Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2STS 15-L TRACKING AND DATA RELAY SATELLITE SYSTEM (TDRSS) AND TDRS-B  The Tracking and Data Relay Satellite (TDRS-B) is the second TDRSS advanced communications spacecraft to be launched from the orbiter Challenger. The first was launched during Challengers maiden flight in April 1983. TDRS-1 is now in geosynchronous orbit over the Atlantic Ocean just east of Brazil (41 degrees west longitude). It initially failed to reach its desired orbit following successful Shuttle deployment because of booster rocket failure. A NASA-industry team conducted a series of delicate spacecraft maneuvers over a 2- month period to place TDRS-1 into the desired 22,300-mile altitude. Following its deployment from the orbiter, TDRS-B will undergo a series of tests prior to being moved to its operational geosynchronous position over the Pacific Ocean south of Hawaii (171 degrees W. longitude). A third TDRSS satellite is scheduled for launch in July 1986, providing the Tracking and Data Relay Satellite System with an on-orbit spare located between the two operational satellites. TDRS-B will be identical to its sister satellite and the two-satellite configuration will support up to 23 user spacecraft simultaneously, providing two basic types of service: a multiple access service which can relay data from as many as 19 low-data-rate user spacecraft at the same time and a single access service which will provide two high-data-rate communications relays from each satellite. TDRS-B will be deployed from the orbiter approximately 10 hours after launch. Transfer to geosynchronous orbit will be provided by the solid propellant Boeing/U.S. Air Force Inertial Upper Stage (IUS). Separation from the IUS occurs approximately 17 hours after launch. The concept of using advanced communication satellites was developed following studies in the early 1970s which showed that a system of communication satellites operated from a single ground terminal could support Space Shuttle and other low Earth-orbit space missions more effectively than a world-wide network of ground stations. NASAs Space Tracking and Data Network ground stations eventually will be phased out. Three of the networks present 12 ground stations ñ Madrid, Spain; Canberra, Australia; and Goldstone, CA ñ have been transferred to the Deep Space Network managed by the Jet Propulsion Laboratory in Pasadena, CA, and the remainder ñ except for two stations considered necessary for Shuttle launch operations ñ will be closed or transferred to other agencies after the successful launch and checkout of the next two TDRS satellites. The ground station network, managed by the Goddard Space Flight Center, Greenbelt, MD, provides communications support for only a small fraction (typically 15-20 percent) of a spacecrafts orbital period. The TDRSS network of satellites, when established, will provide coverage for almost the entire orbital period of user spacecraft (about 85 percent). A TDRSS ground terminal has been built at White Sands, NM, a location that provides a clear view to the TDRSS satellites and weather conditions generally good for communications. The NASA Ground Terminal at White Sands provides the interface between the TDRSS and its network elements, which have their primary tracking and communication facilities at Goddard. Also located at Goddard is the Network Control Center, which provides system scheduling and is the focal point for NASA communications with the TDRSS satellites and network elements. The TDRSS satellites are the largest privately-owned telecommunications spacecraft ever built, each weighing about 5,000 lb. Each satellite spans more than 57 ft., measured across its solar panels. The single access antennas, fabricated of molybdenum and plated with 14k gold, each measure 16 ft. in diameter, and when deployed, span more than 42 ft. from tip to top. The satellite consists of two modules. The equipment module houses the subsystems that operate the satellite. The telecommunications payload module has electronic equipment for linking the user spacecraft with the ground terminal. The spacecraft has seven antennas. The TDRS spacecraft are the first designed to handle communications through S, Ku and C frequency bands. Under contract, NASA has leased the TDRSS service from the Space Communications Co. (Space com), Gaithersburg, MD, the owner, operator and prime contractor for the system. TRW Space and Technology Group, Redondo Beach, CA, and the Harris Government Communications System Division, Melbourne, Fl, are the two primary subcontractors to Space com for spacecraft and ground terminal equipment, respectively. TRW also provided the total software for the ground segment operation and did the integration and testing for the ground terminal and the TDRSS, as well as the systems engineering. Primary users of the TDRSS satellite have been the Space Shuttle, Landsat Earth resources satellites, the Solar Mesosphere Explorer, the Earth Radiation Budget Satellite, the Solar Maximum Mission satellite and Spacelab. Future users include the Hubble Space Telescope, scheduled for launch Oct. 27, 1986; the Gamma Ray Observatory, due to be launched in 1988; and the Upper Atmosphere Research Satellite in 1989  INERTIAL UPPER STAGE  The Inertial Upper Stage (IUS) will be used to place NASAs second Tracking and Data Relay Satellite (TDRS-B) into geosynchronous orbit. The first TDRS was launched by an IUS aboard Challenger in April 1983 during mission STS-6. The 51-L crew will deploy IUS/TDRS-B approximately 10 hours after liftoff from a low-Earth orbit of 153.5 nautical miles. Upper stage airborne support equipment, located in the orbiter payload bay, positions the combined IUS/TDRS-B into the proper deployment attitude ñ an angle of 59 degrees ñ and ejects it into low-Earth orbit. Deployment from the orbiter will be by a spring eject system. Following deployment from the payload bay, the orbiter will move away from the IUS/TDRS-B to a safe distance. The first stage will fire about 55 minutes after deployment. Following the aft (first) stage burn of two minutes, 26 seconds, the solid fuel motor will shut down and the two stages will separate. After coasting for several hours, the forward (second) stage motor will ignite at six hours, 14 minutes after deployment to place the spacecraft into its desired orbit. Following a one-minute, 49-second burn, the forward stage will shut down as the IUS/TDRS-B reaches the predetermined geosynchronous orbit position. Six hours, 54 minutes after deployment from Challenger, the forward stage will separate from TDRS-B and perform an anti-collision maneuver with its onboard reaction control system. After the IUS reaches a safe distance from TDRS-B, the upper stage will relay performance data back to a NASA tracking station and then shut itself down seven hours, five minutes after deployment from the payload bay. As wit the first NASA IUS launched in 1983, the second has a number of features which distinguish it from other previous upper stages. It has the first completely redundant avionics system ever developed for an unmanned space vehicle. The system has the capability to correct in-flight features within milliseconds. Other advanced features include a carbon composite nozzle throat that makes possible the high-temperature, long-duration firing of the IUS motors and a redundant computer system in which the second computer is capable of taking over functions from the primary computer if necessary. The IUS is 17 ft. long, 9 ft. in diameter and weights more than 32,000 lb., including 27,000 lb. of solid fuel propellant. The IUS consists of an aft skirt; an aft stage containing 21,000 lb. of solid propellant fuel, generating 45,000 lb. of thrust; an interstage; a forward stage containing 6,000 lb. of propellant, generating 18,500 lb. of thrust; and an equipment support section. The equipment support section contains the avionics which provide guidance, navigation, telemetry, command and data management, reaction control and electrical power. Solid propellant rocket motors were selected in the design of the IUS because of their compactness, simplicity, inherent safety, reliability and lower cost. The IUS is built by Boeing Aerospace Corp, Seattle, under contract to the U.S. Air Force Systems Command. Marshall Space Flight Center, Huntsville, AL, is NASAs lead center for IUS development and program management of NASA-configured IUSs procured from the Air Force.  SPARTAN-HALLEY MISSION  For the Spartan-Halley mission, NASAs Goddard Space Flight Center and the University of Colorados laboratory for Atmospheric and Space Physics (LASP) have recycled several instruments and designs to produce a low-cost, high-yield spacecraft to watch Halleys Comet when it is too close to the sun for other observatories to do so. IT will record ultraviolet light emitted by the comets chemistry when it is closest to the sun and most active so that scientists may determine how fast water is broken down by sunlight, search for carbon and sulfur atoms and related compounds, and understand how the tail evolves. Principal investigator is Dr. Charles Barth of the University of Colorado LASP. Mission manager is Morgan Windsor of Goddard Space Flight Center. The Instruments Two spectrometers, derived from backups for a Mariner 9 instrument which studied the Martian atmosphere in 1971, have been rebuilt to survey Halley’s Comet in ultraviolet light from 128 to 340 nanometers (nm) wavelength, stopping just above the human eyes limit of about 400 nm. Each spectrometer uses the Ebert-Fastie design: an off-axis reflector telescope, with magnesium fluoride coatings to enhance transmission which focuses light from Halley, via s spherical mirror and a spectral grating, on a coded anode converter with 1,024 detectors in a straight line. The grating is ruled at 2,400 lines per millimeter. The detectors are made of cesium iodide (CsI) for the G-spectrometer (128-168 nm) and cesium telluride (CsTe) for the F-spectrometer (180-340 nm). The system has a focal length of 250 mm and an aperture of 50 mm. The F-spectrometer grating can be rotated to cover its wider range in six 40 nm sections. A slit limits its field of view to a strip of sky 1 by 80 arc-minutes (the apparent diameter of the moon is about 30 arcminutes). The G-spectrometer has a 3 x 80 arc-minute slit because emissions are fainter at shorter wavelengths. With Halley as little as 10 degrees away from the sun, two sets of baffles must be used to reduce stray light. An internal set is part of the Mariner design. A new external set serves both instruments. It has two knife-edge baffles 38.5 inches away from the spectrometer entrances, and 20 secondary baffles to stop earthlight. Together, the two baffle sets reduce stray light by a factor of a trillion. It is this system that will make it possible for Spartan-Halley to observe the comet while so close to the sun. In addition, internal filters reduce solar Lyman-alpha light (121.6 nm), scattered by the Earthís hydrogen corona, which would saturate the instruments. Two film cameras, boresighted with the spectrometers, will photograph Halley to assure pointing accuracy in post-flight analysis and to match changes in the tail with spectral changes. The 35 mm Nikon F3 cameras have 105 mm and 135 mm lenses and are loaded with 65-frame rolls of QX-851 thin-base color film. The cameras will capture large-scale activity such as the separation angle between the dust and ion tails, bursts from the nucleus, and asymmetries in the shape of the coma. The whole instrument package is mounted on a n aluminum optical bench ñ 35 by 37 inches and weighing 175 lb. ñ attached to the Spartan carrier. This provides a clean interface with the carrier and aligns the spectrometers with the Spartan attitude control sensors. A 15-inch-high housing covers the spectrometers and the cameras. The instrument package is controlled by a LASP-developed microprocessor which stores the comet Halley ephemeris and directs the Spartan carrier attitude control system.   MISSION OPERATIONS  Halley’s Comet will be of greatest scientific interest from Jan. 20 to Feb. 22; perihelion is on Feb. 9. At that time, Halley will be 139.5 million miles from Earth and 59.5 million mi. from the sun. The Shuttle will go into an orbit 176 miles high and inclined 28.5 degrees to the equator. This will have Halley visible for more than 3,000 seconds per orbit (about 56 percent of the orbit), including more than 90 seconds with the sun occulted by the Earth. After a pre-deployment health check of Spartan voltages and currents, the Shuttle robot arm will pick up the spacecraft and hold it over the side. Upon release, Spartan will perform a 90-second pirouette to confirm that it is working and the Shuttle will back away to at least five miles so light reflected from the Shuttle does not confuse Spartan’s sensors. After two orbits of preparation, the 40-hour science mission will begin. A backup timer will ensure that the spectrometer doors open 70 minutes after release. Spartan-Halley will conduct 20 orbits of science observations interspersed with five orbits of attitude control updates. A typical science orbit will start with four 100-second calibration scans of Earthís atmosphere, followed by a 900-second tail scan. Observing will be interrupted for 15 minutes of pointing updates and housekeeping. It then resumes with four 200-second scans of the coma, followed by sunset and four coma scans while the sun is occulted. At the end of the mission Spartan-Halley will be retrieved by the Shuttle robot arm and placed in the payload bay. After the mission, the processed film and data tapes will be returned to the University of Colorado team for scientific analysis.  The Science  Current theories hold that comets are dirty snowballs made up largely of water ice and lightweight elements and compounds left over from the creation of the solar system. Remote sensing of the chemistry of Halley’s Comet, by measuring how sunlight is reflected, will help in assaying the comet. The dirt in the snowball is detectable in visible light, and the snow (water ice) and other gases are detectable, indirectly, in ultraviolet. The most important objective of the Spartan-Halley mission is to obtain ultraviolet spectra of comet Halley when it is less than 67 million miles from the sun. As Halley nears the sun, temperatures rise, releasing ices and clathrates, compounds trapped in ice crystals. The highest science priority for Spartan is to determine the rate at which water is broken down (dissociated) by sunlight. This must be measured indirectly from the spectra of hydroxyl radicals (OH) and atomic oxygen which are the primary and secondary products. The hydroxyl coma of the comet will be more compact than the atomic oxygen coma because of its short life when exposed to sunlight. Hydrogen, the other product, will not be detectable because of the Lyman-alpha filters in the spectrometers. Heavier compounds will be sought by measuring spectral lines unique to carbon, carbon monoxide (CO), carbon dioxide (CO2), sulfur, carbon sulfide (CS) molecular sulfur (S2), nitric oxide (NO) and cyanogen (CN), among others. Spartan-Halley’s spectrometers will not produce images, but will reveal the comet’s chemistry thought the ultraviolet spectral lines they record. With these data, scientists will gain a better understanding of how: # Chemical structure of the comet evolves from the coma and proceeds down the tail; # Species change with relation to sunlight and dynamic processes within the comet; and # Dominant atmospheric activities at perihelion relate to the comets long-term evolution. Other observatories will be studying Halley’s comet, but only Spartan can observe near perihelion. Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2STS 15-L TRACKING AND DATA RELAY SATELLITE SYSTEM (TDRSS) AND TDRS-B  The Tracking and Data Relay Satellite (TDRS-B) is the second TDRSS advanced communications spacecraft to be launched from the orbiter Challenger. The first was launched during Challengers maiden flight in April 1983. TDRS-1 is now in geosynchronous orbit over the Atlantic Ocean just east of Brazil (41 degrees west longitude). It initially failed to reach its desired orbit following successful Shuttle deployment because of booster rocket failure. A NASA-industry team conducted a series of delicate spacecraft maneuvers over a 2- month period to place TDRS-1 into the desired 22,300-mile altitude. Following its deployment from the orbiter, TDRS-B will undergo a series of tests prior to being moved to its operational geosynchronous position over the Pacific Ocean south of Hawaii (171 degrees W. longitude). A third TDRSS satellite is scheduled for launch in July 1986, providing the Tracking and Data Relay Satellite System with an on-orbit spare located between the two operational satellites. TDRS-B will be identical to its sister satellite and the two-satellite configuration will support up to 23 user spacecraft simultaneously, providing two basic types of service: a multiple access service which can relay data from as many as 19 low-data-rate user spacecraft at the same time and a single access service which will provide two high-data-rate communications relays from each satellite. TDRS-B will be deployed from the orbiter approximately 10 hours after launch. Transfer to geosynchronous orbit will be provided by the solid propellant Boeing/U.S. Air Force Inertial Upper Stage (IUS). Separation from the IUS occurs approximately 17 hours after launch. The concept of using advanced communication satellites was developed following studies in the early 1970s which showed that a system of communication satellites operated from a single ground terminal could support Space Shuttle and other low Earth-orbit space missions more effectively than a world-wide network of ground stations. NASAs Space Tracking and Data Network ground stations eventually will be phased out. Three of the networks present 12 ground stations ñ Madrid, Spain; Canberra, Australia; and Goldstone, CA ñ have been transferred to the Deep Space Network managed by the Jet Propulsion Laboratory in Pasadena, CA, and the remainder ñ except for two stations considered necessary for Shuttle launch operations ñ will be closed or transferred to other agencies after the successful launch and checkout of the next two TDRS satellites. The ground station network, managed by the Goddard Space Flight Center, Greenbelt, MD, provides communications support for only a small fraction (typically 15-20 percent) of a spacecrafts orbital period. The TDRSS network of satellites, when established, will provide coverage for almost the entire orbital period of user spacecraft (about 85 percent). A TDRSS ground terminal has been built at White Sands, NM, a location that provides a clear view to the TDRSS satellites and weather conditions generally good for communications. The NASA Ground Terminal at White Sands provides the interface between the TDRSS and its network elements, which have their primary tracking and communication facilities at Goddard. Also located at Goddard is the Network Control Center, which provides system scheduling and is the focal point for NASA communications with the TDRSS satellites and network elements. The TDRSS satellites are the largest privately-owned telecommunications spacecraft ever built, each weighing about 5,000 lb. Each satellite spans more than 57 ft., measured across its solar panels. The single access antennas, fabricated of molybdenum and plated with 14k gold, each measure 16 ft. in diameter, and when deployed, span more than 42 ft. from tip to top. The satellite consists of two modules. The equipment module houses the subsystems that operate the satellite. The telecommunications payload module has electronic equipment for linking the user spacecraft with the ground terminal. The spacecraft has seven antennas. The TDRS spacecraft are the first designed to handle communications through S, Ku and C frequency bands. Under contract, NASA has leased the TDRSS service from the Space Communications Co. (Space com), Gaithersburg, MD, the owner, operator and prime contractor for the system. TRW Space and Technology Group, Redondo Beach, CA, and the Harris Government Communications System Division, Melbourne, Fl, are the two primary subcontractors to Space com for spacecraft and ground terminal equipment, respectively. TRW also provided the total software for the ground segment operation and did the integration and testing for the ground terminal and the TDRSS, as well as the systems engineering. Primary users of the TDRSS satellite have been the Space Shuttle, Landsat Earth resources satellites, the Solar Mesosphere Explorer, the Earth Radiation Budget Satellite, the Solar Maximum Mission satellite and Spacelab. Future users include the Hubble Space Telescope, scheduled for launch Oct. 27, 1986; the Gamma Ray Observatory, due to be launched in 1988; and the Upper Atmosphere Research Satellite in 1989  INERTIAL UPPER STAGE  The Inertial Upper Stage (IUS) will be used to place NASAs second Tracking and Data Relay Satellite (TDRS-B) into geosynchronous orbit. The first TDRS was launched by an IUS aboard Challenger in April 1983 during mission STS-6. The 51-L crew will deploy IUS/TDRS-B approximately 10 hours after liftoff from a low-Earth orbit of 153.5 nautical miles. Upper stage airborne support equipment, located in the orbiter payload bay, positions the combined IUS/TDRS-B into the proper deployment attitude ñ an angle of 59 degrees ñ and ejects it into low-Earth orbit. Deployment from the orbiter will be by a spring eject system. Following deployment from the payload bay, the orbiter will move away from the IUS/TDRS-B to a safe distance. The first stage will fire about 55 minutes after deployment. Following the aft (first) stage burn of two minutes, 26 seconds, the solid fuel motor will shut down and the two stages will separate. After coasting for several hours, the forward (second) stage motor will ignite at six hours, 14 minutes after deployment to place the spacecraft into its desired orbit. Following a one-minute, 49-second burn, the forward stage will shut down as the IUS/TDRS-B reaches the predetermined geosynchronous orbit position. Six hours, 54 minutes after deployment from Challenger, the forward stage will separate from TDRS-B and perform an anti-collision maneuver with its onboard reaction control system. After the IUS reaches a safe distance from TDRS-B, the upper stage will relay performance data back to a NASA tracking station and then shut itself down seven hours, five minutes after deployment from the payload bay. As wit the first NASA IUS launched in 1983, the second has a number of features which distinguish it from other previous upper stages. It has the first completely redundant avionics system ever developed for an unmanned space vehicle. The system has the capability to correct in-flight features within milliseconds. Other advanced features include a carbon composite nozzle throat that makes possible the high-temperature, long-duration firing of the IUS motors and a redundant computer system in which the second computer is capable of taking over functions from the primary computer if necessary. The IUS is 17 ft. long, 9 ft. in diameter and weights more than 32,000 lb., including 27,000 lb. of solid fuel propellant. The IUS consists of an aft skirt; an aft stage containing 21,000 lb. of solid propellant fuel, generating 45,000 lb. of thrust; an interstage; a forward stage containing 6,000 lb. of propellant, generating 18,500 lb. of thrust; and an equipment support section. The equipment support section contains the avionics which provide guidance, navigation, telemetry, command and data management, reaction control and electrical power. Solid propellant rocket motors were selected in the design of the IUS because of their compactness, simplicity, inherent safety, reliability and lower cost. The IUS is built by Boeing Aerospace Corp, Seattle, under contract to the U.S. Air Force Systems Command. Marshall Space Flight Center, Huntsville, AL, is NASAs lead center for IUS development and program management of NASA-configured IUSs procured from the Air Force.  SPARTAN-HALLEY MISSION  For the Spartan-Halley mission, NASAs Goddard Space Flight Center and the University of Colorados laboratory for Atmospheric and Space Physics (LASP) have recycled several instruments and designs to produce a low-cost, high-yield spacecraft to watch Halleys Comet when it is too close to the sun for other observatories to do so. IT will record ultraviolet light emitted by the comets chemistry when it is closest to the sun and most active so that scientists may determine how fast water is broken down by sunlight, search for carbon and sulfur atoms and related compounds, and understand how the tail evolves. Principal investigator is Dr. Charles Barth of the University of Colorado LASP. Mission manager is Morgan Windsor of Goddard Space Flight Center. The Instruments Two spectrometers, derived from backups for a Mariner 9 instrument which studied the Martian atmosphere in 1971, have been rebuilt to survey Halley’s Comet in ultraviolet light from 128 to 340 nanometers (nm) wavelength, stopping just above the human eyes limit of about 400 nm. Each spectrometer uses the Ebert-Fastie design: an off-axis reflector telescope, with magnesium fluoride coatings to enhance transmission which focuses light from Halley, via s spherical mirror and a spectral grating, on a coded anode converter with 1,024 detectors in a straight line. The grating is ruled at 2,400 lines per millimeter. The detectors are made of cesium iodide (CsI) for the G-spectrometer (128-168 nm) and cesium telluride (CsTe) for the F-spectrometer (180-340 nm). The system has a focal length of 250 mm and an aperture of 50 mm. The F-spectrometer grating can be rotated to cover its wider range in six 40 nm sections. A slit limits its field of view to a strip of sky 1 by 80 arc-minutes (the apparent diameter of the moon is about 30 arcminutes). The G-spectrometer has a 3 x 80 arc-minute slit because emissions are fainter at shorter wavelengths. With Halley as little as 10 degrees away from the sun, two sets of baffles must be used to reduce stray light. An internal set is part of the Mariner design. A new external set serves both instruments. It has two knife-edge baffles 38.5 inches away from the spectrometer entrances, and 20 secondary baffles to stop earthlight. Together, the two baffle sets reduce stray light by a factor of a trillion. It is this system that will make it possible for Spartan-Halley to observe the comet while so close to the sun. In addition, internal filters reduce solar Lyman-alpha light (121.6 nm), scattered by the Earthís hydrogen corona, which would saturate the instruments. Two film cameras, boresighted with the spectrometers, will photograph Halley to assure pointing accuracy in post-flight analysis and to match changes in the tail with spectral changes. The 35 mm Nikon F3 cameras have 105 mm and 135 mm lenses and are loaded with 65-frame rolls of QX-851 thin-base color film. The cameras will capture large-scale activity such as the separation angle between the dust and ion tails, bursts from the nucleus, and asymmetries in the shape of the coma. The whole instrument package is mounted on a n aluminum optical bench ñ 35 by 37 inches and weighing 175 lb. ñ attached to the Spartan carrier. This provides a clean interface with the carrier and aligns the spectrometers with the Spartan attitude control sensors. A 15-inch-high housing covers the spectrometers and the cameras. The instrument package is controlled by a LASP-developed microprocessor which stores the comet Halley ephemeris and directs the Spartan carrier attitude control system.   MISSION OPERATIONS  Halley’s Comet will be of greatest scientific interest from Jan. 20 to Feb. 22; perihelion is on Feb. 9. At that time, Halley will be 139.5 million miles from Earth and 59.5 million mi. from the sun. The Shuttle will go into an orbit 176 miles high and inclined 28.5 degrees to the equator. This will have Halley visible for more than 3,000 seconds per orbit (about 56 percent of the orbit), including more than 90 seconds with the sun occulted by the Earth. After a pre-deployment health check of Spartan voltages and currents, the Shuttle robot arm will pick up the spacecraft and hold it over the side. Upon release, Spartan will perform a 90-second pirouette to confirm that it is working and the Shuttle will back away to at least five miles so light reflected from the Shuttle does not confuse Spartan’s sensors. After two orbits of preparation, the 40-hour science mission will begin. A backup timer will ensure that the spectrometer doors open 70 minutes after release. Spartan-Halley will conduct 20 orbits of science observations interspersed with five orbits of attitude control updates. A typical science orbit will start with four 100-second calibration scans of Earthís atmosphere, followed by a 900-second tail scan. Observing will be interrupted for 15 minutes of pointing updates and housekeeping. It then resumes with four 200-second scans of the coma, followed by sunset and four coma scans while the sun is occulted. At the end of the mission Spartan-Halley will be retrieved by the Shuttle robot arm and placed in the payload bay. After the mission, the processed film and data tapes will be returned to the University of Colorado team for scientific analysis.  