inside the middeck of the \Jspacecraft\j's pressurized crew cabin. It has an inside diameter of 63 inches, is 83 inches long and has two 40-inch- diameter D-shaped openings that are 36 inches across. It also has two pressure-sealing hatches and a complement of airlock support systems. The airlock's volume is 150 cubic feet.
The airlock is sized to accommodate two fully suited flight crew members simultaneously. Support functions include airlock depressurization and repressurization, extravehicular activity equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. The EVA gear, checkout panel and recharge stations are located on the internal walls of the airlock.
The airlock hatches are mounted on the airlock. The inner hatch is mounted on the exterior of the airlock (orbiter crew cabin middeck side) and opens into the middeck. The inner hatch isolates the airlock from the orbiter crew cabin. The outer hatch is mounted inside the airlock and opens into the airlock. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.
Airlock repressurization is controllable from the orbiter crew cabin middeck and from inside the airlock. It is performed by equalizing the airlock's and cabin's pressure with equalization valves mounted on the inner hatch. The airlock is depressurized from inside the airlock by venting the airlock's pressure overboard. The two D-shaped airlock hatches open toward the primary pressure source, the orbiter crew cabin, to achieve pressure-assist sealing when closed.
Each hatch has six interconnected latches and a gearbox/actuator, a window, a hinge mechanism and hold-open device, a differential pressure gauge on each side and two equalization valves.
The 4-inch diameter window in each airlock hatch is used for crew observation from the cabin/airlock and the airlock/payload bay. The dual window panes are made of polycarbonate plastic and mounted directly to the hatch by means of bolts fastened through the panes. Each hatch window has dual pressure seals, with seal grooves located in the hatch.
Each airlock hatch has dual pressure seals to maintain pressure integrity. One seal is mounted on the airlock hatch and the other on the airlock structure. A leak check quick disconnect is installed between the hatch and the airlock pressure seals to verify hatch pressure integrity before flight.
The gearbox with latch mechanisms on each hatch allows the flight crew to open and close the hatch during transfers and EVA operations. The gearbox and the latches are mounted on the low-pressure side of each hatch; with a gearbox handle installed on both sides to permit operation from either side of the hatch.
Three of the six latches on each hatch are double-acting and have cam surfaces that force the sealing surfaces apart when the latches are opened, thereby acting as crew assist devices. The latches are interconnected with push-pull rods and an idler bell crank that is installed between the rods for pivoting the rods. Self-aligning dual rotating bearings are used on the rods for attachment to the bellcranks and the latches. The gearbox and hatch open support struts are also connected to the latching system by the same rod/bellcrank and bearing system. To latch or unlatch the hatch, the gearbox handle must be rotated 440 degrees.
The hatch actuator/gearbox is used to provide the mechanical advantage to open and close the latches. The hatch actuator lock lever requires a force of 8 to 10 pounds through an angle of 180 degrees to unlatch the actuator. A minimum rotation of 440 degrees with a maximum force of 30 pounds applied to the actuator handle is required to operate the latches to their fully unlatched positions.
The hinge mechanism for each hatch permits a minimum opening sweep into the airlock or the crew cabin middeck. The inner hatch (airlock to crew cabin) is pulled or pushed forward to the crew cabin approximately 6 inches. The hatch pivots up and to the right side. Positive locks are provided to hold the hatch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is also provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The outer hatch (airlock to payload bay) opens and closes to the contour of the airlock wall. The hatch is hinged to be pulled first into the airlock and then forward at the bottom and rotated down until it rests with the low-pressure (outer) side facing the airlock ceiling (middeck floor). The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke.
The hatch has a hold-open hook that snaps into place over a flange when the hatch is fully open. The hook is released by depressing the spring-loaded hook handle and pushing the hatch toward the closed position. To support and protect the hatch against the airlock ceiling, the hatch incorporates two deployable struts. The struts are connected to the hatch linkage mechanism and are deployed when the hatch linkage is rotated open. When the hatch latches are rotated closed, the struts are retracted against the hatch.
The airlock hatches can be removed in flight from the hinge mechanism using pip pins, if required.
The airlock air circulation system provides conditioned air to the airlock during non-EVA periods. The airlock revitalization system duct is attached to the outside airlock wall at launch. Upon airlock hatch opening in flight, the duct is rotated by the flight crew through the cabin/airlock hatch, installed in the airlock and held in place by a strap holder.
The duct has a removable air diffuser cap, installed on the end of the flexible duct, which can adjust the air flow from 216 pounds per hour. The duct must be rotated out of the airlock before the cabin/airlock hatch is closed for airlock depressurization. During the EVA preparation period, the duct is rotated out of the airlock and can be used for supplemental air circulation in the middeck.
To assist the crew member before and after EVA operations, the airlock incorporates handrails and foot restraints. Handrails are located alongside the avionics and ECLSS panels. Aluminum alloy handholds mounted on each side of the hatches have oval configurations 0.75 by 1.32 inches and are painted yellow. They are bonded to the airlock walls with an epoxyphenolic adhesive.
Each handrail has a clearance of 2.25 inches between the airlock wall and the handrail to allow the astronauts to grip it while wearing a pressurized glove. Foot restraints are installed on the airlock floor nearer the payload bay side. The ceiling handhold is installed nearer the cabin side of the airlock. The foot restraints can be rotated 360 degrees by releasing a spring-loaded latch and lock in every 90 degrees.
A rotation release knob on the foot restraint is designed for shirt-sleeve operation and, therefore, must be positioned before the suit is donned. The foot restraint is bolted to the floor and cannot be removed in flight. It is sized for the EMU boot. The crew member first inserts his foot under the toe bar and then rotates his heel from inboard to outboard until the heel of the boot is captured.
There are four floodlights in the airlock.
If the airlock is relocated to the payload bay from the middeck, it will function in the same manner as in the middeck. \JInsulation\j is installed on the airlock's exterior for protection from the extreme temperatures of space.
For Spacelab pressurized module missions, the airlock remains in the crew compartment middeck, and a tunnel adapter that mates with the airlock and the Spacelab tunnel is installed in the payload bay.
The airlock tunnel adapter, hatches, tunnel extension and tunnel permit the flight crew members to transfer from the \Jspacecraft\j's pressurized middeck crew compartment to Spacelab's pressurized shirt-sleeve environment.
In addition, the airlock, tunnel adapter and hatches permit the EVA flight crew members to transfer from the airlock/tunnel adapter in the space suit assembly into the payload bay without depressurizing the crew cabin and Spacelab.
The Spacelab tunnel and Spacelab are accessed via the tunnel adapter, which is located in the payload bay and is attached to the airlock at orbiter station Xo 576 and the tunnel extension at X o 660. The tunnel adapter has an inside diameter of 63 inches at its widest section and tapers in the cone area at each end to two 40-inch- diameter D-shaped openings 36 inches across. A 40-inch- diameter D-shaped opening 36 inches across is located at the top of the tunnel adapter.
Two pressure-sealing hatches are located in the tunnel adapter, one in the upper area of the tunnel adapter and one in the aft end of the tunnel adapter. The tunnel adapter is a welded structure constructed of 2219 aluminum with 2.4- by 2.4-inch exposed structural ribs on the exterior surface and external waffle skin stiffening.
The hatch located on the middeck side of the airlock is mounted on the exterior of the airlock and opens into the middeck. The hatch isolates the airlock from the crew cabin. The hatch located in the tunnel adapter's aft end isolates the tunnel adapter/airlock from the tunnel extension, tunnel and Spacelab. This hatch opens into the tunnel adapter.
The hatch located in the tunnel adapter at the upper D-shaped opening isolates the airlock/tunnel adapter from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock/tunnel adapter to the payload bay when open. This hatch opens into the tunnel adapter.
The hinge mechanism for each hatch permits a minimum opening sweep into the tunnel adapter or the \Jspacecraft\j crew cabin middeck. The airlock crew cabin hatch in the middeck is pulled/pushed forward to the middeck approximately 6 inches.
The hatch pivots up and right. Positive locks are provided to hold the latch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The aft hatch is hinged to be pulled first into the tunnel adapter and then forward at the bottom. The top of the hatch is rotated towards the tunnel and downward until the hatch rests with the Spacelab side facing the tunnel adapter floor. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The upper (EVA) hatch in the tunnel adapter opens and closes to the left wall of the tunnel adapter. The hatch is hinged to be pulled first into the tunnel adapter and then forward at the hinge area and rotated down until it rests against the port wall of the tunnel adapter. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The hatches can be removed in flight from the hinge mechanisms via pip pins, if required.
The crew compartment, bunk sleep stations (if installed), airlock and modular stowage lockers are built by Rockwell's Space Transportation Systems Division, Downey, Calif. The original crew seat contractor was AMI of \JColorado\j Springs, Colo., but later Rockwell's Space Transportation Systems Division. The Spacelab pressurized module tunnel adapter and tunnel contractor is McDonnell Douglas Astronautics, Huntington Beach, Calif.
#
"Orbiter Forward Fuselage and Cabin Windows",278,0,0,0
The orbiter windows provide visibility for entry, landing and on-orbit operations. For atmospheric flight, the flight crew needs forward, left and right viewing areas. On-orbit mission phases require visibility for rendezvous, docking and payload-handling operations.
The six windows located at the forward flight deck commander and pilot stations provide forward, left and right viewing. The two overhead windows and two payload-viewing windows at the aft station location on the flight deck provide rendezvous, docking and payload viewing. There is also a window in the middeck side hatch.
The six planeform-shaped forward windows are the thickest pieces of glass ever produced in the optical quality for see-through viewing. Each consists of three individual panes. The innermost pane is constructed of tempered aluminosilicate glass to withstand the crew compartment pressure. It is 0.625 of an inch thick. Aluminosilicate glass is a low-expansion glass that can be tempered to provide maximum mechanical strength. The exterior of this pane, called a pressure pane, is coated with a red reflector coating to reflect the infrared (heat portion) rays while transmitting the visible spectrum.
The center pane is constructed of low-expansion, fused \Jsilica\j glass because of its high optical quality and excellent thermal shock resistance. This pane is 1.3 inches thick.
The inner and outer panes are coated with a high-efficiency, anti-reflection coating to improve visible light transmission. These windows withstand a proof pressure of 8,600 psi at 240 F and 0.017 relative \Jhumidity\j.
The outer pane is made of the same material as the center pane and is 0.625 of an inch thick. The exterior is uncoated, but the interior is coated with high-efficiency, anti-reflection coating. The outer surface withstands approximately 800 F.
Each of the forward six windows' outer panes measures 42 inches diagonally, and the center and inner panes each measure 35 inches diagonally. The outer panes of the forward six windows are mounted and attached to the forward fuselage. The center and inner panes are mounted and attached to the crew compartment. Redundant seals are employed on each window. No sealing/bonding compounds are used.
The two overhead windows at the flight deck aft station are identical in construction to the six forward windows except for thickness. The inner and center panes are 0.45 of an inch thick, and the outer pane is 0.68 of an inch thick. The outer pane is attached to the forward fuselage, and the center and inner panes are attached to the crew compartment.
The two overhead windows' clear view area is 20 by 20 inches. The left-hand overhead window provides the crew members with a secondary emergency egress. The inner and center panes open into the crew cabin, and the outer pane is jettisoned up and over the top of the orbiter. This provides a secondary emergency exit area of 20 by 20 inches.
On the aft flight deck, each of the two windows for viewing the payload bay consists of only two panes of glass, which are identical to the forward windows' inner and center panes. The outer thermal panes are not installed. Each pane is 0.3 of an inch thick. The windows are 14.5 by 11 inches. Both panes are attached to the crew compartment.
The side hatch viewing window consists of three panes of glass identical to the six forward windows. The inner pane is 11.4 inches in diameter and 0.25 of an inch thick. The center pane is 11.4 inches in diameter and 0.5 of an inch thick. The outer pane is 15 inches in diameter and 0.3 of an inch thick.
During orbital operations, the large window areas of transparency expose the flight crew to sun glare; therefore, window shades and filters are provided to preclude or minimize exposure. Shades are provided for all windows, and filters are supplied for the aft and overhead viewing windows. The window shades and filters are stored in the middeck of the orbiter crew compartment. Attachment mechanisms and devices are provided for their installation at each window on the flight deck.
The forward station window shades (W-1 through W-6) are fabricated from Kevlar/epoxy glass fabric with silver and Inconel-coated Teflon tape on the outside surface and paint on the inside surface. When the shade is installed next to the inner window pane, a silicone rubber seal around the periphery deforms to prevent light leakage. The shade is held in place by the shade installation guide, the hinge plate and the Velcro keeper.
The overhead window shades (W-7 and W-8) are nearly the same as the forward shades; but the rubber seal is deleted, and the shade is sealed and held in place by a separate seal around the window opening, a hinge plate and secondary frame, and Velcro retainer. The overhead window filters are fabricated from Lexan and are used interchangeably with the shades.
The aft window shades (W-9 and W-10) are the same as the overhead window shades except that a 0.63-inch-wide strip of Nomex Velcro has been added around the perimeter of the shade. The shade is attached to the window by pressing the Velcro strip to the pile strip around the window opening. The aft window filters are the same as the overhead window filters except for the addition of the Velcro hook strip. The filters and shades are used interchangeably.
The side hatch window cover is permanently attached to the window frame and is hinged to allow opening and closing.
The contractor for the windows is Corning Glass Co., Corning, N.Y.
#
"Orbiter Wing",279,0,0,0
The wing is an aerodynamic lifting surface that provides conventional lift and control for the orbiter. The left and right wings consist of the wing glove; the intermediate section, which includes the main landing gear well; the \Jtorque\j box; the forward spar for mounting the reusable reinforced carbon-carbon leading edge structure thermal protection system; the wing/elevon interface; the elevon seal panels; and the elevons.
The wing is constructed of conventional aluminum alloy with a multirib and spar arrangement with skin-stringer-stiffened covers or honeycomb skin covers. Each wing is approximately 60 feet long at the fuselage intersection and has a maximum thickness of 5 feet.
The forward wing box is an extension of the basic wing that aerodynamically blends the wing leading edge into the midfuselage wing glove. The forward wing box is a conventional design of aluminum ribs, aluminum tubes and tubular struts. The upper and lower wing skin panels are stiffened aluminum. The leading edge spar is constructed of corrugated aluminum.
The intermediate wing section consists of the conventional aluminum multiribs and aluminum tubes. The upper and lower skin covers are constructed of aluminum honeycomb. A portion of the lower wing surface skin panel includes the main landing gear door.
The intermediate section houses the main landing gear compartment and reacts a portion of the main landing gear loads. A structural rib supports the outboard main landing gear door hinges and the main landing gear trunnion and drag link. The support for the inboard main landing gear trunnion and drag link attachment is provided by the midfuselage. The main landing gear door is conventional aluminum honeycomb.
The four major spars are constructed of corrugated aluminum to minimize thermal loads. The forward spar provides the attachment for the thermal protection system reusable reinforced carbon-carbon leading edge structure. The rear spar provides the attachment interfaces for the elevons, hinged upper seal panels, and associated hydraulic and electrical system components. The upper and lower wing skin panels are stiffened aluminum.
The elevons provide orbiter flight control during atmospheric flight. The two-piece elevons are conventional aluminum multirib and beam construction with aluminum honeycomb skins for compatibility with the acoustic environment and thermal interaction. The elevons are divided into two segments for each wing, and each segment is supported by three hinges. The elevons are attached to the flight control system hydraulic actuators at points along their forward extremities, and all hinge moments are reacted at these points. Each elevon travels 40 degrees up and 25 degrees down.
The transition area on the upper surface between the \Jtorque\j box and the movable elevon consists of a series of hinged panels that provide a closeout of the wing-to-elevon cavity. These panels are of Inconel honeycomb sandwich construction outboard of wing station Y w 312.5 and of \Jtitanium\j honeycomb sandwich construction inboard of wing station Y w 312.5.
The upper leading edge of each elevon incorporates \Jtitanium\j rub strips. The rub strips are of \Jtitanium\j honeycomb construction and are not covered with the thermal protection system reusable surface \Jinsulation\j. They provide the sealing surface area for the elevon seal panels.
The exposed areas of the wings, main landing gear doors and elevons are covered with reusable surface \Jinsulation\j thermal protection system materials except for the elevon seal panels.
Thermal seals are provided on the elevon lower cove area along with thermal spring seals on the upper rub panels. Pressure seals and thermal barriers are provided on the main landing gear doors.
The wing is attached to the fuselage with a tension bolt splice along the upper surface. A shear splice along the lower surface in the area of the fuselage carry-through completes attachment interface.
Prior to the manufacturing of the wings for Discovery (OV-103) and Atlantis (OV-104), a weight reduction program resulted in a redesign of certain areas of the wing structure. An assessment of wing air loads was made from actual flight data that indicated greater loads on the wing structure. As a result, to maintain positive margins of safety during ascent, structural modifications were incorporated into certain areas of the wings. The modifications consisted of the addition of doublers and stiffeners.
The wing, elevon and main landing gear door contractor is Grumman Corp., Bethpage, N.Y.
#
"Orbiter Midfuselage",280,0,0,0
The midfuselage structure interfaces with the forward fuselage, aft fuselage and wings. It supports the payload bay doors, hinges, tie-down fittings, forward wing glove, and various orbiter system components and forms the payload bay area.
The forward and aft ends of the midfuselage are open, with reinforced skin and longerons interfacing with the bulkheads of the forward and aft fuselages. The midfuselage is primarily an aluminum structure 60 feet long, 17 feet wide and 13 feet high. It weighs approximately 13,502 pounds.
The midfuselage skins are integrally machined by numerical control. The panels above the wing glove and the wings for the forward eight bays have longitudinal T-stringers. The five aft bays have aluminum honeycomb panels. The side skins in the shadow of the wing are also numerically control machined but have vertical stiffeners.
Twelve main-frame assemblies stabilize the midfuselage structure. The assemblies consist of vertical side elements and horizontal elements. The side elements are machined; whereas the horizontal elements are boron/aluminum tubes with bonded \Jtitanium\j end fittings, which reduced the weight by 49 percent (approximately 305 pounds).
In the upper portion of the midfuselage are the sill and door longerons. The machined sill longerons not only make up the primary body-bending elements, but also take the longitudinal loads from payloads in the payload bay. The payload bay door longerons and associated structure are attached to the 13 payload bay door hinges. These hinges provide the vertical reaction from the payload bay doors.
Five of the hinges react the payload bay door shears. The sill longeron also provides the base support for the payload bay manipulator arm (if installed) and its stowage provisions, the Ku-band rendezvous antenna, the antenna base support and its stowage provisions, and the payload bay door actuation system.
The side wall forward of the wing carry-through structure provides the inboard support for the main landing gear. The total lateral landing gear loads are reacted by the midfuselage structure.
The midfuselage also supports the two electrical wire trays that contain the wiring between the crew compartment and aft fuselage.
Plumbing and wiring in the lower portion of the midfuselage are supported by fiberglass milk stools.
The remainder of the exposed areas of the midfuselage is covered with the reusable surface \Jinsulation\j thermal protection system.
Because of additional detailed analysis of actual flight data concerning descent stress thermal \Jgradient\j loads, torsional straps were added to the lower midfuselage stringers in bays 1 through 11. The torsional straps tie all stringers together similarly to a box section, which eliminates rotational (torsional) capabilities to provide positive margins of safety.
Also, because of additional detailed analysis of actual flight data during descent, room-temperature vulcanizing silicone rubber material was bonded to the lower midfuselage from bay 4 through 12 to act as a heat sink and distribute temperatures evenly across the bottom of the midfuselage, which will reduce thermal gradients and ensure positive margins of safety.
The contractor for the midfuselage is General Dynamics Corp., Convair Aerospace Division, San Diego, Calif.
#
"Orbiter Payload Bay Doors",281,0,0,0
The payload bay doors are opened shortly after orbit is achieved to allow exposure of the environmental control and life support system radiators for heat rejection of the orbiter's systems. The payload bay doors consist of port and starboard doors hinged at each side of the midfuselage and latched mechanically at the forward and aft fuselage and at the split-top centerline.
Thermal seals on the doors provide a relatively air-tight payload compartment when the doors are closed and latched. During prelaunch and postlanding, the purge, vent and drain system permits purging of undesirable gases and maintains a positive delta pressure for venting of payloads within the payload area when the doors are closed.
The port and starboard doors are 60 feet long with a combined area of approximately 1,600 square feet. Each door is made up of five segments that are interconnected by circumferential expansion joints. Each door hinges on 13 Inconel 718 external hinges (five shear and eight idlers).
The lower half of each hinge attaches to the midfuselage sill longeron. The hinges rotate on bearings with dual rotational surfaces. There are five shear hinges and eight floating hinges. The floating hinges allow fore and aft movement of the door panels for thermal expansion.
Each door actuation system provides the mechanism to drive each door side to the open or closed position. Each mechanism consists of an electromechanical power drive unit and six rotary gear actuators, which are connected by \Jtorque\j tubes to each other and to the power drive unit. Linkages transmit \Jtorque\j from the rotary actuators to the doors.
The forward 30-foot sections of both doors incorporate radiators that can be deployed; they are hinged and latched to the door inner surface in order to reject the excess heat of the Freon-21 coolant loops from both sides of the radiator panels when the doors are open. An electromechanical actuation system on the door unlatches and deploys the radiators when open and latches and stows the radiators when closed.
The radiators may be left in the stowed position for a given flight and will only radiate the excess heat from the one side. Fixed radiator panels are installed on the forward end of the aft payload bay doors and radiate from one side only. Kitted fixed radiator panels may be installed on the aft end of the aft payload bay doors when required by a specific mission; they also will radiate from only one side.
During payload bay door closure, the aft flight deck payload bay door crewman optical alignment sight is used to check door alignment.
When the payload bay doors are closed, they are fixed at the aft fuselage bulkhead and allowed to move longitudinally at the forward fuselage. The doors also accommodate vehicle torsional loads (a force that causes a body, such as a shaft, to twist about its longitudinal axis), aerodynamic pressure loads and payload bay vent lag pressures. The payload bay is not a pressurized area.
Thermal and pressure seals are used to close the gaps at the forward and aft fuselage interface, door centerline and circumferential expansion joints.
The doors are 60 feet long. Each consists of five segments interconnected by expansion joints. The chord of each half of these curved doors is approximately 10 feet, and the doors are 15 feet in diameter.
The doors are constructed of \Jgraphite\j epoxy composite material, which reduces the weight by 23 percent over that of aluminum honeycomb sandwich. This is a reduction of approximately 900 pounds, which brings the weight of the doors down to approximately 3,264 pounds. The payload bay doors are the largest aerospace structure to be constructed from composite material.
The composite doors will withstand 163-decibel acoustic noise and a temperature range of minus 170 to plus 135 F.
The doors are made up of subassemblies consisting of \Jgraphite\j epoxy honeycomb sandwich panels, solid \Jgraphite\j epoxy laminate frames, expansion joint frames, \Jtorque\j box, seal depressor, centerline beam intercostals, gussets, end fittings and clips. There are also aluminum 2024 shear pins, \Jtitanium\j fittings, and Inconel 718 floating and shear hinges. The assembly is joined by mechanical fasteners. \JLightning\j strike protection is provided by aluminum mesh wire bonded to the outer skin.
Extravehicular activity handholds are attached in the \Jtorque\j box areas.
The payload bay doors are covered with reusable surface \Jinsulation\j.
The left door with attached systems weighs approximately 2,375 pounds and the right weighs about 2,535 pounds. The right door contains the centerline latch active mechanisms, which accounts for the weight difference. These weights do not include the radiator panel system, which adds 833 pounds per door.
When closed, the doors are latched to the forward and aft bulkheads and along the upper centerline of the doors. The latching system consists of 16 bulkhead latches (eight aft and eight forward) and 16 payload bay door centerline latches. The forward and aft bulkhead latches are in groups of four ganged latch hooks. The centerline latches are also in groups of four ganged latches. Each centerline latch gang incorporates four latches, bellcranks, push rods, levers, rollers and an electromechanical actuator.
The forward and aft bulkhead latches are arranged in groups of four ganged latches. Each group is opened or closed by an electromechanical actuator with two redundant, three-phase ac reversible motors that receive ac power from mid motor controller assemblies when commanded in the automatic predetermined sequence or by manual keyboard entries. In the automatic mode, the forward and aft bulkhead latches operate simultaneously.
The payload bay door centerline latch groups are controlled automatically in a predetermined sequence or manually by individual latch groups through keyboard entries in a manner similar to the bulkhead latch groups. The 16 centerline latches are arranged into groups of four, similar to the bulkhead latches.
Each centerline latch group consists of two ac reversible electric motors that drive a rotary shaft and bellcrank and four hooks to engage a corresponding passive roller to latch the door closed or disengage the passive roller to unlatch the door. All 16 centerline hook assemblies contain alignment rollers to eliminate payload bay door overlap due to thermal distortion. Passive shear fittings in each centerline latch group align door closure and cause the fore and aft shear loads to react once the doors are closed.
The centerline latch group ac reversible motors are automatically turned off by limit switches when the latches are opened or closed. Each motor has a brake that operates similarly to the brakes in the bulkhead motors. When both motors are operating, the nominal operating time is 20 seconds.
If only one motor is operating, the time is 40 seconds. Each mid motor controller assembly has its own timer set to twice the normal operating time to allow single-motor operation of the centerline latch group without causing a sequence fail signal PLB doors CRT message and SM alert.
The power drive unit drives a 55-foot-long \Jtorque\j shaft. The shaft turns the rotary actuators, which causes the push rod, bell crank and link to push the doors open. The same arrangement pulls the doors closed.
The payload bay door opening and closing sequence is controlled automatically through in a predetermined sequence or manually through keyboard entries. The starboard doors must be opened first and closed last due to the arrangement of the centerline latching mechanism and the structural and seal overlap.
Limit switches on each power drive unit turn the ac motors off when the doors are open or closed. Each ac motor has an associated brake that operates similarly to the bulkhead and centerline latch motors. When both motors are operating, the nominal time for payload bay door opening or closing is 63 seconds.
If only one motor is operating, the time is 126 seconds. Each MMCA has its own timer set to twice the normal operating time to allow single-motor operation of the payload bay doors without causing a sequence fail signal PLB doors CRT message and an SM alert .
Torque limiters are incorporated into the rotary actuators to avoid damaging the drive motors or mechanisms if limit switches fail to turn off an electrical drive motor or the mechanisms jam.
Two bolts on the bellcrank and the bolt connecting the link to the rotary actuator can be EVA disconnect points if the linkage fails when the doors close. The power drive unit can be disengaged manually on the ground or on orbit.
The payload bay doors open through an angle of 175.5 degrees.
Two radiator panels on each forward payload bay door can be deployed when the doors are opened on orbit and stowed when the doors are closed before entry, or they can be left in the stowed position for a given flight. Freon-21 coolant loop 1 flows through the left-hand radiator panels, and the No. 2 loop flows through the right-hand panels.
On orbit, the panels radiate excess heat collected by the Freon-21 coolant loops from heat exchangers and cold plates throughout the orbiter. Coolant flows through the radiators from aft to forward. The radiator panels mounted on the forward end of the aft payload bay doors are fixed to the bay doors.
The radiator deploy and stow operation is controlled manually from the aft flight deck panel R13. The PL bay mech (payload bay mechanisms) pwr, radiator latch and radiator control sys switches control the panels. Four indicators show the radiator latch and deploy status.
#
"Orbiter Aft Fuselage",282,0,0,0
The aft fuselage consists of an outer shell, thrust structure and internal secondary structure. It is approximately 18 feet long, 22 feet wide and 20 feet high.
The aft fuselage supports and interfaces with the left-hand and right-hand aft orbital maneuvering system/reaction control system pods, the wing aft spar, midfuselage, orbiter/external tank rear attachments, space shuttle main engines, aft heat shield, body flap, vertical tail and two T-0 launch umbilical panels.
The aft fuselage provides the load path to the midfuselage main longerons, main wing spar continuity across the forward bulkhead of the aft fuselage, structural support for the body flap, and structural housing around all internal systems for protection from operational environments (pressure, thermal and acoustic) and controlled internal pressures during flight.
