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SOYEZ11.txt

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+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.
+

apollo1.txt

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+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)
+

apollo13.txt

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+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.
+

apollo17.txt

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+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.
+ 
+

challengerspXCESHUTTLE.txt

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+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.
+
+
+
+ 

chandrayaan2.txt

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+The U.S. National Aeronautics and Space Administration (NASA) contributed two instruments, the Moon Mineralogy Mapper (M3) and the Miniature Synthetic Aperture Radar (Mini-SAR), which sought ice at the poles. M3 studied the lunar surface in wavelengths from the visible to the infrared in order to isolate signatures of different minerals on the surface. It found small amounts of water and hydroxyl radicals on the Moon’s surface. M3 also discovered in a crater near the Moon’s equator evidence for water coming from beneath the surface. Mini-SAR broadcast polarized radio waves at the north and south polar regions. Changes in the polarization of the echo measured the dielectric constant and porosity, which are related to the presence of water ice. The European Space Agency (ESA) had two other experiments, an infrared spectrometer and a solar wind monitor. The Bulgarian Aerospace Agency provided a radiation monitor.
+The principal instruments from ISRO—the Terrain Mapping Camera, the HyperSpectral Imager, and the Lunar Laser Ranging Instrument—produced images of the lunar surface with high spectral and spatial resolution, including stereo images with a 5-metre (16-foot) resolution and global topographic maps with a resolution of 10 metres (33 feet). The Chandrayaan Imaging X-ray Spectrometer, developed by ISRO and ESA, was designed to detect magnesium, aluminum, silicon, calcium, titanium, and iron by the X-rays they emit when exposed to solar flares. This was done in part with the Solar X-Ray Monitor, which measured incoming solar radiation.
+Chandrayaan-1 operations were originally planned to last two years, but the mission ended on August 28, 2009, when radio contact was lost with the spacecraft.
+Chandrayaan-2 launched on July 22, 2019, from Sriharikota on a Geosynchronous Satellite Launch Vehicle Mark III. The spacecraft consisted of an orbiter, a lander, and a rover. The orbiter circles the Moon in a polar orbit at a height of 100 km (62 miles) and has a planned mission lifetime of seven and a half years. The mission’s Vikram lander (named after ISRO founder Vikram Sarabhai) was planned to land on September 7. Vikram carried the small (27-kg [60-pound]) Pragyan (Sanskrit: “Wisdom”) rover. Both Vikram and Pragyan were designed to operate for 1 lunar day (14 Earth days). However, just before Vikram was to touch down on the Moon, contact was lost at an altitude of 2 km (1.2 miles).
+The Chandrayaan-2 mission  was successfully launched on 22nd July 2019 at 14:43 hrs by GSLV MkIII-M1 from Satish Dhawan Space Centre (SDSC), Sriharikota. After a series of Earth bound manoeuvres, the spacecraft entered into Lunar Transfer Trajectory (LTT) on August 14th. Lunar Orbit Insertion (LOI) manoeuvre was performed on August 20th, thereby Chandrayaan-2 was successfully inserted into the elliptical orbit around the Moon. This was followed by a series of Lunar bound orbit maneuvers for reducing the orbit to circular polar orbit around the Moon.
+Chandrayaan-2 Orbiter is currently in a 100 km x 100 km orbit around the Moon On September 2nd, Vikram lander separated from the Orbiter and de-orbiting maneuver was performed to reduce the orbit to 35 km x 101 km. Vikram landing was attempted on 7th September and it followed the planned descent trajectory from its orbit of 35 km to around 2 km above the surface. Communication with lander and ground station was lost. All the systems and sensors of the Lander functioned excellently until this point and proved many new technologies such as variable thrust propulsion technology used in the Lander. However, the Orbiter is healthy and all the payloads are operational.
+ 
+Chandrayaan-2 carried eight experiment payload on board for studying surface geology, composition and exospheric measurements of Moon. These measurements will continue to enhance upon the understanding built from previous lunar missions.
+1.	Chandrayaan-2 Large Area Soft X-ray Spectrometer (CLASS)
+CLASS measures the Moon’s X-ray Fluorescence (XRF) spectra to examine the presence of major elements such as Magnesium, Aluminium, Silicon, Calcium, Titanium, Iron, and Sodium. The XRF technique will detect these elements by measuring the characteristic X-rays they emit when excited by solar X-ray emission. CLASS is a non-imaging spectrometer with gold coated copper collimators defining the field of view of each detector as 7 deg X 7 deg Full Width at Half Maximum (FWHM). This translates to a 12.5 km X 12.5 km footprint at the 100 km altitude of the spacecraft. Aluminium door to protect the sensors from the high energy particle flux in the radiation belts en-route to the Moon also houses radioactive isotopes with a Titanium foil for calibration onboard.
+ 
+2.	Solar X-ray Monitor (XSM)
+XSM detects X-rays emitted by the Sun and its corona, measures its intensity, and supports the CLASS payload. It provides the solar X-ray spectrum in the energy range of 1-15 keV incident on the lunar surface. XSM provides high energy resolution and high-cadence measurements (full spectrum every second) as input for analysis of data from CLASS. The XSM is operating in the lunar orbit from 12th September 2019. The XSM provides disk integrated solar spectra in the energy range of 1 – 15 keV with a spectral resolution of better than 180 eV at 5.9 keV, which is the best available among similar instruments that carried out such measurements till now. The XSM also offers the highest time cadence for such instruments: full spectrum every second and light curves in three energy bands every 100 ms. The unique design features of the XSM allows observations over a wide dynamic range of X-ray fluxes from the quiet Sun to X-class flares. Presently, the XSM is the only instrument operational providing soft X-ray spectral measurements of the Sun over a broad energy range.
+ 
+3.	CHandra’s Atmospheric Compositional Explorer 2 (CHACE 2)
+CHACE 2 will expand upon the CHACE experiment on Chandrayaan-1. It is a Quadrupole Mass Spectrometer (QMA) capable of studies of the lunar neutral exosphere in the mass range of 1 to 300 amu with the mass resolution of ~0.5 amu. CHACE 2’s primary objective is to carry out an in-situ study of the composition and distribution of the lunar neutral exosphere and its variability. The CHACE-2 instrument consists of the sensor probe (cylindrical in shape) and the electronics. The sensor probe consists of a built-in electron impact ionizer along with a Bayard-Alpert collector to measure the total pressure; a set of four quadrupole rods and a detector assembly. The detector assembly consists of a Faraday Cup (FC) and a Channel Electron Multiplier (CEM).
+ 
+4.	Dual Frequency Synthetic Aperture Radar (DFSAR)
+The dual frequency (L and S) SAR will provide enhanced capabilities compared to Chandrayaan-1’s S-band mini SAR in areas such as:
+o	L-band for greater depth of penetration (About 5m — twice that of S-band)
+o	Circular and full polarimetry — with a range of resolution options (2-75 m) and incident angles (9°-35°) — for understanding scattering properties of permanently shadowed regions
+o	The main scientific objectives of this payload are:
+	Quantitative estimation of water-ice in the polar regions
+	High-resolution lunar mapping in the polar regions
+	Estimation of regolith thickness and its distribution
+The Dual Frequency Synthetic Aperture Radar (DFSAR) on board Chandrayaan-2 Orbiter is a microwave imaging instrument in L- and S-band frequencies and the first fully-polarimetric SAR to study the Moon. L-band frequency enables double the surface penetration capability with respect to what is obtained using S-band. The instrument is basically two SAR systems (for L & S bands) sharing a common antenna aperture with wide bandwidth.