The Science  Current theories hold that comets are dirty snowballs made up largely of water ice and lightweight elements and compounds left over from the creation of the solar system. Remote sensing of the chemistry of Halley’s Comet, by measuring how sunlight is reflected, will help in assaying the comet. The dirt in the snowball is detectable in visible light, and the snow (water ice) and other gases are detectable, indirectly, in ultraviolet. The most important objective of the Spartan-Halley mission is to obtain ultraviolet spectra of comet Halley when it is less than 67 million miles from the sun. As Halley nears the sun, temperatures rise, releasing ices and clathrates, compounds trapped in ice crystals. The highest science priority for Spartan is to determine the rate at which water is broken down (dissociated) by sunlight. This must be measured indirectly from the spectra of hydroxyl radicals (OH) and atomic oxygen which are the primary and secondary products. The hydroxyl coma of the comet will be more compact than the atomic oxygen coma because of its short life when exposed to sunlight. Hydrogen, the other product, will not be detectable because of the Lyman-alpha filters in the spectrometers. Heavier compounds will be sought by measuring spectral lines unique to carbon, carbon monoxide (CO), carbon dioxide (CO2), sulfur, carbon sulfide (CS) molecular sulfur (S2), nitric oxide (NO) and cyanogen (CN), among others. Spartan-Halley’s spectrometers will not produce images, but will reveal the comet’s chemistry thought the ultraviolet spectral lines they record. With these data, scientists will gain a better understanding of how: # Chemical structure of the comet evolves from the coma and proceeds down the tail; # Species change with relation to sunlight and dynamic processes within the comet; and # Dominant atmospheric activities at perihelion relate to the comets long-term evolution. Other observatories will be studying Halley’s comet, but only Spartan can observe near perihelion. Soyez11:Problems- Cabin Depressurization During Re-entry: The cabin depressurization on Soyuz 11 happened during the re-entry phase of the mission, as the spacecraft was returning to Earth from its stay on the Salyut 1 space station. The depressurization occurred when the Soyuz spacecraft was re-entering Earth's atmosphere, descending through the atmosphere at high speeds. Possible Sequence of Events: While the exact sequence of events leading to the cabin depressurization is not completely documented, here's a plausible scenario based on available information:  Reentry and Atmosphere Compression: During reentry, the spacecraft experiences intense heat and friction as it enters Earth's atmosphere. The capsule's heat shield is designed to protect it from the extreme temperatures generated during this phase.  Temperature and Pressure Fluctuations: The rapid deceleration and compression of air during re-entry cause the exterior of the spacecraft to heat up significantly. As the spacecraft slows down, it transitions from the vacuum of space to the denser atmosphere, leading to changes in temperature and pressure.  Potential Structural Stress: The combination of thermal stresses, aerodynamic forces, and changes in atmospheric pressure during reentry could potentially impact the structural integrity of the spacecraft, including its seals and joints.  Cabin Vent Valve Failure: It's believed that a cabin vent valve, designed to regulate the pressure inside the spacecraft, malfunctioned or failed to close properly during reentry. This could have allowed the cabin's breathable atmosphere to vent into the vacuum of space.  Rapid Depressurization: With the cabin vent valve stuck open or improperly closed, the cabin's air pressure could have rapidly dropped to near-vacuum levels. This would have led to a loss of breathable air within the cabin within a matter of seconds. Solution: Redesign of the Cabin Ventilation System: The redesign and testing of the Cabin Vent Valve (CVV) following the Soyuz 11 tragedy marked a pivotal effort in enhancing spaceflight safety and preventing cabin depressurization incidents. The CVV, a crucial element of a spacecraft's life support system, regulates cabin pressure to ensure astronaut well-being. Post-tragedy, the CVV underwent a comprehensive redesign and testing process to address its shortcomings and fortify its reliability.  The redesign process commenced with a meticulous failure analysis, delving into the causes of the CVV malfunction during Soyuz 11. Engineers scrutinized its mechanical structure, materials, and operational behavior to pinpoint vulnerabilities. Subsequent modifications aimed to bolster the CVV's mechanisms, materials, and design, mitigating the risk of unintended openings or closures.  Testing the redesigned CVV encompassed an array of critical aspects. The valve's resilience to extreme temperature fluctuations and atmospheric pressures encountered during reentry was assessed in thermal testing chambers. Operational testing simulated launch vibrations and reentry forces, mimicking real-world conditions to evaluate the valve's performance under dynamic scenarios.  Ensuring longevity and reliability, the CVV underwent prolonged testing under accelerated aging conditions. This exhaustive process subjected the valve to repeated cycles of simulated space conditions, validating its enduring functionality throughout an entire mission duration. Fail-safe mechanisms were potentially integrated, guaranteeing that in the event of a failure, the valve defaults to a secure position, averting inadvertent cabin depressurization. Emergency Escape Systems: The Soyuz spacecraft was equipped with emergency escape systems, which allowed for the safe and rapid evacuation of the crew in the event of an emergency during launch, re-entry, or landing. These systems provide an additional layer of safety and a means of escape in critical situations.  Enhanced Safety Procedures and Training: Following the Soyuz 11 incident, safety procedures and crew training were further emphasized to ensure that astronauts are adequately prepared to handle emergency situations. This includes training in emergency response, problem-solving, and critical decision-making to enhance crew members' ability to react and mitigate risks.  Improved Flight Control and Monitoring: The development of more advanced flight control systems allowed for better monitoring and communication with spacecraft during missions. Real-time diagnostic capabilities and telemetry systems help detect and address potential problems promptly, minimizing the risks to crew members.  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2STS 15-L TRACKING AND DATA RELAY SATELLITE SYSTEM (TDRSS) AND TDRS-B  The Tracking and Data Relay Satellite (TDRS-B) is the second TDRSS advanced communications spacecraft to be launched from the orbiter Challenger. The first was launched during Challengers maiden flight in April 1983. TDRS-1 is now in geosynchronous orbit over the Atlantic Ocean just east of Brazil (41 degrees west longitude). It initially failed to reach its desired orbit following successful Shuttle deployment because of booster rocket failure. A NASA-industry team conducted a series of delicate spacecraft maneuvers over a 2- month period to place TDRS-1 into the desired 22,300-mile altitude. Following its deployment from the orbiter, TDRS-B will undergo a series of tests prior to being moved to its operational geosynchronous position over the Pacific Ocean south of Hawaii (171 degrees W. longitude). A third TDRSS satellite is scheduled for launch in July 1986, providing the Tracking and Data Relay Satellite System with an on-orbit spare located between the two operational satellites. TDRS-B will be identical to its sister satellite and the two-satellite configuration will support up to 23 user spacecraft simultaneously, providing two basic types of service: a multiple access service which can relay data from as many as 19 low-data-rate user spacecraft at the same time and a single access service which will provide two high-data-rate communications relays from each satellite. TDRS-B will be deployed from the orbiter approximately 10 hours after launch. Transfer to geosynchronous orbit will be provided by the solid propellant Boeing/U.S. Air Force Inertial Upper Stage (IUS). Separation from the IUS occurs approximately 17 hours after launch. The concept of using advanced communication satellites was developed following studies in the early 1970s which showed that a system of communication satellites operated from a single ground terminal could support Space Shuttle and other low Earth-orbit space missions more effectively than a world-wide network of ground stations. NASAs Space Tracking and Data Network ground stations eventually will be phased out. Three of the networks present 12 ground stations ñ Madrid, Spain; Canberra, Australia; and Goldstone, CA ñ have been transferred to the Deep Space Network managed by the Jet Propulsion Laboratory in Pasadena, CA, and the remainder ñ except for two stations considered necessary for Shuttle launch operations ñ will be closed or transferred to other agencies after the successful launch and checkout of the next two TDRS satellites. The ground station network, managed by the Goddard Space Flight Center, Greenbelt, MD, provides communications support for only a small fraction (typically 15-20 percent) of a spacecrafts orbital period. The TDRSS network of satellites, when established, will provide coverage for almost the entire orbital period of user spacecraft (about 85 percent). A TDRSS ground terminal has been built at White Sands, NM, a location that provides a clear view to the TDRSS satellites and weather conditions generally good for communications. The NASA Ground Terminal at White Sands provides the interface between the TDRSS and its network elements, which have their primary tracking and communication facilities at Goddard. Also located at Goddard is the Network Control Center, which provides system scheduling and is the focal point for NASA communications with the TDRSS satellites and network elements. The TDRSS satellites are the largest privately-owned telecommunications spacecraft ever built, each weighing about 5,000 lb. Each satellite spans more than 57 ft., measured across its solar panels. The single access antennas, fabricated of molybdenum and plated with 14k gold, each measure 16 ft. in diameter, and when deployed, span more than 42 ft. from tip to top. The satellite consists of two modules. The equipment module houses the subsystems that operate the satellite. The telecommunications payload module has electronic equipment for linking the user spacecraft with the ground terminal. The spacecraft has seven antennas. The TDRS spacecraft are the first designed to handle communications through S, Ku and C frequency bands. Under contract, NASA has leased the TDRSS service from the Space Communications Co. (Space com), Gaithersburg, MD, the owner, operator and prime contractor for the system. TRW Space and Technology Group, Redondo Beach, CA, and the Harris Government Communications System Division, Melbourne, Fl, are the two primary subcontractors to Space com for spacecraft and ground terminal equipment, respectively. TRW also provided the total software for the ground segment operation and did the integration and testing for the ground terminal and the TDRSS, as well as the systems engineering. Primary users of the TDRSS satellite have been the Space Shuttle, Landsat Earth resources satellites, the Solar Mesosphere Explorer, the Earth Radiation Budget Satellite, the Solar Maximum Mission satellite and Spacelab. Future users include the Hubble Space Telescope, scheduled for launch Oct. 27, 1986; the Gamma Ray Observatory, due to be launched in 1988; and the Upper Atmosphere Research Satellite in 1989  INERTIAL UPPER STAGE  The Inertial Upper Stage (IUS) will be used to place NASAs second Tracking and Data Relay Satellite (TDRS-B) into geosynchronous orbit. The first TDRS was launched by an IUS aboard Challenger in April 1983 during mission STS-6. The 51-L crew will deploy IUS/TDRS-B approximately 10 hours after liftoff from a low-Earth orbit of 153.5 nautical miles. Upper stage airborne support equipment, located in the orbiter payload bay, positions the combined IUS/TDRS-B into the proper deployment attitude ñ an angle of 59 degrees ñ and ejects it into low-Earth orbit. Deployment from the orbiter will be by a spring eject system. Following deployment from the payload bay, the orbiter will move away from the IUS/TDRS-B to a safe distance. The first stage will fire about 55 minutes after deployment. Following the aft (first) stage burn of two minutes, 26 seconds, the solid fuel motor will shut down and the two stages will separate. After coasting for several hours, the forward (second) stage motor will ignite at six hours, 14 minutes after deployment to place the spacecraft into its desired orbit. Following a one-minute, 49-second burn, the forward stage will shut down as the IUS/TDRS-B reaches the predetermined geosynchronous orbit position. Six hours, 54 minutes after deployment from Challenger, the forward stage will separate from TDRS-B and perform an anti-collision maneuver with its onboard reaction control system. After the IUS reaches a safe distance from TDRS-B, the upper stage will relay performance data back to a NASA tracking station and then shut itself down seven hours, five minutes after deployment from the payload bay. As wit the first NASA IUS launched in 1983, the second has a number of features which distinguish it from other previous upper stages. It has the first completely redundant avionics system ever developed for an unmanned space vehicle. The system has the capability to correct in-flight features within milliseconds. Other advanced features include a carbon composite nozzle throat that makes possible the high-temperature, long-duration firing of the IUS motors and a redundant computer system in which the second computer is capable of taking over functions from the primary computer if necessary. The IUS is 17 ft. long, 9 ft. in diameter and weights more than 32,000 lb., including 27,000 lb. of solid fuel propellant. The IUS consists of an aft skirt; an aft stage containing 21,000 lb. of solid propellant fuel, generating 45,000 lb. of thrust; an interstage; a forward stage containing 6,000 lb. of propellant, generating 18,500 lb. of thrust; and an equipment support section. The equipment support section contains the avionics which provide guidance, navigation, telemetry, command and data management, reaction control and electrical power. Solid propellant rocket motors were selected in the design of the IUS because of their compactness, simplicity, inherent safety, reliability and lower cost. The IUS is built by Boeing Aerospace Corp, Seattle, under contract to the U.S. Air Force Systems Command. Marshall Space Flight Center, Huntsville, AL, is NASAs lead center for IUS development and program management of NASA-configured IUSs procured from the Air Force.  