The forward bulkhead closes off the aft fuselage from the midfuselage and is composed of machined and beaded sheet metal aluminum segments. The upper portion of the bulkhead attaches to the front spar of the vertical tail.
The internal thrust structure supports the three SSMEs. The upper section of the thrust structure supports the upper SSME, and the lower section of the thrust structure supports the two lower SSMEs. The internal thrust structure includes the SSMEs, load reaction truss structures, engine interface fittings and the actuator support structure. It supports the SSMEs, the SSME low-pressure turbopumps and propellant lines. The two orbiter/external tank aft attach points interface at the longeron fittings.
The internal thrust structure is composed mainly of 28 machined, diffusion-bonded truss members. In diffusion bonding, \Jtitanium\j strips are bonded together under heat, pressure and time. This fuses the \Jtitanium\j strips into a single hollow, homogeneous mass that is lighter and stronger than a forged part. In looking at the cross section of a diffusion bond, one sees no weld line.
It is a homogeneous parent metal, yet composed of pieces joined by diffusion bonding. (In OV-105, the internal thrust structure is a forging.) In selected areas, the \Jtitanium\j construction is reinforced with boron/epoxy tubular struts to minimize weight and add stiffness. This reduced the weight by 21 percent, approximately 900 pounds.
The upper thrust structure of the aft fuselage is of integral-machined aluminum construction with aluminum frames except for the vertical fin support frame, which is \Jtitanium\j. The skin panels are integrally machined aluminum and attach to each side of the vertical fin to react drag and torsion loading.
The outer shell of the aft fuselage is constructed of integral-machined aluminum. Various penetrations are provided in the shell for access to installed systems. The exposed outer areas of the aft fuselage are covered with reusable thermal protection system.
The secondary structure of the aft fuselage is of conventional aluminum construction except that \Jtitanium\j and fiberglass are used for thermal isolation of equipment. The aft fuselage secondary structures consist of brackets, buildup webs, truss members, and machined fittings, as required by system loading and support constraints.
Certain system components, such as the avionics shelves, are shock-mounted to the secondary structure. The secondary structure includes support provisions for the auxiliary power units, \Jhydraulics\j, ammonia boiler, flash evaporator and electrical wire runs.
The two external tank umbilical areas interface with the orbiter's two aft external tank attach points and the external tank's liquid oxygen and \Jhydrogen\j feed lines and electrical wire runs. The umbilicals are retracted, and the umbilical areas are closed off after external tank separation by an electromechanically operated \Jberyllium\j door at each umbilical. Thermal barriers are employed at each umbilical door. The exposed area of each closed door is covered with reusable surface \Jinsulation\j.
The aft fuselage heat shield and seal provide a closeout of the orbiter aft base area. The aft heat shield consists of a base heat shield of machined aluminum. Attached to the base heat shield are domes of honeycomb construction that support flexible and sliding seal assemblies. The engine-mounted heat shield is of Inconel honeycomb construction and is removable for access to the main engine power heads. The heat shield is covered with a reusable thermal protection system except for the Inconel segments.
#
"Orbiter OMS/RCS Pods",283,0,0,0
The orbital maneuvering system/reaction control system left- and right-hand pods are attached to the upper aft fuselage left and right sides. Each pod is fabricated primarily of \Jgraphite\j epoxy composite and aluminum. Each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet.
Each pod is divided into two compartments: the OMS and the RCS housings. Each pod houses all the OMS and RCS propulsion components and is attached to the aft fuselage with 11 bolts. The pod skin panels are \Jgraphite\j epoxy honeycomb sandwich. The forward and aft bulkhead aft tank support bulkhead and floor truss beam are machined aluminum 2124.
The centerline beam is 2024 aluminum sheet with \Jtitanium\j stiffeners and \Jgraphite\j epoxy frames. The OMS thrust structure is conventional 2124 aluminum construction. The cross braces are aluminum tubing, and the attach fittings at the forward and aft fittings are 2124 aluminum. The intermediate fittings are corrosion-resistant steel.
The RCS housing, which attaches to the OMS pod structure, contains the RCS thrusters and associated propellant feed lines. The RCS housing is constructed of aluminum sheet metal, including flat outer skins. The curved outer skin panels are \Jgraphite\j epoxy honeycomb sandwich. Twenty-four doors in the skins provide access to the OMS and RCS and attach points.
The two \Jgraphite\j epoxy pods per \Jspacecraft\j reduce the weight by 10 percent, approximately 450 pounds. The pods will withstand 162-decibel acoustic noise and a temperature range from minus 170 to plus 135 F.
The exposed areas of the OMS/RCS pods are covered with a reusable thermal protection system, and a pressure and thermal seal is installed at the OMS/RCS pod aft fuselage interface. Thermal barriers are installed, and they interface with the RCS thrusters and reusable thermal protection system.
#
"Orbiter Body Flap",284,0,0,0
The body flap thermally shields the three SSMEs during entry and provides the orbiter with pitch control trim during its atmospheric flight after entry.
The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. The main upper and lower forward honeycomb skin panels are joined to the ribs, spars and honeycomb trailing edge with structural fasteners. The removable upper forward honeycomb skin panels complete the body flap structure.
The upper skin panels aft of the forward spar and the entire lower skin panels are mechanically attached to the ribs. The forward upper skin consists of five removable access panels attached to the ribs with quick-release fasteners. The four integral-machined aluminum actuator ribs provide the aft fuselage interface through self-aligning bearings.
Two bearings are located in each rib for attachment to the four rotary actuators located in the aft fuselage, which are controlled by the flight control system and the hydraulically actuated rotary actuators. The remaining ribs consist of eight stability ribs and two closeout ribs constructed of chemically milled aluminum webs bonded to aluminum honeycomb core.
The forward spar web is of chemically milled sheets with flanged holes and stiffened beads. The spar web is riveted to the ribs. The trailing edge includes the rear spar, which is composed of piano-hinge half-cap angles, chemically milled skins, honeycomb aluminum core, closeouts and plates. The trailing edge attaches to the upper and lower forward panels by the piano-hinge halves and hinge pins. Two moisture drain lines and one hydraulic fluid drain line penetrate the trailing edge honeycomb core for horizontal and vertical drainage.
The body flap is covered with a reusable thermal protection system and an articulating pressure and thermal seal to its forward cover area on the lower surface of the body flap to block heat and air flow from the structures.
The aft fuselage is built by Rockwell's Space Transportation Systems Division, Downey, Calif. The OMS/RCS pods are built by McDonnell Douglas, St. Louis, Mo. The body flap is built by Rockwell's Columbus, Ohio, division.
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"Orbiter Vertical Tail",285,0,0,0
The vertical tail consists of a structural fin surface, the rudder/speed brake surface, a tip and a lower trailing edge. The rudder splits into two halves to serve as a speed brake.
The vertical tail structure fin is made of aluminum. The main \Jtorque\j box is constructed of integral-machined skins and strings, ribs, and two machined spars. The fin is attached by two tension tie bolts at the root of the front spar of the vertical tail to the forward bulkhead of the aft fuselage and by eight shear bolts at the root of the vertical tail rear spar to the upper structural surface of the aft fuselage.
The rudder/speed brake control surface is made of conventional aluminum ribs and spars with aluminum honeycomb skin panels and is attached through rotating hinge parts to the vertical tail fin.
The lower trailing edge area of the fin, which houses the rudder/speed brake power drive unit, is made of aluminum honeycomb skin.
The hydraulic power drive unit/mechanical rotary actuation system drives left- and right-hand drive shafts in the same direction for rudder control of plus or minus 27 degrees. For speed brake control, the drive shafts turn in opposite directions for a maximum of 49.3 degrees each. The rotary drive actions are also combined for joint rudder/speed brake control. The hydraulic power drive unit is controlled by the orbiter flight control system.
The vertical tail structure is designed for a 163-decibel acoustic environment with a maximum temperature of 350 F.
All-Inconel honeycomb conical seals house the rotary actuators and provide a pressure and thermal seal that withstands a maximum of 1,200 F.
The split halves of the rudder panels and trailing edge contain a thermal barrier seal.
The vertical tail and rudder/speed brake are covered with a reusable thermal protection system. A thermal barrier is also employed at the interface of the vertical stabilizer and aft fuselage.
The contractor for the vertical tail and rudder/speed brake is Fairchild Republic, Farmingdale, N.Y.
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"Orbiter Passive Thermal Control",286,0,0,0
A passive thermal control system helps maintain the temperature of the orbiter \Jspacecraft\j, systems and components within their temperature limits. This system uses available orbiter heat sources and heat sinks supplemented by \Jinsulation\j blankets, thermal coatings and thermal isolation methods. Heaters are provided on components and systems in areas where passive thermal control techniques are not adequate. (The heaters are described under the various systems.)
The \Jinsulation\j blankets are of two basic types: fibrous bulk and multilayer. The bulk blankets are fibrous materials with a density of 2 pounds per cubic foot and a sewn cover of reinforced acrylic film Kapton. The cover material has 13,500 holes per square foot for venting. Acrylic film tape is used for cutouts, patching and reinforcements. Tufts throughout the blankets minimize billowing during venting.
The multilayer blankets are constructed of alternate layers of perforated acrylic film Kapton reflectors and Dacron net separators. There are 16 reflector layers in all, the two cover halves counting as two layers. Covers, tufting and acrylic film tape are similar to that used for the bulk blankets.
The contractors are Hi-Temp \JInsulation\j Inc., Camarillo, Calif. (fibrous insulation); Scheldahl, Northfield, Minn. (cover materials and inner layers); Apex Mills, Los Angeles, Calif. (separators).
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"Orbiter Purge, Vent and Drain System",287,0,0,0
The purge, vent and drain system on the orbiter is designed to perform the following functions: provide unpressurized compartments with gas purge for thermal conditioning and prevent accumulation of hazardous gases, vent the unpressurized compartments during ascent and entry, drain trapped fluids (water and hydraulic fluid) and condition window cavities to maintain visibility.
Three purge circuits are connected by the T-0 umbilical to ground equipment before launch during the preflight countdown and postlanding phases. Purge gas (cool, dry air and gaseous nitrogen) is provided to three sets of distribution plumbing: the forward fuselage, orbital maneuvering system/reaction control system pods, wings and vertical stabilizer; the midfuselage; and the aft fuselage. The purge gas makes all the unpressurized volumes inert, maintains constant \Jhumidity\j and temperature, forces out any hazardous gases and ensures that external contaminants cannot enter.
The active vent system provides the flow area to control pressure during purge, depressurization during ascent, molecular venting in orbit and repressurization during entry.
The vent and purge system is controlled exclusively through guidance, navigation and control software. The active ports are positioned by the software on the basis of mission time or mission events during ascent, entry and aborts and by crew inputs on the CRT and keyboard in the crew compartment flight deck.
There are 18 active vents in the orbiter fuselage, nine on each side. Each vent has a door that can be positioned for a specific purpose at various phases of flight. For identification, each door is numbered, starting at the nose of the orbiter. Each compartment has a dedicated vent on the left and right side of the orbiter for redundancy.
Internal vents are used to vent compartments that have no vent doors of their own, such as the nose wheel well, the two main wheel wells and the vertical tail section. Passive vents are used to back up vent 7 of the forward wing compartment, which responds to a delta pressure to open a check valve (passive vent) during ascent to vent the wing to the midbody if vent 7 fails; or, on descent, the midfuselage pressurizes the wing if vent 7 fails at a delta pressure of 0.72 to 1 psid. The aft bulkhead (X o 1307) has 14 one-way check valves that vent the payload bay into the aft fuselage at a delta pressure of 0.004 to 0.04 psid. Vent 8 vents the OMS/RCS pods, which are joined by a duct that enables the pod to vent through the opposite side of the vehicle if vent 8 fails to open.
All vent doors are driven by an electromechanical actuator. Vent doors located near each other share common actuators and controls. Vents 1 and 2, 4 and 7, and 8 and 9 share drive mechanisms on the left and right side. The 18 doors are divided into six groups of four ac motors each and are staggered so that all 24 motors do not run at the same time. All vent doors are driven inward, and each door has a pressure seal and thermal seal. The normal opening or closing time of a door with two motors operating is five seconds.
Vent doors 1, 2, 8 and 9 have purge positions that control flow from the forward and aft volumes, respectively. Vent 6 has two purge positions and a closed position that accommodates the different purge flow rates available to the payloads and payload bay. These doors are in the purge position before launch.
Two minutes 20 seconds before launch, the launch processing system reduces the purge flow in anticipation of closing vent 6. At T minus 35 seconds, vent 6 is closed. At T minus 25 seconds, the onboard general-purpose computers are enabled and take over the sequences.
At T minus 10 seconds, the vents are configured for launch. Vents 3, 4, 5, 6 and 7 are closed to limit sound pressure levels in the payload bay. Vents 1, 2, 8 and 9 are opened. If a launch abort occurs from T minus 10 seconds to T minus zero, the vent doors reposition to the prelaunch configuration. At T minus four seconds, any vent door out of configuration and not overridden causes the onboard GPCs to call a hold.
At T plus 10 seconds, all vent doors are commanded open. At T plus 80 seconds, vent doors 8 and 9 are commanded closed to prevent hazardous gases from entering the aft fuselage; at T plus 122 seconds, vents 8 and 9 are commanded open. All vent doors remain open during the remainder of ascent and on-orbit operations.
In preparation for entry, the onboard operational sequence software (OPS 3) closes all vent doors. The doors remain closed until the velocity of the orbiter reaches 2,400 feet per second, when all vents are opened by the onboard GPCs.
At the end of the mission, after the orbiter stops on the runway, vent doors 1, 2, 6, 8 and 9 are configured to their purge positions for ground cooling.
The purge and vent ducting is now made of Kevlar/epoxy (115 pieces up to 11 inches in diameter), which replaced the fiberglass or aluminum ducts and reduced the weight of the ducts 33 percent, or approximately 200 pounds.
The window cavity conditioning system prevents moisture from entering into the windshields and the cavities of the overhead and payload-viewing windows. It also depressurizes and repressurizes these cavities during flight and supplies the purge conditioning to dry them during ground operations. The side hatch window is self-contained.
A hazardous gas detection system detects hazardous levels of explosive or toxic gases. The onboard orbiter sample lines duct the compartment gases to the ground support equipment at the T-0 right-hand umbilical panel and to the ground-based mass spectrometer for analysis at the launch pad.
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"Orbiter In-Flight Crew Escape System",288,0,0,0
The in-flight crew escape system is provided for use only when the orbiter would be in controlled gliding flight and unable to reach a runway. This condition would normally lead to ditching. The crew escape system provides the flight crew with an alternative to water ditching or to landing on terrain other than a landing site. The probability of the flight crew surviving a ditching is very slim.
The hardware changes required to the orbiters enable the flight crew to equalize the pressurized crew compartment with the outside pressure via the depressurization valve opened by pyrotechnics in the crew compartment aft bulkhead that would be manually activated by a flight crew member in the middeck of the crew compartment; pyrotechnically jettison the crew ingress/egress side hatch manually in the middeck of the crew compartment; and bail out from the middeck through the ingress/egress side hatch opening after manually deploying the escape pole through, outside and down from the side hatch opening.
One by one, each flight crew member attaches a lanyard hook assembly, which surrounds the deployed escape pole, to his or her parachute harness and egresses through the side hatch opening. Attached to the escape pole, the crew member slides down the pole and off the end. The escape pole provides each crew member with a trajectory that takes the crew member below the orbiter's left wing.
Changes were also made in the software of the orbiter's general-purpose computers. The software changes were required for the primary avionics software system and the backup flight system for transatlantic-landing and glide-return-to-launch-site abort modes. The changes provide the orbiter with an automatic-mode input by flight crew members through keyboards at the commander's and/or pilot's panel C3, which provides the orbiter with an automatic stable flight for crew bailout. This software change, which is required to allow the flight crew commander's departure, automatically controls the orbiter's velocity and angle of attack to the desired bailout conditions.
The crew would make the escape decision at an altitude of approximately 60,000 feet and would immediately make an input to the flight control system software autopilot mode.
When the orbiter descends to an altitude of approximately 30,000 feet, its airspeed must be decreased to approximately 200 knots (230 mph). At approximately 25,000 feet, a crew member in the middeck (referred to as the jump master and seated in the forward left seat in the middeck) raises a cover on the left side of the crew compartment middeck at floor level and pulls the T-handle, which activates the pyrotechnics for the depressurization valve at the crew compartment X o 576 aft bulkhead. This equalizes the crew compartment cabin and outside pressure before the side hatch is jettisoned.
At approximately 25,000 feet, the software for the automatic autopilot mode changes the orbiter's angle of attack to approximately 15 degrees. This angle of attack must remain nearly constant for approximately three minutes until the orbiter reaches an altitude of approximately 2,000 feet.
At approximately 25,000 feet, the jump master jettisons the side hatch by pulling the hatch jettison T-handle next to the depressurization T-handle. When the T-handle is pulled, pyrotechnics separate the hatch assembly by severing the side hatch hinge, and three pyrotechnic thrusters jettison the tunnel/hatch from the orbiter at a velocity of approximately 50 feet per second.
The jump master pulls the pip pin on the escape pole and pulls the ratchet handle down, which permits the two telescoping sections of the escape pole to be deployed through the hatch opening by spring tension.
A magazine assembly located near the side hatch contains a lanyard assembly for each flight crew member. Each lanyard assembly consists of a hook attached to a Kevlar strap that surrounds the escape pole. Five roller bearings on each strap surround the pole and permit the lanyard to roll freely down the pole. Each flight crew member positions himself or herself at the hatch opening and attaches himself or herself to the escape pole via the lanyard hook assembly and jumps out the hatch opening.
Each lanyard assembly incorporates an energy absorber rated at 1,000 pounds. The Kevlar strap consists of two sections of permanent Nomex thread stitching and a section of breakaway Kevlar thread stitching. When the crew member exits the side hatch on the escape pole, the breakaway Kevlar thread stitching can break away, providing the crew member with an energy absorber. The crew member slides down the escape pole and off the end into a free-fall. The escape pole extends downward 9.8 feet from the side hatch and provides the crew member with a trajectory that will carry him or her beneath the orbiter's left wing.
It would take approximately 90 seconds for a maximum crew of eight to bail out. After the first crew member bails out from the middeck, the remaining crew members follow at approximately 12-second intervals until all are out by approximately 10,000 feet altitude.
A handhold was added in the middeck next to the side hatch to permit the crew members to position themselves through the side hatch opening for bailout.
The escape pole is constructed of aluminum and steel. The arched housing for the pole is 126.75 inches long and is attached to the middeck ceiling above the airlock hatch and at the 2 o'clock position at the side hatch for deployment during launch and entry. The escape pole telescopes from the middeck housing through the side hatch in two sections. The primary extension is 73 inches long, and the end extension is 32 inches long. The diameter of the housing is 3.5 inches. The two telescoping sections are slightly smaller in diameter. The escape pole weighs approximately 241 pounds-248 pounds with attachments.
On orbit, the escape pole's primary stowage position requires unpinning the escape pole at the starboard and port attachments, rotating the pole so it is flat against the middeck ceiling and strapping it to the ceiling. An alternate on-orbit stowage approach also requires unpinning the escape pole at the starboard and port attachments, rotating it so it is flat against the middeck ceiling and strapping it to the ceiling.
The side hatch water coolant lines for side hatch thermal conditioning were modified to accommodate the installation of the side hatch pyrotechnic separation system.
The flight crew members' seats were also modified to accommodate the seat/crew altitude protection system suit for each crew member.
The pyrotechnically operated crew compartment depressurization valve consists of two flapper valves with debris screens on the crew compartment side and payload bay side that open to depressurize the crew compartment and close when the pressure equalizes.
It is noted that the hatch jettison features could be used in a landing emergency.
The crew member's altitude protection suit includes an emergency oxygen system, pilot and drogue parachutes that are operated automatically and have manual backup, a main parachute that is operated automatically and has manual backup, a seawater activation release system, flotation devices, a life raft and survival equipment. The crew altitude protection suit and its associated equipment weigh approximately 70 pounds.
The side hatch jettison thruster contractor is OEA, \JDenver\j, Colo. The pyrotechnics contractor for the hatch tunnel, hinge and the energy transfer system lines is Explosive Technology, Fairfield, Calif. The escape pole is government-furnished equipment that is supplied by NASA's Johnson Space Center, Houston, \JTexas\j, as is the crew altitude protection suit.
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"Orbiter Emergency Egress Slide",289,0,0,0
The emergency egress slide provides the orbiter flight crew members with a rapid and safe emergency egress through the orbiter middeck ingress/egress side hatch after a normal opening of the side hatch or after jettisoning of the side hatch at the nominal end-of-mission landing site or at a remote or emergency landing site.
The emergency egress slide replaces the emergency egress side hatch bar, which required the flight crew members to drop approximately 10.5 feet to the ground. This drop could cause injury to the flight crew members and prevent an injured flight crew member from moving to a safe distance from the orbiter.
The emergency egress slide will support return- to- launch- site, transatlantic-landing, abort-once-around and normal end-of-mission landings.
The system will be activated manually by the flight crew rotating the slide from the middeck through the egress side hatch opening onto the side hatch if the hatch has not been jettisoned or through the egress side hatch opening if the hatch has been jettisoned. The flight crew pulls a lanyard to inflate the slide with a self-contained air bottle supply.
The slide allows the safe egress of the flight crew members to the ground within 60 seconds after the side hatch is fully opened or jettisoned; accommodates the egress of the flight crew members wearing the launch and entry crew altitude protection system; accommodates the egress of incapacitated crew members; withstands and remains functional in the egress environment for a minimum of six minutes after deployment; and can be released from the side hatch to permit fire truck access.
The slide is installed inside the middeck below the side hatch where it will not inhibit ingress/egress when the system is not required and not interfere with normal on-orbit operations.
The egress slide contractor is Inflatable Systems Inc., a division of OEA, \JDenver\j, Colo.
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"Orbiter Secondary Emergency Egress",290,0,0,0
The left-hand flight deck overhead window provides the flight crew with a secondary emergency egress route. The left overhead window consists of three panes of glass, an inner pane attached to the crew compartment and a center and outer pane attached to the upper forward fuselage.
When the secondary emergency egress path is utilized, pulling the T handle located forward of the flight deck center console (between the commander and pilot) activates the overhead window jettison system. When initiated, the center and outer panes are jettisoned as a unit, upward and aft. A time delay in the pyrotechnic firing circuit delays the initiation of the jettisoning of the inner pane 0.3 of a second after the center and outer panes are jettisoned.
Upon the initiation of the jettisoning of the inner window pane, it rotates downward and aft into the crew compartment aft flight deck on hinges located at the aft portion of the window frame. A capture device attenuates the opening rate and holds the window in position.
The overhead window jettison system consists primarily of expanding tube assemblies, mild detonating fuses, frangible bolts and associated initiators.
The left overhead window jettison system can be initiated from the outside of the orbiter on the right side of the forward fuselage by ground personnel.
Egress steps are mounted at the aft flight deck station (left side) to assist the flight crew up through the window.
Emergency ground descent devices are stowed on the overhead aft flight deck adjacent to the left overhead window. One device is provided for each flight crew member. The emergency ground descent device enables flight crew members to lower themselves to the ground over the side of the orbiter.
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"Orbiter Side Hatch Jettison",291,0,0,0
The middeck ingress/egress side hatch was modified to provide the capability of pyrotechnically jettisoning the side hatch for emergency egress on the ground. In addition, a crew compartment pressure equalization valve provided at the crew compartment aft bulkhead, X o 576, is also pyrotechnically activated to equalize cabin/outside pressure before the jettisoning of the side hatch.
A panel on the left side of the middeck of the crew compartment contains two T-handles. One T-handle controls the initiation of the pyrotechnic pressure equalization valves, which equalize the cabin pressure with outside pressure.
The other T-handle in the same panel in the middeck jettisons the side hatch pyrotechnically. When this T-handle is activated, pyrotechnics sever the hinges of the side hatch and three pyrotechnic tunnel/hatch thrusters are initiated, which jettisons the side hatch from the orbiter.
The side hatch jettison thruster contractor is OEA, \JDenver\j, Colo. The pyrotechnics contractor for the hatch tunnel, hinges and the energy transfer system lines is Explosive Technology, Fairfield, Calif.
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"Shuttle Food System and Dining",292,0,0,0
The middeck of the orbiter is equipped with facilities for food stowage, preparation, and dining for each crew member. The food supply is categorized as either menu food or pantry food. Menu food consists of three daily meals per crew member and provides an average energy intake of approximately 2,700 calories per crew member per day. The pantry food is a two-day contingency food supply that also contains food for snacks and beverages between meals and for individual menu changes.
It provides an average energy intake of 2,100 calories per crew member per day. The types of food include fresh, thermostabilized, rehydratable, irradiated, intermediate-moisture, and natural-form food and beverages.
If a payload is installed in the middeck in lieu of the galley, the food preparation system is limited. It consists of the water dispenser, food warmer, food trays and food system accessories.
The water dispenser provides the flight crew with ambient and chilled water for drinking and reconstituting food. The water dispenser consists of a housing assembly, rehydration station, hygiene water quick disconnect and water lines. Two flex lines 10 feet long connect the housing assembly to the ambient and chilled potable water system. Both lines have quick disconnects. A 12-foot-long flex line with a quick disconnect and water-dispensing valve supplies water for personal hygiene.
The water selector valve "amb" position provides ambient water to the rehydration station between 65 and 75 F. The "off" position prevents water from flowing to the rehydration station (it does not shut off water flow to the personal hygiene water outlet quick disconnect). The "chd" water position provides chilled water to the rehydration station between 45 and 55 F.
Depressing the hygiene water valve handle allows a constant flow of ambient water. Releasing the handle prevents water flow. The locked-open position allows a constant flow of ambient water without holding the handle.
The rehydration station is an electronic dispensing system that interfaces directly with food and beverage packages to provide rehydration capability and drinking water for flight crew members. The system dispenses 2, 3, 4 and 8 ounces of water through a replaceable needle. A spare needle is stowed at the rear of the rehydration unit and another in the in-flight maintenance middeck locker. The needles are removed and installed with a 3/8-inch open-end wrench. Depressing the "pwr" push button at the rehydration station provides power to the electronic rehydration system and an indicating light is illuminated within the switch upon activation.
Depressing the "pwr" push button again deactivates the system. The water quantity rotary switch's 2, 3, 4 and 8 positions provide 2, 3, 4 and 8 ounces of water, respectively. The needle must be inserted into the package before depressing the "fill" push button to prevent free water from being dispensed into the crew cabin environment. Depressing the "fill" push button activates the electronic filling mechanism when the water quantity selection has been made.
A light comes on within the "fill" switch during filling and goes out when filling is complete. The operation is automatically deactivated. The bypass valve provides a continuous flow of water to the food rehydration unit when the handle is depressed or lifted to the "up locked-open" position.
The rehydratable food container is inserted into the rehydration station, the water dispenser needle penetrates the rubber septum on the rehydratable container, and the specified amount of water is discharged into the container. The rehydrated food is mixed and heated, if required.
The rehydrated food container is opened by grasping the center portion of the lid liner with the fingers, piercing the liner with a knife or scissors and pulling the liner up to aspirate air. While grasping the center of the liner, the \Jastronaut\j swings container in a gentle forward and backward semicircular motion to place food contents at the bottom of the container. The inside edge of the lid liner (three sides) is cut with a knife or scissors to expose the food.
The rehydratable beverage container is inserted into the rehydration station, the water dispenser needle penetrates the rubber septum on the rehydratable container, and the specified amount of water is discharged into the container. The rehydratable beverage is mixed and heated, if required. A plastic clip is affixed to the straw in the closed position, the probe end of the straw is inserted into the container rubber septum, the straw is placed in the mouth, the clip is released, and the beverage is drunk. All straws are color-coded for each crew member.
Food trays are kept in a middeck stowage locker (or in the galley, if installed) at launch and are removed and installed in the use locations during preparations for the first meal. The tray is a clear, anodized aluminum sheet that restrains food and accessories during dining. The trays are color-coded for each crew member. Velcro on the bottom of the food trays allows them to be attached to the front of the middeck lockers (or the galley door, if installed) for food preparation or dining.