+ 
+Various instrument features like high-efficiency transmitter, low-noise high-gain receiver, onboard range-compression (first for any ISRO SAR mission) have enabled a highly sensitive instrument with polarimetric capability. Its best resolution (2m in slant-range) is one order better than the previously flown SARs to the Moon. The backscattered signals from the targets are coherently measured by the DFSAR in different polarizations to enable studies of physical and dielectric properties of the lunar surface/ shallow-subsurface. With these polarimetric measurements, the instrument primarily aims to unambiguously address the presence of water-ice in permanently shadowed regions (PSRs), characterizing the physical and dielectric properties of lunar surface, volcanic features, impact craters and their associated ejecta.
+5.	Imaging Infra-Red Spectrometer (IIRS)
+Imaging Infra-Red Spectrometer (IIRS) is a hyper-spectral optical imaging instrument. This instrument maps geomorphology and mineralogy of lunar surface. The mission is intended to cover the Moon surface. The prime objectives of IIRS are:
+o	Global mineralogical and volatile mapping of the Moon in the spectral range of ~0.8-5.0 μm for the first time, at the high resolution of ~20 nm
+o	Complete characterization of water/hydroxyl feature near 3.0 μm for the first time at high spatial (~80 m) and spectral (~20 nm) resolutions
+IIRS measures the reflected solar radiation along with the emissions from the lunar surface at an altitude of 100 km on a polar circular orbit at a spatial resolution of ~80 m and spectral resolution of ~20-25 nm across the spectral range of 0.8-5.0 μm in ~250 spectrally contiguous bands. The diagnostic absorption features of major and minor lunar minerals are found to occur in the spectral domain of ~0.75-2.5 μm that fall well within the spectral range of IIRS thereby making their detection possible by the spectrometer. On the other hand, the spectral range of ~2.5-3.3 μm is being dedicatedly used to detect the presence of lunar OH/H2O features having fundamental absorptions around 3.0 μm.
+ 
+6.	Terrain Mapping Camera (TMC 2)
+TMC 2 is a miniature version of the Terrain Mapping Camera on Chandrayaan-1 mission. Its primary objective is to map the lunar surface in the panchromatic spectral band (0.5-0.8 microns) with a high spatial resolution of 5 m and a swath of 20 km from 100 km lunar polar orbit. The data collected by TMC 2 will give us clues about the Moon’s evolution and help us prepare 3D maps of the lunar surface. This camera enables in preparing global high resolution image mosaic and Digital Elevation Model (DEM).
+ 
+7.	Orbiter High Resolution Camera (OHRC)
+OHRC provides high-resolution images of the landing site which ensure the Lander’s safe touchdown by detecting any craters or boulders, prior to separation. The images it captures, taken from two different look angles, serve dual purposes. First, these images are used to generate DEMs (Digital Elevation Models) of the landing site. Second, they are used for scientific research after its initial role in the landing phase. OHRC’s images can capture the same area on the lunar surface from two different orbits. The coverage area in this case is of 12 km x 3 km with ground resolution of 0.32 m. OHRC is an optical camera system based on Time Delay Integration (TDI) imaging sensors with 12000 detectors. It has 4 TDI settings and 7 different integration times.
+ 
+8.	Dual Frequency Radio Science (DFRS) Experiment
+To study the temporal evolution of electron density in the Lunar ionosphere. Two coherent signals at X (8496 MHz), and S (2240 MHz) band are transmitted simultaneously from satellite, and received at ground-based receivers. DFRS is a radio science experiment used to analyze the planetary/lunar atmosphere ionosphere. It uses two highly correlated radio frequencies in X- and S-bands. Since there is no need for any specific data to be sent for the DFRS experiment, the signals allocated for tele-command and ranging can also be used to carry out the said experiment.
+ 
+Major results from Chandrayaan-2
+Science results from Chandrayaan-2 payloads were documented and released to public on the occasion of two-year completion of the mission. In addition to this, few science results from payloads are provided below.
+CHandra’s Atmospheric Composition Explorer-2 (CHACE-2) onboard Chandrayaan-2 orbiter is a quadrupole based neutral mass spectrometer aimed at observing the tenuous Lunar exospheric composition from spacecraft altitude. Argon-40 (Ar-40) is a noble gas in the lunar exosphere, understood to be originated from the radiogenic potassimum-40 and it had been detected by several previous missions, mostly covering the equatorial and low-latitude regions of Moon. The CHACE-2 not only made observations over the low-latitude regions, but also covered the other latitude regions as well, in-situ from a polar orbit, for the first time. Figure 2 shows the map of the surface densities of Ar-40 estimated from CHACE-2 observations. The Ar-40 distribution is depicted both in terms of Solar longitudes (Fig. 1a) and Selenographic longitudes (Fig. 1b), covering both low- and mid-latitude regions. The observations show that the diurnal trend agrees with LACE/Apollo observations from the low latitude region on the lunar surface. In addition, CHACE-2 observations show for the first time that these features extend to the mid-latitude regions. Further, the number density of Ar-40 is seen to exhibit significant spatial heterogeneity. CHACE-2 observations showed Ar-40 enhancements over certain longitude sectors including KREEP (Potassium, Rare Earth Elements and Phosphorous rich region on the Moon) region and the South Pole Aitken (SPA) terrain. These observations call for a deeper understanding of the surface-exosphere interactions and source distribution.
+ 
+Lunar lobate scarps are relatively small-scale tectonic landforms that are interpreted to be the surface expression of low angle thrust faults. Lobate scarps are believed to be young lunar landforms and are observed both in the mare and highland regions. A plausible lobate scarp (Length is 1416 m and average relief across this scarp is 24 m) was mapped by the Terrain Mapping Camera (TMC-2). This NW-SE oriented scarp is located between Dorsa Geike and Dorsa Mawson. It is estimated that this lobate scarp could have been formed in the Copernican period. Image acquisition at low Sun elevation angle provides opportunity to map the features having smaller dimensions such as lobate scarps.
+ 
+a.	Ortho image of TMC-2 of lobate scarp region. The plausible Lobate scarp (yellow colored line), craters (red colored circles) used for surface age estimation and Hanging Wall (HW), Foot Wall (FW) are indicated by blue polygons. Profiles P1-P4 are shown in the red colored lines.
+b.	Topographic cross-sections along the profiles P1-P4.
+XSM is carrying out broadband spectroscopy of the Sun from lunar orbit. Currently XSM is the only X-ray spectrometer in the world which regularly measures the soft X-ray spectrum of the Sun with the highest time cadence. This has yielded very interesting observations of the microflares occurring outside active region as well as elemental abundances in the quiet Sun corona. XSM has also observed number of B-class flares and their analysis has yielded unprecedented observations of variation of the elemental abundances during such flares
+XSM observed nine B-class flares ranging from B1.3 to B4.5 during the minimum phase of Solar Cycle 24. The evolution of temperature, emission measure, and absolute elemental abundances of four elements Mg, Al, Si, and S are examined. These are the first measurements of absolute abundances during such small flares and this study offers a unique insight into the evolution of absolute abundances as the flares evolve. The results demonstrate that the abundances of these four elements decrease towards their photospheric values during the peak phase of the flares. During the decay phase, the abundances are observed to quickly return to their pre-flare coronal values as shown in below figure.
+ 
+The six panels show the results of the time resolved X-ray spectroscopy for a representative flare 
+