SPARTAN-HALLEY MISSION  For the Spartan-Halley mission, NASAs Goddard Space Flight Center and the University of Colorados laboratory for Atmospheric and Space Physics (LASP) have recycled several instruments and designs to produce a low-cost, high-yield spacecraft to watch Halleys Comet when it is too close to the sun for other observatories to do so. IT will record ultraviolet light emitted by the comets chemistry when it is closest to the sun and most active so that scientists may determine how fast water is broken down by sunlight, search for carbon and sulfur atoms and related compounds, and understand how the tail evolves. Principal investigator is Dr. Charles Barth of the University of Colorado LASP. Mission manager is Morgan Windsor of Goddard Space Flight Center. The Instruments Two spectrometers, derived from backups for a Mariner 9 instrument which studied the Martian atmosphere in 1971, have been rebuilt to survey Halley’s Comet in ultraviolet light from 128 to 340 nanometers (nm) wavelength, stopping just above the human eyes limit of about 400 nm. Each spectrometer uses the Ebert-Fastie design: an off-axis reflector telescope, with magnesium fluoride coatings to enhance transmission which focuses light from Halley, via s spherical mirror and a spectral grating, on a coded anode converter with 1,024 detectors in a straight line. The grating is ruled at 2,400 lines per millimeter. The detectors are made of cesium iodide (CsI) for the G-spectrometer (128-168 nm) and cesium telluride (CsTe) for the F-spectrometer (180-340 nm). The system has a focal length of 250 mm and an aperture of 50 mm. The F-spectrometer grating can be rotated to cover its wider range in six 40 nm sections. A slit limits its field of view to a strip of sky 1 by 80 arc-minutes (the apparent diameter of the moon is about 30 arcminutes). The G-spectrometer has a 3 x 80 arc-minute slit because emissions are fainter at shorter wavelengths. With Halley as little as 10 degrees away from the sun, two sets of baffles must be used to reduce stray light. An internal set is part of the Mariner design. A new external set serves both instruments. It has two knife-edge baffles 38.5 inches away from the spectrometer entrances, and 20 secondary baffles to stop earthlight. Together, the two baffle sets reduce stray light by a factor of a trillion. It is this system that will make it possible for Spartan-Halley to observe the comet while so close to the sun. In addition, internal filters reduce solar Lyman-alpha light (121.6 nm), scattered by the Earthís hydrogen corona, which would saturate the instruments. Two film cameras, boresighted with the spectrometers, will photograph Halley to assure pointing accuracy in post-flight analysis and to match changes in the tail with spectral changes. The 35 mm Nikon F3 cameras have 105 mm and 135 mm lenses and are loaded with 65-frame rolls of QX-851 thin-base color film. The cameras will capture large-scale activity such as the separation angle between the dust and ion tails, bursts from the nucleus, and asymmetries in the shape of the coma. The whole instrument package is mounted on a n aluminum optical bench ñ 35 by 37 inches and weighing 175 lb. ñ attached to the Spartan carrier. This provides a clean interface with the carrier and aligns the spectrometers with the Spartan attitude control sensors. A 15-inch-high housing covers the spectrometers and the cameras. The instrument package is controlled by a LASP-developed microprocessor which stores the comet Halley ephemeris and directs the Spartan carrier attitude control system.   MISSION OPERATIONS  Halley’s Comet will be of greatest scientific interest from Jan. 20 to Feb. 22; perihelion is on Feb. 9. At that time, Halley will be 139.5 million miles from Earth and 59.5 million mi. from the sun. The Shuttle will go into an orbit 176 miles high and inclined 28.5 degrees to the equator. This will have Halley visible for more than 3,000 seconds per orbit (about 56 percent of the orbit), including more than 90 seconds with the sun occulted by the Earth. After a pre-deployment health check of Spartan voltages and currents, the Shuttle robot arm will pick up the spacecraft and hold it over the side. Upon release, Spartan will perform a 90-second pirouette to confirm that it is working and the Shuttle will back away to at least five miles so light reflected from the Shuttle does not confuse Spartan’s sensors. After two orbits of preparation, the 40-hour science mission will begin. A backup timer will ensure that the spectrometer doors open 70 minutes after release. Spartan-Halley will conduct 20 orbits of science observations interspersed with five orbits of attitude control updates. A typical science orbit will start with four 100-second calibration scans of Earthís atmosphere, followed by a 900-second tail scan. Observing will be interrupted for 15 minutes of pointing updates and housekeeping. It then resumes with four 200-second scans of the coma, followed by sunset and four coma scans while the sun is occulted. At the end of the mission Spartan-Halley will be retrieved by the Shuttle robot arm and placed in the payload bay. After the mission, the processed film and data tapes will be returned to the University of Colorado team for scientific analysis.  The Science  Current theories hold that comets are dirty snowballs made up largely of water ice and lightweight elements and compounds left over from the creation of the solar system. Remote sensing of the chemistry of Halley’s Comet, by measuring how sunlight is reflected, will help in assaying the comet. The dirt in the snowball is detectable in visible light, and the snow (water ice) and other gases are detectable, indirectly, in ultraviolet. The most important objective of the Spartan-Halley mission is to obtain ultraviolet spectra of comet Halley when it is less than 67 million miles from the sun. As Halley nears the sun, temperatures rise, releasing ices and clathrates, compounds trapped in ice crystals. The highest science priority for Spartan is to determine the rate at which water is broken down (dissociated) by sunlight. This must be measured indirectly from the spectra of hydroxyl radicals (OH) and atomic oxygen which are the primary and secondary products. The hydroxyl coma of the comet will be more compact than the atomic oxygen coma because of its short life when exposed to sunlight. Hydrogen, the other product, will not be detectable because of the Lyman-alpha filters in the spectrometers. Heavier compounds will be sought by measuring spectral lines unique to carbon, carbon monoxide (CO), carbon dioxide (CO2), sulfur, carbon sulfide (CS) molecular sulfur (S2), nitric oxide (NO) and cyanogen (CN), among others. Spartan-Halley’s spectrometers will not produce images, but will reveal the comet’s chemistry thought the ultraviolet spectral lines they record. With these data, scientists will gain a better understanding of how: # Chemical structure of the comet evolves from the coma and proceeds down the tail; # Species change with relation to sunlight and dynamic processes within the comet; and # Dominant atmospheric activities at perihelion relate to the comets long-term evolution. Other observatories will be studying Halley’s comet, but only Spartan can observe near perihelion. Soyez11:Problems- Cabin Depressurization During Re-entry: The cabin depressurization on Soyuz 11 happened during the re-entry phase of the mission, as the spacecraft was returning to Earth from its stay on the Salyut 1 space station. The depressurization occurred when the Soyuz spacecraft was re-entering Earth's atmosphere, descending through the atmosphere at high speeds. Possible Sequence of Events: While the exact sequence of events leading to the cabin depressurization is not completely documented, here's a plausible scenario based on available information:  Reentry and Atmosphere Compression: During reentry, the spacecraft experiences intense heat and friction as it enters Earth's atmosphere. The capsule's heat shield is designed to protect it from the extreme temperatures generated during this phase.  Temperature and Pressure Fluctuations: The rapid deceleration and compression of air during re-entry cause the exterior of the spacecraft to heat up significantly. As the spacecraft slows down, it transitions from the vacuum of space to the denser atmosphere, leading to changes in temperature and pressure.  Potential Structural Stress: The combination of thermal stresses, aerodynamic forces, and changes in atmospheric pressure during reentry could potentially impact the structural integrity of the spacecraft, including its seals and joints.  Cabin Vent Valve Failure: It's believed that a cabin vent valve, designed to regulate the pressure inside the spacecraft, malfunctioned or failed to close properly during reentry. This could have allowed the cabin's breathable atmosphere to vent into the vacuum of space.  Rapid Depressurization: With the cabin vent valve stuck open or improperly closed, the cabin's air pressure could have rapidly dropped to near-vacuum levels. This would have led to a loss of breathable air within the cabin within a matter of seconds. Solution: Redesign of the Cabin Ventilation System: The redesign and testing of the Cabin Vent Valve (CVV) following the Soyuz 11 tragedy marked a pivotal effort in enhancing spaceflight safety and preventing cabin depressurization incidents. The CVV, a crucial element of a spacecraft's life support system, regulates cabin pressure to ensure astronaut well-being. Post-tragedy, the CVV underwent a comprehensive redesign and testing process to address its shortcomings and fortify its reliability.  The redesign process commenced with a meticulous failure analysis, delving into the causes of the CVV malfunction during Soyuz 11. Engineers scrutinized its mechanical structure, materials, and operational behavior to pinpoint vulnerabilities. Subsequent modifications aimed to bolster the CVV's mechanisms, materials, and design, mitigating the risk of unintended openings or closures.  Testing the redesigned CVV encompassed an array of critical aspects. The valve's resilience to extreme temperature fluctuations and atmospheric pressures encountered during reentry was assessed in thermal testing chambers. Operational testing simulated launch vibrations and reentry forces, mimicking real-world conditions to evaluate the valve's performance under dynamic scenarios.  Ensuring longevity and reliability, the CVV underwent prolonged testing under accelerated aging conditions. This exhaustive process subjected the valve to repeated cycles of simulated space conditions, validating its enduring functionality throughout an entire mission duration. Fail-safe mechanisms were potentially integrated, guaranteeing that in the event of a failure, the valve defaults to a secure position, averting inadvertent cabin depressurization. Emergency Escape Systems: The Soyuz spacecraft was equipped with emergency escape systems, which allowed for the safe and rapid evacuation of the crew in the event of an emergency during launch, re-entry, or landing. These systems provide an additional layer of safety and a means of escape in critical situations.  Enhanced Safety Procedures and Training: Following the Soyuz 11 incident, safety procedures and crew training were further emphasized to ensure that astronauts are adequately prepared to handle emergency situations. This includes training in emergency response, problem-solving, and critical decision-making to enhance crew members' ability to react and mitigate risks.  Improved Flight Control and Monitoring: The development of more advanced flight control systems allowed for better monitoring and communication with spacecraft during missions. Real-time diagnostic capabilities and telemetry systems help detect and address potential problems promptly, minimizing the risks to crew members.  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2STS 15-L TRACKING AND DATA RELAY SATELLITE SYSTEM (TDRSS) AND TDRS-B  The Tracking and Data Relay Satellite (TDRS-B) is the second TDRSS advanced communications spacecraft to be launched from the orbiter Challenger. The first was launched during Challengers maiden flight in April 1983. TDRS-1 is now in geosynchronous orbit over the Atlantic Ocean just east of Brazil (41 degrees west longitude). It initially failed to reach its desired orbit following successful Shuttle deployment because of booster rocket failure. A NASA-industry team conducted a series of delicate spacecraft maneuvers over a 2- month period to place TDRS-1 into the desired 22,300-mile altitude. Following its deployment from the orbiter, TDRS-B will undergo a series of tests prior to being moved to its operational geosynchronous position over the Pacific Ocean south of Hawaii (171 degrees W. longitude). A third TDRSS satellite is scheduled for launch in July 1986, providing the Tracking and Data Relay Satellite System with an on-orbit spare located between the two operational satellites. TDRS-B will be identical to its sister satellite and the two-satellite configuration will support up to 23 user spacecraft simultaneously, providing two basic types of service: a multiple access service which can relay data from as many as 19 low-data-rate user spacecraft at the same time and a single access service which will provide two high-data-rate communications relays from each satellite. TDRS-B will be deployed from the orbiter approximately 10 hours after launch. Transfer to geosynchronous orbit will be provided by the solid propellant Boeing/U.S. Air Force Inertial Upper Stage (IUS). Separation from the IUS occurs approximately 17 hours after launch. The concept of using advanced communication satellites was developed following studies in the early 1970s which showed that a system of communication satellites operated from a single ground terminal could support Space Shuttle and other low Earth-orbit space missions more effectively than a world-wide network of ground stations. NASAs Space Tracking and Data Network ground stations eventually will be phased out. Three of the networks present 12 ground stations ñ Madrid, Spain; Canberra, Australia; and Goldstone, CA ñ have been transferred to the Deep Space Network managed by the Jet Propulsion Laboratory in Pasadena, CA, and the remainder ñ except for two stations considered necessary for Shuttle launch operations ñ will be closed or transferred to other agencies after the successful launch and checkout of the next two TDRS satellites. The ground station network, managed by the Goddard Space Flight Center, Greenbelt, MD, provides communications support for only a small fraction (typically 15-20 percent) of a spacecrafts orbital period. The TDRSS network of satellites, when established, will provide coverage for almost the entire orbital period of user spacecraft (about 85 percent). A TDRSS ground terminal has been built at White Sands, NM, a location that provides a clear view to the TDRSS satellites and weather conditions generally good for communications. The NASA Ground Terminal at White Sands provides the interface between the TDRSS and its network elements, which have their primary tracking and communication facilities at Goddard. Also located at Goddard is the Network Control Center, which provides system scheduling and is the focal point for NASA communications with the TDRSS satellites and network elements. The TDRSS satellites are the largest privately-owned telecommunications spacecraft ever built, each weighing about 5,000 lb. Each satellite spans more than 57 ft., measured across its solar panels. The single access antennas, fabricated of molybdenum and plated with 14k gold, each measure 16 ft. in diameter, and when deployed, span more than 42 ft. from tip to top. The satellite consists of two modules. The equipment module houses the subsystems that operate the satellite. The telecommunications payload module has electronic equipment for linking the user spacecraft with the ground terminal. The spacecraft has seven antennas. The TDRS spacecraft are the first designed to handle communications through S, Ku and C frequency bands. Under contract, NASA has leased the TDRSS service from the Space Communications Co. (Space com), Gaithersburg, MD, the owner, operator and prime contractor for the system. TRW Space and Technology Group, Redondo Beach, CA, and the Harris Government Communications System Division, Melbourne, Fl, are the two primary subcontractors to Space com for spacecraft and ground terminal equipment, respectively. TRW also provided the total software