The straps will also hold the trays on the crew member's leg for dining. A cutout on the tray allows three rehydratable food packages to be secured to the tray. Another cutout with rubber strips adapts to various-sized food packages, including cans, pouches and rehydratable food packages.
Two magnetic strips hold eating utensils and two 0.75-inch-wide binder clips on the tray retain such things as condiment packets and wipes. Accessories used during food preparation and dining include condiments, gum and candy, vitamins, wet wipes, dry wipes, drinking containers, drinking straws, utensils and a re-entry kit that contains salt tablets and long straws.
Condiments include salt, pepper, taco sauce, hot pepper sauce, catsup, mayonnaise and mustard. The salt and pepper are liquids stored in small plastic squeeze bottles. The remaining condiments are packaged in individual, sealed, flexible plastic pouches. Vitamin tablets supplement dietary requirements. Wet wipes are packaged in 21 individual packets per dispenser for cleaning utensils after dining.
A light spring action retains and positions wipes for dispensing. Empty beverage containers of rigid plastic for drinking and storing water are carried in crew members' clothing and can be filled at the water dispenser. Approximately five to 10 different color-coded straws are provided for each crew member (depending on the flight's duration) for drinking beverages and water.
Additional straws are kept in the pantry beverages and various menu locker trays. Color-coded utensils include a knife, two spoons (large and small), a fork and a can opener for each crew member. They are stowed with a soft plastic holder that has a Velcro snap cover. Dry wipes are packaged in a 30-wipe container that can be attached to the crew cabin wall with Velcro for cleanup after dining.
The re-entry kit consists of one package containing eight salt tablets for each crew member and long straws (four per crew member). Two salt tablets are to be taken with 8 ounces of water or other beverage by each crew member four times before entry. The re-entry kit may be stowed in one of three locations depending on space available-with the accessories, near the last meal to be consumed on orbit or with the pantry beverages.
The food warmer is a portable heating unit that can warm a meal for at least four crew members within one hour when the galley is not flown. It is stowed in a middeck locker at launch and is removed and installed during meal preparation activity. The food is heated by thermal \Jconduction\j on a hot plate (element). The warmer is thermostatically controlled between 165 and 175 F.
The case is constructed of aluminum with an exterior envelope of 13 by 18 by 6 inches. It has latches and is lined with clear urethane \Jfoam\j \Jinsulation\j coated with room-temperature vulcanizing compound. The case has straps for handling and on-orbit installation. The exterior contains controls and displays, a power connector that interfaces with the power cable and Velcro attachment. A hinged element is sandwiched between two aluminum plates and is contained by a fiberglass frame.
The aluminum plates have spring-bungee restraints for foil-backed food packages on one side. An on/off switch provides two-phase ac power to the unit and a light indicates the warmer is operating. The power cable is 156 inches long and attaches to a middeck ac utility outlet. The cable is stowed inside the case at launch.
A maximum of 14 packages can be installed on the side of the spring bungees and eight on the other side. Rehydratable beverages should be placed on the side opposite the spring bungees, and the \Jfoam\j on the other side is additionally relieved to prevent the packages from popping out in zero gravity.
When 14 rehydratable packages are heated, no foil-backed food pouches can be heated. A maximum of six foil-backed food pouches can be heated in conjunction with 12 rehydratable packages. When foil-backed pouches are heated, only four rehydratable packages can be heated on the side of the spring bungees. The foil-backed pouches are stacked three deep. Four rehydratable packages are inserted in the outer recessed \Jfoam\j cutouts. At the bottom of the cutouts, 0.5-inch-thick uncoated \Jfoam\j absorbs moisture or spilled liquids.
The galley is a multipurpose facility that provides a centralized location for one individual to handle all food preparation activities for a meal. The galley has facilities for heating food, rehydrating food and stowing food system accessories and food trays. The galley consists of a rehydration station, oven, food trays and food system accessories.
The oven is divided into two principal compartments-a lower compartment designed for heating at least 14 rehydratable food containers inserted on tracks and an upper compartment designed to accept a variety of food packages, including the rehydratable containers. At least seven food pouches can be heated in the upper compartment and are held against the heat sink by four spring-loaded plates. The oven has a heating range of 145 to 185 F. During launch and entry, the oven door is held closed by a restraining strap, which is removed from the door by releasing the snap for on-orbit operations. An on/off switch enables and removes power to three fans.
The rehydration station dispensing system interfaces directly with food and beverage packages, providing rehydration capability and drinking water for crew members. A gauge indicates hot water temperatures of 100 to 220 F. The volume/ounces switch selects the volume of water to be dispensed by the rehydration station in 0.5-ounce increments from 0.5 of an ounce to 8 ounces.
The yellow hot push button indicator allows hot water to be dispensed when it is depressed and is illuminated when energized. When the selected volume of water has been dispensed, the push button will begin to flash on and off. The light will be extinguished when the food package is retracted, releasing the \Jhydration\j station lever arm/limit switch.
The rehydration station lever arm/limit switch serves as an interlock so water can be dispensed only when a food package is connected to the needle. The food package makes contact with the rehydration station lever, which activates the limit switch (note that the flight crew does not physically actuate the lever).
The blue cold push button indicator allows cold water to be dispensed when it is depressed and is illuminated when energized. When the selected volume of water has been dispensed, the push button will begin to flash on and off. The light will be extinguished when the food package is retracted, releasing the rehydration station lever arm/limit switch.
The galley light is located on the upper left-hand side of the galley structure surface and has a single light brightness control. Moving the knob clockwise from off applies power to the light and provides variable brightness control.
Two condiment dispensers are attached to the galley by Velcro tabs on the back of the dispensers. The dispensers are available for holding individual packets, such as catsup, taco sauce, mayonnaise and mustard, on the front panel below the oven. The dispensers are open-ended boxes designed to hold the stack of packets together so they may be individually removed as needed. A slide plate keeps the packets from becoming loose as the items are depleted.
A single dispenser for holding individual packets of wet wipes is located on the front panel below the oven and is slightly different in design than the condiment dispensers.
Dispensers for liquid salt and pepper and vitamins can be restrained by clips conveniently located below the rehydration station.
Food trays and food system accessories are the same as those used on flights without the galley.
On the upper left-hand corner of the galley behind a Teflon cloth panel is an MV3 valve that has "emer off" and "on" positions. The "on" position serves as the nominal open position of the manual shutoff valve, and "emer off" serves as a manual shutoff valve for the ambient temperature water supply to the galley.
On the upper right-hand corner of the galley behind a Teflon cloth panel are test connectors, a dc power bus A and B switch and a flush port quick-disconnect test port. The two test connectors serve as a hookup for ground support equipment. The dc power bus A switch's "on" position activates the galley oven heaters, rehydration station system and one of six water tank strip heaters; the "off" position deactivates the heaters. The dc power bus B switch's "on" position activates five of the six galley water tank strip heaters, and the "off" position deactivates them. The flush port quick disconnect serves as the galley water system GSE flush port.
On the lower left-hand side of the galley is an auxiliary port water quick disconnect that allows the crew members to obtain ambient potable water when the MV3 valve is off or hygiene water when the 12-foot flex line and water dispensing valve are attached to the quick disconnect.
Three one-hour meal periods are scheduled for each day of the mission. This hour includes actual eating time and the time required to clean up. Breakfast, lunch and dinner are scheduled as close to the usual hours as possible. Dinner is scheduled at least two to three hours before crew members begin preparations for their sleep period.
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"Shuttle Orbiter Medical System",293,0,0,0
The shuttle orbiter medical system is required to provide medical care in flight for minor illnesses and injuries. It also provides support for stabilizing severely injured or ill crew members until they are returned to Earth. The SOMS consists of two separate packages: the medications and bandage kit and the emergency medical kit. The MBK is blue and the EMK is also blue with red Velcro.
The medical kits are stowed in a modular locker in the middeck of the crew compartment. If the kits are required on orbit, they are unstowed and installed on the locker doors with Velcro.
Each kit contains pallets. The MBK pallet designators are D, E and F. The D pallet contains oral medications consisting of pills, capsules and suppositories. The E pallet contains bandage materials for covering or immobilizing body parts. The F pallet contains medications to be administered by topical application.
The EMK pallet designators are A, B, C and G. The A pallet contains medications to be administered by injection. The B pallet contains items for performing minor surgeries. The C pallet contains diagnostic/therapeutic items consisting of instruments for measuring and inspecting the body. The G pallet contains a microbiological test kit for testing for bacterial infections.
The diagnostic equipment on board and information from the flight crew will allow diagnosis and treatment of injuries and illnesses through consultation with flight surgeons in the Mission Control Center in Houston.
The operational bioinstrumentation system provides an amplified electrocardiograph analog signal from either of two designated flight crew members to the orbiter avionics system, where it is converted to digital tape and transmitted to the ground in real time or stored on tape for dump at a later time. The designated flight crew members wear the OBS during the ascent and entry phases. On-orbit use will be limited to contingency situations.
The OBS electrodes are attached to the skin with electrode paste to establish electrical contact. The electrode is composed of a plastic housing containing a non-polarizable pressed pellet. The housing is attached to the skin with double-sided adhesive tape and the pellet contacts the skin. There are three electrodes on the harness marked LC (lower chest), UC (upper chest) and G (ground).
The ECG signal conditioner is a hybrid microcircuit with variable gain (adjusted for each crew member before flight). It provides a zero- to 5-volt output and has an on/off switch within the input plug, which is actuated when the intravehicular activity biomed cable is plugged in. The unit has batteries that will not be replaced in flight.
The IVA biomed cable connects to the signal conditioner and is routed under the IVA clothing to connect to the biomed seat cable. The biomed seat cable is routed to one of the biomed input connectors located on panel A11, A15 or M062M. Rotary control switches on panel R10 provide circuits from the biomed outlets to the orbiter's network signal processor for downlink or recording. The two rotary switches on panel R10 are biomed channel 1 and channel 2 . Extravehicular activity positions provide circuits for the EVA UHF transceiver.
The electrode application kit contains components to aid in the application of electrodes. The components include wet wipes, double-sided adhesive tape, overtapes, electrode paste and a cue card illustrating electrode placement.
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"Shuttle Radiation Equipment",295,0,0,0
The harmful biological effects of radiation must be minimized through mission planning based on calculated predictions and monitoring of dosage exposures. Preflight requirements include a projection of mission radiation dosage, an assessment of the probability of solar flares during the mission and a radiation exposure history of flight crew members. In-flight requirements include the carrying of passive dosimeters by the flight crew members and, in the event of solar flares or other radiation contingencies, the readout and reporting of the active dosimeters.
There are four types of active dosimeters: pocket dosimeter high, pocket dosimeter low, pocket dosimeter FEMA and high-rate dosimeter. All four function in the same manner and contain a \Jquartz\j fiber positioned to zero by electrostatic charging before flight. The unit discharges according to the amount of radiation received; and as the unit discharges, the \Jquartz\j moves. The position of the fiber along a scale is noted visually. The PDH unit's range is zero to 100 rads. The PDF and PDL units' ranges are zero to 200 millirads and the HRD unit's range is zero to 600 rads.
The rad is a unit based on the amount of energy absorbed and is defined as any type of radiation that is deposited in the absorbing media, and radiation absorbed by man is expressed in roentgen equivalent in man, or rems. The rem is determined by multiplying rads times a qualifying factor that is a variable depending on wavelength, source, etc. For low-inclination orbits (35 degrees and lower), the qualifying factor is approximately equal to one; therefore, the rem is approximately equal to the rad. In space transportation system flights, the doses received have ranged from 0.05 to 0.07 rem, well below flight crew exposure limits.
The flight crew's passive dosimeters are squares of fine-ground photo film sandwiched between plastic separators in a light-proof package. Radiation striking the silver halide causes spots on the film, which can be analyzed after the flight. Included in the badge dosimeters are thermoluminescent dosimeter chips, which are analyzed on Earth.
Passive radiation dosimeters are placed in the crew compartment before launch by ground support personnel and removed after landing for laboratory analysis. Each flight crew member carries a passive dosimeter at all times during the mission. The remaining dosimeters are stowed in a pouch in a middeck modular locker. If a radiation contingency arises, the PDL, PDH, HRD and PDF active dosimeters will be unstowed, read, and recorded for downlink to the ground.
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"Shuttle Crew Apparel",296,0,0,0
During launch and entry, crew members wear the crew altitude protection system, which consists of a helmet; communications cap; pressure garment; anti-exposure, anti-gravity suit; gloves; and boots.
The crew wears escape equipment over the CAPS during launch and entry. It consists of an emergency oxygen system; parachute harness, parachute pack with automatic opener, pilot chute, drogue chute and main canopy; a life raft; 2 liters of emergency drinking water; flotation devices; and survival vest pockets containing a radio/beacon, signal mirror, shroud cutter, pen gun flare kit, sea dye marker, smoke flare and beacon. Manual activation of the parachute automatic opening sequences is provided, as well as manual release of the parachute main canopy.
On orbit, optional clothing and equipment include underwear, urine collection devices, eyeglasses, communications headset, emesis bag, flashlight, Swiss army knife, kneeboard, pens and pencils, stowage bags, watches and food and drink containers.
Crew clothing and equipment used during on-orbit activities include flight suits, IVA trousers, IVA jackets, IVA shirts, sleep shorts, IVA soft slippers, underwear, scissors/lanyard, pocket dosimeter and pocket food.
Crew clothing is designed for use by 90 percent of the male and female population, the 5th to 95th percentile.
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"Shuttle Sleeping Provisions",297,0,0,0
Sleeping provisions for flight crew members consist of sleeping bags, sleep restraints or rigid sleep stations. The sleeping arrangements can consist of a mix of bags and sleep restraints or rigid sleep stations on a given mission. During a mission with one shift, all crew members sleep simultaneously. If all crew members sleep simultaneously, at least one crew member will wear a communication headset to ensure reception of ground calls and orbiter caution and warning alarms.
If sleeping bags are used, they are installed on the starboard middeck wall and deployed for use on orbit.
If the rigid sleep station is used on a mission, it is installed on the starboard side of the middeck. There are two types of rigid sleep stations. One sleep station type accommodates three crew members and the other accommodates four.
If the rigid sleep station is not installed for a mission, a sleeping bag is furnished each crew member. Each sleeping bag contains a support pad with adjustable restraining straps and a reversible/removable pillow and head restraint. Apollo sleeping bags may be provided for the crew members on request. The Apollo sleeping bag is constructed of beta material and is perforated for thermal comfort.
Six adjustable straps permit the sleeping bag to be adjusted to its proper configuration. Three helical springs above the adjustable straps on one side of the bay relieve loads exerted by the crew member on the crew compartment structure. Six pip pins allow the bag to be attached to the middeck locker face in either a horizontal or vertical configuration. Two elastic adjustable straps restrain the upper and lower parts of the body in the bag. Velcro strips on the ends of both sides of the head restraint attach it to the pillow.
A double zipper arrangement permits the sleeping bag to be opened and closed from the bottom to the top of the bag. One zipper on each side of the sleeping bag allows the bag to be attached to a support pad for better rigidity.
The Apollo beta cloth sleeping bag has four adjustable straps with pip pins that are connected to any two lockers in the middeck separated by a distance equal to a four-tiered locker configuration. For torso restraint, a single two-piece strap is provided and a single zipper opens the bag. The bags are stowed in a middeck locker during launch and entry.
A sleep kit is provided for each crew member and is stowed in the crew member's clothing locker during launch and entry. Each kit contains eye covers and ear plugs for use as required during the sleep period.
The three- or four-tier rigid sleep stations contain a sleeping bag, personal stowage provisions, a light and a ventilation inlet and outlet in each of the tiers. The cotton sleeping bag is installed on the ground in each tier and held in place by six spring clips. The light in each tier is a single fluorescent fixture with a brightness control knob and an off position. The air ventilation inlet duct is an air diffuser similar to an \Jautomobile\j ventilation duct. It is adjusted by moving the vane control knob.
The air ventilation outlet duct is located in the fixed panel at each tier and is opened or closed by moving the vane control knob. The air inlet is located at the crew member's head. The outlet is at the feet. All crew members' heads are toward the airlock and their feet toward the avionics bay.
In the three-tier configuration, the upper and middle crew members face the ceiling and the lower tier crew member faces the floor. The fixed panel at the lower sleep station is removable to provide access to the cabin debris trap door for cleaning the cabin filter, to gain access to floor locker MD76C and to enter the forward portion of the lower equipment bay to clean the avionics bay fan filter.
In the four-tier configuration, the bottom tier sleep restraint hookup provision allows the crew member to position himself at a 15-degree angle, which provides more room, or in the normal horizontal position. The sleeping bag, personal stowage provisions, light and ventilation inlet and outlet are the same as in the three-tier configuration. The head and feet orientations of the crew members are also the same as in the three-tier configuration. The lowest tier is removable so access can be obtained to the cabin debris trap door to clean the cabin filter, gain access to floor locker MD76C and enter the forward portion of the lower equipment bay to clean the avionics bay fan filter.
The three-tier rigid sleep station is made of plastic honeycomb panel and weighs approximately 205 pounds. The four-tier rigid sleep station is made of metal and weighs 173 pounds.
A 24-hour period is normally divided into an eight-hour sleep period and a 16-hour wake period for each crew member. Forty-five minutes are allocated for the crew members to prepare for the sleep period and another 45 minutes when they awake to wash and get ready for the day.
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"Shuttle Personal Hygiene Provisions",298,0,0,0
To maintain good hygiene and appearance, personal hygiene and grooming provisions are furnished for both male and female flight crew members. Water is provided by the water dispensing system.
A personal hygiene kit is furnished each crew member for brushing teeth, hair care, shaving, nail care, etc. A kit is also furnished with articles essential to female hygiene and grooming.
Two washcloths and one towel per crew member per day are provided in addition to two paper tissue dispensers per crew member for each seven days. The washcloths are 12 by 12 inches and the towels 16 by 27 inches. The tissues are absorbent, multi-ply, low-linting paper. Rubber restraints with a Velcro base allow the crew members to restrain their towels and washcloths on the waste management door or middeck walls.
The personal hygiene provisions are stowed in middeck stowage lockers at launch and are removed for use on orbit.
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"Shuttle Housekeeping",299,0,0,0
In addition to time scheduled for sleep periods and meals, each crew member has housekeeping tasks that require from five to 15 minutes of his time at intervals throughout the day. These include cleaning the waste management compartment, the dining area and equipment, floors and walls (as required), the cabin air filters; trash collection and disposal; and changeout of the crew compartment carbon dioxide (lithium hydroxide) absorber canisters.
The materials and equipment available for cleaning operations are biocidal cleanser, disposable gloves, general-purpose wipes and a vacuum cleaner. The cleaning materials and vacuum are stowed in middeck lockers. The vacuum cleaner is powered by the orbiter's electrical power system.
The biocidal cleanser is a liquid detergent formulation in a container approximately 2 inches in diameter and 6 inches long. The container has a built-in bladder, dispensing valve and nozzle. The cleanser is sprayed on the surface to be cleaned and wiped off with dry general-purpose wipes. It is used for periodic cleansing of the waste collection system urinal and seat and the dining area and equipment. It is also used, as required, to clean walls and floors. Disposable plastic gloves are worn while using the biocidal cleanser.
General-purpose wipes are also used for general-purpose cleaning.
The vacuum cleaner is provided for general housekeeping and cleaning of the crew compartment air filters and Spacelab filters (on Spacelab missions). It has a normal hose, extension hose and several attachments. It is powered by the orbiter dc electrical power system.
Trash management operations include routine stowage and daily collection of wet and dry trash, such as expended wipes, tissues and food containers. Wet trash includes all items that could offgas. The equipment available for trash management includes trash bags, trash bag liners, wet trash containers and the stowable wet trash vent hose.
Three trash bags are located in the crew compartment. Each bag contains a disposable trash bag liner. Two bags are designated for dry trash and one for wet trash. At a scheduled time each day, the trash bag liner for dry trash is removed from its trash bag. The liner is closed with a strip of Velcro and stowed in an empty locker.
When more than 8 cubic feet of wet trash is expected, the trash bag liners for wet trash are removed at a scheduled time each day and placed in a wet trash container. The container is then closed with a zipper and the unit is stowed. If expansion due to offgassing is evident, the container is connected to a vent in the waste management system for overboard venting of the gas.
The wet trash container is made of airtight fabric and is closed with a seal-type slide fastener. The container has a volume of approximately 0.7 cubic foot and has an air inlet valve on one end and a quick disconnect on the other end. It is attached to the waste management vent system beneath the commode, enabling air to flow through the wet trash container and then overboard. It is attached through a 41-inch- long vent hose filter. When the container is full, it is removed and stowed in a modular locker.
An 8-cubic-foot wet trash stowage compartment is available under the middeck floor. Each day, the trash bag liners for wet trash are removed from the trash bags and stowed in the wet trash stowage compartment, which is vented overboard. If the compartment becomes full, the trash bag liners for wet trash are stowed in wet trash containers.
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"Shuttle Sighting Aids",300,0,0,0
Sighting aids include all items used to aid the flight crew within and outside the crew compartment. The sighting aids include the crewman optical alignment sight, binoculars, adjustable mirrors, spotlights and eyeglasses.
The COAS is a collimator device similar to an \Jaircraft\j gunsight. Two are installed in the crew compartment flight deck. One COAS is mounted during launch and entry over the positive X commander's forward window and on orbit is removed and mounted next to the aft flight deck overhead right negative Z window. The other COAS is mounted at the aft flight deck station for checking the alignment of the payload bay doors.
When the COAS is mounted at the commander's station, it allows the viewers to reassure themselves of proper attitude orientation during the ascent and deorbit thrusting periods. When the COAS is removed from the commander's station to the aft flight deck for on-orbit operations, it provides a backup to the orbiter star trackers for inertial measurement unit alignment. It is also used as the primary optical instrument for measuring range and rotational rates and allows the flight crew members to align the vehicles and dock.
The COAS consists of a lamp with an intensity control, a reticle, a barrel-shaped housing, a mount, a combiner assembly and a power cable. The reticle consists of a 10-degree circle, vertical and horizontal cross hairs with 1-degree marks, and an elevation scale on the right side of minus 10 degrees to 31.5 degrees.
For IMU alignments, the flight crew member at the aft flight deck station maneuvers the orbiter using the COAS at the right overhead negative Z window until the selected star is in the field of view. The crew member continues maneuvering the orbiter until the star crosses the center of the reticle.
At the instant of crossing, the crew member makes a mark, which means he depresses the "att ref" (attitude reference) push button. At the time of the mark, software stores the gimbal angles of the three IMUs. The mark can be taken again if it is felt the star was not centered as well as it could have been. When the crew member feels a good mark was taken, the software is notified to accept it. Good marks for two stars are required for an IMU alignment.
By knowing the star being sighted and the COAS location and mounting relationship in the orbiter, software can determine a line-of-sight vector from the COAS to the star in an inertial coordinate system. Line-of-sight vectors to two stars define the attitude of the orbiter in inertial space. This attitude can be compared to the attitude defined by the IMUs, and if the IMUs are in error, they can be realigned to the more correct orientation by the COAS sightings.
The COAS requires 115-volt ac power for reticle illumination. The COAS is 9.5 by 6 by 4.3 inches and weighs 2.5 pounds.
The 10-by-40 binoculars are a space-modified version of the commercial Leitz Trinovid binocular noted especially for its small size, high \Jmagnification\j, wide field of view, and rugged sealed construction. The 7-by-35 binoculars are noted for close focal distance at high \Jmagnification\j. The 14-by-40 gyrostabilized binoculars contain a gyrostabilized system that enhances target acquisition and retention.
When the crew member is subjected to ambient vibrations or hand tremor while using the gyrostabilized binoculars, the target image remains clear and stable. The gyrostabilized binoculars are electrically powered by six alkaline-type AA batteries and will operate continuously up to three hours on one battery pack.
Adjustable mirrors are installed before launch on handholds located between windows 2 and 3 for the commander and windows 4 and 5 for the pilot. During ascent and entry, the commander and pilot use the adjustable mirrors to better see controls that are in obscured areas of their vision. On orbit, the mirrors can be removed and stowed if desired. Each mirror is approximately 3 by 5 inches and weighs approximately 1 pound.
The spotlight is a high-intensity, hand-held flashlight powered by a battery pack consisting of five 1.2-volt one-half D size nickel-cadmium batteries. The spotlight produces a 20,000-candlepower output with a continuous running time of 1.5 hours. The lamp is a 6-volt \Jtungsten\j filament and cannot be replaced in flight. A spare battery pack is available on board.
For those crew members requiring them, two pairs of eyeglasses are available on board.
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"Shuttle Microcassette Recorder",301,0,0,0
The microcassette recorder is flown primarily for voice recording of data but may also be used to play prerecorded tapes. A microcassette tape has a recording time of 30 minutes per side. It is powered by two 1.5-volt AAA alkaline batteries.
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"Shuttle Photographic Equipment",302,0,0,0
Three camera systems-16mm, 35mm and 70mm-are used by the flight crew to document activities inside and outside the orbiter. All three camera systems are used to document on-orbit operations. The 16mm camera is also used during the launch and landing phases of the flight.
The 16mm camera is like a motion picture camera with independent shutter speeds and frame rates. The camera can be operated in one of three modes: pulse, cine, or time exposure. In the pulse mode, the camera operates at a continuous frame rate of two, six or 12 frames per second. In the cine mode, the camera operates at 24 frames per second. In the time exposure mode, the first switch actuation opens the shutter and the second actuation closes it. The camera uses 140-foot film magazines and has 5mm, 10mm and 18mm lenses.
The 35mm camera is a motorized, battery-operated Nikon camera with reflex viewing, through-the-lens coupled light metering and automatic film advancement. The camera has the standard manual operation and three automatic (electrically controlled) modes-single exposure, continuous and time. It uses an f/1.4 lens.
The 70mm camera system is a modified battery-powered, motor-driven, single-reflex Hasselblad camera that has 80mm and 250mm lenses and film magazines. Each magazine contains approximately 80 exposures. This camera has only one mode of operation, automatic; however, there are five automatic-type camera functions from which to select. The camera has a fixed viewfinder for through-the-lens viewing.
Interdeck light shades are provided to minimize light leakage between the flight deck and middeck during in-cabin photography. The light shade is attached with Velcro to the middeck ceiling around the interdeck access. Adjustable louvers are provided to regulate the amount of light between the flight deck and middeck.
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"Shuttle Wicket Tabs",303,0,0,0
Wicket tabs are devices that help the crew member activate controls when his vision is degraded. The tabs provide the crew member with tactile cues to the location of controls to be activated as well as a memory aid to their function, sequence of activation and other pertinent information. Wicket tabs are found on controls that are difficult to see during the ascent and entry flight phases on panels O8, C3 and R2.
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"Shuttle Reach Aid",304,0,0,0
The reach aid, sometimes known as the ''swizzle stick,'' is a short adjustable bar with a multipurpose end effector that is used to actuate controls that are out of the reach of seated crew members. The reach aid is used to push in and pull out circuit breakers and move toggle switches. It may be used during any phase of flight, but is not recommended for use during ascent because of the attenuation and switch-cueing difficulties resulting from acceleration forces.
Operation of the reach aid consists of extending it and actuating controls with the end effector. To extend the reach aid, one depresses the spring-loaded extension tab and pulls the end effector out to the desired length.
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"Shuttle Restraints and Mobility Aids",305,0,0,0
Restraints and mobility aids are provided in the orbiter to enable the flight crew to perform all tasks safely and efficiently during ingress (1-g, orbiter vertical), egress (1-g, orbiter horizontal) and orbital flight (orbiter orientation arbitrary). Restraints and mobility aids consist of foot loop restraints, the airlock foot restraint platform and the work/dining table. In-flight restraints consist of temporary stowage bags, Velcro, tape, snaps, cable restraints, clips, bungees and tethers.
Mobility aids and devices consist of handholds, footholds, handrails, ladders and the ingress-egress platform.
Foot loop restraints are cloth loops attached to the crew compartment decks by adhesive to secure crew members to the deck. Before launch, the foot loop restraints are installed on the floor areas of the aft flight deck work stations, middeck lockers, waste collection system and galley (if installed).