columbia.txt

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+
+
+COLUMBIA:-
+Physical Characteristics The Space Shuttle was the most complex space vehicle design of its time. It was comprised of four main components: the External Tank (ET); three Space Shuttle Main Engines; two Solid Rocket Boosters (SRBs); and the Orbiter vehicle. It was the first side-mounted space system dictated by the need to have a large winged vehicle for cross-range capability for re-entry into Earth’s atmosphere and the ability to land a heavyweight payload. These four components provided the shuttle with the ability to accomplish a diverse set of missions over its flight history. The Orbiter’s heavy cargo/payload carrying capability, along with the crew habitability and flexibility to operate in space, made this vehicle unique. Because of its lift capability and due-East inclination, the shuttle was able to launch a multitude of satellites, Spacelab modules, science platforms, interplanetary probes, Department of Defense payloads, and components/modules for the assembly of the International Space Station (ISS). The shuttle lift capability or payload decreased with increased operational altitude or orbit inclination because more fuel was required to reach the higher altitude or inclination. Shuttle lift capability was also limited by total vehicle landing weight— different limits for different cases (nominal or abort landing). An abort landing was required if a system failure
+during ascent caused the shuttle not to have enough energy to reach orbit or was a hazard to crew or mission. Abort landing sites were located around the world, with the prime abort landing sites being Kennedy Space Center (KSC) in Florida, Dryden Flight Research Center on the Edwards Air Force Base in California, and Europe. The entire shuttle vehicle, fully loaded, weighed about 2 million kg (4.4 million pounds) and required a combined thrust of about 35 million newtons (7.8 million pounds-force) to reach orbital altitude. Thrust was provided by the boosters for the first 2 minutes and the main engines for the approximately 8 minutes and 30 seconds ascent required for the vehicle to reach orbital speed at the requisite altitude range of 185 to about 590 km (100 to 320 nautical miles). Once in orbit, the Orbital Maneuvering System engines and Reaction Control System thrusters were used to perform all orbital operations, Orbiter maneuvers, and deorbit. Re-entry required orbital velocity decelerations of about 330 km/hr (204 mph) depending on orbital altitude, which caused the Orbiter to slow and fall back to Earth. The Orbiter Thermal Protection System, which covered the entire vehicle, provided the protection needed to survive the extreme high temperatures experienced during re-entry. Primarily friction between the Orbiter and the Earth’s atmosphere generated temperatures ranging from 927°C (1,700°F) to 1,600°C (3,000°F). The highest temperatures experienced were on the wing leading edge and nose cone. The time it took the Orbiter to start its descent from orbital velocity of about 28,160 km/hr (17,500 mph) to a landing speed of about 346 km/hr (215 mph) was 1 hour and 5 minutes. During re-entry, the Orbiter was essentially a glider. It did not have any propulsion capability, except for the Reaction Control System thrusters required for roll control to adjust its trajectory early during re-entry. Management of the Orbiter energy from its orbital speed was critical to allow the Orbiter to reach its desired runway target. The Orbiter’s limited cross-range capability of about 1,480 km (800 nautical miles) made management of the energy during final phases of re-entry close to the ground—otherwise called terminal area energy management—critical for a safe landing. The Orbiter performed as a glider during re-entry, thus its mass properties had to be well understood to ensure that the Flight Control System could control the vehicle and reach the required landing site with the right amount of energy for landing. One of the critical components of its aerodynamic flight was to ensure that the Orbiter center of gravity was correctly calculated and entered into the Orbiter flight design process. Because of the tight center of gravity constraints, the cargo bay payloads were placed in the necessary cargo bay location to protect the down weight and center of gravity of the Orbiter for landing. Considering the Orbiter’s size, the center of gravity box was only 91 cm (36 in.) long, 5 cm (2 in.) wide, and 5 cm (2 in.) high. External Tank The ET was 46.8 m (153.6 ft) in length with a diameter of 8.4 m (27.6 ft), which made it the largest component of the shuttle. The ET contained two internal tanks—one for the storage of liquid hydrogen 72% during first stage to preclude having the vehicle exceed structural limits during high dynamic pressure as well as close to main engine shutdown to preclude the vehicle from exceeding 3 gravitational force (3g) limits. The only manual main engine control capability available to the crew was the manual throttle control, which allowed the crew to decrease engine performance from 104.5% to a level of 72% if required for vehicle control. The main engines had the capability to gimbal about 10.5 degrees up and down and 8.5 degrees to either side to change the thrust direction required for changes in trajectory parameters.
+ ORBITER The Orbiter was the primary component of the shuttle; it carried the crew members and mission cargo/payload hardware to orbit. The Orbiter was about 37.1 m (122 ft) long with a wingspan of about 23.8 m (78 ft). The cargo/payload carrying capacity was limited by the 18.3-m- (60-ft)-long by 4.6-m- (15-ft)- wide payload bay. The cargo/payload weighed up to 29,000 kg (65,000 pounds), depending on the desired orbital inclination. The Orbiter payload bay doors, which were constructed of graphite epoxy composite material, were 18.3 m (60 ft) in length and 4.5 m (15 ft) in diameter and rotated through an angle of 175 degrees. A set of radiator panels, affixed to each door, dissipated heat from the crew cabin avionic systems. The first vehicle, Columbia, was the heaviest Orbiter fabricated due to the installation of additional test instrumentation required to gather data on vehicle performance. As each Orbiter was fabricated, the test instrumentation was deleted and system changes implemented, resulting in each subsequent vehicle being built lighter. The Orbiter crew cabin consisted of the flight deck and the middeck andand the other for the storage of liquid oxygen. The hydrogen tank, which was the bigger of the two internal tanks, held 102,737 kg (226,497 pounds) of hydrogen. The oxygen tank, located at the top of the ET, held 619,160 kg (1,365,010 pounds) of oxygen. Both tanks provided the fuel to the main engines required to provide the thrust for the vehicle to achieve a safe orbit. During powered flight and ascent to orbit, the ET provided about 180,000 L/min (47,000 gal/min) of hydrogen and about 67,000 L/min (18,000 gal/min) of oxygen to all three Space Shuttle Main Engines with a 6-to-1 mixture ratio of liquid hydrogen to liquid oxygen. Solid Rocket Boosters The two SRBs provided the main thrust to lift the shuttle off the launch pad. Each booster provided about 14.7 meganewtons (3,300,000 pounds-force) of thrust at launch, and they were only ignited once the three main engines reached the required 104.5% thrust level for launch. Once the SRBs were ignited, they provided about 72% of the thrust required of the entire shuttle at liftoff and through the first stage, which ended at SRB separation. The SRB thrust vector control system enabled the nozzles to rotate, allowing the entire shuttle to maneuver to the required ascent trajectory during first stage. Two minutes after launch, the spent SRBs were jettisoned, having taken the vehicle to an altitude of about 45 km (28 miles). Not only were the boosters reusable, they were also the largest solid propellant motors in use then. Each measured about 45.4 m (149 ft) long and about 3.6 m (12 ft) in diameter.