for the ground segment operation and did the integration and testing for the ground terminal and the TDRSS, as well as the systems engineering. Primary users of the TDRSS satellite have been the Space Shuttle, Landsat Earth resources satellites, the Solar Mesosphere Explorer, the Earth Radiation Budget Satellite, the Solar Maximum Mission satellite and Spacelab. Future users include the Hubble Space Telescope, scheduled for launch Oct. 27, 1986; the Gamma Ray Observatory, due to be launched in 1988; and the Upper Atmosphere Research Satellite in 1989  INERTIAL UPPER STAGE  The Inertial Upper Stage (IUS) will be used to place NASAs second Tracking and Data Relay Satellite (TDRS-B) into geosynchronous orbit. The first TDRS was launched by an IUS aboard Challenger in April 1983 during mission STS-6. The 51-L crew will deploy IUS/TDRS-B approximately 10 hours after liftoff from a low-Earth orbit of 153.5 nautical miles. Upper stage airborne support equipment, located in the orbiter payload bay, positions the combined IUS/TDRS-B into the proper deployment attitude ñ an angle of 59 degrees ñ and ejects it into low-Earth orbit. Deployment from the orbiter will be by a spring eject system. Following deployment from the payload bay, the orbiter will move away from the IUS/TDRS-B to a safe distance. The first stage will fire about 55 minutes after deployment. Following the aft (first) stage burn of two minutes, 26 seconds, the solid fuel motor will shut down and the two stages will separate. After coasting for several hours, the forward (second) stage motor will ignite at six hours, 14 minutes after deployment to place the spacecraft into its desired orbit. Following a one-minute, 49-second burn, the forward stage will shut down as the IUS/TDRS-B reaches the predetermined geosynchronous orbit position. Six hours, 54 minutes after deployment from Challenger, the forward stage will separate from TDRS-B and perform an anti-collision maneuver with its onboard reaction control system. After the IUS reaches a safe distance from TDRS-B, the upper stage will relay performance data back to a NASA tracking station and then shut itself down seven hours, five minutes after deployment from the payload bay. As wit the first NASA IUS launched in 1983, the second has a number of features which distinguish it from other previous upper stages. It has the first completely redundant avionics system ever developed for an unmanned space vehicle. The system has the capability to correct in-flight features within milliseconds. Other advanced features include a carbon composite nozzle throat that makes possible the high-temperature, long-duration firing of the IUS motors and a redundant computer system in which the second computer is capable of taking over functions from the primary computer if necessary. The IUS is 17 ft. long, 9 ft. in diameter and weights more than 32,000 lb., including 27,000 lb. of solid fuel propellant. The IUS consists of an aft skirt; an aft stage containing 21,000 lb. of solid propellant fuel, generating 45,000 lb. of thrust; an interstage; a forward stage containing 6,000 lb. of propellant, generating 18,500 lb. of thrust; and an equipment support section. The equipment support section contains the avionics which provide guidance, navigation, telemetry, command and data management, reaction control and electrical power. Solid propellant rocket motors were selected in the design of the IUS because of their compactness, simplicity, inherent safety, reliability and lower cost. The IUS is built by Boeing Aerospace Corp, Seattle, under contract to the U.S. Air Force Systems Command. Marshall Space Flight Center, Huntsville, AL, is NASAs lead center for IUS development and program management of NASA-configured IUSs procured from the Air Force.  SPARTAN-HALLEY MISSION  For the Spartan-Halley mission, NASAs Goddard Space Flight Center and the University of Colorados laboratory for Atmospheric and Space Physics (LASP) have recycled several instruments and designs to produce a low-cost, high-yield spacecraft to watch Halleys Comet when it is too close to the sun for other observatories to do so. IT will record ultraviolet light emitted by the comets chemistry when it is closest to the sun and most active so that scientists may determine how fast water is broken down by sunlight, search for carbon and sulfur atoms and related compounds, and understand how the tail evolves. Principal investigator is Dr. Charles Barth of the University of Colorado LASP. Mission manager is Morgan Windsor of Goddard Space Flight Center. The Instruments Two spectrometers, derived from backups for a Mariner 9 instrument which studied the Martian atmosphere in 1971, have been rebuilt to survey Halley’s Comet in ultraviolet light from 128 to 340 nanometers (nm) wavelength, stopping just above the human eyes limit of about 400 nm. Each spectrometer uses the Ebert-Fastie design: an off-axis reflector telescope, with magnesium fluoride coatings to enhance transmission which focuses light from Halley, via s spherical mirror and a spectral grating, on a coded anode converter with 1,024 detectors in a straight line. The grating is ruled at 2,400 lines per millimeter. The detectors are made of cesium iodide (CsI) for the G-spectrometer (128-168 nm) and cesium telluride (CsTe) for the F-spectrometer (180-340 nm). The system has a focal length of 250 mm and an aperture of 50 mm. The F-spectrometer grating can be rotated to cover its wider range in six 40 nm sections. A slit limits its field of view to a strip of sky 1 by 80 arc-minutes (the apparent diameter of the moon is about 30 arcminutes). The G-spectrometer has a 3 x 80 arc-minute slit because emissions are fainter at shorter wavelengths. With Halley as little as 10 degrees away from the sun, two sets of baffles must be used to reduce stray light. An internal set is part of the Mariner design. A new external set serves both instruments. It has two knife-edge baffles 38.5 inches away from the spectrometer entrances, and 20 secondary baffles to stop earthlight. Together, the two baffle sets reduce stray light by a factor of a trillion. It is this system that will make it possible for Spartan-Halley to observe the comet while so close to the sun. In addition, internal filters reduce solar Lyman-alpha light (121.6 nm), scattered by the Earthís hydrogen corona, which would saturate the instruments. Two film cameras, boresighted with the spectrometers, will photograph Halley to assure pointing accuracy in post-flight analysis and to match changes in the tail with spectral changes. The 35 mm Nikon F3 cameras have 105 mm and 135 mm lenses and are loaded with 65-frame rolls of QX-851 thin-base color film. The cameras will capture large-scale activity such as the separation angle between the dust and ion tails, bursts from the nucleus, and asymmetries in the shape of the coma. The whole instrument package is mounted on a n aluminum optical bench ñ 35 by 37 inches and weighing 175 lb. ñ attached to the Spartan carrier. This provides a clean interface with the carrier and aligns the spectrometers with the Spartan attitude control sensors. A 15-inch-high housing covers the spectrometers and the cameras. The instrument package is controlled by a LASP-developed microprocessor which stores the comet Halley ephemeris and directs the Spartan carrier attitude control system.   MISSION OPERATIONS  Halley’s Comet will be of greatest scientific interest from Jan. 20 to Feb. 22; perihelion is on Feb. 9. At that time, Halley will be 139.5 million miles from Earth and 59.5 million mi. from the sun. The Shuttle will go into an orbit 176 miles high and inclined 28.5 degrees to the equator. This will have Halley visible for more than 3,000 seconds per orbit (about 56 percent of the orbit), including more than 90 seconds with the sun occulted by the Earth. After a pre-deployment health check of Spartan voltages and currents, the Shuttle robot arm will pick up the spacecraft and hold it over the side. Upon release, Spartan will perform a 90-second pirouette to confirm that it is working and the Shuttle will back away to at least five miles so light reflected from the Shuttle does not confuse Spartan’s sensors. After two orbits of preparation, the 40-hour science mission will begin. A backup timer will ensure that the spectrometer doors open 70 minutes after release. Spartan-Halley will conduct 20 orbits of science observations interspersed with five orbits of attitude control updates. A typical science orbit will start with four 100-second calibration scans of Earthís atmosphere, followed by a 900-second tail scan. Observing will be interrupted for 15 minutes of pointing updates and housekeeping. It then resumes with four 200-second scans of the coma, followed by sunset and four coma scans while the sun is occulted. At the end of the mission Spartan-Halley will be retrieved by the Shuttle robot arm and placed in the payload bay. After the mission, the processed film and data tapes will be returned to the University of Colorado team for scientific analysis.  The Science  Current theories hold that comets are dirty snowballs made up largely of water ice and lightweight elements and compounds left over from the creation of the solar system. Remote sensing of the chemistry of Halley’s Comet, by measuring how sunlight is reflected, will help in assaying the comet. The dirt in the snowball is detectable in visible light, and the snow (water ice) and other gases are detectable, indirectly, in ultraviolet. The most important objective of the Spartan-Halley mission is to obtain ultraviolet spectra of comet Halley when it is less than 67 million miles from the sun. As Halley nears the sun, temperatures rise, releasing ices and clathrates, compounds trapped in ice crystals. The highest science priority for Spartan is to determine the rate at which water is broken down (dissociated) by sunlight. This must be measured indirectly from the spectra of hydroxyl radicals (OH) and atomic oxygen which are the primary and secondary products. The hydroxyl coma of the comet will be more compact than the atomic oxygen coma because of its short life when exposed to sunlight. Hydrogen, the other product, will not be detectable because of the Lyman-alpha filters in the spectrometers. Heavier compounds will be sought by measuring spectral lines unique to carbon, carbon monoxide (CO), carbon dioxide (CO2), sulfur, carbon sulfide (CS) molecular sulfur (S2), nitric oxide (NO) and cyanogen (CN), among others. Spartan-Halley’s spectrometers will not produce images, but will reveal the comet’s chemistry thought the ultraviolet spectral lines they record. With these data, scientists will gain a better understanding of how: # Chemical structure of the comet evolves from the coma and proceeds down the tail; # Species change with relation to sunlight and dynamic processes within the comet; and # Dominant atmospheric activities at perihelion relate to the comets long-term evolution. Other observatories will be studying Halley’s comet, but only Spartan can observe near perihelion. Apollo 1  The launch simulation on January 27, 1967, on pad 34, was a "plugs-out" test to determine whether the spacecraft would operate nominally on (simulated) internal power while detached from all cables and umbilicals. Passing this test was essential to making the February 21 launch date. The test was considered non-hazardous because neither the launch vehicle nor the spacecraft was loaded with fuel or cryogenics and all pyrotechnic systems (explosive bolts) were disabled.[11] At 1:00 pm EST (1800 GMT) on January 27, first Grissom, then Chaffee, and White entered the command module fully pressure-suited, and were strapped into their seats and hooked up to the spacecraft's oxygen and communication systems. Grissom immediately noticed a strange odor in the air circulating through his suit which he compared to "sour buttermilk", and the simulated countdown was put on hold at 1:20 pm, while air samples were taken. No cause of the odor could be found, and the countdown was resumed at 2:42 pm. The accident investigation found this odor not to be related to the fire.[11] Three minutes after the count was resumed the hatch installation was started. The hatch consisted of three parts: a removable inner hatch which stayed inside the cabin; a hinged outer hatch which was part of the spacecraft's heat shield; and an outer hatch cover which was part of the boost protective cover enveloping the entire command module to protect it from aerodynamic heating during launch and from launch escape rocket exhaust in the event of a launch abort. The boost hatch cover was partially, but not fully, latched in place because the flexible boost protective cover was slightly distorted by some cabling run under it to provide the simulated internal power (the spacecraft's fuel cell reactants were not loaded for this test). After the hatches were sealed, the air in the cabin was replaced with pure oxygen at 16.7 psi (115 kPa), 2 psi (14 kPa) higher than atmospheric pressure.[11][17]: Enclosure V-21, [181]  Movement by the astronauts was detected by the spacecraft's inertial measurement unit and the astronauts' biomedical sensors, and also indicated by increases in oxygen spacesuit flow, and sounds from Grissom's stuck-open microphone. The stuck microphone was part of a problem with the communications loop connecting the crew, the Operations and Checkout Building, and the Complex 34 blockhouse control room. The poor communications led Grissom to remark: "How are we going to get to the Moon if we can't talk between two or three buildings?" The simulated countdown was put on hold again at 5:40 pm while attempts were made to troubleshoot the communications problem. All countdown functions up to the simulated internal power transfer had been successfully completed by 6:20 pm, and at 6:30 the count remained on hold at T minus 10 minutes.[11]   The crew members were using the time to run through their checklist again, when a momentary increase in AC Bus 2 voltage occurred. Nine seconds later (at 6:31:04.7), one of the astronauts (some listeners and laboratory analysis indicate Grissom) exclaimed "Hey!", "Fire!",[17]: 5–8  or "Flame!";[23] this was followed by two seconds of scuffling sounds through Grissom's open microphone. This was immediately followed at 6:31:06.2 (23:31:06.2 GMT) by someone (believed by most listeners, and supported by laboratory analysis, to be Chaffee) saying, "[I've, or We've] got a fire in the cockpit." After 6.8 seconds of silence, a second, badly garbled transmission was heard by various listeners as: •	"They're fighting a bad fire—Let's get out ... Open 'er up", •	"We've got a bad fire—Let's get out ... We're burning up", or •	"I'm reporting a bad fire ... I'm getting out ..." The transmission lasted 5.0 seconds and ended with a cry of pain.[17]: 5–8, 5–9  Some blockhouse witnesses said that they saw White on the television monitors, reaching for the inner hatch release handle[11] as flames in the cabin spread from left to right.