Spares will be stowed in the modular lockers. To install a foot restraint, the protective backing on the underside of the restraint is removed and the restraint is placed in its desired location. The foot loop restraints are easily used by placing one or both feet in the loop.
The temporary stowage bag is used to restrain, stow or transport loose equipment temporarily. It is snapped or attached with Velcro to the crew station standard Velcro and snap patterns.
Mobility aids and devices are located in the crew compartment for movement of the flight crew members during ingress, egress and orbital flight. These devices consist of handholds for ingress and egress to and from crew seats in the launch and landing configuration, handholds in the primary interdeck access opening for ingress and egress in the launch and landing configuration, a platform in the middeck for ingress and egress to and from the middeck when the orbiter is in the launch configuration, and an interdeck access ladder to enter the flight deck from the middeck in the launch configuration and go from the flight deck to the middeck in the launch and landing configuration.
The flight data file is a flight reference data file that is readily available to crew members aboard the orbiter. It consists of the onboard complement of documentation and related crew aids and includes documentation, such as procedural checklists (normal, backup and emergency procedures), malfunction procedures, crew activity plans, schematics, photographs, cue cards, star charts, Earth maps and crew notebooks; FDF stowage containers; and FDF ancillary equipment, such as tethers, clips, tape and erasers.
Four permanently mounted containers are located to the left and right side of the commander's and pilot's seats for stowing FDFs on the flight deck. The remaining FDF items are stowed in a middeck modular stowage locker.
The flight data file quantity and stowage locations are similar for all flights. The baseline stowage volume is sufficient to contain all FDF items for all orbiter configurations except the pallet-mounted payload. In this case, a larger flight data file and, consequently, additional locker space are required because all payload operations are performed in the orbiter.
FDF items are used throughout the flight-from prelaunch use of the ascent checklist through crew use of the entry checklist.
Flight data files are packaged and stowed on an individual flight basis. FDF items will be stowed in five types of stowage containers: lockers, the flight deck module, the commander's and pilot's seat-back FDF assemblies, the middeck FDF assembly and the map bag. The portable containers are stowed in a middeck modular locker for launch and entry.
If the flight carries a Spacelab module, all Spacelab books are stowed for launch in a portable container on the middeck and transferred in flight to the Spacelab. The FDF stowage is flexible and easily accessible.
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"Shuttle Crew Equipment Stowage",306,0,0,0
Crew equipment on board the orbiter is stowed in lockers with insertable trays. The trays can be adapted to accommodate a wide variety of soft goods, loose equipment and food. The lockers are interchangeable and attach to the orbiter with crew fittings. The lockers can be removed or installed in flight by the crew members. There are two sizes of trays: a half-size tray (two of which fit inside a locker) and a full-size tray. Approximately 150 cubic feet of stowage space is available, almost 95 percent of it on the middeck.
The lockers are made of either epoxy- or polyimide-coated Kevlar honeycomb material joined at the corners with aluminum channels. Inside dimensions are approximately 10 by 17 by 20 inches. The honeycomb material is approximately 0.25 of an inch thick and was chosen for its strength and light weight. The lockers contain about 2 cubic feet of space and can hold up to 60 pounds.
Dividers are used in the trays to provide a friction fit for zero-g retention. This will reduce the necessity for the straps, bags, Velcro snaps and other cumbersome attach devices previously used. Soft containers will be used in orbiter spaces too small for the fixed lockers.
The trays are packed with gear so that no item covers another type of gear. This method of packing will reduce the confusion usually associated with finding loose equipment and maintaining a record of the equipment.
Stowage areas in the orbiter crew compartment are located in the forward flight deck, the aft flight deck, the middeck, the equipment bay and the airlock module.
In the aft flight deck, stowage lockers are located below the rear payload control panels in the center of the deck. Container modules can be mounted to the right and left of the payload control station. Since these side containers are interchangeable, they may not be carried on every mission, depending on any payload-unique installed electronic gear.
In the middeck, container modules can be inserted in the forward avionics bay. Provisions for 42 containers are available in this area. In addition, there is an area to the right side of the airlock module where nine containers can be attached.
Harness stowage bags stowed in a middeck stowage locker or airlock are used on orbit to stow flight crew members' launch equipment, such as helmets, harnesses, boots and waste/trash materials.
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"Shuttle Exercise Equipment",307,0,0,0
The only exercise equipment presently being flown is a treadmill. The exact stowage location in the crew compartment middeck for launch, orbit and entry depends on the mission.
The treadmill is used with a restraint system to allow a crew member to run or jog in orbit. The treadmill kit is stowed on top of the treadmill and contains the waist belt, two shoulder straps, four extender hooks and a physiological monitor. The treadmill kit is restrained by four force cords that are used to restrain the body during exercise. The treadmill attaches to four middeck quick disconnects.
The quick disconnects contain several metal hooks that are hinged within the quick disconnect and actuated by the knurled lock ring. To release the quick disconnects, the lock push button is depressed and the knurled lock ring is pushed up, releasing the metal hooks. When the lock ring is pushed down, the metal hooks converge and capture the top of the middeck stud.
The treadmill has a speed control knob, which controls a rapid onset braking system. When the preset speed is reached, the brake engages and produces increased drag on the running track.
The physiological monitor provides heart rate, the time run and the distance run. The heart rate is determined by an ear clip, which has an infrared sensor that detects increased blood flow (pulses) in the ear lobe. Distance run is determined by connecting a mechanical sensor wire on the side of the treadmill to the physiological monitor.
The mechanical sensor detects the number of revolutions of the track and sends an electrical signal to the physiological monitor, where the distance is computed and shown on the display along with the heart rate. The monitor is stowed on the treadmill handle while the crew member runs.
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"Shuttle Sound Level Meter",308,0,0,0
The sound level meter is provided to determine on-orbit acoustical noise levels in the cabin. Depending on the requirements for each flight, the flight crew is required to take meter readings at specified crew compartment and equipment locations. The data obtained by the flight crew is logged and/or voice recorded. The meter is operated by four 1.5-volt batteries.
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"Shuttle Air Sampling System",309,0,0,0
The air sampling system consists of air bottles that are stowed in a modular locker. They are removed for sampling and restowed for entry.
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"Shuttle Main Propulsion System",310,0,0,0
The main propulsion system, assisted by the two solid rocket boosters during the initial phases of the ascent trajectory, provides the velocity increment from lift-off to a predetermined velocity increment before orbit insertion. The two SRBs are jettisoned after their fuel has been expended, but the MPS continues to thrust until the predetermined velocity is achieved. At that time, main engine cutoff is initiated.
The external tank is jettisoned, and the orbital maneuvering system is ignited to provide the final velocity increment for orbital insertion. The magnitude of the velocity increment supplied by the OMS depends on payload weight, mission trajectory and system limitations.
Coincident with the start of the OMS thrusting maneuver (which settles the MPS propellants), the remaining liquid oxygen propellant in the orbiter feed system and space shuttle main engines is dumped through the nozzles of the three SSMEs. At the same time, the remaining liquid \Jhydrogen\j propellant in the orbiter feed system and SSMEs is dumped overboard through the \Jhydrogen\j fill and drain valves for six seconds.
Then the \Jhydrogen\j inboard fill and drain valve is closed, and the \Jhydrogen\j recirculation valve is opened, continuing the dump. The \Jhydrogen\j flows through the engine \Jhydrogen\j bleed valves to the orbiter \Jhydrogen\j MPS line between the inboard and outboard \Jhydrogen\j fill and drain valves, and the remaining \Jhydrogen\j is dumped through the outboard fill and drain valve for approximately 120 seconds.
During on-orbit operations, the flight crew vacuum inerts the MPS by opening the liquid oxygen and liquid \Jhydrogen\j fill and drain valves, which allows the remaining propellants to be vented to space.
Before entry, the flight crew repressurizes the MPS propellant lines with \Jhelium\j to prevent contaminants from being drawn into the lines during entry and to maintain internal positive pressure. MPS \Jhelium\j is also used to purge the \Jspacecraft\j's aft fuselage. The last activity involving the MPS occurs at the end of the landing rollout. At that time, the \Jhelium\j remaining in onboard \Jhelium\j storage tanks is released into the MPS to provide an inert atmosphere for safety.
The MPS consists of the following major subsystems: three SSMEs, three SSME controllers, the external tank, the orbiter MPS propellant management subsystem and \Jhelium\j subsystem, four ascent thrust vector control units, and six SSME hydraulic servoactuators.
The main engines are reusable, high-performance, liquid-propellant rocket engines with variable thrust. The propellant fuel is liquid \Jhydrogen\j and the oxidizer is liquid oxygen. The propellant is carried in separate tanks in the external tank and supplied to the main engines under pressure. Each engine can be gimbaled plus or minus 10.5 degrees in the yaw axis and plus or minus 10.5 degrees in the pitch axis for thrust vector control by hydraulically powered gimbal actuators.
The main engines can be throttled over a range of 65 to 109 percent of their rated power level in 1-percent increments. A value of 100 percent corresponds to a thrust level of 375,000 pounds at sea level and 470,000 pounds in a vacuum. A value of 104 percent corresponds to 393,800 pounds at sea level and 488,800 pounds in a vacuum; 109 percent corresponds to 417,300 pounds at sea level and 513,250 pounds in a vacuum.
At sea level, the engine throttling range is reduced due to flow separation in the nozzle, prohibiting operation of the engine at its 65-percent throttle setting, referred to as minimum power level. All three main engines receive the same throttle command at the same time. Normally, these come automatically from the orbiter general-purpose computers through the engine controllers. During certain contingency situations, manual control of engine throttling is possible through the speed brake/thrust controller handle. The throttling ability reduces vehicle loads during maximum aerodynamic pressure and limits vehicle acceleration to 3 g's maximum during boost.
Each engine is designed for 7.5 hours of operation over a life span of 55 starts. Throughout the throttling range, the ratio of the liquid oxygen-liquid \Jhydrogen\j mixture is 6-to-1. Each nozzle area ratio is 77.5-to-1. The engines are 14 feet long and 7.5 feet in diameter at the nozzle exit.
The SSME controllers are digital, computer system, electronic packages mounted on the SSMEs. They operate in conjunction with engine sensors, valve actuators and spark igniters to provide a self-contained system for monitoring engine control, checkout and status. Each controller is attached to the forward end of the SSME.
Engine data and status collected by each controller are transmitted to the engine interface unit, which is mounted in the orbiter. There is one EIU for each main engine. The EIU transmits commands from the orbiter GPCs to the main engine controller. When engine data and status are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the computers.
Three orbiter hydraulic systems provide hydraulic pressure to position the SSME servoactuators for thrust vector control during the ascent phase of the mission in addition to performing other functions in the main propulsion system. The three orbiter auxiliary power units provide mechanical shaft power through a gear train to drive the hydraulic pumps that provide hydraulic pressure to their respective hydraulic systems.
The ascent thrust vector control units receive commands from the orbiter GPCs and send commands to the engine gimbal actuators. The units are \Jelectronics\j packages (four in all) mounted in the orbiter's aft fuselage avionics bays. Hydraulic isolation commands are directed to engine gimbal actuators that indicate faulty servovalve position. In conjunction with this, a servovalve isolation signal is transmitted to the computers.
The SSME hydraulic servoactuators are used to gimbal the main engine. There are two actuators per engine, one for pitch motion and one for yaw motion. They convert electrical commands received from the orbiter GPCs and position servovalves, which direct hydraulic pressure to a piston that converts the pressure into a mechanical force that is used to gimbal the SSMEs. The hydraulic pressure status of each servovalve is transmitted to the ATVC units.
The orbiter MPS propellant management subsystem consists of the manifolds, distribution lines and valves by which the liquid propellants pass from the external tank to the main engines and the gaseous propellants pass from the main engines to the external tank. The SSMEs' gaseous propellants are used to pressurize the external tank. All the valves in the propellant management subsystem are under direct control of the orbiter GPCs and are either electrically or pneumatically actuated.
The orbiter MPS \Jhelium\j subsystem consists of a series of \Jhelium\j supply tanks and regulators, check valves, distribution lines and control valves. The subsystem supplies the \Jhelium\j used within the engine to purge the high-pressure oxidizer turbopump intermediate seal and preburner oxidizer domes and to actuate valves during emergency pneumatic shutdown. The balance of the \Jhelium\j is used to actuate all the pneumatically operated valves within the propellant management subsystem and to pressurize the propellant lines before re-entry.
#
"Orbiter Main Propulsion System Helium Subsystem",311,0,0,0
The MPS \Jhelium\j subsystem consists of seven 4.7-cubic-foot \Jhelium\j supply tanks; three 17.3-cubic-foot \Jhelium\j supply tanks; and associated regulators, check valves, distribution lines and control valves. Four of the 4.7-cubic-foot \Jhelium\j supply tanks are located in the aft fuselage, and the other three are located below the payload bay liner in the midfuselage in the area originally reserved for the cryogenic storage tanks of the power reactant storage and distribution system. The three 17.3-cubic-foot \Jhelium\j supply tanks are also located below the payload bay liner in the midfuselage.
The tanks are composite structures consisting of a \Jtitanium\j liner with a fiberglass structural overwrap. The large tanks are 40.3 inches in diameter and have a dry weight of 272 pounds. The smaller tanks are 26 inches in diameter and have a dry weight of 73 pounds. The tanks are serviced before lift-off to a pressure of 4,500 psi.
Each of the larger supply tanks is plumbed to two of the smaller supply tanks (one in the midbody, the other in the aft body), forming three sets of three tanks for the engine \Jhelium\j pneumatic supply system. Each set of tanks normally provides \Jhelium\j to only one engine and is commonly referred to as left, center, or right engine \Jhelium\j, depending on the engine serviced. Each set normally provides \Jhelium\j to its designated engine for in-flight purges and provides pressure for actuating engine valves during emergency pneumatic shutdown.
The remaining 4.7-cubic-foot \Jhelium\j tank is referred to as the pneumatic \Jhelium\j supply tank. It normally provides pressure to actuate all of the pneumatically operated valves in the propellant management subsystem.
There are eight \Jhelium\j supply tank isolation valves grouped in pairs. One pair of valves is connected to each engine \Jhelium\j supply tank cluster, and one pair is connected to the pneumatic supply tank. In the engine \Jhelium\j supply tank system, each pair of isolation valves is connected in parallel, with each valve in the pair controlling \Jhelium\j flow through one leg of a dual-redundant \Jhelium\j supply circuit. Each \Jhelium\j supply circuit contains two check valves, a filter, an isolation valve, a regulator and a relief valve.
The two isolation valves connected to the pneumatic supply tanks are also connected in parallel; however, the rest of the pneumatic supply system consists of a filter, the two isolation valves, a regulator, a relief valve and a single check valve. Each engine \Jhelium\j supply isolation valve can be individually controlled by its He isolation A left , ctr , right open , GPC , close and He isolation B left , ctr , right , open , GPC, close switches on panel R2. The two pneumatic \Jhelium\j supply isolation valves are controlled by a single pneumatic He isol , open, GPC, close switch on panel R2.
All of the valves in the \Jhelium\j subsystem (with the exception of the supply tank isolation valves) are spring loaded to one position and electrically actuated to the other position. The supply tank isolation valves are spring loaded to the closed position and pneumatically actuated to the open position. Valve position is controlled via electrical signals from either the onboard GPCs or manually by the flight crew. All of the valves can be controlled automatically by the GPCs, and the flight crew can control some of the valves.
The \Jhelium\j source pressure of the pneumatic, left, center and right supply systems can be monitored on the \Jhelium\j , pneu , l (left), c (center), r (right) meters on panel F7 by positioning the tank, reg (regulator) switch below the meters to tank . In addition, the regulated pressure of the pneumatic, left, center and right systems can be monitored on the same meters by placing the switch to reg .
Each of the four \Jhelium\j supply systems operates independently until after main engine cutoff. Each engine \Jhelium\j supply has two interconnect (crossover) valves associated with it, and each valve in the pair of interconnect valves is connected in series with a check valve. The check valves allow \Jhelium\j to flow through the interconnect valves in one direction only.
One check valve associated with one interconnect valve controls \Jhelium\j flow in one direction, and the other interconnect valve and its associated check valve permit \Jhelium\j flow in the opposite direction. The in interconnect valve controls \Jhelium\j flow into the associated engine \Jhelium\j distribution system from the pneumatic \Jhelium\j supply tank. The out interconnect valve controls \Jhelium\j flow out of the associated engine \Jhelium\j supply system to the pneumatic distribution system.
Each pair of interconnect valves is controlled by a single switch on panel R2. Each He interconnect , left , ctr , right switch has three positions- in open/out close , GPC , and in close/out open. With the switch in the in open/out close position, the in interconnect valve is open and the out interconnect valve is closed. The in close/out open position does the reverse. With the switch in GPC, the out interconnect valve opens automatically at the beginning of the liquid oxygen dump and closes automatically at the end of the liquid \Jhydrogen\j dump.
In a return-to-launch-site abort, the GPC position will cause the in interconnect valve to open automatically at MECO and close automatically 20 seconds later. If an engine was shut down before MECO, its in interconnect valve will remain closed at MECO. At any other time, placing the switch in GPC results in both interconnect valves being closed.
An additional interconnect valve between the left engine \Jhelium\j supply and pneumatic \Jhelium\j supply would be used if the pneumatic \Jhelium\j supply regulator failed. This crossover valve would be opened and the pneumatic \Jhelium\j supply tank isolation valves would be closed, allowing the left engine \Jhelium\j supply system to supply \Jhelium\j to the pneumatic \Jhelium\j supply. The crossover \Jhelium\j valve is controlled by its own three-position switch on panel R2. The pneumatics l (left) eng He xovr (crossover) switch positions are open, GPC and close. The GPC position allows the valve to be controlled by the flight software loaded in the GPCs.
Manifold pressurization valves located downstream of the pneumatic \Jhelium\j pressure regulator are used to control the flow of \Jhelium\j to propellant manifolds during a nominal propellant dump and manifold repressurization. There are four of these valves grouped in pairs. One pair controls \Jhelium\j pressure to the liquid oxygen propellant manifolds, and the other pair controls \Jhelium\j pressure to the liquid \Jhydrogen\j propellant manifold.
The liquid \Jhydrogen\j RTLS dump pressurization valves located downstream of the pneumatic \Jhelium\j pressure regulator are used to control the pressurization of the liquid \Jhydrogen\j propellant manifolds during an RTLS liquid \Jhydrogen\j dump. There are two of these valves connected in series.
Unlike the liquid \Jhydrogen\j manifold pressurization valves, the liquid \Jhydrogen\j RTLS dump pressurization valves cannot be controlled by flight deck switches. During an RTLS abort, these valves are opened and closed automatically by GPC commands. An additional difference between the nominal and the RTLS liquid \Jhydrogen\j dumps is in the routing of the \Jhelium\j and the place where it enters the liquid \Jhydrogen\j feed line manifold.
For the nominal liquid \Jhydrogen\j dump, \Jhelium\j passes through the liquid \Jhydrogen\j manifold pressurization valves and enters the feed line manifold in the vicinity of the liquid \Jhydrogen\j feed line disconnect valve. For the liquid \Jhydrogen\j RTLS dump, \Jhelium\j passes through the RTLS liquid \Jhydrogen\j dump pressurization valves and enters the feed line manifold in the vicinity of the liquid \Jhydrogen\j inboard fill and drain valve on the inboard side. There is no RTLS liquid oxygen dump pressurization valve since the liquid oxygen manifold is not pressurized during the RTLS liquid oxygen dump.
Each engine \Jhelium\j supply tank has two pressure regulators operating in parallel. Each regulator controls pressure in one leg of a dual-redundant \Jhelium\j supply circuit and is capable of providing all of the \Jhelium\j needed by the main engines.
The pressure regulator for the pneumatic \Jhelium\j supply system is not redundant and is set to provide outlet pressure between 715 to 770 psig. Downstream of the regulator are two more regulators: the liquid \Jhydrogen\j manifold pressure regulator and the liquid oxygen manifold pressure regulator. These regulators are used only during MPS propellant dumps and manifold pressurization. Both regulators are set to provide outlet pressure between 20 to 25 psig. Flow through the regulators is controlled by the appropriate set of two normally closed manifold pressurization valves.
Downstream of each pressure regulator, with the exception of the two manifold repressurization regulators, is a relief valve. The valve protects the downstream \Jhelium\j distribution lines from overpressurization if the associated regulator fails fully open. The two relief valves in each engine \Jhelium\j supply are set to relieve at 785 to 850 psig and reseat at 785 psig. The relief valve in the pneumatic \Jhelium\j supply circuit also relieves at 785 to 850 psig and reseats at 785 psig.
There is one pneumatic control assembly on each of the three space shuttle main engines. The PCA is essentially a manifold pressurized by one of the engine \Jhelium\j supply systems and contains \Jsolenoid\j valves to control and direct pressure to perform various essential functions. The valves are energized by discrete on/off commands from the output \Jelectronics\j of the associated SSME controller. Functions controlled by the PCA include the high-pressure oxidizer turbopump intermediate seal cavity and preburner oxidizer dome purge, pogo system postcharge and pneumatic shutdown.
#
"Shuttle Main Propulsion System Propellant Management Subsystem",312,0,0,0
Within the orbiter aft fuselage, liquid \Jhydrogen\j and liquid oxygen pass through the manifolds, distribution lines and valves of the propellant management subsystem.
During prelaunch activities, this subsystem is used to control the loading of liquid oxygen and liquid \Jhydrogen\j in the external tank. During SSME thrusting periods, propellants from the external tank flow into this subsystem and to the three SSMEs. The subsystem also provides a path that allows gases tapped from the three SSMEs to flow back to the external tank through two gas umbilicals to maintain pressure in the external tank's liquid oxygen and liquid \Jhydrogen\j tanks. After MECO, this subsystem controls MPS dumps, vacuum inerting and MPS repressurization for entry.
All the valves in the MPS are either electrically or pneumatically operated. Pneumatic valves are used where large loads are encountered, such as in the control of liquid propellant flows. Electrical valves are used for lighter loads, such as in the control of gaseous propellant flows.
The pneumatically actuated valves are divided into two types: those that require pneumatic pressure to open and close the valve (type 1) and those that are spring loaded to one position and require pneumatic pressure to move to the other position (type 2).
Each type 1 valve actuator is equipped with two electrically actuated \Jsolenoid\j valves. Each \Jsolenoid\j valve controls \Jhelium\j pressure to an ''open'' or ''close'' port on the actuator. Energizing the \Jsolenoid\j valve on the open port allows \Jhelium\j pressure to open the pneumatic valve.
Energizing the \Jsolenoid\j on the close port allows \Jhelium\j pressure to close the pneumatic valve. Removing power from a \Jsolenoid\j valve removes \Jhelium\j pressure from the corresponding port of the pneumatic actuator and allows the \Jhelium\j pressure trapped in that side of the actuator to vent overboard.
Removing power from both solenoids allows the pneumatic valve to remain in the last commanded position. This type of valve is used for the liquid oxygen and liquid \Jhydrogen\j feed line 17-inch umbilical disconnect valves (two), the liquid oxygen and liquid \Jhydrogen\j prevalves (six), the three liquid \Jhydrogen\j and liquid oxygen inboard and outboard fill and drain valves (four), and the liquid \Jhydrogen\j 4-inch recirculation disconnect valves.
Each type 2 valve is a single electrically actuated \Jsolenoid\j valve that controls \Jhelium\j pressure to either an open or a close port on the actuator. Removing power from the \Jsolenoid\j valve removes \Jhelium\j pressure from the corresponding port of the pneumatic actuator and allows \Jhelium\j pressure trapped in that side of the actuator to vent overboard. Spring force takes over and drives the valve to the opposite position.
If the spring force drives the valve to the open position, the valve is referred to as a normally open valve. If the spring force drives the valve to a closed position, the valve is referred to as a normally closed valve.
This type of valve is used for the liquid \Jhydrogen\j RTLS inboard dump valve (NC), the liquid \Jhydrogen\j RTLS outboard dump valve (NC), the liquid \Jhydrogen\j feed line relief shutoff valve (NO), the liquid oxygen feed line relief shutoff valve (NO), the three liquid \Jhydrogen\j engine recirculation valves (NC), the two liquid oxygen pogo recirculation valves (NO), the liquid \Jhydrogen\j topping valve (NC), the liquid \Jhydrogen\j high-point bleed valve (NC), and the liquid oxygen overboard bleed valve (NO).
The electrically actuated \Jsolenoid\j valves are spring loaded to one position and move to the other position when electrical power is applied. These valves also are referred to as either normally open or normally closed, based on their position in the de-energized state. Electrically actuated \Jsolenoid\j valves are the gaseous \Jhydrogen\j pressurization line vent valve (NC), the three gaseous \Jhydrogen\j pressurization flow control valves (NO) and the three gaseous oxygen pressurization flow control valves (NO).
There are two 17-inch-diameter MPS propellant feed line manifolds in the orbiter aft fuselage, one for liquid oxygen and one for liquid \Jhydrogen\j. Each manifold has an outboard and inboard fill and drain valve in series that interface with the respective port (left) and starboard (right) T-0 umbilical. The port T-0 umbilical is for liquid \Jhydrogen\j; the starboard, for liquid oxygen. In addition, each manifold connects the orbiter to the external tank in the lower aft fuselage through a port 17-inch liquid \Jhydrogen\j disconnect valve umbilical and a starboard 17-inch liquid oxygen disconnect valve umbilical.
There are three outlets in both the liquid oxygen and liquid \Jhydrogen\j 17-inch manifolds between the orbiter-external tank 17-inch umbilical disconnect valves and the inboard fill and drain valve. The outlets in the manifolds provide liquid oxygen and liquid \Jhydrogen\j to each SSME in 12-inch-diameter feed lines.
Each 17-inch liquid \Jhydrogen\j and liquid oxygen manifold has a 1-inch-diameter line that is routed to a feed line relief isolation valve and feed line relief valve in the respective liquid \Jhydrogen\j and liquid oxygen system. The LO 2 and LH 2 feed line rlf (relief) isol (isolation) switches on panel R4 have open , GPC and close positions. When a feed line relief isolation valve is opened, the corresponding manifold can relieve excessive pressure overboard through its relief valve.
The liquid \Jhydrogen\j feed line manifold has another outlet directed to the two liquid \Jhydrogen\j RTLS dump valves in series. Both valves are controlled by the MPS prplt dump LH 2 valve switch on panel R2, which has backup LH 2 vlv open , GPC , close positions. When opened, these valves enable the liquid \Jhydrogen\j dump during RTLS aborts or provide a backup to the normal liquid \Jhydrogen\j dump after a nominal main engine cutoff. In an RTLS abort dump, liquid \Jhydrogen\j is dumped overboard through a port at the outer aft fuselage's left side between the orbital maneuvering system/reaction control system pod and the upper surface of the wing.
The MPS propellant management subsystem also contains two 2-inch-diameter manifolds, one for gaseous oxygen and one for gaseous \Jhydrogen\j. Each manifold individually permits ground support equipment servicing with \Jhelium\j through the respective T-0 umbilical and provides initial pressurization of the external tank's liquid oxygen and liquid \Jhydrogen\j orbiter/external tank disconnect umbilicals. Self-sealing quick disconnects are provided at the T-0 umbilical and the orbiter/external tank umbilical.
Six 0.63-inch-diameter pressurization lines, three for gaseous oxygen and three for gaseous \Jhydrogen\j, are used after SSME start to pressurize the external tank's liquid oxygen and liquid \Jhydrogen\j tanks.
In each SSME, a small portion of liquid oxygen is diverted into the engine's oxidizer heat exchanger, and the heat generated by the engine's high-pressure oxidizer turbopump converts the liquid oxygen into gaseous oxygen and directs it through a check valve to two orifices and a flow control valve for each engine.
During SSME thrusting periods, liquid oxygen tank pressure is maintained between 20 and 22 psig by the orifices in the two lines and the action of the flow control valve from each SSME. The flow control valve is controlled by one of three liquid oxygen pressure transducers. When tank pressure decreases below 20 psig, the valve opens. If the tank pressure is greater than 24 psig, it is relieved through the liquid oxygen tank's vent and relief valve.