+
+
+
+SPACE SHUTTLE MAIN ENGINES After SRB separation, the main engines provided the majority of thrust required for the shuttle to reach orbital velocity. Each main engine weighed about 3,200 kg (7,000 pounds). With a total length of 4.3 m (14 ft), each engine, operating at the 104.5% power level, provided a thrust level of about 1.75 meganewtons (394,000 pounds-force) at sea level and about 2.2 meganewtons (492,000 pounds-force) at vacuum throughout the entire 8 minutes and 30 seconds of powered flight. The engine nozzle by itself was 2.9 m (9.4 ft) long with a nozzle exit diameter of 2.4 m (7.8 ft). Due to the high heat generated by the engine thrust, each engine contained 1,082 tubes throughout its entire diameter, allowing circulation of liquid hydrogen to cool the nozzle during powered flight. The main engines were a complex piece of machinery comprised of high- and low-pressure fuel and oxidizer pumps, engine controllers, valves, etc. The engines were under constant control by the main engine controllers. These consisted of an electronics package mounted on each engine to control engine operation under strict and critical performance parameters. The engines ran at 104.5% performance for much of the entire operation, except when they were throttled down to about could be configured for a maximum crew size of seven astronauts, including their required equipment to accomplish the mission objectives. The flight deck contained the Orbiter cockpit and aft station where all the vehicle and systems controls were located. The crew used six windows in the forward cockpit, two windows overhead, and two windows looking aft for orbit operations and viewing. The middeck was mostly the crew accommodations area, and it housed all the crew equipment required to live and work in space. The middeck also contained the three avionic bays where the Orbiter electronic boxes were installed. Due to their limited power generation capability, the Orbiter fuel cells consumables (power generation cryogenics) provided mission duration capability on the order of about 12 to 14 days, dependent on vehicle configuration. In 2006, NASA put into place the Station-to-Shuttle Power Transfer System, which allowed the ISS to provide power to the Orbiter vehicle, thereby allowing the Orbiter to have a total mission duration of about 16 days. The Orbiter configuration (amount of propellant loaded in the forward and aft propellant tanks, payload mounting hardware in the payload bay, loading of cryogenic tanks required for power generation, crew size, etc.) was adjusted and optimized throughout the pre-mission process. Because of its payload size and robotic arm capability, the Orbiter could be configured to perform as a platform for different cargo/payload hardware configurations. In the total 132 Space Shuttle missions (as of October 2010) over a period of 29 years, the Orbiter deployed a multitude of satellites for Earth observation and telecommunications; interplanetary probes such as Galileo/Jupiter spacecraft and Magellan/Venus Radar Mapper; and great observatories that included the Hubble Space Telescope, Compton Gamma Ray Observatory, and Chandra X-ray Observatory. The Orbiter even functioned as a science platform/laboratory; e.g., Spacelab, Astronomy Ultraviolet Telescope, US Microgravity Laboratory, US Microgravity Payload, etc. Aside from the experiments and satellite deployments the shuttle performed, its most important accomplishment was the delivery and assembly of the ISS.
+SPACE SHUTTLE REUSABILITY All components of the Space Shuttle vehicle, except for the ET, were designed to be reusable flight after flight. The ET, once jettisoned from the Orbiter, fell to Earth where atmospheric heating caused the tank to break up over the ocean. The SRBs, once jettisoned from the tank, parachuted back to the ocean where they were recovered by special ships and brought back to KSC. With their solid propellant spent, the boosters were de-stacked and shipped back to aerospace and defense company Thiokol in Utah for refurbishment and reuse. The SRBs were thoroughly inspected after every mission to ensure that the components were not damaged and could be refurbished for another flight. Any damage found was either repaired or the component was discarded. The Orbiter was the only fully reusable component of the shuttle system. Each Orbiter was designed and certified for 100 space missions and required about 5 months, once it landed, to service the different systems and configure the payload bay to support requirements for its next mission. NASA replaced the components only when they sustained a system failure and could not be repaired. Even though certified for 100 missions, Discovery, Atlantis, and Endeavour completed 39, 32, and 25 missions, respectively, by October 2010. Challenger flew 10 missions and Columbia flew 28 missions before their loss on January 28, 1986, and February 1, 2003, respectively
+AUTOMATION, AUTONOMY, AND REDUNDANCY 
+The Space Shuttle was the first space vehicle to use the fly-by-wire computerized digital flight control system. Except for manual switch throws for system power-up and certain valve actuations, control of the Orbiter systems was through the general purpose computers installed in the forward avionics bay in the middeck. Each Orbiter had five hardwareidentical general purpose computers; four functioned as the primary means to control the Orbiter systems, and one was used as a backup should a software anomaly or problem cause the loss of the four primary computers. During ascent and re-entry—the critical phases of flight—four general purpose computers were used to control the spacecraft. The primary software, called the Primary Avionics Software System, was divided into two major systems: system software, responsible for computer operation, synchronization, and management of input and output operations; and applications software, which performed the actual duties required to fly the vehicle and operate the vehicle systems. Even though simple in their architecture compared to today’s computers, the general purpose computers had a complex redundancy management scheme in which all four primary computers were tightly coupled together and processed the same information at the same time. This tight coupling was achieved through synchronization steps and cross-check results of their processes about 440 times per second. The original International Business Machines computers had only about 424 kilobytes of memory each. The central processing unit could process about 400,000 instructions per second and did not have a hard disk drive capability. These computers were replaced in April 1991 (first flight was STS-37) with an upgraded model that had about 2.5 times the memory capacity and three times the processor speed. To protect against corrupt software, the general purpose computers had a backup computer that operated with a completely different code independent of the Primary Avionics Software System. This fifth computer, called the Backup Flight System, operated in the background, processing the same critical ascent/re-entry functions in case the four general purpose computers failed or were corrupted by problems with their software. The Backup Flight System could be engaged at any moment only by manual crew command, and it also performed oversight and management of Orbiter noncritical functions. For the first 132 flights of the Space Shuttle Program, the Backup Flight System computer was never engaged and, therefore, was not used for Orbiter control. The overall avionics system architecture that used the general purpose computer redundancy was developed with a redundancy requirement for fail-operational/fail-safe capability. These redundancy schemes allowed for the loss of redundancy in the avionics systems and still allowed continuation of the mission or safe landing of the Orbiter. All