[17]: 5–3  The heat of the fire fed by pure oxygen caused the pressure to rise to 29 psi (200 kPa), which ruptured the command module's inner wall at 6:31:19 (23:31:19 GMT, initial phase of the fire). Flames and gases then rushed outside the command module through open access panels to two levels of the pad service structure. The intense heat, dense smoke, and ineffective gas masks designed for toxic fumes rather than smoke, hampered the ground crew's attempts to rescue the men. There were fears the command module had exploded, or soon would, and that the fire might ignite the solid fuel rocket in the launch escape tower above the command module, which would have likely killed nearby ground personnel, and possibly have destroyed the pad.[11] As the pressure was released by the cabin rupture, the rush of gases within the module caused flames to spread across the cabin, beginning the second phase. The third phase began when most of the oxygen was consumed and was replaced with atmospheric air, essentially quenching the fire, but causing high concentrations of carbon monoxide and heavy smoke to fill the cabin, and large amounts of soot to be deposited on surfaces as they cooled.[11][17]: 5–3, 5–4  It took five minutes for the pad workers to open all three hatch layers, and they could not drop the inner hatch to the cabin floor as intended, so they pushed it out of the way to one side. Although the cabin lights remained on, they were unable to see the astronauts through the dense smoke. As the smoke cleared they found the bodies, but were not able to remove them. The fire had partly melted Grissom's and White's nylon space suits and the hoses connecting them to the life support system. Grissom had removed his restraints and was lying on the floor of the spacecraft. White's restraints were burned through, and he was found lying sideways just below the hatch. It was determined that he had tried to open the hatch per the emergency procedure, but was not able to do so against the internal pressure. Chaffee was found strapped into his right-hand seat, as procedure called for him to maintain communication until White opened the hatch. Because of the large strands of melted nylon fusing the astronauts to the cabin interior, removing the bodies took nearly 90 minutes.[11] Deke Slayton was possibly the first NASA official to examine the spacecraft's interior.[24] His testimony contradicted the official report concerning the position of Grissom's body. Slayton said of Grissom and White's bodies, "it is very difficult for me to determine the exact relationships of these two bodies. They were sort of jumbled together, and I couldn't really tell which head even belonged to which body at that point. I guess the only thing that was real obvious is that both bodies were at the lower edge of the hatch. They were not in the seats. They were almost completely clear of the seat areas.     As a result of the in-flight failure of the Gemini 8 mission on March 17, 1966, NASA Deputy Administrator Robert Seamans wrote and implemented Management Instruction 8621.1 on April 14, 1966, defining Mission Failure Investigation Policy And Procedures. This modified NASA's existing accident procedures, based on military aircraft accident investigation, by giving the Deputy Administrator the option of performing independent investigations of major failures, beyond those for which the various Program Office officials were normally responsible. It declared, "It is NASA policy to investigate and document the causes of all major mission failures which occur in the conduct of its space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations."[26] Immediately after the fire NASA Administrator James E. Webb asked President Lyndon B. Johnson to allow NASA to handle the investigation according to its established procedure, promising to be truthful in assessing blame, and to keep the appropriate leaders of Congress informed.[27] Seamans then directed establishment of the Apollo 204 Review Board chaired by Langley Research Center director Floyd L. Thompson, which included astronaut Frank Borman, spacecraft designer Maxime Faget, and six others. On February 1, Cornell University professor Frank A. Long left the board,[28] and was replaced by Robert W. Van Dolah of the U.S. Bureau of Mines.[29] The next day North American's chief engineer for Apollo, George Jeffs, also left.[30] Seamans ordered all Apollo 1 hardware and software impounded, to be released only under control of the board. After thorough stereo photographic documentation of the CM-012 interior, the board ordered its disassembly using procedures tested by disassembling the identical CM-014 and conducted a thorough investigation of every part. The board also reviewed the astronauts' autopsy results and interviewed witnesses. Seamans sent Webb weekly status reports of the investigation's progress, and the board issued its final report on April 5, 1967   According to the Board, Grissom suffered severe third-degree burns on over one-third of his body and his spacesuit was mostly destroyed. White suffered third-degree burns on almost half of his body and a quarter of his spacesuit had melted away. Chaffee suffered third-degree burns over almost a quarter of his body and a small portion of his spacesuit was damaged. The autopsy report determined that the primary cause of death for all three astronauts was cardiac arrest caused by high concentrations of carbon monoxide. Burns suffered by the crew were not believed to be major factors, and it was concluded that most of them had occurred postmortem. Asphyxiation occurred after the fire melted the astronauts' suits and oxygen tubes, exposing them to the lethal atmosphere of the cabin.  Major causes of accident[edit] The review board identified several major factors which combined to cause the fire and the astronauts' deaths:[11] •	An ignition source most probably related to "vulnerable wiring carrying spacecraft power" and "vulnerable plumbing carrying a combustible and corrosive coolant" •	A pure oxygen atmosphere at higher than atmospheric pressure •	A cabin sealed with a hatch cover which could not be quickly removed at high pressure •	An extensive distribution of combustible materials in the cabin •	Inadequate emergency preparedness (rescue or medical assistance, and crew escape)  Apollo 13: Problems-  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 oxygen tanks were highly insulated spherical tanks which held liquid oxygen with a fill line and heater running down the centre. The no. 2 oxygen tank used in Apollo 13 (North American Rockwell; serial number 10024X-TA0008) 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 Centre. 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 pre-flight testing, tank no. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. (On the ground, the tanks were emptied by forcing oxygen gas into the tank and forcing the liquid oxygen out, in space there was no need to empty the tanks.) The heaters in the tanks were normally used for very short periods to heat the interior slightly, increasing the pressure to keep the oxygen flowing. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65-volt DC power. This probably damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches welded shut, allowing the temperature within the tank to rise locally to over 1000 degrees F. The gauges measuring the temperature inside the tank were designed to measure only to 80 F, so the extreme heating was not noticed. The high temperature emptied the tank, but also resulted in serious damage to the Teflon insulation 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 for the third "cryo-stir" of the mission, a procedure to stir the liquid oxygen inside the tank which would tend to stratify. The exposed fan wires shorted and the teflon insulation caught fire in the pure oxygen environment. This fire rapidly heated and increased the pressure of the oxygen inside the tank, and may have spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the pressure, 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. Solution: The lunar module had charged batteries and full oxygen tanks for use on the lunar surface, so Kranz directed that the astronauts power up the LM and use it as a "lifeboat"– a scenario anticipated but considered unlikely. Procedures for using the LM in this way had been developed by LM flight controllers after a training simulation for Apollo 10 in which the LM was needed for survival, but could not be powered up in time. Had Apollo 13's accident occurred on the return voyage, with the LM already jettisoned, the astronauts would have died, as they would have following an explosion in lunar orbit, including one while Lovell and Haise walked on the Moon.  A key decision was the choice of return path. A "direct abort" would use the SM's main engine (the Service Propulsion System or SPS) to return before reaching the Moon. However, the accident could have damaged the SPS, and the fuel cells would have to last at least another hour to meet its power requirements, so Kranz instead decided on a longer route: the spacecraft would swing around the Moon before heading back to Earth. Apollo 13 was on the hybrid trajectory which was to take it to Fra Mauro; it now needed to be brought back to a free return. The LM's Descent Propulsion System (DPS), although not as powerful as the SPS, could do this, but new software for Mission Control's computers needed to be written by technicians as it had never been contemplated that the CSM/LM spacecraft would have to be maneuverer from the LM. As the CM was being shut down, Lovell copied down its guidance system's orientation information and performed hand calculations to transfer it to the LM's guidance system, which had been turned off; at his request Mission Control checked his figures. At 61:29:43.49 the DPS burn of 34.23 seconds took Apollo 13 back to a free return trajectory.  Challenges during this solution: 1.Life Support Challenges and Carbon Dioxide Scrubber: Problem: With the oxygen tank explosion, the lunar module was repurposed as a lifeboat, but it had limited resources to support three astronauts for an extended period. Solution: The team on Earth devised a solution to manage the increasing levels of carbon dioxide in the lunar module. They developed a procedure to construct a makeshift carbon dioxide scrubber using materials available on the spacecraft. The procedure involved fitting square lithium hydroxide canisters, designed for the command module's round connectors, using duct tape and plastic bags. The crew followed these instructions, effectively reducing the levels of carbon dioxide to safe levels and maintaining breathable air.  2.Course Corrections and Safe Re-entry: Problem: With the loss of propulsion in the service module, Apollo 13 needed a series of precise engine burns to adjust its trajectory for a safe return to Earth. Solution: Mission control calculated a series of engine burns using the lunar module's descent engine. The crew had to manually input these calculations into the guidance system using paper and pencil. The calculated burns were essential to ensure that the spacecraft would hit the re-entry corridor in Earth's atmosphere, which was a small target from such a distance.  3.Power-up of Command Module and Re-entry: Problem: The command module had been shut down to conserve power, and it needed to be safely powered up for re-entry. Solution: Mission control developed a procedure to guide the crew through the process of safely powering up the command module. This included reactivating the command module's systems and ensuring that critical components were operational for re-entry. The crew followed these instructions successfully, allowing the command module to be ready for the re-entry process.  Apollo 17  Command Module The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module. The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, an auxiliary battery, three cryogenic oxygen and three cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, the subsystems the main propulsion unit, and a Scientific Instrument Module (SIM) bay which held a package of science instruments and cameras to be operated from lunar orbit. Two helium tanks were mounted in the central cylinder. Electrical power system radiators were at the top of the cylinder and environmental control radiator panels spaced around the bottom.   Spacecraft and Subsystems As the name implies, the Command and Service Module (CSM) comprised two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. Apollo 17 was the third of the Apollo J-series spacecraft. The CSM mass of 30,320 kg was the launch mass including propellants and expendables, of this the Command Module (CM-114) had a mass of 5960 kg and the Service Module (SM-114) 24,360 kg. Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.    Lunar Module Spacecraft and Subsystems The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 16456 kg was the total mass of the LM ascent and descent stages including propellants (fuel and oxidizer). The dry mass of the ascent stage was 2260 kg and it held 2387 kg of propellant. The descent stage dry mass (including stowed surface equipment) was 2935 kg and 8874 kg of propellant were onboard initially. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage served as a platform for launching the ascent stage and was left behind on the Moon. The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was launched from the Moon at the end of lunar surface operations and returned the astronauts to the CSM. The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer. Apollo Lunar Surface Experiments Package (ALSEP) The Apollo Lunar Surface Experiments Package (ALSEP) consisted of a set of scientific instruments emplaced at the landing site by the astronauts. The instruments were arrayed around a central station which supplied power to run the instruments and communications so data collected by the experiments could be relayed to Earth. The central station was a 25 kg box with a stowed volume of 34,800 cubic cm. Thermal control was achieved by passive elements (insulation, reflectors, thermal coatings) as well as power dissipation resistors and heaters. Communications with Earth were achieved through a 58 cm long, 3.8 cm diameter modified axial-helical antenna mounted on top of the central station and pointed towards Earth by the astronauts. Transmitters, receivers, data processors and multiplexers were housed within the central station. Data collected from the instruments were converted into a telemetry format and transmitted to Earth. The ALSEP system and instruments were controlled by commands from Earth. The uplink frequency for all Apollo mission ALSEP's was 2119 MHz, the downlink frequency for the Apollo 17 ALSEP was 2275.5 MHz. Radioisotope Thermoelectric Generator (RTG) The SNAP-27 model RTG produced the power to run the ALSEP operations. The generator consisted of a 46 cm high central cylinder and eight radiating rectangular fins with a total tip-to-tip diameter of 40 cm. The central cylinder had a thinner concentric inner cylinder inside, and the two cylinders were attached along their surfaces by 442 spring-loaded lead-telluride thermoelectric couples mounted radially along the length of the cylinders. The generator assembly had a total mass of 17 kg. The power source was an approximately 4 kg fuel capsule in the shape of a long rod which contained plutonium-238 and was placed in the inner cylinder of the RTG by the astronauts on deployment. Plutonium-238 decays with a half-life of 89.6 years and produces heat. This heat would conduct from the inner cylinder to the outer via the thermocouples which would convert the heat directly to electrical power. Excess heat on the outer cylinder would be radiated to space by the fins. The RTG produced approximately 70 W DC at 16 V. (63.5 W after one year.) The electricity was routed through a cable to a power conditioning unit and a power distribution unit in the central station to supply the correct voltage and power to each instrument. ALSEP Scientific Instruments All ALSEP instruments were deployed on the surface by the astronauts and attached to the central station by cables. The Apollo 17 ALSEP instruments consisted of: (1) a heat flow experiment, designed to measure the rate of heat loss from the lunar interior and the thermal properties of lunar material; (2) a lunar surface gravimeter, designed to measure the lunar surface gravity and its temporal variations at a selected point on the surface; (3) a lunar mass spectrometer, designed to measure the composition of the tenuous lunar atmosphere; (4) a lunar seismic profiling experiment, to study the physical properties of lunar surface and subsurface materials and the structure of the local near-surface layers; and (5) a lunar ejecta and meteorites experiment, designed to measure the speed, direction, energy, and momentum of cosmic dust particles and lunar ejecta. The central station, located at 20.1921 N latitude, 30.7649 E longitude, was turned on at 02:53 UT on 12 December 1972 and shut down along with the other ALSEP stations on 30 September 1977.    