In each SSME, gaseous \Jhydrogen\j from the low-pressure fuel turbopump is directed through two check valves to two orifices and a flow control valve for each engine. During the main engine thrusting period, the liquid \Jhydrogen\j tank's pressure is maintained between 32 and 34 psia by the orifices and the action of the flow control valve from each SSME.
The flow control valve is controlled by one of three liquid \Jhydrogen\j pressure transducers. When tank pressure decreases below 32 psia, the valve opens; and when tank pressure increases to 33 psia, the valve closes. If the tank pressure is greater than 35 psia, the pressure is relieved through the liquid \Jhydrogen\j tank's vent and relief valve. If the pressure falls below 32 psia, the LH 2 ullage press switch on panel R2 is positioned from auto to open , which will cause all three flow control valves to go to full open and remain in the full-open position.
The single gaseous \Jhydrogen\j manifold repressurization line connects to the \Jhydrogen\j line vent valve, which is controlled by the H 2 press line vent switch on panel R4. This valve is normally closed, and the switch is positioned to open when vacuum inerting the gaseous \Jhydrogen\j pressurization lines after MECO and the liquid \Jhydrogen\j dump. The gnd position allows the launch processing system to control the valve during ground operations.
#
"Shuttle MPS External Tank",313,0,0,0
The external tank is attached to the orbiter at one forward and two aft attach points. At the two aft attach points are the two external tank/orbiter umbilicals for the fluid, gas, signal and electrical power connections between the orbiter and the external tank. Each external tank umbilical plate mates with a corresponding umbilical plate on the orbiter. The umbilical plates help maintain alignment of the various connecting components. The corresponding umbilical plates are bolted together; and when external tank separation is commanded, the bolts are severed by pyrotechnics.
At the forward end of each external tank propellant tank is a vent and relief valve that can be opened by GSE-supplied \Jhelium\j before launch for venting or by excessive tank pressure for relief. The vent function is available only before launch; after lift-off only the relief function is operable.
The liquid oxygen tank relieves at an ullage pressure of 25 psig, while the liquid \Jhydrogen\j tank relieves at an ullage pressure of 38 psi. The flight crew has no control over the position of the vent and relief valves before launch or during ascent. Normal range of the tank ullage pressure of the liquid \Jhydrogen\j tank during ascent is 32 to 39 psia.
During prelaunch activities, the liquid \Jhydrogen\j tank is pressurized to 44.1 psi to meet the start requirement of the main engine LPFT. The liquid oxygen and liquid \Jhydrogen\j tanks' ullage pressures are monitored on the panel F7 eng manf LO2 and LH2 meters as well as on a \Jcathode\j ray tube display.
In addition to the vent and relief valve, the liquid oxygen tank has a tumble vent valve that is opened during the external tank separation sequence. The thrust force provided by opening the valve imparts an angular velocity to the external tank to assist in the separation maneuver and provide more positive control of the external tank's re-entry aerodynamics.
There are eight propellant depletion sensors. Four of them sense fuel depletion and four sense oxidizer depletion. The oxidizer depletion sensors are mounted in the external tank's liquid oxygen feed line manifold downstream of the tank. The fuel depletion sensors are located in the liquid \Jhydrogen\j tank. During prelaunch activities, the launch processing system tests each propellant depletion sensor.
If any are found to be in a failed condition, the LPS sets a flag in the computer's SSME operational sequence, sequence logic that will instruct the computer to ignore the output of the failed sensor or sensors. During main engine thrusting, the computer constantly computes the instantaneous mass of the vehicle, which constantly decreases due to propellant usage from the external tank.
When the computed vehicle mass matches a predetermined initialized-loaded value, the computer arms the propellant depletion sensors. After this time, if any two of the good fuel depletion sensors (those not flagged before launch) or any two of the good oxidizer depletion sensors indicate a dry condition, the computers command main engine cutoff. This type of MECO is a backup to the nominal MECO, which is based on vehicle velocity.
The oxidizer sensors sense propellant depletion before the fuel sensors to ensure that all depletion cutoffs are fuel-rich since an oxidizer-rich cutoff can cause burning and severe erosion of engine components. To ensure that the oxidizer sensors sense depletion first, a plus 700-pound bias is included in the amount of liquid \Jhydrogen\j loaded in the external tank.
This amount is in excess of that dictated by the 6-1 ratio of oxidizer to fuel. The position of the oxidizer propellant depletion sensors allows the maximum amount of oxidizer to be consumed in the engines and allows sufficient time to cut off the engines before the oxidizer turbopumps cavitate (run dry).
Four ullage pressure transducers are located at the top end of each propellant tank (liquid oxygen and liquid hydrogen). One of the four is considered a spare and is normally off-line. Before launch, GSE normally checks out the four transducers; and if one of the three active transducers is determined to be bad, it can be taken off-line and the output of the spare \Jtransducer\j selected.
The flight crew can also perform this operation after lift-off via the computer keyboard; however, because of the time involved from lift-off to MECO, this would probably be impractical. The three active ullage pressure sensors provide outputs for CRT display and control of ullage pressure within their particular propellant tanks. For CRT display, computer processing selects the middle value output of the three transducers and displays this single value. For ullage pressure control, all three outputs are used.
The external tank/orbiter aft umbilicals have five propellant disconnects: two for the liquid oxygen tank and three for the liquid \Jhydrogen\j tank. One of the liquid oxygen propellant umbilicals carries liquid oxygen and the other carries gaseous oxygen.
The liquid \Jhydrogen\j tank has two disconnects that carry liquid \Jhydrogen\j and one that carries gaseous \Jhydrogen\j. The external tank liquid \Jhydrogen\j recirculation disconnect is the smaller of the two disconnects that carry liquid \Jhydrogen\j and is used only during the liquid \Jhydrogen\j chill-down sequence before launch.
In addition, the external tank/orbiter umbilicals contain two electrical umbilicals, each made of many smaller electrical cables. These cables carry electrical power from the orbiter to the external tank and the two solid rocket boosters and bring telemetry back to the orbiter from the SRBs and external tank.
The operational instrumentation telemetry that comes back from the SRBs is conditioned, digitized and multiplexed in the SRBs themselves. The external tank OI measurements that return to the orbiter are raw \Jtransducer\j outputs and must be processed within the orbiter telemetry system.
The external tank's liquid oxygen tank is serviced at the launch pad before prelaunch from ground support equipment through the starboard T-0 umbilical of the orbiter, the MPS outboard and inboard fill and drain valves, the MPS 17-inch liquid oxygen line, and the orbiter/external tank 17-inch umbilical disconnect valves.
Once the liquid oxygen is loaded and ready for main engine ignition, the liquid oxygen tank's vent and relief valve is closed, and the tank is pressurized to 21 psig by GSE-supplied \Jhelium\j. During SSME thrusting, liquid oxygen flows out of the external tank through the orbiter/external tank umbilical into the orbiter MPS and to each SSME. Pressurization in the tank is maintained by gaseous oxygen tapped from the three main engines and supplied to the liquid oxygen tank through the orbiter/external tank gaseous oxygen umbilical.
The external tank's liquid \Jhydrogen\j tank is serviced before launch from GSE at the launch pad similarly to the liquid oxygen tank but through the port T-0 umbilical and port orbiter/external tank umbilical. When the liquid \Jhydrogen\j is loaded and ready for main engine ignition, the liquid \Jhydrogen\j tank's vent and relief valve is closed, and the tank is pressurized to 42.5 psia by GSE-supplied \Jhelium\j.
Approximately 45 minutes after loading starts, three electrically powered liquid \Jhydrogen\j pumps in the orbiter begin to circulate the liquid \Jhydrogen\j in the external tank through the three SSMEs and back to the external tank through a special recirculation umbilical. This recirculation chills down the liquid \Jhydrogen\j lines between the external tank and the high-pressure fuel turbopump in the SSMEs so that the path is free of any gaseous \Jhydrogen\j bubbles and is at the proper temperature for engine start.
Recirculation ends approximately six seconds before engine start. During engine thrusting, liquid \Jhydrogen\j flows from the external tank and through the orbiter/external tank liquid \Jhydrogen\j umbilical into the orbiter MPS and to the main engines. Tank pressurization is maintained by gaseous \Jhydrogen\j tapped from the three SSMEs and supplied to the liquid \Jhydrogen\j tank through the orbiter/external tank gaseous \Jhydrogen\j umbilical.
#
"Space Shuttle Main Engines",314,0,0,0
Oxidizer from the external tank enters the orbiter at the orbiter/external tank umbilical disconnect and then the orbiter's main propulsion system liquid oxygen feed line. There it branches out into three parallel paths, one to each engine. In each branch, a liquid oxygen prevalve must be opened to permit flow to the low-pressure oxidizer turbopump.
The LPOT is an axial-flow pump driven by a six-stage \Jturbine\j powered by liquid oxygen. It boosts the liquid oxygen's pressure from 100 psia to 422 psia. The flow from the LPOT is supplied to the high-pressure oxidizer turbopump. During engine operation, the pressure boost permits the HPOT to operate at high speeds without cavitating. The LPOT operates at approximately 5,150 rpm. The LPOT, which is approximately 18 by 18 inches, is connected to the vehicle propellant ducting and supported in a fixed position by the orbiter structure.
The HPOT consists of two single-stage centrifugal pumps (a main pump and a preburner pump) mounted on a common shaft and driven by a two-stage, hot-gas \Jturbine\j. The main pump boosts the liquid oxygen's pressure from 422 psia to 4,300 psia while operating at approximately 28,120 rpm.
The HPOT discharge flow splits into several paths, one of which is routed to drive the LPOT \Jturbine\j. Another path is routed to and through the main oxidizer valve and enters into the main combustion chamber. Another small flow path is tapped off and sent to the oxidizer heat exchanger.
The liquid oxygen flows through an anti-flood valve that prevents it from entering the heat exchanger until sufficient heat is present to convert the liquid oxygen to gas. The heat exchanger utilizes the heat contained in the discharge gases from the HPOT \Jturbine\j to convert the liquid oxygen to gas.
The gas is sent to a manifold and is then routed to the external tank to pressurize the liquid oxygen tank. Another path enters the HPOT second-stage preburner pump to boost the liquid oxygen's pressure from 4,300 psia to 7,420 psia. It passes through the oxidizer preburner oxidizer valve into the oxidizer preburner and through the fuel preburner oxidizer valve into the fuel preburner. The HPOT is approximately 24 by 36 inches. It is attached by flanges to the hot-gas manifold.
Fuel enters the orbiter at the liquid \Jhydrogen\j feed line disconnect valve, then flows into the orbiter gaseous \Jhydrogen\j feed line manifold and branches out into three parallel paths to each engine. In each liquid \Jhydrogen\j branch, a prevalve permits liquid \Jhydrogen\j to flow to the low-pressure fuel turbopump when the prevalve is open.
The LPFT is an axial-flow pump driven by a two-stage \Jturbine\j powered by gaseous \Jhydrogen\j. It boosts the pressure of the liquid \Jhydrogen\j from 30 psia to 276 psia and supplies it to the high-pressure fuel turbopump. During engine operation, the pressure boost provided by the LPFT permits the HPFT to operate at high speeds without cavitating. The LPFT operates at approximately 16,185 rpm. The LPFT is approximately 18 by 24 inches. It is connected to the vehicle propellant ducting and is supported in a fixed position by the orbiter structure 180 degrees from the LPOT.
The HPFT is a three-stage centrifugal pump driven by a two-stage, hot-gas \Jturbine\j. It boosts the pressure of the liquid \Jhydrogen\j from 276 psia to 6,515 psia. The HPFT operates at approximately 35,360 rpm. The discharge flow from the turbopump is routed to and through the main valve and then splits into three flow paths. One path is through the jacket of the main combustion chamber, where the \Jhydrogen\j is used to cool the chamber walls. It is then routed from the main combustion chamber to the LPFT, where it is used to drive the LPFT \Jturbine\j. A small portion of the flow from the LPFT is then directed to a common manifold from all three engines to form a single path to the external tank to maintain liquid \Jhydrogen\j tank pressurization.
The remaining \Jhydrogen\j passes between the inner and outer walls to cool the hot-gas manifold and is discharged into the main combustion chamber. The second \Jhydrogen\j flow path from the main fuel valve is through the engine nozzle (to cool the nozzle). It then joins the third flow path from the chamber coolant valve. The combined flow is then directed to the fuel and oxidizer preburners. The HPFT is approximately 22 by 44 inches. It is attached by flanges to the hot-gas manifold.
The oxidizer and fuel preburners are welded to the hot-gas manifold. The fuel and oxidizer enter the preburners and are mixed so that efficient combustion can occur. The augmented spark igniter is a small combination chamber located in the center of the injector of each preburner.
The two dual-redundant spark igniters, which are activated by the engine controller, are used during the engine start sequence to initiate combustion in each preburner. They are turned off after approximately three seconds because the combustion process is then self-sustaining. The preburners produce the fuel-rich hot gas that passes through the turbines to generate the power to operate the high-pressure turbopumps. The oxidizer preburner's outflow drives a \Jturbine\j that is connected to the HPOT and the oxidizer preburner pump. The fuel preburner's outflow drives a \Jturbine\j that is connected to the HPFT.
The HPOT \Jturbine\j and HPOT pumps are mounted on a common shaft. Mixing of the fuel-rich hot gas in the \Jturbine\j section and the liquid oxygen in the main pump could create a hazard. To prevent this, the two sections are separated by a cavity that is continuously purged by the MPS engine \Jhelium\j supply during engine operation. Two seals minimize leakage into the cavity. One seal is located between the \Jturbine\j section and the cavity, and the other is between the pump section and cavity. Loss of \Jhelium\j pressure in this cavity results in an automatic engine shutdown.
The speed of the HPOT and HPFT turbines depends on the position of the corresponding oxidizer and fuel preburner oxidizer valves. These valves are positioned by the engine controller, which uses them to throttle the flow of liquid oxygen to the preburners and, thus, control engine thrust.
The oxidizer and fuel preburner oxidizer valves increase or decrease the liquid oxygen flow, thus increasing or decreasing preburner chamber pressure, HPOT and HPFT \Jturbine\j speed, and liquid oxygen and gaseous \Jhydrogen\j flow into the main combustion chamber, which increases or decreases engine thrust, thus throttling the engine. The oxidizer and fuel preburner valves operate together to throttle the engine and maintain a constant 6-1 propellant mixture ratio.
The main oxidizer valve and the main fuel valve control the flow of liquid oxygen and liquid \Jhydrogen\j into the engine and are controlled by each engine controller. When an engine is operating, the main valves are fully open.
A coolant control valve is mounted on the combustion chamber coolant bypass duct of each engine. The engine controller regulates the amount of gaseous \Jhydrogen\j allowed to bypass the nozzle coolant loop, thus controlling its temperature. The chamber coolant valve is 100 percent open before engine start. During engine operation, it will be 100 percent open for throttle settings of 100 to 109 percent for minimum cooling. For throttle settings between 65 to 100 percent, its position will range from 66.4 to 100 percent open for maximum cooling.
Each engine main combustion chamber receives fuel-rich hot gas from a hot-gas manifold cooling circuit. The gaseous \Jhydrogen\j and liquid oxygen enter the chamber at the injector, which mixes the propellants. A small augmented spark igniter chamber is located in the center of the injector.
The dual-redundant igniter is used during the engine start sequence to initiate combustion. The igniters are turned off after approximately three seconds because the combustion process is self-sustaining. The main injector and dome assembly is welded to the hot-gas manifold. The main combustion chamber also is bolted to the hot-gas manifold.
The inner surface of each combustion chamber, as well as the inner surface of each nozzle, is cooled by gaseous \Jhydrogen\j flowing through coolant passages. The nozzle assembly is a bell-shaped extension bolted to the main combustion chamber.
The nozzle is 113 inches long, and the outside diameter of the exit is 94 inches. A support ring welded to the forward end of the nozzle is the engine attach point to the orbiter-supplied heat shield. Thermal protection for the nozzles is necessary because of the exposure that portions of the nozzles experience during the launch, ascent, on-orbit and entry phases of a mission. The \Jinsulation\j consists of four layers of metallic batting covered with a metallic foil and screening.
The five propellant valves on each engine (oxidizer preburner oxidizer, fuel preburner oxidizer, main oxidizer, main fuel, and chamber coolant) are hydraulically actuated and controlled by electrical signals from the engine controller. They can be fully closed by using the MPS engine \Jhelium\j supply system as a backup actuation system.
The low-pressure oxygen and low-pressure fuel turbopumps are mounted 180 degrees apart on the orbiter's aft fuselage thrust structure. The lines from the low-pressure turbopumps to the high-pressure turbopumps contain flexible bellows that enable the low-pressure turbopumps to remain stationary while the rest of the engine is gimbaled for thrust vector control. The liquid \Jhydrogen\j line from the LPFT to the HPFT is insulated to prevent the formation of liquid air.
The main oxidizer valve and fuel bleed valve are used after shutdown. The main oxidizer valve is opened during a propellant dump to allow residual liquid oxygen to be dumped overboard through the engine, and the fuel bleed valve is opened to allow residual liquid \Jhydrogen\j to be dumped through the liquid \Jhydrogen\j fill and drain valves overboard. After the dump is completed, the valves close and remain closed for the remainder of the mission.
The gimbal bearing is bolted to the main injector and dome assembly and is the thrust interface between the engine and orbiter. The bearing assembly is approximately 11.3 by 14 inches.
Overall, a space shuttle main engine weighs approximately 7,000 pounds.
#
"Shuttle Pogo Suppression System",315,0,0,0
A pogo suppression system prevents the transmission of low-frequency flow oscillations into the high-pressure turbopump and, ultimately, prevents main combustion chamber pressure (engine thrust) \Joscillation\j. Flow oscillations transmitted from the space shuttle vehicle are suppressed by a partially filled gas accumulator, which is attached by flanges to the high-pressure oxidizer turbopump's inlet duct.
The system consists of a 0.6-cubic-foot accumulator with an internal standpipe, \Jhelium\j precharge valve package, gaseous oxygen supply valve package and two recirculation isolation valves (one located on the orbiter).
During engine start, the accumulator is charged with \Jhelium\j 2.4 seconds after the start command to provide pogo protection until the engine heat exchanger is operational and gaseous oxygen is available.
The accumulator is partially chilled by liquid oxygen during the engine chill-down operation. It fills to the overflow standpipe line inlet level, which is sufficient to preclude gas ingestion at engine start.
During engine operation, the accumulator is charged with a continuous gaseous oxygen flow maintained at a rate governed by the engine operation point.
The liquid level in the accumulator is controlled by the overflow standpipe line in the accumulator, which is orificed to regulate the gaseous oxygen overflow over the engine's operating power level. The system is sized to provide sufficient replenishment of gaseous oxygen at the minimum flow rate and to permit sufficient gaseous oxygen overflow at the maximum decreasing pressure transient in the low-pressure oxidizer turbopump discharge duct. At all other conditions, excess gaseous and liquid oxygen are recirculated to the low-pressure oxidizer turbopump inlet through the engine oxidizer bleed duct. The pogo accumulator is charged (pressurized) at engine shutdown to provide a positive pressure at the HPOT inlet, which prevents HPOT overspeed in the zero-gravity environment.
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"Space Shuttle Main Engine Controllers",316,0,0,0
The controller is an \Jelectronics\j package mounted on each SSME. It contains two digital computers and the associated \Jelectronics\j to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings.
Each controller operates in conjunction with engine sensors, valves, actuators and spark igniters to provide a self-contained system for engine control, checkout and monitoring. The controller provides engine flight readiness verification; engine start and shutdown sequencing; closed-loop thrust and propellant mixture ratio control; sensor excitation; valve actuator and spark igniter control signals; engine performance limit monitoring; onboard engine checkout, response to vehicle commands and transmission of engine status; and performance and maintenance data.
Each engine controller receives engine commands transmitted by the orbiter's general-purpose computers through its own engine interface unit. The engine controller provides its own commands to the main engine components. Engine data are sent to the engine controller, where they are stored in a vehicle data table in the controller's computer memory.
Data on the controller's status compiled by the engine controller's computer are also added to the vehicle data table. The vehicle data table is periodically output by the controller to the EIU for transmission to the orbiter's GPCs.
The engine interface unit is a specialized multiplexer/demultiplexer that interfaces with the GPCs and with the engine controller. When engine commands are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the GPCs. The EIU then sends data to each GPC. Each EIU is dedicated to one space shuttle main engine and communicates only with the engine controller that controls its SSME. The EIUs have no interface with each other.
The controller provides responsive control of engine thrust and propellant mixture ratio throughout the digital computer in the controller, updating the instructions to the engine control elements 50 times per second (every 20 milliseconds). Engine reliability is enhanced by a dual-redundant system that allows normal operation after the first failure and a fail-safe shutdown after a second failure. High-reliability electronic parts are used throughout the controller.
The digital computer is programmable, allowing engine control equations and constants to be modified by changing the stored program (software). The controller is packaged in a sealed, pressurized chassis and is cooled by convection heat transfer through pin fins as part of the main chassis. The \Jelectronics\j are distributed on functional modules with special thermal and vibration protection.
The controller is divided into five subsystems: input \Jelectronics\j, output \Jelectronics\j, computer interface \Jelectronics\j, digital computer and power supply \Jelectronics\j. Each subsystem is duplicated to provide dual-redundant capability.
The input \Jelectronics\j receive data from all engine sensors, condition the signals and convert them to digital values for processing by the digital computer. Engine control sensors are dual-redundant, and maintenance data sensors are non-redundant.
The output \Jelectronics\j convert computer digital control commands into voltages suitable for powering the engine spark igniters, the off/on valves and the engine propellant valve actuators.
The computer interface \Jelectronics\j control the flow of data within the controller, data input to the computer and computer output commands to the output \Jelectronics\j. They also provide the controller interface with the vehicle engine \Jelectronics\j interface unit for receiving engine commands that are triple-redundant channels from the vehicle and for transmitting engine status and data through dual-redundant channels to the vehicle. The computer interface \Jelectronics\j include the watchdog timers that determine which channel of the dual-redundant \Jmechanization\j is in control.
The digital computer is an internally stored, general-purpose computer that provides the computational capability necessary for all engine control functions. The memory has a program storage capacity of 16,384 data and instruction words (17-bit words; 16 bits for program use, one bit for parity).
The power supply \Jelectronics\j convert the 115-volt, three-phase, 400-hertz vehicle ac power to the individual power supply voltage levels required by the engine control system and monitor the level of power supply channel operation to ensure it is within satisfactory limits.
Each orbiter GPC, operating in a redundant set, issues engine commands to the engine interface units for transmission to their corresponding engine controllers. Each orbiter GPC has SSME subsystem operating program applications software residing in it. Engine commands are output over the engine's assigned flight-critical data bus (a total of four GPCs outputting over four FC data buses). Therefore, each EIU will receive four commands. The nominal ascent configuration has GPCs 1, 2, 3 and 4 outputting on FC data buses 5, 6, 7 and 8, respectively. Each FC data bus is connected to one multiplexer interface adapter in each EIU.
The EIU checks the received engine commands for transmission errors. If there are none, the EIU passes the validated engine commands on to the controller interface assemblies, which output the validated engine commands to the engine controller. An engine command that does not pass validation is not sent to the controller interface assembly. Instead, it is dead-ended in the EIU's multiplexer interface adapter. Commands that come through MIAs 1 and 2 are sent to CIAs 1 and 2, respectively.
Commands that come to MIAs 3 and 4 pass through a \JCIA\j 3 data-select logic. This logic outputs the command that arrives at the logic first, from either MIA 3 or 4. The other command is dead-ended in the \JCIA\j 3 select logic. The selected command is output through \JCIA\j 3. In this manner, the EIU reduces the four commands sent to the EIU to three commands output by the EIU.
The engine controller vehicle interface \Jelectronics\j receive the three engine commands output by its EIU, check for transmission errors (hardware validation), and send controller hardware-validated engine commands to the controller A and B \Jelectronics\j. Normally, channel A \Jelectronics\j are in control, with channel B \Jelectronics\j active, but not in control. If channel A fails, channel B will assume control.
If channel B subsequently fails, the engine controller will shut down the engine pneumatically. If two or three commands pass voting, the engine controller will issue its own commands to accomplish the function commanded by the orbiter GPCs. If command voting fails and two or all three commands fail, the engine controller will maintain the last command that passed voting.
The engine controller provides all the main engine data to the GPCs. Sensors in the engine supply pressures, temperatures, flow rates, turbopump speeds, valve position and engine servovalve actuator positions to the engine controller. The engine controller assembles these data into a vehicle data table and adds status data of its own to the vehicle data table.
The vehicle data tables output channels A and B to the vehicle interface \Jelectronics\j for transmission to the EIUs. The vehicle interface \Jelectronics\j output over both data paths. The data paths are called primary and secondary. The channel A vehicle data table is normally sent over both primary and secondary control (channel A has failed); then the vehicle interface \Jelectronics\j output the channel B vehicle data table over both the primary and secondary data paths.
The vehicle data table is sent by the controller to the EIU. There are only two data paths versus three command paths between the engine controller and the EIU. The data path that interfaces with \JCIA\j 1 is called primary data. The path that interfaces with \JCIA\j 2 is called secondary data. Primary and secondary data are held in buffers until the GPCs send a data request command to the EIUs. The GPCs request both primary and secondary data. Primary data is output only through MIA 1 on each EIU. Secondary data is output only through MIA 4 on each EIU.
During prelaunch, the orbiter's computers look at both primary and secondary data. Loss of either primary or secondary data will result in data path failure and either an engine ignition inhibit or a launch pad shutdown of all three main engines.
At T minus zero, the orbiter GPCs request both primary and secondary data from each EIU. For no failures, only primary data are looked at. If there is a loss of primary data (which can occur between the engine controller channel A \Jelectronics\j and the SSME SOP), the secondary data are looked at.
There are two primary written engine controller computer software programs: the flight operational program and the test operational program. The flight operational program is an on-line, real-time, process-control program that processes inputs from engine sensors; controls the operation of the engine servovalves, actuators, solenoids and spark igniters; accepts and processes vehicle commands; provides and transmits data to the vehicle; and provides checkout and monitoring capabilities.
The test operational program supports engine testing. Functionally, it is similar to the flight operational program but differs with respect to implementation. The computer software programs are modular and are defined as computer program components, which consist of a data base organized into tables and 15 computer program components.
During application of the computer program components, the programs perform data processing for failure detection and status to the vehicle. As system operation progresses through an operating phase, different combinations of control functions are operative at different times. These combinations within a phase are defined as operating modes.
The checkout phase initiates active control monitoring or checkout. The standby mode in this phase is a waiting mode of controller operation while active control sequence operations are in process. Monitoring functions that do not affect engine hardware status are continually active during the mode. Such functions include processing of vehicle commands, status update and controller self-test.
During checkout, data and instructions can be loaded into the engine controller's computer memory. This permits updating of the software program and data as necessary to proceed with engine-firing operations or checkout operations. Also in this phase, component checkout, consisting of checkout or engine leak tests, is performed on an individual engine system component.
The start preparation phase consists of system purges and propellant conditioning, which are performed in preparation for engine start. The purge sequence 1 mode is the first purge sequence, including oxidizer system and intermediate seal purge operation. The purge sequence 2 mode is the second purge sequence, including fuel system purge operation and the continuation of purges initiated during purge sequence 1.
The purge sequence 3 mode includes propellant recirculation (bleed valve operation). The purge sequence 4 mode includes fuel system purge and the indication engine is ready to enter the start phase. The engine-ready mode occurs when proper engine thermal conditions for start have been attained and other criteria for start have been satisfied, including a continuation of the purge sequence 4 mode.
The start phase covers operations involved with starting or firing the engines, beginning with scheduled open-loop operation of propellant valves. The start initiation mode includes all functions before ignition confirmed and the closing of the thrust control loop. The thrust buildup mode detects ignition by monitoring main combustion chamber pressure and verifying that closed-loop thrust buildup sequencing is in progress.
The main stage phase is automatically entered upon successful completion of the start phase. The normal control mode has initiated mixture ratio control, and thrust control is operating normally. In case of a malfunction, the electrical lock mode will be activated. In that mode, engine propellant valves are electrically held in a fixed configuration, and all control loop communications are suspended. There is also the hydraulic lockup mode, in which all fail-safe valves are deactivated to hydraulically hold the propellant valves in a fixed configuration and all control loop functions are suspended.