re-entry critical avionics functions, such as general purpose computers, aero surface actuators, rate gyro assemblies, accelerometer assemblies, air data transducer assemblies, etc., were designed with four levels of redundancy. This meant that each of these functions was controlled by four avionic boxes that performed the same specific function. The loss of the first box allowed for safe continuation of the mission. The loss of the second box still allowed the function to work properly with only two remaining boxes, which subsequently allowed for safe re-entry and landing of the Orbiter. Other critical functions were designed with only triple redundancy, which meant that fail-operational/fail-safe reliability allowed the loss of two of the boxes before the function was lost. The avionics systems redundancy management scheme was essentially controlled via computer software that operated within the general purpose computers. This scheme was to select the middle value of the avionics components when the systems had three or four avionics boxes executing the same function. On loss of the first box, the redundancy management scheme would down mode to the “average value” of the input received from the functioning boxes. Upon the second box failure, the scheme would further down mode to the “use value,” which essentially meant that the function was performed by using input data from only one remaining unit in the system. This robust avionics architecture allowed the loss of avionics redundancy within a function without impacting the ability of the Orbiter to perform its required mission. 
+MANEUVERABILITY, RENDEZVOUS, AND DOCKING CAPABILITY :-
+MANEUVERABILITY The Orbiter was very maneuverable and could be tightly controlled in its pointing accuracy, depending on the objective it was trying to achieve. The Orbiter controllability and pointing capability was performed by the use of 44 Reaction Control System thrusters installed both in the forward and the aft portions of the vehicle. Of the 44 thrusters, six were Reaction Control Systems and each had a thrust level of only 111 newtons (25 pounds-force). The remaining 38 thrusters were considered primary thrusters and each had a thrust level of 3,825 newtons(860 pounds-force). The total thruster complement was divided between the forward thrusters located forward of the crew cabin, and the aft thrusters located on the two Orbital Maneuvering System pods in the tail of the Orbiter. The forward thrusters (total of 16) consisted of 14 primary thrusters and two vernier thrusters. Of the 28 thrusters in the aft, 24 were primary thrusters and four were vernier thrusters. The thrusters were installed on the Orbiter in such a way that both the rotational and the translational control was provided to each of the Orbiter’s six axes of control with each axis having either two or three thrusters available for control. The Orbital Maneuvering System provided propulsion for the shuttle. During the orbit phase of the flight, it was used for the orbital maneuvers needed to achieve orbit after the Main Propulsion System had shut down. It was also the primary propulsion system for orbital transfer maneuvers and the deorbit maneuver. The general purpose computers also controlled the tight Orbiter attitude and pointing capability via the Orbiter Digital Auto Pilot—a key piece of application software within the computers. During orbit operations, the Digital Auto Pilot was the primary means for the crew to control Orbiter pointing by the selection of different attitude and attitude rate deadbands, which varied between +/-1.0 and 5.0 degrees for attitude and +/-0.02 and 0.2 deg/sec for attitude rate. The Digital Auto Pilot could perform three-axis automatic maneuver, attitude tracking, and rotation about any axis or body vector. Crew interface to the Digital Auto Pilot was via the Orbiter cathode ray tubes/keyboard interface, which allowed the crew to control parameters in the software. With very accurate control of its orientation, the Orbiter could provide a pointing capability to any part of the celestial sky as required to accomplish its mission objectives.
+RENDEZVOUS AND DOCKING 
+The shuttle docked to, grappled, deployed, retrieved, and otherwise serviced a more diverse set of orbiting objects than any other spacecraft in history. It became the world’s first general purpose space rendezvous vehicle. Astronauts retrieved payloads no larger than a refrigerator and docked to targets as massive as the ISS, despite the shuttle being designed without specific rendezvous targets in mind. In fact, the shuttle wasn’t designed to physically dock with anything; it was intended to reach out and grapple objects with its robotic arm. A rendezvous period lasted up to 4 days and could be divided into three phases: ground targeted; on-board targeted; and human-piloted proximity operations. The first phase began with launch into a lower orbit, which lagged the target vehicle. The Orbiter phased toward the target vehicle due to the different orbital rates caused by orbital altitude. Mission Control at Johnson Space Center tracked the shuttle via ground assets and computed orbital burn parameters to push the shuttle higher toward the target vehicle. As the shuttle neared the target, it transitioned to on-board targeting using radar and star trackers. These sensors provided navigation data that allowed on-board computers to calculate subsequent orbital burns to reach the target vehicle. The final stage of rendezvous operations—proximity operations— began with the Orbiter’s arrival within thousands of meters (feet) of the target orbital position. During proximity operations, the crew used their highest fidelity sensors (laser, radar, or direct measurement out the window with a camera) to obtain the target vehicle’s relative position. The crew then transitioned to manual control and used the translational hand controller to delicately guide the Orbiter in for docking or grappling operations. The first rendezvous missions targeted satellite objects less massive than the shuttle and grappled these objects with its robotic arm. During the proximity operations phase, the commander only had a docking camera view and accompanying radar information to guide the vehicle. Other astronauts aimed payload bay cameras at the target and recorded elevation angles, which were charted on paper to give the commander awareness of the Orbiter’s position relative to the target. Once the commander maneuvered into a position where the target was above the payload bay, a mission specialist grappled the target with the robotic arm. This method proved highly reliable and applicable to a wide array of rendezvous missions. Shuttle rendezvous needed a new strategy to physically dock with large vehicles: the Russian space station Mir and the ISS. Rendezvous with larger space stations required more precise navigation, stricter thruster plume limitations, and tighter tolerances during docking operations. New tools such as the laser sensors provided highly accurate range and range rate information for the crew. The laser was mounted in the payload bay and its data were routed into the shuttle cabin but could not be incorporated directly into the shuttle guidance, navigation, and control software. Instead, data were displayed on and controlled by a laptop computer mounted in the aft cockpit. This laptop hosted software called the Rendezvous Proximity Operations Program that displayed the Orbiter’s position relative to the target for increased crew situational awareness. This display was used extensively by the commander to manually fly the vehicle from 610 m (2,000 ft) to docking. This assembly of hardware and software aptly met the increased accuracy required by delicate docking mechanisms and enabled crews to pilot the massive shuttle within amazing tolerances. In fact, during the final 0.9 m (3 ft) of docking with the ISS, the Orbiter had to maintain a 7.62-cm (3-in.) lateral alignment cylinder and the closing rate had to be controlled to within 0.02 m/sec (0.06 ft/sec). The commander could control this with incredibly discrete pulses of the Reaction Control System thrusters. Both the commander and the pilot were trained extensively in the art of shuttle proximity operations, learning techniques that allowed them to pilot the Orbiter to meet tolerances. The shuttle was never meant to be piloted to this degree of accuracy, but innovative engineering and training made these dockings uneventful and even routine. The success of shuttle rendezvous missions was remarkable considering its operational complexity. Spacecraft rendezvous is an art requiring the highly scripted choreography of hardware systems, astronauts, and members of Mission Control. It is a precise and graceful waltz of billions of dollars of hardware and human decision making.