Challenger Space Shuttle Disaster (STS-51-L): Problem- 1. O-ring seal failure occurred in one of the solid rocket boosters (SRBs) used to propel the Space Shuttle into orbit. The failure of the O-ring allowed hot gases to escape, resulting in the structural failure of the SRB and the subsequent breakup of the Space Shuttle. The O-rings were designed to seal the joints between sections of the SRBs and prevent leakage of hot gases during launch. However, in the case of the Challenger mission, the O-rings failed to maintain a proper seal due to the cold temperatures on the day of the launch. The low temperatures caused the rubber material of the O-rings to harden, compromising their ability to form a reliable seal. The failure of the O-ring seal was a critical design flaw that ultimately led to  the catastrophic failure of the Challenger Space Shuttle.  2.PoorCommunication among different teams The Challenger disaster exposed several communication issues within NASA that contributed to the tragic outcome. Some of the communication issues were as follows: 1.	Lack of Information Sharing: There was a failure to effectively share and communicate crucial information regarding the concerns about the O-ring seals. Despite engineers raising concerns about the performance of the O-rings in colder temperatures, this information was not adequately communicated to decision-makers, resulting in a lack of awareness about the potential risks. 2.	Siloed Decision-Making: Decision-making within NASA was siloed, meaning that information and decisions were not effectively shared across different teams and departments. This led to a lack of comprehensive understanding of the risks associated with the mission and hindered the ability to make informed decisions. 3.	Deficient Communication Channels: The existing communication channels were insufficient for facilitating effective communication and information exchange. This hindered the flow of critical information and contributed to a lack of awareness about the risks and challenges associated with the Space Shuttle launch. 4.	Inadequate Feedback Loops: There was a lack of effective feedback loops where concerns and information from lower-level engineers could reach decision-makers in a timely manner. This prevented decision-makers from having access to complete and accurate information needed to make informed decisions.   Solution developed : 1.O-Ring Improvement: Key Improvements in  O-Ring: 1.Redesign of the Joint: The joint design of the solid rocket boosters (SRBs) was modified to enhance the reliability and performance of the O-ring seals. The redesigned joint included additional features and mechanisms to ensure a more robust and secure seal, reducing the risk of failure.  2.Improved Sealing Materials: NASA worked on developing and using improved sealing materials for the O-rings. The goal was to select materials that could better withstand extreme temperatures and provide a reliable seal even in challenging conditions.  3.Enhanced Inspection and Testing Procedures: More rigorous inspection and testing procedures were implemented to ensure the integrity of the O-ring seals. This included more extensive pre-launch testing, which involved subjecting the O-rings to various environmental conditions and evaluating their performance under simulated launch conditions.  4.Thorough Analysis of Failure Modes: NASA conducted detailed analyses to better understand the failure modes and vulnerabilities of the O-ring seals. This involved studying the performance of the seals under different conditions, as well as investigating the factors that contributed to the failure in the Challenger disaster. These analyses helped identify the necessary improvements needed to enhance the reliability of the seals.  5.Quality Control Measures: Improved quality control measures were implemented to ensure the consistency and reliability of the O-ring seals. This involved strict adherence to manufacturing processes, as well as enhanced inspection and verification checks throughout the production process.  Redundancy and Contingency Planning: NASA implemented measures to provide redundancy and contingency plans for the O-ring seals. This involved designing backup systems and alternative methods for sealing the joints to ensure a reliable seal in case of any failures or anomalies. 2.Improving Communication: The failures in communication and decision-making that contributed to the Challenger disaster prompted NASA to prioritize open and transparent communication among teams and to ensure that decisions are made based on reliable data and thorough analysis. 3.Developing Crew space system: A crew escape system is designed to enable the safe evacuation of astronauts in the event of an emergency during launch or ascent. By providing a means of escape, it can significantly increase the chances of survival in the event of a catastrophic failure.  Following the Challenger disaster, NASA implemented the Crew Escape System on the Space Shuttle fleet. This system included the launch escape system (LES), which could propel the crew module away from the Space Shuttle in case of an emergency during launch or the early stages of ascent.  The crew escape system adds an extra layer of safety and redundancy to the overall design of a spacecraft.      absl-py==1.4.0 accelerate==0.22.0 aiohttp==3.8.5 aiosignal==1.3.1 altgraph==0.17.3 anyio==3.7.1 argon2-cffi==21.3.0 argon2-cffi-bindings==21.2.0 arrow==1.2.3 asttokens==2.2.1 astunparse==1.6.3 async-lru==2.0.4 async-timeout==4.0.3 attrs==23.1.0 Babel==2.12.1 backcall==0.2.0 beautifulsoup4==4.12.2 bleach==6.0.0 bs4==0.0.1 cachetools==5.3.1 certifi==2023.5.7 cffi==1.15.1 charset-normalizer==3.2.0 click==8.1.3 colorama==0.4.6 comm==0.1.3 debugpy==1.6.7 decorator==5.1.1 defusedxml==0.7.1 distlib==0.3.6 EasyProcess==1.1 entrypoint2==1.1 et-xmlfile==1.1.0 exceptiongroup==1.1.2 executing==1.2.0 fastjsonschema==2.18.0 filelock==3.10.3 Flask==2.2.3 flatbuffers==23.5.26 fqdn==1.5.1 frozenlist==1.4.0 fsspec==2023.6.0 gast==0.4.0 google-auth==2.22.0 google-auth-oauthlib==1.0.0 google-pasta==0.2.0 grpcio==1.57.0 h11==0.14.0 h5py==3.9.0 html5lib==1.1 huggingface-hub==0.16.4 idna==3.4 ipykernel==6.25.0 ipython==8.14.0 ipython-genutils==0.2.0 ipywidgets==8.1.0 isoduration==20.11.0 itsdangerous==2.1.2 jedi==0.19.0 Jinja2==3.1.2 json5==0.9.14 jsonpointer==2.4 jsonschema==4.18.4 jsonschema-specifications==2023.7.1 jupyter==1.0.0 jupyter-console==6.6.3 jupyter-events==0.7.0 jupyter-lsp==2.2.0 jupyter_client==8.3.0 jupyter_core==5.3.1 jupyter_server==2.7.0 jupyter_server_terminals==0.4.4 jupyterlab==4.0.3 jupyterlab-pygments==0.2.2 jupyterlab-widgets==3.0.8 jupyterlab_server==2.24.0 keras==2.13.1 libclang==16.0.6 Markdown==3.4.4 MarkupSafe==2.1.2 matplotlib-inline==0.1.6 mistune==3.0.1 MouseInfo==0.1.3 mpmath==1.3.0 mss==9.0.1 multidict==6.0.4 nbclient==0.8.0 nbconvert==7.7.3 nbformat==5.9.2 nest-asyncio==1.5.7 networkx==3.1 notebook==7.0.1 notebook_shim==0.2.3 numpy==1.24.3 oauthlib==3.2.2 openai==0.27.9 openpyxl==3.1.2 opt-einsum==3.3.0 outcome==1.2.0 overrides==7.3.1 packaging==23.1 pandas==2.0.3 pandocfilters==1.5.0 parso==0.8.3 pefile==2023.2.7 pickleshare==0.7.5 Pillow==10.0.0 platformdirs==3.1.1 prometheus-client==0.17.1 prompt-toolkit==3.0.39 protobuf==4.24.1 psutil==5.9.5 pure-eval==0.2.2 pyasn1==0.5.0 pyasn1-modules==0.3.0 PyAutoGUI==0.9.54 pycparser==2.21 PyGetWindow==0.0.9 Pygments==2.15.1 pyinstaller==5.13.0 pyinstaller-hooks-contrib==2023.6 PyMsgBox==1.0.9 pyperclip==1.8.2 PyRect==0.2.0 pyscreenshot==3.1 PyScreeze==0.1.29 PySocks==1.7.1 python-dateutil==2.8.2 python-json-logger==2.0.7 pytweening==1.0.7 pytz==2023.3 pywin32==306 pywin32-ctypes==0.2.2 pywinpty==2.0.11 PyYAML==6.0.1 pyzmq==25.1.0 qtconsole==5.4.3 QtPy==2.3.1 referencing==0.30.0 regex==2023.8.8 requests==2.31.0 requests-oauthlib==1.3.1 rfc3339-validator==0.1.4 rfc3986-validator==0.1.1 rpds-py==0.9.2 rsa==4.9 safetensors==0.3.3 selenium==4.10.0 Send2Trash==1.8.2 six==1.16.0 sniffio==1.3.0 sortedcontainers==2.4.0 soupsieve==2.4.1 stack-data==0.6.2 sympy==1.12 tensorboard==2.13.0 tensorboard-data-server==0.7.1 tensorflow==2.13.0 tensorflow-estimator==2.13.0 tensorflow-intel==2.13.0 tensorflow-io-gcs-filesystem==0.31.0 termcolor==2.3.0 terminado==0.17.1 tinycss2==1.2.1 tokenizers==0.13.3 torch==2.0.1 tornado==6.3.2 tqdm==4.66.1 traitlets==5.9.0 transformers==4.32.0 trio==0.22.2 trio-websocket==0.10.3 typing_extensions==4.5.0 tzdata==2023.3 uri-template==1.3.0 urllib3==1.26.16 virtualenv==20.21.0 wcwidth==0.2.6 webcolors==1.13 webencodings==0.5.1 websocket-client==1.6.1 Werkzeug==2.2.3 widgetsnbextension==4.0.8 wrapt==1.15.0 wsproto==1.2.0 yarl==1.9.2STS 15-L TRACKING AND DATA RELAY SATELLITE SYSTEM (TDRSS) AND TDRS-B  The Tracking and Data Relay Satellite (TDRS-B) is the second TDRSS advanced communications spacecraft to be launched from the orbiter Challenger. The first was launched during Challengers maiden flight in April 1983. TDRS-1 is now in geosynchronous orbit over the Atlantic Ocean just east of Brazil (41 degrees west longitude). It initially failed to reach its desired orbit following successful Shuttle deployment because of booster rocket failure. A NASA-industry team conducted a series of delicate spacecraft maneuvers over a 2- month period to place TDRS-1 into the desired 22,300-mile altitude. Following its deployment from the orbiter, TDRS-B will undergo a series of tests prior to being moved to its operational geosynchronous position over the Pacific Ocean south of Hawaii (171 degrees W. longitude). A third TDRSS satellite is scheduled for launch in July 1986, providing the Tracking and Data Relay Satellite System with an on-orbit spare located between the two operational satellites. TDRS-B will be identical to its sister satellite and the two-satellite configuration will support up to 23 user spacecraft simultaneously, providing two basic types of service: a multiple access service which can relay data from as many as 19 low-data-rate user spacecraft at the same time and a single access service which will provide two high-data-rate communications relays from each satellite. TDRS-B will be deployed from the orbiter approximately 10 hours after launch. Transfer to geosynchronous orbit will be provided by the solid propellant Boeing/U.S. Air Force Inertial Upper Stage (IUS). Separation from the IUS occurs approximately 17 hours after launch. The concept of using advanced communication satellites was developed following studies in the early 1970s which showed that a system of communication satellites operated from a single ground terminal could support Space Shuttle and other low Earth-orbit space missions more effectively than a world-wide network of ground stations. NASAs Space Tracking and Data Network ground stations eventually will be phased out. Three of the networks present 12 ground stations ñ Madrid, Spain; Canberra, Australia; and Goldstone, CA ñ have been transferred to the Deep Space Network managed by the Jet Propulsion Laboratory in Pasadena, CA, and the remainder ñ except for two stations considered necessary for Shuttle launch operations ñ will be closed or transferred to other agencies after the successful launch and checkout of the next two TDRS satellites. The ground station network, managed by the Goddard Space Flight Center, Greenbelt, MD, provides communications support for only a small fraction (typically 15-20 percent) of a spacecrafts orbital period. The TDRSS network of satellites, when established, will provide coverage for almost the entire orbital period of user spacecraft (about 85 percent). A TDRSS ground terminal has been built at White Sands, NM, a location that provides a clear view to the TDRSS satellites and weather conditions generally good for communications. The NASA Ground Terminal at White Sands provides the interface between the TDRSS and its network elements, which have their primary tracking and communication facilities at Goddard. Also located at Goddard is the Network Control Center, which provides system scheduling and is the focal point for NASA communications with the TDRSS satellites and network elements. The TDRSS satellites are the largest privately-owned telecommunications spacecraft ever built, each weighing about 5,000 lb. Each satellite spans more than 57 ft., measured across its solar panels. The single access antennas, fabricated of molybdenum and plated with 14k gold, each measure 16 ft. in diameter, and when deployed, span more than 42 ft. from tip to top. The satellite consists of two modules. The equipment module houses the subsystems that operate the satellite. The telecommunications payload module has electronic equipment for linking the user spacecraft with the ground terminal. The spacecraft has seven antennas. The TDRS spacecraft are the first designed to handle communications through S, Ku and C frequency bands. Under contract, NASA has leased the TDRSS service from the Space Communications Co. (Space com), Gaithersburg, MD, the owner, operator and prime contractor for the system. TRW Space and Technology Group, Redondo Beach, CA, and the Harris Government Communications System Division, Melbourne, Fl, are the two primary subcontractors to Space com for spacecraft and ground terminal equipment, respectively. TRW also provided the total software for the ground segment operation and did the integration and testing for the ground terminal and the TDRSS, as well as the systems engineering. Primary users of the TDRSS satellite have been the Space Shuttle, Landsat Earth resources satellites, the Solar Mesosphere Explorer, the Earth Radiation Budget Satellite, the Solar Maximum Mission satellite and Spacelab. Future users