The shutdown phase covers operations to reduce main combustion chamber pressure and drive all valves closed to effect full engine shutdown. Throttling to minimum power level is the portion of the shutdown in progress at a programmed shutdown thrust reference level above the MPL.
The valve schedule throttling mode is the stage in the shutdown sequence at which the programmed thrust reference has decreased below the MPL. Propellant valves closed is the stage in the shutdown sequence after all liquid propellant valves have been closed, the shutdown purge has been activated, and verification sequences are in progress. The fail-safe pneumatic mode is when the fail-safe pneumatic shutdown is used.
The post-shutdown phase represents the state of the SSME and engine controller at the completion of engine firing. The standby mode is a waiting mode of controller operations whose functions are identical to those of standby during checkout. It is the normal mode that is entered after completion of the shutdown phase. The terminate sequence mode terminates a purge sequence by a command from the vehicle. All propellant valves are closed, and all \Jsolenoid\j and \Jtorque\j motor valves are de-energized.
Each controller utilizes ac power provided by the MPS engine power left, ctr, right switches on panel R2.
Each controller has internal electrical heaters that provide environmental temperature control and are powered by main bus power through a remote power controller. The RPC is controlled by the main propulsion system engine cntrl htr left, ctr, right switches on panel R4. The heaters are not normally used until after main engine cutoff and are only turned on if environmental control is required during the mission.
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"Shuttle Malfunction Detection",317,0,0,0
There are three separate means of detecting malfunctions within the main propulsion system: the engine controllers, the caution and warning system and the GPCs.
The engine controller, through its network of sensors, has access to numerous engine operating parameters. A group of these parameters has been designated critical operating parameters, and special limits defined for these parameters are hard-wired and limit sensed within the caution and warning system. If a violation of any limit is detected, the caution and warning system will illuminate the red MPS caution and warning light on panel F7.
The light will be illuminated by an MPS engine liquid oxygen manifold pressure above 249 psia; an MPS engine liquid \Jhydrogen\j manifold pressure below 28 psia or above 60 psia; an MPS center, left or right \Jhelium\j pressure below 1,150 psia; an MPS center, left or right \Jhelium\j regulated pressure above 820 psia; or an MPS left, center or right \Jhelium\j delta pressure/delta time above 29 psia.
Note that the flight crew can monitor the MPS press \Jhelium\j pneu, l, c, r meter on panel F7 when the switch is placed in the tank or reg position. The MPS press eng manf LO 2 , LH2 meter can also be monitored on panel F7. A number of the conditions will require crew action.
For example, an MPS engine liquid \Jhydrogen\j manifold pressure below the minimum setting will require the flight crew to pressurize the external liquid \Jhydrogen\j tank by setting the LH2 ullage press switch on panel R2 to open , and a low \Jhelium\j pressure may require the flight crew to interconnect the pneumatic \Jhelium\j tank and the engine \Jhelium\j tanks using the MPS He interconnect valve switches on panel R2 for the engine \Jhelium\j system that is affected.
The engine controller also has a self-test feature that allows it to detect certain malfunctions involving its own sensors and control devices. For each of the three engines, a yellow main engine status left, ctr, right light (lower half) on panel F7 will be illuminated when the corresponding engine \Jhelium\j pressure is below 1,150 psia or regulated \Jhelium\j pressure is above 820 psia.
The lower half of the main engine status left, ctr, right light on panel F7 may also be illuminated by the SSME SOP (GPC- detected malfunctions). The yellow light may be illuminated due to an electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the GPC of the corresponding engine.
In an electronic hold for the affected SSME, loss of data from both pairs of the four fuel flow rate sensors and the four chamber pressure sensors will result in the propellant valve actuators being maintained electronically in the positions existing at the time the second sensor failed. (To fail both sensors in a pair, it is only necessary to fail one sensor.) In the case of either the hydraulic lockup or an electronic hold, all engine-throttling capability for the affected engine is lost; thus, subsequent throttling commands to that engine will not change the thrust level.
The red upper half of the main engine status left, ctr, right light on panel F7 will be illuminated if the corresponding engine's high-pressure oxidizer \Jturbine\j's discharge temperature is above 1,760 degrees R, the main combustion chamber's pressure is below 1,000 psia, the high-pressure oxidizer turbopump's intermediate seal purge pressure is below 170 or above 650 psia, or the high-pressure oxidizer turbopump's secondary seal purge pressure is below 5 or above 85 psia.
Because of the rapidity with which it is possible to exceed these limits, the engine controller has been programmed to sense the limits and automatically cut off the engine if the limits are exceeded. Although a shutdown as a result of violating operating limits is normally automatic, the flight crew can, if necessary, inhibit an automatic shutdown through the use of the main engine limit shut dn switch on panel C3. The switch has three positions: enable, auto and inhibit.
The enable position allows only the first engine that violates operating limits to be shut down automatically. If either of the two remaining engines subsequently violates operating limits, it would be inhibited from automatically shutting down. The inhibit position inhibits all automatic shutdowns. The main engine shutdown left, ctr, right push buttons on panel C3 have spring-loaded covers (guards). When the guard is raised and the push button is depressed, the corresponding engine shuts down immediately.
The backup caution and warning processing of the orbiter GPCs can detect certain specified out-of-limit or fault conditions of the MPS. The backup C/W alarm light on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm sounds if the MPS engine liquid oxygen manifold pressure is zero or above 29 psia; the MPS engine liquid \Jhydrogen\j manifold pressure is below 30 or above 46 psia; the MPS left, center or right \Jhelium\j pressure is below 1,150 psia; or the MPS regulated left, center or right \Jhelium\j pressure is below 680 or above 820 psia. This is identical to the parameter limit sensed by the caution and warning system; thus, the MPS red light on panel F7 will also be illuminated.
The SM alert indicator on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm is sounded when MPS malfunctions/conditions are detected by the SSME SOP or special systems-monitoring processing.
The first four conditions are detected by the SSME SOP and are identical to those that illuminate the yellow lower light of the respective main engine status light on panel F7 due to electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the orbiter GPC.
The last four conditions are special systems-monitoring processing and illuminate the SM alert light on panel F7, sound an audio alarm and provide a fault message on all CRTs because of an external tank liquid \Jhydrogen\j ullage pressure below 30 psia or above 46 psia or an external tank liquid oxygen ullage pressure of zero or above 29 psia. (Note that the main engine status lights on panel F7 will not be illuminated.)
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"Orbiter Hydraulic Systems",318,0,0,0
The three orbiter hydraulic systems supply hydraulic pressure to the main propulsion system for providing thrust vector control and actuating engine valves on each SSME.
The three hydraulic supply systems are distributed to the MPS TVC valves. These valves are controlled by \Jhydraulics\j MPS/TVC 1, 2, 3 switches on panel R4. A valve is opened by positioning its respective switch to open. The talkback indicator above each switch indicates op or cl for open and close.
When the three MPS TVC hydraulic isolation valves are opened, hydraulic pressure actuates the engine main fuel valve, the main oxidizer valve, the fuel preburner oxidizer valve, the oxidizer preburner oxidizer valve and the chamber coolant valve. All hydraulically actuated engine valves on an engine receive hydraulic pressure from the same hydraulic system.
The left engine valves are actuated by hydraulic system 2, the center engine valves are actuated by hydraulic system 1, and the right engine valves are actuated by hydraulic system 3. Each engine valve actuator is controlled by dual-redundant signals: channel A/engine servovalve 1 and channel B/engine servovalve 2 from that engine controller \Jelectronics\j.
As a backup, all of the hydraulically actuated engine valves on an engine are supplied with \Jhelium\j pressure from the \Jhelium\j subsystem left, center and right engine \Jhelium\j tank supply system. In the event of a hydraulic lockup in an engine, \Jhelium\j pressure is used to actuate the engine's propellant valves to their fully closed position when the engine is shut down.
Hydraulic lockup is a condition in which all of the propellant valves on an engine are hydraulically locked in a fixed position. This is a built-in protective response of the MPS propellant valve actuator/control circuit. It takes effect any time low hydraulic pressure or loss of control of one or more propellant valve actuators renders closed-loop control of engine thrust or propellant mixture ratio impossible.
Hydraulic lockup allows an engine to continue to thrust in a safe manner under conditions that normally would require that the engine be shut down; however, the affected engine will continue to operate at approximately the throttle level in effect at the time hydraulic lockup occurred. Once an engine is in a hydraulic lockup, any subsequent shutoff commands, whether nominal or premature, will cause a pneumatic \Jhelium\j shutdown.
Hydraulic lockup does not affect the capability of the engine controller to monitor critical operating parameters or issue an automatic shutdown if an operating limit is out of tolerance; however, the engine shutdown would be accomplished pneumatically.
The three MPS thrust vector control valves must also be opened to supply hydraulic pressure to the six main engine TVC actuators. There are two servoactuators per SSME: one for yaw and one for pitch. Each actuator is fastened to the orbiter thrust structure and to the powerhead of one of the three SSMEs. The two actuators per engine provide attitude control and trajectory shaping by gimbaling the SSMEs in conjunction with the solid rocket boosters during first-stage ascent and without the SRBs during second-stage ascent.
Each SSME servoactuator receives hydraulic pressure from two of the three orbiter hydraulic systems; one system is the primary system and the other is a standby system. Each servoactuator has its own hydraulic switching valve. The switching valve receives hydraulic pressure from two of the three orbiter hydraulic systems and provides a single source to the actuator. Normally, the primary hydraulic supply is directed to the actuator; however, if the primary system were to fail and lose hydraulic pressure, the switching valve would automatically switch over to the standby system, and the actuator would continue to function on the standby system.
The left engine's pitch actuator utilizes hydraulic system 2 as the primary and hydraulic system 1 as the standby. The engine's yaw actuator utilizes hydraulic system 1 as the primary and hydraulic system 2 as the standby. The center engine's pitch actuator utilizes hydraulic system 1 as the primary and hydraulic system 3 as the standby, and the yaw actuator utilizes hydraulic system 3 as the primary and hydraulic system 1 as the standby. The right engine's pitch actuator utilizes hydraulic system 3 as the primary and hydraulic system 2 as the standby. Its yaw actuator utilizes hydraulic system 2 as the primary and hydraulic system 3 as the standby.
The hydraulic systems are distributed among the actuators and engine valves to equalize the hydraulic work load among the three systems.
The hydraulic MPS/TVC isol vlv sys1, sys2, sys3 switches on panel R4 are positioned to close during on-orbit operations to protect against hydraulic leaks downstream of the isolation valves. In addition, there is no requirement to gimbal the main engines from the stow position. During on-orbit operations when the MPS TVC valves are closed, the hydraulic pressure supply and return lines within each MPS TVC component are interconnected to enable hydraulic fluid to circulate for thermal conditioning.
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"Shuttle MPS Thrust Vector Control",319,0,0,0
The space shuttle ascent thrust vector control portion of the flight control system directs the thrust of the three main engines and two solid rocket boosters to control attitude and trajectory during lift-off and first-stage ascent and the main engines alone during second-stage ascent.
Ascent thrust vector control is provided by avionics hardware packages that supply gimbal commands and fault detection for each hydraulic gimbal actuator. The MPS ATVC packages are located in the three aft avionics bays in the orbiter aft fuselage and are cooled by cold plates and the Freon-21 system. The associated flight aft multiplexers/demultiplexers are also located in the aft avionics bays.
The MPS TVC command flow starts in the general-purpose computers, in which the flight control system generates the TVC position commands, and terminates at the SSME servoactuators, where the actuators gimbal the SSMEs in response to the commands. All the MPS TVC position commands generated by the flight control system are issued to the MPS TVC command subsystem operating program, which processes and disburses them to their corresponding flight aft MDMs.
The flight aft MDMs separate these linear discrete commands and disburse them to ATVC channels, which generate equivalent command analog voltages for each command issued. These voltages are, in turn, sent to the servoactuators, commanding the SSME hydraulic actuators to extend or retract, thus gimbaling the main engines to which they are fastened.
Six MPS TVC actuators respond to the command voltages issued by four ATVC channels. Each ATVC channel has six MPS drivers and four SRB drivers. Each actuator receives four identical command voltages from four different MPS drivers, each located in different ATVC channels.
Each main engine servoactuator consists of four independent, two-stage servovalves, which receive signals from the drivers. Each servovalve controls one power spool in each actuator, which positions an actuator ram and the engine to control thrust direction.
The four servovalves in each actuator provide a force-summed majority voting arrangement to position the power spool. With four identical commands to the four servovalves, the actuator's force-sum action prevents a single erroneous command from affecting power ram motion.
If the erroneous command persists for more than a predetermined time, differential pressure sensing activates an isolation driver, which energizes an isolation valve that isolates the defective servovalve and removes hydraulic pressure, permitting the remaining channels and servovalves to control the actuator ram spool provided the FCS channel 1, 2, 3, 4 switch on panel C3 is in the auto position. A second failure would isolate the defective servovalve and remove hydraulic pressure in the same manner as the first failure, leaving only two channels remaining.
Failure monitors are provided for each channel on the CRT and backup caution and warning light to indicate which channel has been bypassed for the MPS and/or SRB. If the FCS channel 1, 2, 3, or 4 switch on panel C3 is positioned to off, that ATVC channel is isolated from its servovalve on all MPS and SRB actuators. The override position of the FCS channel 1, 2, 3, 4 switch inhibits the isolation valve driver from energizing the isolation valve for its respective channel and provides the capability of resetting a failed or bypassed channel.
The ATVC 1, 2, 3, 4 power switch is located on panel O17. The on position enables the ATVC channel selected; off disables the channel.
Each actuator ram is equipped with transducers for position feedback to the TVC system.
The SSME servoactuators change each main engine's thrust vector direction as needed during the flight sequence. The three pitch actuators gimbal the engine up or down a maximum of 10 degrees 30 minutes from the installed null position. The three yaw actuators gimbal the engine left or right a maximum of 8 degrees 30 minutes from the installed position.
The installed null position for the left and right main engines is 10 degrees up from the X axis in a negative Z direction and 3 degrees 30 minutes outboard from an engine centerline parallel to the X axis. The center engine's installed null position is 16 degrees above the X axis for pitch and on the X axis for yaw. When any engine is installed in the null position, the other engines cannot collide with it.
The minimum gimbal rate is 10 degrees per second; the maximum rate is 20 degrees per second.
There are three actuator sizes for the main engines. The piston area of the one upper pitch actuator is 24.8 square inches, its stroke is 10.8 inches, it has a peak flow of 50 gallons per minute, and it weighs 265 pounds. The piston area of the two lower pitch actuators is 20 square inches, their stroke is 10.8 inches, their peak flow is 45 gallons per minute, and they weigh 245 pounds. All three yaw actuators have a piston area of 20 square inches, a stroke of 8.8 inches and a peak flow of 45 gallons per minute and weigh 240 pounds.
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"Orbiter/External Tank Separation System",320,0,0,0
The orbiter/external tank separation system consists of the oxygen and \Jhydrogen\j umbilical disconnects located at the lower left and right aft fuselage, one forward structural attach point just aft of the nose landing gear doors and two structural attach points located in the orbiter/external tank umbilical disconnect cavities. An umbilical retraction system retracts the orbiter umbilicals within the orbiter aft fuselage, and umbilical doors close over each of the umbilical cavities after separation.
The 17-inch liquid oxygen and liquid \Jhydrogen\j disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three space shuttle main engines. The respective 17-inch disconnects also provide the capability for external tank fill and drain of oxygen and \Jhydrogen\j through the orbiter main propulsion system and the T-0 umbilicals.
The liquid \Jhydrogen\j interface between the orbiter and the ground storage tank is provided by a T-0 umbilical located on the left side of the aft fuselage. The liquid oxygen interface between the orbiter and the ground storage tank is provided by a T-0 umbilical on the right side of the aft fuselage.
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"Orbiter 17-Inch Disconnect",321,0,0,0
Each mated pair of 17-inch disconnects contains two flapper valves, one on the orbiter side of the interface and one on the external tank side of the interface. Both valves in each disconnect pair are opened to permit propellant flow between the orbiter and the external tank. Before the separation of the external tank, both valves in each mated pair of disconnects are commanded closed by pneumatic (helium) pressure from the main propulsion system. The closure of both valves in each disconnect pair prevents propellant discharge from the external tank or orbiter at separation. Valve closure on the orbiter side of each disconnect also prevents contamination of the orbiter main propulsion system during landing and ground operations.
Inadvertent closure of either valve in a 17-inch disconnect during space shuttle main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.
To prevent inadvertent closure of the 17-inch disconnect valves during the main engine thrusting, a latch mechanism was added in the orbiter half of the disconnects. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so it overlaps both flappers and obstructs closure for any reason.
In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic (helium) pressure from the main propulsion system causes the latch actuator to rotate the latch shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.
If the latch pneumatic actuator malfunctions, a backup mechanical separation capability is provided. When the orbiter umbilical initially moves away from the external tank umbilical, the mechanical latch disengages from the external tank flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.
During ground mating of the external tank to the orbiter, the latch engagement mechanism in each 17-inch disconnect provides a go/no-go verification that flapper angle rigging is within stability limits. Misrigged flappers will prevent full engagement of latch. The angle of each flapper in each disconnect is still carefully rigged within specific tolerances to assure basic stability independently of the latch safety feature.
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"Orbiter External Tank Separation System",322,0,0,0
The external tank is separated from the orbiter at three structural attach points. Separation from the orbiter occurs before orbit insertion and is automatically controlled by the orbiter's general-purpose computers. External tank separation can be manually initiated by the flight crew using the same jettison circuits as the automatic sequence.
Separation is controlled by the ET separation "auto", "man" switch on panel C3 and the "sep" push button on panel C3. In the auto position, the onboard GPCs initiate separation. To manually initiate separation, the ET separation switch is positioned to "man" and the "sep" push button is depressed.
The forward structural attachment consists of a shear bolt unit mounted in a spherical bearing. The bolt separates at a break area when two pressure cartridges are initiated. The pressure from one or both cartridges drives one of a pair of pistons to shear the bolt, with the second piston acting as a hole plugger to fill the cavity left by the sheared bolt. A centering mechanism rotates the unit from the displacement position to a centered position, aligning the bearing flush with the adjacent thermal protection system mold line.
The aft structural attachment consists of two special bolts and pyrotechnically actuated frangible nuts that attach the external tank strut hemisphere to the orbiter's left- and right-side cavities. At separation the frangible nuts are split by a booster cartridge initiated by a \Jdetonator\j cartridge. The attach bolts are driven by the separation forces and a spring into a cavity in the tank strut. The frangible nut, cartridge fragments and hot gases are contained within a cover assembly, and a hole plugger isolates the fragments in the container.
The aft separation involves right and left umbilical assemblies. Each assembly contains three dual-detonator frangible nut and bolt combinations that hold the orbiter and external tank umbilical plates together during mated flight. Each bolt has a retraction spring that, after release of the nut, retracts the bolt to the external tank side of the interface. On the orbiter side, each frangible nut and its detonators are enclosed in a debris container that captures nut fragments and hot gases generated by the operation of the detonators, either of which will fracture the nut.
The right aft umbilical assembly consists of an electrical disconnect, the gaseous oxygen 2-inch pressurization disconnect used for pressurization of the external tank's oxygen tank and the 17-inch liquid oxygen disconnect.
The left aft umbilical assembly consists of an electrical disconnect plate, the gaseous \Jhydrogen\j 2-inch pressurization disconnect used for pressurization of the external tank's \Jhydrogen\j tank, the 4-inch recirculation disconnect used during prelaunch to precondition the main engine and the 17-inch liquid \Jhydrogen\j disconnect.
After release of the three frangible nuts and bolts at each aft umbilical, three lateral support arms at each orbiter umbilical plate hold the plates in the lateral position when the external tank separates from the umbilical plates. Each 17-inch disconnect has been commanded closed. The orbiter umbilical plates are retracted inside the orbiter aft fuselage approximately 2.5 inches by three hydraulic actuators and locked to permit closure of the umbilical doors in the bottom of the aft fuselage.
Hydraulic system 1 source pressure is supplied to one actuator at each umbilical, hydraulic system 2 source pressure is supplied to the second actuator at each umbilical, and hydraulic system 3 source pressure is supplied to a third actuator at each umbilical.
The retraction of each umbilical disconnects the external tank and orbiter electrical umbilical in the first 0.5 of an inch of travel and releases any fluids trapped between the 17-inch disconnect flappers.
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"Orbiter Umbilical Doors",323,0,0,0
An electromechanical actuation system on each umbilical door closes the left and right umbilical cavities after the external tank is jettisoned and the umbilical plates retracted inside the orbiter's aft fuselage. Each umbilical door is approximately 50 inches square.
The doors are held in the full-open position by two centerline latches, one forward and one aft. They are opened before the mating of the orbiter to the external tank in the Vehicle Assembly Building.
The orbiter umbilical doors normally are controlled by the flight crew with switches on panel R2. In return-to-launch-site aborts, the doors are controlled automatically. The ET umbilical door mode switch on panel R2 positioned to GPC enables automatic control of the doors. The GPC/man position enables manual flight crew control of the doors.
The ET umbilical door centerline latch switch on panel R2 positioned to gnd permits ground control of the door centerline latches during ground turnaround operations. The stow position, enables flight crew manual control of the door centerline latches. The talkback indicator above the switch indicates "sto" when the door centerline latches are stowed, which permits closure of the doors, and barberpole when the latches are latched or the doors are in transit.
The ET umbilical door left and right latch, off, release switches on panel R2 are used by the flight crew to unlatch the corresponding centerline latches during normal operations. Positioning the respective switch to release provides electrical power to redundant ac reversible motors which operate an electromechanical actuator for each centerline latch that causes the latch to rotate and retract the latch blade flush with the reusable thermal protection system mold line.
It takes approximately six seconds for the latches to complete their motion. The talkback indicator above the respective switch indicates rel when the corresponding latches are released. The latch position of each switch is used during ground turnaround operations to latch the respective door open, and the talkback indicator indicates lat when the latches have latched the doors in the open position. The talkback indicators indicate barberpole when the latches are in transit. The off position of the switches removes power from the motors, which stops the latches.
The ET umbilical door left and right , open , off , latch switches on panel R2 normally are used by the flight crew to close the umbilical doors. Positioning the switches to close provides electrical power to redundant ac reversible motors, which position the doors closed through a system of bellcranks and push rods. It takes approximately 24 seconds for the doors to close; and when they are within 2 inches of the closed position, ready-to-latch indicators activate the door uplatch system.
Three uplatch hooks for each door engage three corresponding rollers near the outboard edge of the door and lock the door in preparation for entry. The motors are automatically turned off. The talkback indicator above the respective switch indicates cl when two of the three ready-to-latch switches for that door have sensed door closure. The open position of the switches is used during ground turnaround operations to open the doors. The talkback indicator indicates op when the doors are open and barberpole when they are in transit. The off position removes power from the motors, which stops the doors' movement.
The ET umbilical door switch on panel R2 positioned to GPC provides a backup method of releasing the centerline latches and closing the umbilical doors through guidance, navigation and control software through \Jcathode\j ray tube display item entry during an RTLS abort. The operation of the centerline latches and closing of the umbilical doors are completely automated after external tank separation when the ET umbilical door switch on panel R2 is positioned to GPC .
Two seconds after external tank separation, the centerline latches release the doors and the latches are stowed. The ET umbilical door centerline latch talkback indicator indicates "sto" when the centerline latches complete their motion eight seconds after external tank separation. The left and right umbilical doors are closed, and the ET umbilical door left and right talkback indicates cl 32 seconds after separation. The left and right umbilical door latches latch the doors closed, and the ET umbilical door left and right talkback indicates lat 38 seconds after separation.
Each umbilical door is covered with reusable thermal protection system in addition to an aerothermal barrier that required approximately 6 psi to compress to seal the door with adjacent thermal protection system tiles.
A closeout curtain is installed at each of the orbiter/external tank umbilicals. After external tank separation, the residual liquid oxygen in the main propulsion system is dumped through the three space shuttle main engines and the residual liquid \Jhydrogen\j is dumped overboard. The umbilical curtain prevents hazardous gases (gaseous oxygen and hydrogen) from entering the orbiter aft fuselage through the umbilical openings before the umbilical doors are closed.
The curtain also acts as a seal during the ascent phase of the mission to permit the aft fuselage to vent through the orbiter purge and vent system, thereby protecting the orbiter aft bulkhead at station Xo 1307. The curtain is designed to operate in range of minus 200 F to plus 250 F. The umbilical doors are opened when the orbiter has stopped at the end of landing rollout.
Various parameters are monitored and displayed on the flight deck control panel and CRT and transmitted by telemetry.
Contractors for the separation system include Hoover Electric, Los Angeles, Calif. (external tank umbilical centerline latch and actuator; umbilical door actuator and umbilical door latch actuator); U.S. Bearing, Chatsworth, Calif. (external tank/orbiter spherical bearing); Bertea Corp., Irvine, Calif. (umbilical retractor actuator); Space Ordnance Systems Division, Trans Technology Corp., Saugus, Calif. (orbiter/external tank separation bolt/cartridge \Jdetonator\j assembly, 0.75-inch frangible nut orbiter/external tank umbilical separation and 2.5-inch frangible nut/pyro components in orbiter/external tank aft attach separation system).
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"Orbital Maneuvering System (OMS)",324,0,0,0
The orbital maneuvering system provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around and can provide up to 1,000 pounds of propellant to the aft reaction control system. The OMS is housed in two independent pods located on each side of the orbiter's aft fuselage.
The pods also house the aft RCS and are referred to as the OMS/RCS pods. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers. The two pods provide redundancy for the OMS. The vehicle velocity required for orbital adjustments is approximately 2 feet per second for each nautical mile of altitude change.
The ascent profile of a mission determines if one or two OMS thrusting periods are used and the interactions of the RCS. After main engine cutoff, the RCS thrusters in the forward and aft RCS pods are used to provide attitude hold until external tank separation. At ET separation, the RCS provides a minus (negative) Z translation maneuver of about minus 4 feet per second to maneuver the orbiter away from the ET. Upon completion of the translation, the RCS provides orbiter attitude hold until time to maneuver to the OMS-1 thrusting attitude.
The targeting data for the OMS-1 thrusting period is selected before launch; however, the target data in the onboard general-purpose computers can be modified by the flight crew via the \Jcathode\j ray tube keyboard, if necessary, before the OMS thrusting period.
During the first OMS thrusting period, both OMS engines are used to raise the orbiter to a predetermined elliptical orbit. During the thrusting period, vehicle attitude is maintained by gimbaling (swiveling) the OMS engines. The RCS will not normally come into operation during an OMS thrusting period. If, during an OMS thrusting period, the OMS gimbal rate or gimbal limits are exceeded, RCS attitude control is required. If only one OMS engine is used during an OMS thrusting period, RCS roll control is required.
During the OMS-1 thrusting period, the liquid oxygen and liquid \Jhydrogen\j trapped in the main propulsion system ducts are dumped. The liquid oxygen is dumped out through the space shuttle main engines' combustion chambers and the liquid \Jhydrogen\j is dumped through the starboard (right) side T-0 umbilical overboard fill and drain. This velocity was precomputed in conjunction with the OMS-1 thrusting period.
Upon completion of the OMS-1 thrusting period, the RCS is used to null any residual velocities, if required. The flight crew uses the rotational hand controller and/or translational hand controller to command the applicable RCS thrusters to null the residual velocities. The RCS then provides attitude hold until time to maneuver to the OMS-2 thrusting attitude.
The second OMS thrusting period using both OMS engines occurs near the apogee of the orbit established by the OMS-1 thrusting period and is used to circularize the predetermined orbit for that mission. The targeting data for the OMS-2 thrusting period is selected before launch; however, the target data in the onboard GPCs can be modified by the flight crew via the CRT keyboard, if necessary, before the OMS thrusting period.
Upon completion of the OMS-2 thrusting period, the RCS is used to null any residual velocities, if required, in the same manner as during OMS-1. The RCS is then used to provide attitude hold and minor translation maneuvers as required for on-orbit operations. The flight crew can select primary or vernier RCS thrusters for attitude control on orbit. Normally, the vernier RCS thrusters are selected for on-orbit attitude hold.