+
+ROBOTIC ARM/OPERATIONAL CAPABILITY The Canadian Space Agency provided the Shuttle Robotic Arm. It was designed, built, and tested by Spar Aerospace Ltd., a Canadian Company. The electromechanical arm measured about 15 m (50 ft) long and 0.4 m (15 in.) in diameter with a six-degreeof-freedom rotational capability, and it consisted of a manipulator arm that was under the control of the crew via displays and control panels located in the Orbiter aft flight deck. The Shuttle Robotic Arm was comprised of six joints that corresponded roughly to the joints of a human arm and could handle a payload weighing up to 29,000 kg (65,000 pounds). An end effector was used to grapple a payload or any other fixture and/or component that had a grapple fixture for handling by the arm. Even though NASA used the Shuttle Robotic Arm primarily for handling payloads, it could also be used as a platform for extravehicular activity (EVA) crew members to attach themselves via a portable foot restraint. The EVA crew member, affixed to the portable foot restraint grappled by the end effector, could then be maneuvered around the Orbiter vehicle as required to accomplish mission objectives. Following the Return to Flight after the loss of Columbia, the Shuttle Robotic Arm was used to move around the Orbiter Boom Sensor System, which allowed the flight crew to inspect the Thermal Protection System around the entire Orbiter or the reinforced carbon-carbon panels installed on the leading edge of the wings. During buildup of the ISS, the Shuttle Robotic Arm was instrumental in the handling of modules carried by the Orbiter—a task that would not have been possible without the use of this robotic capability.
+
+EXTRAVEHICULAR ACTIVITY CAPABILITY The Space Shuttle Program provided a dramatic expansion in EVA capability for NASA, including the ability to perform tasks in the space environment and ways to best protect and accommodate a crew member in that environment. The sheer number of EVAs performed during the course of the program resulted in a significant increase in knowledge of how EVA systems and EVA crew members perform. Prior to the start of the program, a total of 38 EVAs were performed by all US space programs combined, including Gemini, Apollo, and Skylab. During previous programs, EVAs focused primarily on simple tasks, such as the jettison of expended hardware or the collection of geology samples. The Space Shuttle Program advanced EVA capability to construction of massive space structures, high-strength maneuvers, and repair of complicated engineering components requiring a combination of precision and gentle handling of sensitive materials and structures. As of October 2010, the shuttle accomplished about 157 EVAs in 132 flights. Of those EVAs, 105 were dedicated to ISS assembly and repair tasks. Shuttle EVA crews succeeded in handling and manipulating elements as large as 9,000 kg (20,000 pounds); relocating and installing large replacement parts; capturing and repairing failed satellites; and performing surgical-like repairs of delicate solar arrays, rotating joints, and much more. The Orbiter’s EVA capability consisted of several key engineering components and equipment. For a crew member to step out of the shuttle and safely enter the harsh environment of space, that crew member had to use the integrated airlock, an extravehicular mobility unit spacesuit, a variety of EVA tools, and EVA translation and attachment aids attached to the vehicle or payload. EVA tools consisted of a suite of components that assisted in handling and translating cargo, translating and stabilizing at the work site, operating manual mechanisms, and attaching bolts and fasteners, often with relatively precise torque requirements. Photo and television operations provided documentation of the results for future troubleshooting, when necessary. Extravehicular Mobility Unit The extravehicular mobility unit was a fully self-sufficient individual spacecraft providing critical life support systems and protection from the harsh space environment. Unlike previous suits, the shuttle suit was designed specifically for EVA and was the cornerstone component for safe conduct of EVA during the shuttle era. It operated at 0.03 kgf/cm2 (4.3 psi) pressure in the vacuum environment and provided thermal protection for interfacing with environments and components from -73°C (-100°F) to 177°C (350°F). It provided oxygen and removed carbon dioxide during an EVA, and it supplied battery power to run critical life support and ancillary extravehicular mobility unit systems, including support lights, cameras, and radio. The suit, which also provided crew members with critical feedback on system operations during EVA, was the first spacesuit controlled by a computer. Future space programs will benefit tremendously from NASA’s EVA experience during the shuttle flights. To ensure success, the goal has been and always will be to design for EVAs that are as simple and straightforward as possible. Fewer and less-complicated provisions will be required for EVA interfaces on spacecraft, and functions previously thought to require complicated and automated systems can now rely on EVA instead. During the shuttle era, NASA took the training wheels off of EVA capability and now has a fully developed and highly efficient operational resource in support of both scheduled and contingency EVA tasks.
+CREW COMPARTMENT ACCOMMODATION FOR CREW AND PAYLOADS The Orbiter’s crew cabin had a habitable volume of 71.5 m3 (2,525 ft3) and consisted of three levels: flight deck, middeck, and utility area. The flight deck, located on the top level, accommodated the commander, pilot, and two mission specialists behind them. The Orbiter was flown and controlled from the flight deck. The middeck, located directly below the flight deck, accommodated up to three additional crew members and included a galley, toilet, sleep locations, storage lockers, and the side hatch for entering and exiting the vehicle. The Orbiter airlock was also located in the middeck area; it allowed up to three astronauts, wearing extravehicular mobility unit spacesuits, to perform an EVA in the vacuum of space. The standard practice was for only two crew members to perform an EVA. Most of the day-to-day mission operations took place on the middeck. The majority of hardware required for crew members to live, work, and perform their mission objectives was stowed in stowage lockers and bags within the middeck volume. The entire middeck stowage capability was equivalent to 127.5 middeck lockers in which each locker was about 0.06 m3 (2 ft3) in volume. This volume could accommodate all required equipment and supplies for a crew of seven for as many as 16 days