include the Hubble Space Telescope, scheduled for launch Oct. 27, 1986; the Gamma Ray Observatory, due to be launched in 1988; and the Upper Atmosphere Research Satellite in 1989  INERTIAL UPPER STAGE  The Inertial Upper Stage (IUS) will be used to place NASAs second Tracking and Data Relay Satellite (TDRS-B) into geosynchronous orbit. The first TDRS was launched by an IUS aboard Challenger in April 1983 during mission STS-6. The 51-L crew will deploy IUS/TDRS-B approximately 10 hours after liftoff from a low-Earth orbit of 153.5 nautical miles. Upper stage airborne support equipment, located in the orbiter payload bay, positions the combined IUS/TDRS-B into the proper deployment attitude ñ an angle of 59 degrees ñ and ejects it into low-Earth orbit. Deployment from the orbiter will be by a spring eject system. Following deployment from the payload bay, the orbiter will move away from the IUS/TDRS-B to a safe distance. The first stage will fire about 55 minutes after deployment. Following the aft (first) stage burn of two minutes, 26 seconds, the solid fuel motor will shut down and the two stages will separate. After coasting for several hours, the forward (second) stage motor will ignite at six hours, 14 minutes after deployment to place the spacecraft into its desired orbit. Following a one-minute, 49-second burn, the forward stage will shut down as the IUS/TDRS-B reaches the predetermined geosynchronous orbit position. Six hours, 54 minutes after deployment from Challenger, the forward stage will separate from TDRS-B and perform an anti-collision maneuver with its onboard reaction control system. After the IUS reaches a safe distance from TDRS-B, the upper stage will relay performance data back to a NASA tracking station and then shut itself down seven hours, five minutes after deployment from the payload bay. As wit the first NASA IUS launched in 1983, the second has a number of features which distinguish it from other previous upper stages. It has the first completely redundant avionics system ever developed for an unmanned space vehicle. The system has the capability to correct in-flight features within milliseconds. Other advanced features include a carbon composite nozzle throat that makes possible the high-temperature, long-duration firing of the IUS motors and a redundant computer system in which the second computer is capable of taking over functions from the primary computer if necessary. The IUS is 17 ft. long, 9 ft. in diameter and weights more than 32,000 lb., including 27,000 lb. of solid fuel propellant. The IUS consists of an aft skirt; an aft stage containing 21,000 lb. of solid propellant fuel, generating 45,000 lb. of thrust; an interstage; a forward stage containing 6,000 lb. of propellant, generating 18,500 lb. of thrust; and an equipment support section. The equipment support section contains the avionics which provide guidance, navigation, telemetry, command and data management, reaction control and electrical power. Solid propellant rocket motors were selected in the design of the IUS because of their compactness, simplicity, inherent safety, reliability and lower cost. The IUS is built by Boeing Aerospace Corp, Seattle, under contract to the U.S. Air Force Systems Command. Marshall Space Flight Center, Huntsville, AL, is NASAs lead center for IUS development and program management of NASA-configured IUSs procured from the Air Force.  SPARTAN-HALLEY MISSION  For the Spartan-Halley mission, NASAs Goddard Space Flight Center and the University of Colorados laboratory for Atmospheric and Space Physics (LASP) have recycled several instruments and designs to produce a low-cost, high-yield spacecraft to watch Halleys Comet when it is too close to the sun for other observatories to do so. IT will record ultraviolet light emitted by the comets chemistry when it is closest to the sun and most active so that scientists may determine how fast water is broken down by sunlight, search for carbon and sulfur atoms and related compounds, and understand how the tail evolves. Principal investigator is Dr. Charles Barth of the University of Colorado LASP. Mission manager is Morgan Windsor of Goddard Space Flight Center. The Instruments Two spectrometers, derived from backups for a Mariner 9 instrument which studied the Martian atmosphere in 1971, have been rebuilt to survey Halley’s Comet in ultraviolet light from 128 to 340 nanometers (nm) wavelength, stopping just above the human eyes limit of about 400 nm. Each spectrometer uses the Ebert-Fastie design: an off-axis reflector telescope, with magnesium fluoride coatings to enhance transmission which focuses light from Halley, via s spherical mirror and a spectral grating, on a coded anode converter with 1,024 detectors in a straight line. The grating is ruled at 2,400 lines per millimeter. The detectors are made of cesium iodide (CsI) for the G-spectrometer (128-168 nm) and cesium telluride (CsTe) for the F-spectrometer (180-340 nm). The system has a focal length of 250 mm and an aperture of 50 mm. The F-spectrometer grating can be rotated to cover its wider range in six 40 nm sections. A slit limits its field of view to a strip of sky 1 by 80 arc-minutes (the apparent diameter of the moon is about 30 arcminutes). The G-spectrometer has a 3 x 80 arc-minute slit because emissions are fainter at shorter wavelengths. With Halley as little as 10 degrees away from the sun, two sets of baffles must be used to reduce stray light. An internal set is part of the Mariner design. A new external set serves both instruments. It has two knife-edge baffles 38.5 inches away from the spectrometer entrances, and 20 secondary baffles to stop earthlight. Together, the two baffle sets reduce stray light by a factor of a trillion. It is this system that will make it possible for Spartan-Halley to observe the comet while so close to the sun. In addition, internal filters reduce solar Lyman-alpha light (121.6 nm), scattered by the Earthís hydrogen corona, which would saturate the instruments. Two film cameras, boresighted with the spectrometers, will photograph Halley to assure pointing accuracy in post-flight analysis and to match changes in the tail with spectral changes. The 35 mm Nikon F3 cameras have 105 mm and 135 mm lenses and are loaded with 65-frame rolls of QX-851 thin-base color film. The cameras will capture large-scale activity such as the separation angle between the dust and ion tails, bursts from the nucleus, and asymmetries in the shape of the coma. The whole instrument package is mounted on a n aluminum optical bench ñ 35 by 37 inches and weighing 175 lb. ñ attached to the Spartan carrier. This provides a clean interface with the carrier and aligns the spectrometers with the Spartan attitude control sensors. A 15-inch-high housing covers the spectrometers and the cameras. The instrument package is controlled by a LASP-developed microprocessor which stores the comet Halley ephemeris and directs the Spartan carrier attitude control system.   MISSION OPERATIONS  Halley’s Comet will be of greatest scientific interest from Jan. 20 to Feb. 22; perihelion is on Feb. 9. At that time, Halley will be 139.5 million miles from Earth and 59.5 million mi. from the sun. The Shuttle will go into an orbit 176 miles high and inclined 28.5 degrees to the equator. This will have Halley visible for more than 3,000 seconds per orbit (about 56 percent of the orbit), including more than 90 seconds with the sun occulted by the Earth. After a pre-deployment health check of Spartan voltages and currents, the Shuttle robot arm will pick up the spacecraft and hold it over the side. Upon release, Spartan will perform a 90-second pirouette to confirm that it is working and the Shuttle will back away to at least five miles so light reflected from the Shuttle does not confuse Spartan’s sensors. After two orbits of preparation, the 40-hour science mission will begin. A backup timer will ensure that the spectrometer doors open 70 minutes after release. Spartan-Halley will conduct 20 orbits of science observations interspersed with five orbits of attitude control updates. A typical science orbit will start with four 100-second calibration scans of Earthís atmosphere, followed by a 900-second tail scan. Observing will be interrupted for 15 minutes of pointing updates and housekeeping. It then resumes with four 200-second scans of the coma, followed by sunset and four coma scans while the sun is occulted. At the end of the mission Spartan-Halley will be retrieved by the Shuttle robot arm and placed in the payload bay. After the mission, the processed film and data tapes will be returned to the University of Colorado team for scientific analysis.  The Science  Current theories hold that comets are dirty snowballs made up largely of water ice and lightweight elements and compounds left over from the creation of the solar system. Remote sensing of the chemistry of Halley’s Comet, by measuring how sunlight is reflected, will help in assaying the comet. The dirt in the snowball is detectable in visible light, and the snow (water ice) and other gases are detectable, indirectly, in ultraviolet. The most important objective of the Spartan-Halley mission is to obtain ultraviolet spectra of comet Halley when it is less than 67 million miles from the sun. As Halley nears the sun, temperatures rise, releasing ices and clathrates, compounds trapped in ice crystals. The highest science priority for Spartan is to determine the rate at which water is broken down (dissociated) by sunlight. This must be measured indirectly from the spectra of hydroxyl radicals (OH) and atomic oxygen which are the primary and secondary products. The hydroxyl coma of the comet will be more compact than the atomic oxygen coma because of its short life when exposed to sunlight. Hydrogen, the other product, will not be detectable because of the Lyman-alpha filters in the spectrometers. Heavier compounds will be sought by measuring spectral lines unique to carbon, carbon monoxide (CO), carbon dioxide (CO2), sulfur, carbon sulfide (CS) molecular sulfur (S2), nitric oxide (NO) and cyanogen (CN), among others. Spartan-Halley’s spectrometers will not produce images, but will reveal the comet’s chemistry thought the ultraviolet spectral lines they record. With these data, scientists will gain a better understanding of how: # Chemical structure of the comet evolves from the coma and proceeds down the tail; # Species change with relation to sunlight and dynamic processes within the comet; and # Dominant atmospheric activities at perihelion relate to the comets long-term evolution. Other observatories will be studying Halley’s comet, but only Spartan can observe near perihelion. Soyez11:Problems- Cabin Depressurization During Re-entry: The cabin depressurization on Soyuz 11 happened during the re-entry phase of the mission, as the spacecraft was returning to Earth from its stay on the Salyut 1 space station. The depressurization occurred when the Soyuz spacecraft was re-entering Earth's atmosphere, descending through the atmosphere at high speeds. Possible Sequence of Events: While the exact sequence of events leading to the cabin depressurization is not completely documented, here's a plausible scenario based on available information:  Reentry and Atmosphere Compression: During reentry, the spacecraft experiences intense heat and friction as it enters Earth's atmosphere. The capsule's heat shield is designed to protect it from the extreme temperatures generated during this phase.  Temperature and Pressure Fluctuations: The rapid deceleration and compression of air during re-entry cause the exterior of the spacecraft to heat up significantly. As the spacecraft slows down, it transitions from the vacuum of space to the denser atmosphere, leading to changes in temperature and pressure.  Potential Structural Stress: The combination of thermal stresses, aerodynamic forces, and changes in atmospheric pressure during reentry could potentially impact the structural integrity of the spacecraft, including its seals and joints.  Cabin Vent Valve Failure: It's believed that a cabin vent valve, designed to regulate the pressure inside the spacecraft, malfunctioned or failed to close properly during reentry. This could have allowed the cabin's breathable atmosphere to vent into the vacuum of space.  Rapid Depressurization: With the cabin vent valve stuck open or improperly closed, the cabin's air pressure could have rapidly dropped to near-vacuum levels. This would have led to a loss of breathable air within the cabin within a matter of seconds. Solution: Redesign of the Cabin Ventilation System: The redesign and testing of the Cabin Vent Valve (CVV) following the Soyuz 11 tragedy marked a pivotal effort in enhancing spaceflight safety and preventing cabin depressurization incidents. The CVV, a crucial element of a spacecraft's life support system, regulates cabin pressure to ensure astronaut well-being. Post-tragedy, the CVV underwent a comprehensive redesign and testing process to address its shortcomings and fortify its reliability.  The redesign process commenced with a meticulous failure analysis, delving into the causes of the CVV malfunction during Soyuz 11. Engineers scrutinized its mechanical structure, materials, and operational behavior to pinpoint vulnerabilities. Subsequent modifications aimed to bolster the CVV's mechanisms, materials, and design, mitigating the risk of unintended openings or closures.  Testing the redesigned CVV encompassed an array of critical aspects. The valve's resilience to extreme temperature fluctuations and atmospheric pressures encountered during reentry was assessed in thermal testing chambers. Operational testing simulated launch vibrations and reentry forces, mimicking real-world conditions to evaluate the valve's performance under dynamic scenarios.  Ensuring longevity and reliability, the CVV underwent prolonged testing under accelerated aging conditions. This exhaustive process subjected the valve to repeated cycles of simulated space conditions, validating its enduring functionality throughout an entire mission duration. Fail-safe mechanisms were potentially integrated, guaranteeing that in the event of a failure, the valve defaults to a secure position, averting inadvertent cabin depressurization. Emergency Escape Systems: The Soyuz spacecraft was equipped with emergency escape systems, which allowed for the safe and rapid evacuation of the crew in the event of an emergency during launch, re-entry, or landing. These systems provide an additional layer of safety and a means of escape in critical situations.  Enhanced Safety Procedures and Training: Following the Soyuz 11 incident, safety procedures and crew training were further emphasized to ensure that astronauts are adequately prepared to handle emergency situations. This includes training in emergency response, problem-solving, and critical decision-making to enhance crew members' ability to react and mitigate risks.  Improved Flight Control and Monitoring: The development of more advanced flight control systems allowed for better monitoring and communication with spacecraft during missions. Real-time diagnostic capabilities and telemetry systems help detect and address potential problems promptly, minimizing the risks to crew members.