If the ascent profile for a mission uses a single OMS thrusting maneuver, it is referred to as direct insertion. In a direct-insertion ascent profile, the OMS-1 thrusting period after main engine cutoff is eliminated and is replaced with a 5-feet- per-second RCS translation maneuver to facilitate the main propulsion system dump. The RCS provides attitude hold after the translation maneuver. The OMS-2 thrusting period is then used to achieve orbit insertion. The direct-insertion ascent profile allows the MPS to provide more energy to orbit insertion and permits easier use of onboard software.
Additional OMS thrusting periods using both or one OMS engine are performed on orbit according to the mission's requirements to modify the orbit for rendezvous, payload deployment or transfer to another orbit.
The two OMS engines are used to deorbit. Target data for the deorbit maneuver is computed by the ground and loaded in the onboard GPCs via uplink. This data is also voiced to the flight crew for verification of loaded values. After verification of the deorbit data, the flight crew initiates an OMS gimbal test on the CRT keyboard unit.
Before the deorbit thrusting period, the flight crew maneuvers the \Jspacecraft\j to the desired deorbit thrusting attitude using the rotational hand controller and RCS thrusters. Upon completion of the OMS thrusting period, the RCS is used to null any residual velocities, if required. The \Jspacecraft\j is then maneuvered to the proper entry interface attitude using the RCS. The remaining propellants aboard the forward RCS are dumped by burning the propellants through the forward RCS thrusters before the entry interface if it is necessary to control the orbiter's center of gravity.
The aft RCS plus X jets can be used to complete any planned OMS thrusting period in the event of an OMS engine failure. In this case, the OMS-to-aft-RCS interconnect would feed OMS propellants to the aft RCS.
From entry interface at 400,000 feet, the orbiter is controlled in roll, pitch and yaw with the aft RCS thrusters. The orbiter's ailerons become effective at a dynamic pressure of 10 pounds per square foot, and the aft RCS roll jets are deactivated. At a dynamic pressure of 20 pounds per square foot, the orbiter's elevons become effective, and the aft RCS pitch jets are deactivated. The rudder is activated at Mach 3.5, and the aft RCS yaw jets are deactivated at Mach 1 and approximately 45,000 feet.
The OMS in each pod consists of a high-pressure gaseous \Jhelium\j storage tank, \Jhelium\j isolation valves, dual pressure regulation systems, vapor isolation valves for only the oxidizer regulated \Jhelium\j pressure path, quad check valves, a fuel tank, an oxidizer tank, a propellant distribution system consisting of tank isolation valves, crossfeed valves, and an OMS engine. Each OMS engine also has a gaseous \Jnitrogen\j storage tank, gaseous \Jnitrogen\j pressure isolation valve, gaseous \Jnitrogen\j accumulator, bipropellant \Jsolenoid\j control valves and actuators that control bipropellant ball valves, and purge valves.
In each of the OMS pods, gaseous \Jhelium\j pressure is supplied to \Jhelium\j isolation valves and dual pressure regulators, which supply regulated \Jhelium\j pressure to the fuel and oxidizer tanks. The fuel is monomethyl hydrazine and the oxidizer is \Jnitrogen\j tetroxide. The propellants are Earth-storable liquids at normal temperatures.
They are pressure-fed to the propellant distribution system through tank isolation valves to the OMS engines. The OMS engine propellant ball valves are positioned by the gaseous \Jnitrogen\j system and control the flow of propellants into the engine. The fuel is directed first through the engine combustion chamber walls and provides regenerative cooling of the chamber walls; it then flows into the engine injector. The oxidizer goes directly to the engine injector. The propellants are sprayed into the combustion chamber, where they atomize and ignite upon contact with each other (hypergolic), producing a hot gas and, thus, thrust.
The gaseous \Jnitrogen\j system is also used after the OMS engines are shut down to purge residual fuel from the injector and combustion chamber, permitting safe restarting of the engines. The nozzle extension of each OMS engine is radiation-cooled and is constructed of columbium alloy.
Each OMS engine produces 6,000 pounds of thrust. The oxidizer-to-fuel ratio is 1.65-to-1. The expansion ratio of the nozzle exit to the throat is 55-to-1. The chamber pressure of the engine is 125 psia. The dry weight of each engine is 260 pounds.
Each OMS engine can be reused for 100 missions and is capable of 1,000 starts and 15 hours of cumulative firing. The minimum duration of an OMS engine firing is two seconds. The OMS may be utilized to provide thrust above 70,000 feet. For vehicle velocity changes of between 3 and 6 feet per second, normally only one OMS engine is used.
Each engine has two electromechanical gimbal actuators, which control the OMS engine thrust direction in pitch and yaw (thrust vector control). The OMS engines can be used singularly by directing the thrust vector through the orbiter center of gravity or together by directing the thrust vector of each engine parallel to the other.
During a two-OMS-engine thrusting period, the RCS will come into operation only if the OMS gimbal rate or gimbal limits are exceeded and should not normally come into operation during the OMS thrust period. However, during a one-OMS-engine thrusting period, roll RCS control is required. The pitch and yaw actuators are identical except for the stroke length and contain redundant electrical channels (active and standby), which couple to a common mechanical drive assembly.
The OMS/RCS pods are designed to be reused for up to 100 missions with only minor repair, refurbishment and maintenance. The pods are removable to facilitate orbiter turnaround, if required.
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"OMS Helium Pressurization",325,0,0,0
Each pod pressurization system consists of a \Jhelium\j tank, two \Jhelium\j isolation valves, two dual pressure regulator assemblies, parallel vapor isolation valves on the regulated \Jhelium\j pressure to the oxidizer tank only, dual series-parallel check valve assemblies and pressure relief valves.
The \Jhelium\j storage tank in each pod has a \Jtitanium\j liner with a fiberglass structural overwrap. This increases safety and decreases the weight of the tank 32 percent over that of conventional tanks. The \Jhelium\j tank is 40.2 inches in diameter and has a volume of 17.03 cubic feet minimum. Its dry weight is 272 pounds. The \Jhelium\j tank's operating pressure range is 4,800 to 460 psia with a maximum operating limit of 4,875 psia at 200 F.
A pressure sensor downstream of each \Jhelium\j tank in each pod monitors the \Jhelium\j source pressure and transmits it to the N 2 , He , kit He switch on panel F7. When the switch is in the He position, the \Jhelium\j pressure of the left and right OMS is displayed on the OMS press left, right meters. This pressure also is transmitted to the CRT and displayed.
The two \Jhelium\j pressure isolation valves in each pod permit \Jhelium\j source pressure to the propellant tanks or isolate the \Jhelium\j from the propellant tanks. The parallel paths in each pod assure \Jhelium\j flow to the propellant tanks of that pod. The \Jhelium\j valves are continuous-duty, solenoid-operated. They are energized open and spring loaded closed.
The OMS He press/vapor isol switches on panel O8 permit automatic or manual control of the valves. With the switches in the GPC position, the valves are automatically controlled by the general-purpose computer during an engine thrusting sequence. The valves are controlled manually by placing the switches to open or close.
The pressure regulators reduce the \Jhelium\j source pressure to the desired working pressure. Pressure is regulated by assemblies downstream of each \Jhelium\j pressure isolation valve. Each assembly contains primary and secondary regulators in series and a flow limiter. Normally, the primary regulator is the controlling regulator. The secondary regulator is normally open during a dynamic flow condition.
It will not become the controlling regulator until the primary regulator allows a higher pressure than normal. All regulator assemblies are in reference to a bellows assembly that is vented to ambient. The primary regulator outlet pressure at normal flow is 252 to 262 psig and 247 psig minimum at high abort flow, with lockup at 266 psig maximum.
The secondary regulator outlet pressure at normal flow is 259 to 269 psig and 254 psig minimum at high abort flow, with lockup at 273 psig maximum. The flow limiter restricts the flow to a maximum of 1,040 standard cubic feet per minute and to a minimum of 304 standard cubic feet per minute.
The vapor isolation valves in the oxidizer pressurization line to the oxidizer tank prevent oxidizer vapor from migrating upstream and over into the fuel system. These are low-pressure, two-position, two-way, solenoid-operated valves that are energized open and spring loaded closed.
They can be commanded manually or automatically by the positioning of the He press/vapor isol switches on panel O8. When either of the A or B switches is in the open position, both vapor isolation valves are energized open; and when both switches are in the close position, both vapor isolation valves are closed. When the switches are in the GPC position, the GPC opens and closes the valves automatically.
The check valve assembly in each parallel path contains four independent check valves connected in a series-parallel configuration to provide a positive checking action against a reverse flow of propellant liquid or vapor, and the parallel path permits redundant paths of \Jhelium\j to be directed to the propellant tanks. Filters are incorporated into the inlet of each check valve assembly.
Two pressure sensors in the \Jhelium\j pressurization line upstream of the fuel and oxidizer tanks monitor the regulated tank pressure and transmit it to the RCS/OMS press rotary switch on panel O3. When the switch is in the OMS prplnt position, the left and right fuel and oxidizer pressure is displayed. If the tank pressure is lower than 234 psia or above 284 psia, the left or right OMS red caution and warning light on panel F7 will be illuminated. These pressures also are transmitted to the CRT and displayed.
The relief valves in each pressurization path limit excessive pressure in the propellant tanks. Each pressure relief valve also contains a burst diaphragm and filter. If excessive pressure is caused by \Jhelium\j or propellant vapor, the diaphragm will rupture and the relief valve will open and vent the excessive pressure overboard. The filter prevents particulates from the non-fragmentation-type diaphragm from entering the relief valve seat.
The relief valve will close and reset after the pressure has returned to the operating level. The burst diaphragm is used to provide a more positive seal of \Jhelium\j and propellant vapors than the relief valve. The diaphragm ruptures between 303 and 313 psig. The relief valve opens at a minimum of 286 psig and a maximum of 313 psig. The relief valve's minimum reseat pressure is 280 psig. The maximum flow capacity of the relief valve at 60 F and 313 psig is 520 cubic feet per minute.
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"OMS Propellant Storage and Distribution",326,0,0,0
The propellant storage and distribution system consists of one fuel tank and one oxidizer tank in each pod. It also contains propellant feed lines, interconnect lines, isolation valves and crossfeed valves.
The OMS propellant tanks of both pods enable the orbiter to reach a 1,000-foot- per-second velocity change with a 65,000-pound payload in the payload bay. An OMS pod crossfeed line allows the propellants in the pods to be used to operate either OMS engine.
The propellant is contained in domed cylindrical \Jtitanium\j tanks within each pod. Each propellant tank is 96.38 inches long with a diameter of 49.1 inches and a volume of 89.89 cubic feet unpressurized. The dry weight of each tank is 250 pounds. The propellant tanks are pressurized by the \Jhelium\j system.
Each tank contains a propellant acquisition and retention assembly in the aft end and is divided into forward and aft compartments. The propellant acquisition and retention assembly is located in the aft compartment and consists of an intermediate bulkhead with communication screen and an acquisition system.
The propellant in the tank is directed from the forward compartment through the intermediate bulkhead through the communication screen into the aft compartment during OMS velocity maneuvers. The communication screen retains propellant in the aft compartment during zero-gravity conditions.
The acquisition assembly consists of four stub galleries and a collector manifold. The stub galleries acquire wall-bound propellant at OMS start and during RCS velocity maneuvers to prevent gas ingestion. The stub galleries have screens that allow propellant flow and prevent gas ingestion.
The collector manifold is connected to the stub galleries and also contains a gas arrestor screen to further prevent gas ingestion, which permits OMS engine ignition without the need of a propellant-settling maneuver employing RCS thrusters. The propellant tank's nominal operating pressure is 250 psi, with a maximum operating pressure limit of 313 psia.
A capacitance gauging system in each OMS propellant tank measures the propellant in the tank. The system consists of a forward and aft probe and a totalizer. The forward and aft fuel probes use fuel (which is a conductor) as one plate of the capacitor and a glass tube that is metallized on the inside as the other.
The forward and aft oxidizer probes use two concentric nickel tubes as the capacitor plates and oxidizer as the dielectric. (Helium is also a dielectric, but has a different dielectric constant than the oxidizer.) The aft probes in each tank contain a resistive temperature-sensing element to correct variations in fluid density. The fluid in the area of the communication screens cannot be measured.
The totalizer receives OMS valve operation information and inputs from the forward and aft probes in each tank and outputs total and aft quantities and a low level quantity. The inputs from the OMS valves allow control logic in the totalizer to determine when an OMS engine is thrusting and which tanks are being used. The totalizer begins an engine flow rate/time \Jintegration\j process at the start of the OMS thrusting period, which reduces the indicated amount of propellants by a preset estimated rate for the first 14.8 seconds.
After 14.8 seconds of OMS thrusting, which settles the propellant surface, the probe capacitance gauging system outputs are enabled, which permits the quantity of propellant remaining to be displayed. The totalizer outputs are displayed on the OMS/RCS prplnt qty meters on panel O3 when the rotary switch is positioned to the OMS fuel or oxid positions.
When the wet or dry analog comparator indicates the forward probe is dry, the ungaugeable propellant in the region of the intermediate bulkhead is added to the aft probe output quantity, decreasing the total quantity at a preset rate for 98.15 seconds, and updates from the aft probes are inhibited. After 98.15 seconds of thrusting, the aft probe output inhibit is removed, and the aft probe updates the total quantity. When the quantity decreases to 5 percent, the low-level signal is output.
There are 64 ac -motor-operated valve actuators in the OMS/RCS \Jnitrogen\j tetroxide and monomethyl hydrazine propellant systems. Each valve actuator was modified to incorporate a 0.25-inch-diameter stainless steel sniff line from the actuator to the mold line of the orbiter. The sniff line permits the monitoring of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of each valve actuator during ground operations.
There are sniff lines in the 12 ac -motor-operated valve actuators in the forward RCS and in the 44 actuators in the aft left and aft right RCS. The remaining 0.25-inch-diameter sniff lines are in the eight OMS tank isolation and crossfeed ac-motor-operated valve actuators in the left and right orbital maneuvering systems. The 44 aft left and right RCS sniff lines and the eight OMS left and right sniff lines are routed to the respective left and right OMS/RCS pod Y web access servicing panels.
During ground operations, an interscan can be connected to the sniff ports to check for the presence of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of the ac-motor-operated valve actuators.
An electrical microswitch in each of the ac-motor-operated valve actuators signals the respective valves' position (open or closed) to the onboard flight crew displays and controls as well as telemetry. An extensive improvement program was implemented to reduce the probability of floating particulates in the electrical microswitch portion of each ac-motor-operated valve actuator. Particulates could affect the operation of the microswitch in each valve and, thus, the position indication of the valves to the onboard displays and controls and telemetry.
Each OMS engine receives pressure-fed propellants at its bipropellant valve assembly. The bipropellant ball valve assembly is controlled by its gaseous \Jnitrogen\j system. The \Jnitrogen\j system consists of a storage tank, engine pressure isolation valve, regulator, relief valve, check valve, accumulator, engine purge valves, bipropellant \Jsolenoid\j control valves and actuators that control the bipropellant ball valves.
A gaseous \Jnitrogen\j spherical storage tank is mounted next to the combustion chamber to supply pressure to its engine pressure isolation valve. The tank contains enough \Jnitrogen\j to operate the ball valves and purge the engine 10 times. Nominal tank capacity is 60 cubic inches. The maximum tank operating pressure is 3,000 psi, with a proof pressure of 6,000 psig.
Each tank's pressure is monitored by two pressure sensors. One sensor transmits the tank pressure to the N 2 , He, kit He switch on panel F7. When the switch is positioned to N 2 , tank pressure is displayed on the OMS press N 2 tank left, right meters on panel F7. The other sensor transmits pressure to telemetry.
A dual-coil, solenoid-operated engine pressure isolation valve is located in each gaseous \Jnitrogen\j system. The valve is energized open and spring-loaded closed. The engine pressure isolation valve permits gaseous \Jnitrogen\j flow from the tank to the regulator, accumulator, the bipropellant ball valve control valves and purge valves 1 and 2 when energized open and isolates the \Jnitrogen\j tank from the gaseous \Jnitrogen\j supply system when closed. The engine pressure isolation valves in each system are controlled by the OMS eng left, right switches on panel C3. When the OMS eng left switch is placed in the arm press position, the left OMS engine pod's pressure isolation valve is energized open.
When the OMS eng right switch is placed in the arm press position, the right OMS engine pod's pressure isolation valve is energized open. The gaseous \Jnitrogen\j engine pressure isolation valve, when energized open, allows gaseous \Jnitrogen\j supply pressure to be directed into a regulator, through a check valve, an in-line accumulator and to a pair of engine bipropellant control valves. The engine bipropellant control valves are controlled by the OMS thrust on/off commands from the GPCs.
A single-stage regulator is installed in each gaseous \Jnitrogen\j pneumatic control system between the gaseous \Jnitrogen\j engine pressure isolation valve and the engine bipropellant control valves. The regulator reduces the gaseous \Jnitrogen\j service pressure to a desired working pressure of 315 to 360 psig.
A pressure relief valve downstream of the gaseous \Jnitrogen\j regulator limits the pressure to the engine bipropellant control valves and actuators if a gaseous \Jnitrogen\j regulator malfunctions. The relief valve relieves between 450 and 500 psig and resets at 400 psig minimum.
A pressure sensor downstream of the regulator monitors the regulated pressure and transmits it to the CRT display and to telemetry.
The check valve located downstream of the gaseous \Jnitrogen\j regulator will close if gaseous \Jnitrogen\j pressure is lost on the upstream side of the check valve and will isolate the remaining gaseous \Jnitrogen\j pressure on the downstream side of the check valve.
The 19-cubic- inch gaseous \Jnitrogen\j accumulator downstream of the check valve and upstream of the bipropellant control valves provides enough pressure to operate the engine bipropellant control valves one time with the engine pressure isolation valve closed or in the event of loss of pressure on the upstream side of the check valve.
Two solenoid-operated, three-way, two-position bipropellant control valves on each OMS engine control the bipropellant control valve actuators and bipropellant ball valves. Control valve 1 controls the No. 1 actuator and the fuel and oxidizer ball valves. Control valve 2 controls the No. 2 actuator and two ball valves, one fuel and oxidizer ball valve in series to the No. 1 system. Each control valve contains two \Jsolenoid\j coils, either of which, when energized, opens the control valve.
The right OMS engine gaseous \Jnitrogen\j \Jsolenoid\j control valves 1 and 2 are energized open by computer commands if the right OMS eng switch on panel C3 is in the arm or arm/press position and the right OMS eng vlv switch on panel O16 is on; the valves are de-energized normally when thrust off is commanded or if the right OMS eng switch is positioned to off . The left OMS engine gaseous \Jnitrogen\j \Jsolenoid\j control valves 1 and 2 are controlled in the same manner, but through the left OMS eng switch on panel C3 and the left OMS eng vlv switch on panel O14.
When the gaseous \Jnitrogen\j \Jsolenoid\j control valves are energized open, pressure is directed into the two actuators in each engine. The \Jnitrogen\j acts against the piston in each actuator, overcoming the spring force on the opposite side of the actuators. Each actuator has a rack-and-pinion gear; and the linear motion of the actuator connecting arm is converted into rotary motion, which drives two ball valves, one fuel and one oxidizer, to the open position. Each pair of ball valves opens simultaneously. Fuel and oxidizer are then directed to the combustion chamber of the engine, where the propellants atomize and ignite upon contact. The hypergolic propellants produce a hot gas, thus thrust.
The chamber pressure of each engine is monitored by a pressure sensor and is transmitted to the OMS press left and right Pc (chamber pressure) meter on panel F7.
When the computer commands thrust off or an engine's OMS eng switch on panel C3 or eng vlv switch on panel O14/O16 is positioned off, the \Jsolenoid\j control valves are de-energized, removing gaseous \Jnitrogen\j pressure from the actuators; and the gaseous \Jnitrogen\j pressure in the actuators is vented overboard through the \Jsolenoid\j control valve. The spring in the actuator forces the actuator's piston to move in the opposite direction, and the actuator drives the fuel and oxidizer ball valves closed simultaneously. The series-redundant arrangement of ball valves ensures engine thrusting is terminated.
Each actuator incorporates a linear position \Jtransducer\j, which supplies ball valve position to a CRT.
Check valves are installed in the vent port outlet of each gaseous \Jnitrogen\j \Jsolenoid\j control valve on the spring pressure side of each actuator to protect the seal of these components from atmospheric contamination.
Each engine has two gaseous \Jnitrogen\j purge valves in series. These valves are solenoid-operated open and spring-loaded closed. They are normally energized open after each thrusting period by the GPCs unless inhibited by a crew entry on the maneuver CRT display. The two purge valves of an engine are energized open 0.36 second after OMS engine thrust off has been commanded and permit gaseous \Jnitrogen\j to flow through the valves and check valve into the fuel line downstream of the ball valves and out through the combustion chamber and engine injector to space for two seconds. This purges the residual fuel from the combustion chamber and injector of the engine, permitting safe engine restart. The purge valves are then de-energized and spring-loaded closed. When the purge is completed, the gaseous \Jnitrogen\j tank pressure isolation valve is closed by placing the respective OMS eng switch (panel C3) to off. The check valve downstream of the purge valves prevents fuel from flowing to the engine purge valves during engine thrusting.
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"OMS Engine Thrust Chamber Assembly",328,0,0,0
When the fuel reaches the thrust chamber, it is directed through 102 coolant channels in the combustion chamber wall, providing regenerative cooling to the combustion chamber walls, and then to the injector of the engine. The oxidizer is routed directly to the injector. The platelet injector assembly consists of a stack of plates, each with an etched pattern that provides proper distribution and propellant injection velocity vector. The stack is diffusion-bonded and welded to the body of the injector. The fuel and oxidizer orifices are positioned so that the propellants will impinge and atomize, causing the fuel and oxidizer to ignite because of hypergolic reaction.
The contoured nozzle extension is bolted to the aft flange of the combustion chamber. The nozzle extension is made of a columbium alloy and is radiantly cooled.
The nominal flow rate of oxidizer and fuel to each engine is 11.93 pounds per second and 7.23 pounds per second, respectively, producing 6,000 pounds of thrust at a vacuum specific impulse of 313 seconds.
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"OMS Thrusting Sequence",329,0,0,0
The OMS thrusting sequence commands the OMS engines on or off and commands the engine purge function. The flight crew can select, via item entry on the maneuver display, a one- or two-engine thrusting maneuver and can inhibit the OMS engine purge.
The sequence determines which engines are selected and then provides the necessary computer commands to open the appropriate \Jhelium\j vapor isolation valves and the engine gaseous \Jnitrogen\j \Jsolenoid\j control valves and sets an engine-on indicator. The sequence will monitor the OMS engine fail flags and, if one or both engines have failed, issue the appropriate OMS cutoff commands as soon as the crew has confirmed the failure by placing the OMS eng switch in the off position. This will then terminate the appropriate engine's control valve commands.
In a normal OMS thrusting period, when the OMS cutoff flag is true, the sequence terminates commands to the \Jhelium\j pressurization, \Jhelium\j vapor isolation valves and two gaseous \Jnitrogen\j engine control valves. If the engine purge sequence is not inhibited, the sequence will check for the left and right engine arm press signals and after 0.36 second open the engine gaseous \Jnitrogen\j purge valves for two seconds for the engines that have the arm press signals present.
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"OMS Engine Thrust Vector Control System",330,0,0,0
The engine TVC system consists of a gimbal ring assembly, two gimbal actuator assemblies and two gimbal actuator controllers. The engine gimbal ring assembly and gimbal actuator assemblies provide OMS TVC by gimbaling the engines in pitch and yaw. Each engine has a pitch actuator and a yaw actuator. Each actuator is extended or retracted by one of a pair of dual-redundant electric motors and is actuated by general-purpose computer control signals.
The gimbal ring assembly contains two mounting pads to attach the engine to the gimbal ring and two pads to attach the gimbal ring to the orbiter. The ring transmits engine thrust to the pod and orbiter.
The pitch and yaw gimbal actuator assembly for each OMS engine provides the force to gimbal the engines. Each actuator contains a primary and secondary motor and drive gears. The primary and secondary drive systems are isolated and are not operated concurrently. Each actuator consists of two redundant brushless dc motors and gear trains, a single jackscrew and nut-tube assembly and redundant linear position feedback transducers.
A GPC position command signal from the primary electronic controller energizes the primary dc motor, which is coupled with a reduction gear and a no-back device. The output from the primary power train drives the jackscrew of the drive assembly, causing the nut-tube to translate (with the secondary power train at idle), which causes angular engine movement.
If the primary power train is inoperative, a GPC position command from the secondary electronic controller energizes the secondary dc motor, providing linear travel by applying \Jtorque\j to the nut-tube through the spline that extends along the nut-tube for the stroke length of the unit. Rotation of the nut-tube about the stationary jackscrew causes the nut-tube to move along the screw. A no-back device in each drive system prevents backdriving of the standby system.
The electrical interface, power and electronic control elements for active and standby control channels are assembled in separate enclosures designated the active actuator controller and standby actuator controller. These are mounted on the OMS/RCS pod structure. The active and standby actuator controllers are electrically and mechanically interchangeable.
The gimbal assembly provides control angles of plus or minus 6 degrees in pitch and plus or minus 7 degrees in yaw with clearance provided for an additional 1 degree for snubbing and tolerances. The engine null position is with the engine nozzles up 15 degrees 49 seconds (as projected in the orbiter XZ plane) and outboard 6 degrees 30 seconds (measured in the 15-degree 49-second plane).
The thrust vector control command subsystem operating program processes and outputs pitch and yaw OMS engine actuator commands and the actuator power selection discretes. The OMS TVC command SOP is active during operational sequences, orbit insertion (OMS-1 and OMS-2), orbit coast, deorbit, deorbit coast and return-to-launch-site abort.
The OMS TVC feedback SOP monitors the primary and secondary actuator selection discretes from the maneuver display and performs compensation on the selected pitch and yaw actuator feedback data. This data is output to the OMS actuator fault detection and identification and to the maneuver display. The OMS TVC feedback SOP is active during orbit insertion (OMS-1 and OMS-2), orbit coast, deorbit maneuver and deorbit maneuver coast. The present OMS gimbal positions can be monitored on the maneuver CRT display when this SOP is active and the primary or secondary actuator motors are selected.
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"OMS Thermal Control",331,0,0,0
OMS thermal control is achieved by \Jinsulation\j on the interior surface of the pods that enclose the OMS hardware components and the use of strip heaters. Wrap around heaters and \Jinsulation\j condition the crossfeed lines. The heaters prevent propellant from freezing in the tanks and lines. The heater system is divided into two areas: the OMS/RCS pods and the aft fuselage crossfeed and bleed lines. Each heater system has two redundant heater systems, A and B, and is controlled by the RCS/OMS heaters switches on panel A14.
Each OMS/RCS pod is divided into eight heater areas. Each of the heater areas in the pods contains an A and B element, and each element has a thermostat that controls the temperature from 55 to 75 F. These heater elements are controlled by the left pod and right pod switches on panel A14. Sensors located throughout the pods supply temperature information to the propellant thermal CRT display and telemetry.
The crossfeed line thermal control in the aft fuselage is divided into 11 heater areas. Each area is heated in parallel by heater systems A and B, and each area has a control thermostat to maintain temperature at 55 F minimum to 75 F maximum. Each circuit also has an overtemperature thermostat to protect against a failed-on heater switch. These heater elements are controlled by the respective crsfd lines switch on panel A14. Temperature sensors near the control thermostats on the crossfeed and bleed lines supply temperature information on the propellant thermal CRT display and telemetry.
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"OMS-RCS Interconnect",332,0,0,0
An interconnect between the OMS crossfeed line and the aft RCS manifolds provides the capability to operate the aft RCS using 1,000 pounds per pod of OMS propellant for orbital maneuvers. The aft RCS may use OMS propellant from either OMS pod in orbit.
The orbital interconnect sequence is available during orbit operations and on-orbit checkout.