+Performance Capabilities and Limitations Throughout the history of the program, the versatile shuttle vehicle was configured and modified to accomplish a variety of missions, including: the deployment of Earth observation and communication satellites, interplanetary probes, and scientific observatories; satellite retrieval and repair; assembly; crew rotation; science and logistics resupply of both the Russian space station Mir and the ISS, and scientific research and operations. Each mission type had its own capabilities and limitations. Deploying and Servicing Satellites The largest deployable payload launched by the shuttle in the life of the program was the Chandra X-ray Observatory. Deployed in 1999 at an inclination of 28.45 degrees and an altitude of about 241 km (130 nautical miles), Chandra—and the support equipment deployed with it—weighed 22,800 kg (50,000 pounds). In 1990, NASA deployed the Hubble Space Telescope into a 28.45-degree inclination and a 555-km (300-nautical-mile) altitude. Hubble weighed 13,600 kg (30,000 pounds). Five servicing missions were conducted over the next 19 years to upgrade Hubble’s science instrumentation, thereby enhancing its scientific capabilities. These subsequent servicing missions were essential in
+correcting the Hubble mirror spherical aberration, thereby extending the operational life of the telescope and upgrading its science capability. 
+ASSEMBLING THE INTERNATIONAL SPACE STATION 
+The ISS Node 1/Unity module was launched on STS-88 (1998), thus beginning the assembly of the ISS, which required a total of 36 shuttle missions to assemble and provide logistical support for ISS vehicle operations. As of October 2010, Discovery had flown 12 missions and Atlantis and Endeavour had flown 11 missions to the ISS, with each mission carrying 12,700 to 18,600 kg (28,000 to 41,000 pounds) of cargo in the cargo bay and another 3,000 to 4,000 kg (7,000 to 9,000 pounds) of equipment stowed in the crew cabin. The combined total of ISS structure, logistics, crew, water, oxygen, nitrogen, and avionics delivered to the station for all shuttle visits totaled more than 603,300 kg (1,330,000 pounds). No other launch vehicle in the world could deliver these large 4.27-m- (14-ft)- diameter by 15.24-m- (50-ft)-long structures or have this much capability. ISS missions required modifications to the three vehicles cited above— Discovery, Atlantis, and Endeavour— to dock to the space station. The docking requirement resulted in the Orbiter internal airlock being moved externally in the payload bay. This change, along with the inclusion of the docking mechanism, added about 1,500 kg (3,300 pounds) of mass to the vehicle weight.
+A Platform for Scientific Research 
+The Orbiter was configured to accommodate many different types of scientific equipment, ranging from large pressurized modules called Spacelab or Spacehab where the crew conducted scientific research in a shirt-sleeve environment to the radars and telescopes for Earth mapping, celestial observations, and the study of solar, atmospheric, and space plasma physics. The shuttle was often used to deploy and retrieve science experiments and satellites. These science payloads were: deployed using the Shuttle Robotic Arm; allowed to conduct free-flight scientific operations; and then retrieved using the arm for return to Earth for further data analysis. This was a unique capability that only the Orbiter could perform. The Orbiter was also unique because it was an extremely stable platform on which to conduct microgravity research studies in material, fundamental physics, combustion science, crystal growth, and biotechnology that required minimal movement or disturbance from the host vehicle. NASA studied the effect of space adaptation on both humans and animals. Crews of seven worked around the clock conducting research in these pressurized modules/laboratories that were packed with scientific equipment. Much research was conducted with the international community. These missions brought together international academic, industrial, and governmental partners to obtain maximum benefits and results. The facilities included middeck glove boxes for conducting research and testing science procedures and for developing new technologies in microgravity. These boxes enabled crew members to handle, transfer, and manipulate experiment hardware and material that were not approved for use in the shuttle. There were furnaces to study diffusion, and combustion modules for conducting research on the single most important chemical process in our everyday lives. The shuttle had freezers for sample return as well as the capability to store large amounts of data for further analysis back on Earth. Scientists used spin tables to conduct biological and physiological research on the crew members. The Orbiter provided all the power and active cooling for the laboratories. A typical Spacelab was provided approximately 6.3 kW (8.45 hp) of power, with peak power as high as 8.1 kW (10.86 hp). To cool the laboratories’ electronics, the modules were tied into the Orbiter’s cooling system so thermal control of the payload was the same as thermal control for the Orbiter avionics. In an effort to share this national resource with industry and academia, NASA developed the Get Away Special Program, designed to provide inexpensive access to space for both novices and professionals to explore new concepts at little risk. In total, over 100 Get Away Special payloads were flown aboard the shuttle, and each payload often consisted of several individual experiments. The cylindrical payload canisters in which these experiments were flown measured 0.91 m (3 ft) in length with a 0.46-m (1.5-ft) diameter. They were integrated into the Orbiter cargo bay on the sill/sidewall and required minimal space and cargo integration engineering. The experiments could be confined inside a sealed canister, or the canister could be configured with a lid that could be opened for experiment pointing or deployment. The shuttle was also an extremely accurate platform for precise pointing of scientific payloads at the Earth and celestial targets. These unpressurized payloads were also integrated into the cargo bay; however, unlike the Spacelab and Spacehab science modules, these payloads were not accessible by the crew, but rather were exposed to the space environment. The crew activated and operated these experiments from the pressurized confines of the Orbiter flight deck. The Shuttle Radar Topography Mission was dedicated to mapping the Earth’s topography between 60° North and 58° South, including the ocean floor. The result of the mission was a threedimensional digital terrain map of 90% of the Earth’s surface. The Orbiter provided about 10 kW (13.4 hp) of power to the Shuttle Radar Topography Mission payload during on-orbit operations and all of the cooling for the payloads’ electronics.
+
+
+
+The Columbia disaster occurred On Feb. 1, 2003, when NASA’s space shuttle Columbia broke up as it returned to Earth, killing the seven astronauts on board. NASA suspended space shuttle flights for more than two years as it investigated the cause of the Columbia disaster.
+An investigation board determined that a large piece of foam fell from the shuttle's external tank and breached the spacecraft wing. This problem with foam had been known for years, and NASA came under intense scrutiny in Congress and in the media for allowing the situation to continue.
+The Columbia mission was the second space shuttle disaster after Challenger, which saw a catastrophic failure during its launch in 1986. The Columbia disaster directly led to the retirement of the space shuttle fleet in 2011. Now, astronauts from the US fly to the International Space Station on Russian Soyuz rockets or aboard commercial spacecraft, like the SpaceX Crew Dragon capsules which began a "space taxi" service to the ISS in 2020.  
+Columbia was the first space shuttle to fly in space; its first flight took place in April 1981, and it successfully completed 27 missions before the disaster. On its 28th flight, Columbia left Earth for the last time on Jan. 16, 2003. At the time, the shuttle program was focused on building the International Space Station. However, Columbia's final mission, known as STS-107, emphasized pure research. 