The left OMS \Jhelium\j pressure vapor isolation valve A will be commanded open when the left OMS tank (ullage) pressure decays to 236 psig, and the open commands will be terminated 30 seconds later. If the left OMS tank (ullage) pressure remains below 236 psia, the sequence will set an OMS/RCS valve miscompare flag and will set a Class 3 alarm and a CRT fault message. The sequence also will enable the OMS-to-RCS gauging sequence at the same time.
The flight crew can terminate the sequence and inhibit the OMS-to-RCS gauging sequence by use of the OMS press ena-off item entry on the RCS SPEC display. The valves can then be reconfigured to their normal position on panels O7 and O8. The OMS-to-aft-RCS interconnect sequence is not available in the backup flight control system.
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"OMS-To-RCS Gauging Sequence",333,0,0,0
The OMS-to-aft-RCS propellant quantities are calculated by burn time \Jintegration\j. Once each cycle, the accumulated aft RCS thruster cycles are used to compute the OMS propellant used since the initiation of gauging. The number of RCS thruster cycles is provided by the RCS command subsystem operating program to account for minimum-impulse firing of the RCS thrusters. The gauging sequence is initiated by item entry of the OMS right or OMS left interconnect on the RCS SPEC CRT display and is terminated by the return to normal item entry.
The gauging sequence maintains a cumulative total of left and right OMS propellant used during OMS-to-aft-RCS interconnects and displays the cumulative totals as percentage of left and right OMS propellant on the RCS SPEC display. The flight crew will be alerted by a Class 3 alarm and a fault message when the total quantity used from either OMS pod exceeds 1,000 pounds or 8.37 percent.
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"OMS Abort Control Sequences",334,0,0,0
The abort control sequence is the software that manages, among other items, the OMS and aft RCS configuration and thrusting periods during ascent aborts to improve performance or to consume OMS and aft RCS propellants for orbiter center-of-gravity control.
Premission-determined parameters are provided for the OMS and aft RCS thrusting periods during aborts since the propellant loading and orbiter center of gravity vary with each mission.
The premission-determined parameters for the abort-to-orbit thrusting period are modified during flight, based on the vehicle velocity at abort initiation. The premission-determined parameters for abort once around are grouped with different values for early or late AOA. The return-to-launch-site parameters are contained in a single table.
The abort control sequence is available in OPS 1 and 6 and is initiated at SRB separation if selected before then or at the time of selection if after SRB separation.
ATO and AOA Aborts. The OMS and aft RCS begin thrusting as soon as an ATO or AOA is initiated with one main engine out.
For some aborts, an OMS-to-aft-RCS interconnect is not desired. A parallel aft RCS plus X thrusting period using aft RCS propellant and the four aft RCS plus X thrusters will be performed during the OMS-1 thrusting period to achieve the desired orbit.
If a plus X aft RCS thrusting period is required before main engine cutoff, the abort control sequence will command the four aft plus X RCS jets on if vehicle acceleration is greater than 0.8 g and will monitor the RCS cutoff time to terminate the thrusting period. If an RCS propellant dump (burn) is required before MECO and vehicle acceleration is greater than 1.8 g, the abort control sequence will command an eight-aft-RCS-jet null thrust and monitor the RCS cutoff time to terminate the thrusting period.
In other abort cases, an OMS-to-aft-RCS interconnect is desired. This thrusting is performed with the OMS and four aft RCS plus X thrusters to consume OMS propellant for orbiter center-of-gravity control. More aft RCS jets can be commanded if needed to increase OMS propellant usage. For example, for an OMS propellant dump (burn), 14 aft RCS null jets can be commanded to thrust to improve orbiter center-of-gravity location.
If the amount of OMS propellant used before MECO leaves less than 28 percent of OMS propellants, a 15-second aft RCS ullage thrust is performed after MECO to provide a positive OMS propellant feed to start the OMS-1 thrusting period.
The OMS-to-aft-RCS interconnect sequence provides for an automatic interconnect of the OMS propellant to the aft RCS when required and reconfigures the propellant feed from the OMS and aft RCS tanks to their normal state after the thrusting periods have ended. The interconnect sequence is initiated by the abort control sequence.
In order to establish a known configuration of the valves, the interconnect sequence terminates the GPC commands to the following valves if they have not been terminated before honoring a request from the abort control sequence: left and right OMS crossfeed A and B valves, aft RCS crossfeed valves and aft RCS tank isolation valves.
A request from the abort control sequence for an OMS-to-aft-RCS interconnect will sequentially configure the OMS/RCS valves as follows: close the left and right aft RCS propellant tank isolation valves, open the left and right OMS crossfeed A and B valves, and open the left and right aft RCS crossfeed valves. The OMS-to-aft-RCS interconnect complete flag is then set to true.
When the abort control sequence requests a return to normal configuration, all affected OMS/RCS propellant valve commands are removed to establish a known condition; and the interconnect sequence will then sequentially configure the valves as follows: close aft RCS crossfeed valves, close left and right OMS crossfeed valves and open aft RCS propellant tank isolation valves. The OMS-to-aft-RCS reconfiguration complete flag is then set to false, and the sequence is terminated.
Return-to-Launch-Site Abort. An RTLS abort requires the dumping of OMS propellant by burning the OMS propellant through both OMS engines and through the 24 aft RCS thrusters to improve abort performance and to achieve an acceptable entry orbiter vehicle weight and center-of-gravity location. The thrusting period is premission-determined and depends on the OMS propellant load.
The OMS engines start the thrusting sequence; and after the OMS-to-aft-RCS interconnect is complete, the aft RCS thrusters are commanded on. The OMS engines and RCS thrusters then continue their burn for a predetermined period. The interconnect sequence is the same for ATO and AOA aborts. The OMS and aft RCS will begin thrusting at SRB staging if the abort is initiated during the first stage of flight or immediately upon abort initiation during second stage.
Contingency Abort. A contingency abort is selected automatically at the loss of a second main engine or manually by the flight crew using an item entry on the RTLS TRAJ or RTLS TRANS CRT displays. For the contingency aborts, the OMS-to-aft-RCS interconnect is performed in a modified manner to allow continuous flow of propellants to the aft RCS jets for vehicle control and to allow contingency rapid dump (burning) of OMS and RCS propellants. The abort control sequence tracks the total time the OMS and aft RCS are on to determine the amount of propellants used.
The request for an interconnect will cause the interconnect sequence to configure the valves sequentially as follows: open the aft RCS crossfeed valves, open the left OMS crossfeed valves A, open the right OMS crossfeed valves B, close the left and right aft RCS tank isolation valves, open the left OMS crossfeed valves B and open the right OMS crossfeed valves A. The OMS-to-aft-RCS interconnect complete flag will then be set to true.
If the rapid dump is selected before MECO, the OMS-to-aft-RCS interconnect occurs, and both OMS engines and the 24 aft RCS jets are commanded to thrust until the desired amount of propellant has been consumed. The rapid dump will be interrupted during external tank separation if the thrusting period is not completed before MECO; otherwise, the thrusting period terminates when thrusting time equals zero or if the normal acceleration exceeds a threshold value.
Upon completion of the thrusting period, the OMS-to-aft-RCS configuration flag will be set to false, and the sequence will be terminated. A return-to-normal-configuration request by the abort control sequence will cause the interconnect sequence to configure the valves sequentially as follows: open aft RCS propellant tank isolation valves, close the aft RCS crossfeed valves, and close the left and right OMS crossfeed A and B valves. The OMS-to-aft-RCS interconnect complete flag will be set to false, and the sequence will be terminated.
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"OMS Engine Fault Detection and Identification",335,0,0,0
The OMS engine FDI function detects and identifies off-nominal performance of the OMS engine, such as off-failures during OMS thrusting periods, on-failures after or before a thrusting period and high or low engine chamber pressures.
Redundancy management software performs OMS engine FDI. It is assumed that the flight crew arms only the OMS engine to be used; the OMS engine not armed cannot be used for thrusting. FDI will be initialized at SRB ignition and terminated after the OMS-1 thrusting period or, in the case of an RTLS abort, at the transition from RTLS entry to the RTLS landing sequence program. The FDI also will be initiated before each OMS burn and will be terminated after the OMS thrusting period is complete.
The OMS engine FDI uses both a velocity comparison and a chamber pressure comparison method to determine a failed-on or failed-off engine. The velocity comparison is used only after MECO since the OMS thrust is small compared to main propulsion thrust before MECO.
The measured velocity increment is compared to a predetermined one-engine and two-engine acceleration threshold value by the redundancy management software to determine the number of engines actually firing. This information, along with the assumption that an armed engine is to be used, allows the software to determine if the engine has low thrust or has shut down prematurely.
The chamber pressure comparison test compares a predetermined threshold chamber pressure level to the measured chamber pressure to determine a failed engine (on, off or low thrust).
The engine-on command and the chamber pressure are used before MECO to determine a failed engine. The velocity indication and the chamber pressure indication are used after MECO to determine a failed engine. If the engine fails the chamber pressure test but passes the velocity test after MECO, the engine will be considered failed.
Such a failure would illuminate the red right OMS or left OMS caution and warning light on panel F7 and the master alarm and produce a fault message. In addition, if an engine fails the chamber pressure and velocity tests, a down arrow is displayed on the maneuver CRT next to the failed engine.
When the flight crew disarms a failed engine by turning the arm/press switch on panel C3 to off , a signal is sent to the OMS thrusting sequence to shut down the engine and to signal guidance to reconfigure. Guidance reconfigures and downmodes from two OMS engines, to one OMS engine, to four plus X RCS jets.
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"OMS Gimbal Actuator FDI",336,0,0,0
The OMS gimbal actuator FDI detects and identifies off-nominal performance of the pitch and yaw gimbal actuators of the OMS engines.
The OMS gimbal actuator FDI is divided into two processes. The first determines if the actuators should move from their present position. If the actuators must move, the second part determines how much they should move and whether the desired movement has occurred.
The first part checks the actuators' gimbal deflection error (which is the difference between the commanded new position and the actuators' last known position) and determines whether the actuators should extend or retract or if they are being driven against a stop. If the actuators are in the desired position or being driven against a stop, the first part of the process will be repeated. If the first part determines that the actuator should move, the second part of the actuator FDI process is performed.
The second part of the actuator FDI process checks the present position of each actuator against its last known position to determine whether the actuators have moved more than a threshold amount. If the actuators have not moved more than this amount, an actuator failure is incremented by one.
Each time an actuator fails this test, the failure is again incremented by one. When the actuator failure counter reaches an I-loaded value of four, the actuator is declared failed and a fault message is output. The actuator failure counter is reset to zero any time the actuator passes the threshold test.
The first and second parts of the actuator FDI process continue to perform in this manner. The actuator FDI process can detect full-off gimbal failures and full-on failures indirectly. The full-on failure determines that the gimbal has extended or retracted too far and commands reverse motion. If no motion occurs, the actuator will be declared failed. The flight crew's response to a failed actuator is to select the secondary actuator \Jelectronics\j by item entry on the maneuver CRT display.
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"Shuttle Electrical Power System",337,0,0,0
The EPS consists of three subsystems: power reactant storage and distribution, fuel cell power plants (electrical power generation) and electrical power distribution and control.
The PRSD subsystem stores the reactants (cryogenic \Jhydrogen\j and oxygen) and supplies them to the three fuel cell power plants, which generate all the electrical power for the vehicle during all mission phases. In addition, cryogenic oxygen is supplied to the environmental control and life support system for crew cabin pressurization. The \Jhydrogen\j and oxygen are stored in their respective storage tanks at cryogenic temperatures and supercritical pressures. The storage temperature of liquid oxygen is minus 285 F and minus 420 F for liquid \Jhydrogen\j.
The three fuel cell power plants, through a chemical reaction, generate all of the 28-volt direct-current electrical power for the vehicle from launch through landing rollout. Before launch, electrical power is provided by ground power supplies and the onboard fuel cell power plants until T minus three minutes and 30 seconds.
Each fuel cell power plant consists of a power section, where the chemical reaction occurs, and a compact accessory section attached to the power section, which controls and monitors the power section's performance. The three fuel cell power plants are individually coupled to the reactant (hydrogen and oxygen) distribution subsystem, the heat rejection subsystem, the potable water storage subsystem and the EPDC subsystem.
The fuel cell power plants generate heat and water as by-products of electrical power generation. The excess heat is directed to fuel cell heat exchangers, where the excess heat is rejected to Freon coolant loops. The water is directed to the potable water storage subsystem.
The EPDC subsystem distributes the 28 volts dc generated by each of the three fuel cell power plants to a three-bus system that distributes dc power to the forward, mid-, and aft sections of the orbiter for equipment in those areas. The three main dc buses-MNA, MNB and MNC-are the prime sources of power for the vehicle's dc loads.
Each of the three dc main buses supplies power to three solid-state (static), single-phase inverters, which constitute one three-phase alternating-current bus; thus, the nine inverters convert dc power to 115-volt, 400-hertz ac power for distribution to three ac buses-AC1, AC2 and AC3-for the vehicle's ac loads.
The EPDC subsystem controls and distributes electrical power (ac and dc) to the orbiter subsystems, the solid rocket boosters, the external tank and payloads. Power is controlled and distributed by assemblies. Each assembly is a housing for electrical components, such as remote switching devices, buses, resistors, diodes and fuses. Each assembly usually contains a power bus or buses and remote switching devices for distributing bus power to subsystems located in its area.
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"Shuttle Power Reactant Storage and Distribution",338,0,0,0
Cryogenic \Jhydrogen\j and oxygen are stored in a supercritical condition in double-walled, thermally insulated spherical tanks with a vacuum annulus between the inner pressure vessel and outer shell of the tank. Each tank has heaters to add energy to the reactants during depletion to control pressure. Each tank is capable of measuring quantity remaining.
The tanks are grouped in sets consisting of one \Jhydrogen\j and one oxygen tank. The number of tank sets installed depends on the specific mission requirement. Up to five tank sets can be installed. The five tank sets are all installed in the midfuselage under the payload bay liner.
The oxygen tanks are identical and consist of inner pressure vessels of Inconel 718 and outer shells of aluminum 2219. The inner vessel is 33.43 inches in diameter and the outer shell is 36.8 inches in diameter. Each tank has a volume of 11.2 cubic feet and stores 781 pounds of oxygen. The dry weight of each tank is 201 pounds. The initial temperature of the stored oxygen is minus 285 F. Maximum fill time is 45 minutes.
The \Jhydrogen\j tanks also are identical. Both the inner pressure vessel and the outer shell are constructed of aluminum 2219. The inner vessel's diameter is 41.51 inches and the outer shell's is 45.5 inches. The volume of each tank is 21.39 cubic feet, and each stores 92 pounds of \Jhydrogen\j. Each tank weighs 216 pounds dry. The initial storage temperature is minus 420 F. Maximum fill time is 45 minutes.
The inner pressure vessels are kept supercold by minimizing conductive, convective and radiant heat transfer. Twelve low-conductive supports suspend the inner vessel within the outer shell. Radiant heat transfer is reduced by a shield between the inner vessel and outer shell (hydrogen tanks only), and convective heat transfer is minimized by maintaining a vacuum between the vessel and shell. A vacuum ion pump maintains the required vacuum level and is also used as a vacuum gauge to determine the vacuum's integrity.
Dual-mode heater operation is available for pairs of oxygen and \Jhydrogen\j tanks. If the heaters of both tanks 1 and 2 or tanks 3 and 4 are placed in the automatic mode, the tank heater logic is interconnected. In this case, the heater controllers of both tanks must send a low pressure signal to the heater logic before the heaters will turn on. Once the heaters are on, a high pressure signal from either tank will turn off the heaters in both tanks.
In the manual mode, the flight crew controls the heaters by using the on/off positions for each heater switch on panel R1 or A11. High or low pressure in each tank is shown on the CRT display or the gauges on panel O2. The specific tank is selected by setting the rotary switch on panel O2.
Before lift-off, the oxygen and \Jhydrogen\j tank 1 and 2 heater switches are set on auto. After SRB separation, all the \Jhydrogen\j and oxygen tank 1 and 2 heater switches are positioned to auto, and the tank 3 and 4 heaters remain off.
On orbit, the tank 3 and 4 heater switches are positioned to auto. Because the tank 3 and 4 heater controller pressure limits are higher than those of tanks 1 and 2, tanks 3 and 4 supply the reactants to the fuel cells. For entry, the tank 3 and 4 heater switches are set to off, and tanks 1 and 2 supply the reactants to the fuel cells.
The cryo oxygen htr assy temp meter on panel O2, in conjunction with the rotary switch tk1 1-2, tk2 1-2, tk3 1-2, tk4 1-2, selects one of the two heaters in each tank and permits the temperature of the heater element to be displayed. The range of the display is from minus 425 F to plus 475 F. The temperature sensor in each heater also is hard-wired directly to the yellow O 2 heater temp caution and warning light on panel F7.
This light is illuminated if the temperature is at or above 349 F. A signal also is sent to the computers, where software checks the limit; and if the temperature is at or above 349 F, the backup C/W alarm light on panel F7 is illuminated. This signal also is transmitted to the CRT and telemetry.
Each tank set (one \Jhydrogen\j and one oxygen tank) has a hydrogen/oxygen control box that contains the electrical logic for the \Jhydrogen\j and oxygen heaters and controllers. The control box is located on cold plates in the midbody under the payload bay envelope.
The reactants from the tanks flow through two relief valve/filter package modules and valve modules and then to the fuel cells through a common manifold. Oxygen is supplied to the manifold from the tank at a pressure of 815 to 881 psia, and \Jhydrogen\j is supplied at a pressure of 200 to 243 psia. The pressure of the reactants will be essentially the same at the fuel cell interface as it is in the tanks since only a small decrease in pressure occurs in the manifolds.
The relief valve/filter package module contains the tank relief valve and a 12-micron filter. The filter removes contaminants that could affect the performance of components within the power reactant storage and distribution subsystem and fuel cells. The valve relieves excessive pressure that builds up in the tank, and a manifold valve relieves pressure in the manifold lines. The oxygen tank relief valve relieves at 1,005 psia, and the \Jhydrogen\j tank relief valve relieves at 310 psia.
The reactants flow from the relief valve/filter packages through four reactant valve modules: two \Jhydrogen\j (hydrogen valve modules 1 and 2) and two oxygen (oxygen modules 1 and 2). Each valve module contains a check valve for each cryogenic tank line to prevent the reactants from flowing from one tank to another tank in the event of a tank leak. This prevents a total loss of reactants. The oxygen valve modules also contain the environmental control and life support system atmosphere pressure control system 1 and 2 oxygen supply. Each module also contains a manifold valve and fuel cell reactant valves.
Each fuel cell reactant valve consists of two valves-one for \Jhydrogen\j and one for oxygen. The valves are controlled by the fuel cell 1, 2, 3 reac open/close switches on panel R1. When the switch is positioned to open, the \Jhydrogen\j and oxygen reactant valves for that fuel cell are opened, and reactants are allowed to flow from the manifold into the fuel cell.
When the switch is positioned to close, the \Jhydrogen\j and oxygen reactant valves for that fuel cell are closed, isolating the reactants from the fuel cell and rendering that fuel cell inoperative. Each fuel cell reac switch on panel R1 also has a talkback indicator. The corresponding talkback indicator indicates op when both valves are open and cl when either valve is closed.
Because it is critical to have reactants available to the fuel cells, the red fuel cell reac light on panel F7 is illuminated when any fuel cell reactant valve is closed, a caution/warning tone is sounded, and the computers sense the closed valve, which causes the backup C/W alarm light on panel F7 to be illuminated, an SM alert to occur, and the data to be displayed on the CRT. This alerts the flight crew that the fuel cell will be inoperative within approximately 20 seconds for a \Jhydrogen\j valve closure and 130 seconds for an oxygen valve closure.
The manifold relief valves are a built-in safety device in the event a manifold valve and fuel cell reactant valves are closed because of a malfunction. The reactants trapped in the manifold lines would be warmed up by the internal heat of the orbiter and overpressurize. The manifold relief valve will open at 290 psi for \Jhydrogen\j and 975 psi for oxygen to relieve pressure and allow the trapped reactants to flow back to their tanks.
Two pressure sensors located in the respective \Jhydrogen\j and oxygen valve modules transmit data to the CRT. This data is also sent to the systems management computer, where its lower limit is checked; and if the respective \Jhydrogen\j and oxygen manifold pressures are below 150 psia and 200 psia, respectively, an SM alert will occur.
If cryogenic tank set 5 is added to an orbiter, the displays and controls associated with controlling the tank set will be added to panel A15.
During prelaunch operations, the onboard fuel cell reactants (oxygen and hydrogen) are supplied by ground support equipment to assure a full load of onboard reactants before lift-off. At T minus two minutes 35 seconds, the GSE filling operation is terminated. The GSE supply pressure is 300 to 320 psia for \Jhydrogen\j and 1,000 to 1,020 psia for oxygen, which is higher than the onboard PRSD pressures. The GSE supply valves close automatically to transfer to onboard reactants.
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"Shuttle Fuel Cell Power Plants",339,0,0,0
Each of the three fuel cell power plants is reusable and restartable. The fuel cells are located under the payload bay area in the forward portion of the orbiter's midfuselage.
The three fuel cells operate as independent electrical power sources, each supplying its own isolated, simultaneously operating 28-volt dc bus. The fuel cell consists of a power section, where the chemical reaction occurs, and an accessory section that controls and monitors the power section's performance. The power section, where \Jhydrogen\j and oxygen are transformed into electrical power, water and heat, consists of 96 cells contained in three substacks.
Manifolds run the length of these substacks and distribute \Jhydrogen\j, oxygen and coolant to the cells. The cells contain \Jelectrolyte\j consisting of \Jpotassium\j \Jhydroxide\j and water, an oxygen electrode (cathode) and a \Jhydrogen\j electrode (anode).
The accessory section monitors the reactant flow, removes waste heat and water from the chemical reaction and controls the temperature of the stack. The accessory section consists of the \Jhydrogen\j and oxygen flow system, the coolant loop and the electrical control unit.
Oxygen is routed to the fuel cell's oxygen electrode, where it reacts with the water and returning electrons to produce hydroxyl ions. The hydroxyl ions then migrate to the \Jhydrogen\j electrode, where they enter into the \Jhydrogen\j reaction. \JHydrogen\j is routed to the fuel cell's \Jhydrogen\j electrode, where it reacts with the hydroxyl ions from the \Jelectrolyte\j.
This electrochemical reaction produces electrons (electrical power), water and heat. The electrons are routed through the orbiter's EPDC subsystem to perform electrical work. The oxygen and \Jhydrogen\j are reacted (consumed) in proportion to the orbiter's electrical power demand.
Excess water vapor is removed by an internal circulating \Jhydrogen\j system. \JHydrogen\j and water vapor from the reaction exits the cell stack, is mixed with replenishing \Jhydrogen\j from the storage and distribution system, and enters a condenser, where waste heat from the \Jhydrogen\j and water vapor is transferred to the fuel cell coolant system.
The resultant temperature decrease condenses some of the water vapor to water droplets. A centrifugal water separator extracts the liquid water and pressure-feeds it to potable tanks in the lower deck of the pressurized crew cabin. Water from the potable water storage tanks can be used for crew consumption and cooling the Freon-21 coolant loops. The remaining circulating \Jhydrogen\j is directed back to the fuel cell stack.
The fuel cell coolant system circulates a liquid fluorinated hydrocarbon and transfers the waste heat from the cell stack through the fuel cell heat exchanger of the fuel cell power plant to the Freon-21 coolant loop system in the midfuselage. Internal control of the circulating fluid maintains the cell stack at a normal operating temperature of approximately 200 F.
When the reactants enter the fuel cells, they flow through a preheater (where they are warmed from a cryogenic temperature to 40 F or greater); a 6-micron filter; and a two-stage, integrated dual gas regulator module. The first stage of the regulator reduces the pressure of the \Jhydrogen\j and oxygen to 135 to 150 psia.
The second stage reduces the oxygen pressure to a range of 62 to 65 psia and maintains the \Jhydrogen\j pressure at 4.5 to 6 psia differential below the oxygen pressure. The regulated oxygen lines are connected to the accumulator, which maintains an equalized pressure between the oxygen and the fuel cell coolant. If the oxygen's and \Jhydrogen\j's pressure decreases, the coolant's pressure is also decreased to prevent a large differential pressure inside the stack that could deform the cell stack structural elements.
Upon leaving the dual gas regulator module, the incoming \Jhydrogen\j mixes with the hydrogen-water vapor exhaust from the fuel cell stack. This saturated gas mixture is routed through a condenser, where the temperature of the mixture is reduced, condensing a portion of the water vapor to form liquid water droplets. The liquid water is then separated from the hydrogen-water mixture by the \Jhydrogen\j pump/water separator.
The \Jhydrogen\j pump circulates the \Jhydrogen\j gas back to the fuel cell stack, where some of the \Jhydrogen\j is consumed in the reaction. The remainder flows through the fuel cell stack, removing the product water vapor formed at the \Jhydrogen\j electrode. The hydrogen-water vapor mixture then combines with the regulated \Jhydrogen\j from the dual gas generator module, and the loop begins again.
The oxygen from the dual gas regulator module flows directly through two ports into a closed-end manifold in the fuel cell stack, achieving optimum oxygen distribution in the cells. All oxygen that flows into the stack is consumed, except during purge operations.
Reactant consumption is directly related to the electrical current produced: if there are no internal or external loads on the fuel cell, no reactants will be used. Because of this direct proportion, leaks may be detected by comparing reactant consumption and current produced. An appreciable amount of excess reactants used indicates a probable leak.
Water and electricity are the products of the chemical reaction of oxygen and \Jhydrogen\j that takes place in the fuel cells. The water must be removed or the cells will become saturated with water, decreasing reaction efficiency. With an operating load of about 7 kilowatts, it takes only a few minutes to flood the fuel cell with produced water, thus effectively halting power generation. \JHydrogen\j is pumped through the stack, reacting with oxygen and picking up and removing water vapor on the way. After being condensed, the liquid water is separated from the \Jhydrogen\j by the \Jhydrogen\j pump/water separator and discharged from the fuel cell to be stored in the ECLSS potable water storage tanks.
If the water tanks are full or there is line blockage, the water relief valves open at 45 psia to allow the water to vent overboard through the water relief line and nozzle. Check valves prevent water tanks from discharging through an open relief valve. An alternate water delivery path is also available to deliver water to the ECLSS tanks if the primary path is lost.
For redundancy, there are two thermostatically activated heaters wrapped around the discharge and relief lines to prevent blockage caused by the formation of ice in the lines. Two switches on panel R12, fuel cell H\D2\dO line htr and H\D2\dO relief htr , provide the flight crew with the capability to select either auto A or auto B for the fuel cell water discharge line heaters and the water relief line and vent heaters, respectively.
Thermostatically controlled heaters will maintain the water line temperature above 53 F, when required. The normal temperature of product water is approximately 140 to 150 F. The thermostatically controlled heaters maintain the water relief valve's temperature when in use between 70 to 100 F. Temperature sensors located on the fuel cell water discharge line, relief valve, relief line and vent nozzle are displayed on the CRT.
If the \Jpotassium\j \Jhydroxide\j \Jelectrolyte\j in the fuel cell migrates into the product water, a \JpH\j sensor located downstream of the \Jhydrogen\j pump/water separator will sense the presence of the \Jelectrolyte\j, and the crew will be alerted by an SM alert and display on the CRT.
During normal fuel cell operation, the reactants are present in a closed-loop system and are 100 percent consumed in the production of electricity. Any inert gases or other contaminants will accumulate in and around the porous electrodes in the cells and reduce the reaction efficiency and electrical load support capability. Purging, therefore, is required at least twice daily to cleanse the cells.
When a purge is initiated by opening the purge valves, the oxygen and \Jhydrogen\j systems become open-loop systems; and increased flows allow the reactants to circulate through the stack, pick up the contaminants and blow them out overboard through the purge lines and vents. Electrical power is produced throughout the purge sequence, although no more than 10 kilowatts should be required from a fuel cell being purged because of the increased reactant flow and preheater limitations.
Fuel cell purge can be activated automatically or manually by the use of fuel cell switches on panel R12. In the automatic mode, the fuel cell purge heater switch is positioned to GPC . The purge line hea