+The seven-member crew — Rick Husband, commander; Michael Anderson, payload commander; David Brown, mission specialist; Kalpana Chawla, mission specialist; Laurel Clark, mission specialist; William McCool, pilot; and Ilan Ramon, payload specialist from the Israeli Space Agency — had spent 24 hours a day doing science experiments in two shifts. They performed around 80 experiments in life sciences, material sciences, fluid physics and other matters before beginning their return to Earth's surface.
+During the crew's 16 days in space, NASA investigated a foam strike that took place during launch. About 82 seconds after Columbia left the ground, a piece of foam fell from a "bipod ramp" that was part of a structure that attached the external tank to the shuttle. Video from the launch appeared to show the foam striking Columbia's left wing. It was later found that a hole on the left wing allowed atmospheric gases to bleed into the shuttle as it went through its fiery re-entry, leading to the loss of the sensors and eventually, Columbia itself and the astronauts inside.
+On Feb. 1, 2003, the shuttle made its usual landing approach to the Kennedy Space Center. Just before 9 a.m. EST, however, abnormal readings showed up at Mission Control. Temperature readings from sensors located on the left wing were lost. Then, tire pressure readings from the left side of the shuttle also vanished.
+
+
+
+ The Capcom, or spacecraft communicator, called up to Columbia to discuss the tire pressure readings. At 8:59:32 a.m., Husband called back from Columbia: "Roger," followed by a word that was cut off in mid-sentence.
+At that point, Columbia was near Dallas, traveling 18 times the speed of sound and still 200,700 feet (61,170 meters) above the ground. Mission Control made several attempts to get in touch with the astronauts, with no success.
+Twelve minutes later, when Columbia should have been making its final approach to the runway, a mission controller received a phone call. The caller said a television network was showing a video of the shuttle breaking up in the sky.
+Shortly afterward, NASA declared a space shuttle 'contingency' and sent search and rescue teams to the suspected debris sites in Texas and later, Louisiana. Later that day, NASA declared the astronauts lost.
+"This is indeed a tragic day for the NASA family, for the families of the astronauts who flew on STS-107, and likewise is tragic for the nation," stated NASA's administrator at the time, Sean O'Keefe.
+The search for debris took weeks, as it was shed over a zone of some 2,000 square miles (5,180 square kilometers) in east Texas alone. NASA eventually recovered 84,000 pieces, representing nearly 40 percent of Columbia by weight. Among the recovered material were crew remains, which were identified with DNA.
+Much later, in 2008, NASA released a crew survival report detailing the Columbia crew's last few minutes. The astronauts probably survived the initial breakup of Columbia, but lost consciousness in seconds after the cabin lost pressure. The crew died as the shuttle disintegrated.
+
+
+
+
+
+The Columbia Accident Investigation Board (CAIB) was established to investigate the disaster and recommend changes to prevent such incidents in the future. Their report highlighted several key issues that contributed to the disaster and provided recommendations for improvement:
+1.	Culture and Communication: The CAIB emphasized the importance of open communication and a strong safety culture within NASA. It recommended changes to ensure that safety concerns from lower-level employees were effectively communicated to management and addressed.
+2.	Inspection and Repair: The CAIB recommended improving the inspection and repair capabilities for the shuttle's thermal protection system. This included developing techniques to repair damaged tiles while in orbit and improving the ability to assess potential damage to the shuttle's heat shield.
+3.	Risk Assessment: The CAIB suggested enhancing the risk assessment processes to better understand and manage potential hazards during shuttle missions. This involved more accurately assessing the risks associated with potential damage to the shuttle's thermal protection system.
+4.	Improve Imaging and Data Sharing: The CAIB recommended improvements to the ability to image and assess the shuttle's condition during flight. This would help in identifying potential damage and making informed decisions about reentry.
+5.	Emergency Preparedness: The CAIB emphasized the importance of preparing for potential in-flight emergencies. The crew should have had more comprehensive training on how to respond to a variety of scenarios, and there should have been better protocols for dealing with potential damage to the shuttle.
+6.	Vehicle Design: The CAIB suggested reviewing the design of the shuttle's thermal protection system and other critical components to enhance their safety and resilience.
+7.	Contingency Plans: The CAIB recommended developing contingency plans for addressing potential damage that could occur during launch. This would involve having procedures in place for repairing damage or seeking safe haven on the International Space Station if necessary.
+
+
+
+1.	@2 Anomaly Detection: I could assist in analyzing data from various sensors and instruments on the space shuttle to detect anomalies or deviations from expected behavior. If any unusual patterns or readings are identified, they could be flagged for further investigation.
+2.	Data Integration: By integrating data from different systems and sources, I could help create a comprehensive picture of the shuttle's status. This could include data related to thermal protection system integrity, propulsion, navigation, and more.
+3.	Automated Alerts: I could be programmed to generate automated alerts when certain predefined conditions are met. For example, if certain temperature or pressure thresholds are exceeded, an alert could be sent to the mission control team.
+1.	@4 Image Enhancement: I could assist in enhancing images captured by spacecraft cameras by applying advanced image processing techniques. This could involve reducing noise, enhancing contrast, and sharpening details to provide clearer and more informative images.
+2.	Data Compression: Space missions often deal with limited bandwidth for transmitting data back to Earth. I could help develop efficient data compression algorithms that reduce the size of image and data files while preserving essential information. This would allow more data to be transmitted within the available bandwidth.
+3.	Real-time Analysis: I could process and analyze images in real-time, identifying noteworthy features or anomalies as soon as they are captured. This information could be quickly relayed to mission control or relevant teams for immediate attention.
+4.	Automated Anomaly Detection: By continuously analyzing images, I could automatically identify and flag potential anomalies or deviations from expected conditions. This could include identifying structural issues, equipment malfunctions, or unexpected phenomena.
+5.	Pattern Recognition: I could assist in recognizing patterns in large sets of images, such as tracking changes over time, mapping surface features, or identifying potential geological or atmospheric patterns.
+6.	Data Fusion: I could integrate data from various sensors and sources to create a more comprehensive understanding of the environment being observed. For instance, combining images with data from spectrometers could provide insights into the composition of planetary surfaces.
+7.	Data Prioritization: Not all data is equally important. I could help prioritize the transmission of critical or high-priority data, ensuring that mission-critical information is sent back to Earth promptly.
+

config.json

@@ -0,0 +1,39 @@
+{
+  "_name_or_path": "gpt2",
+  "activation_function": "gelu_new",
+  "architectures": [
+    "GPT2LMHeadModel"
+  ],
+  "attn_pdrop": 0.1,
+  "bos_token_id": 50256,
+  "embd_pdrop": 0.1,
+  "eos_token_id": 50256,
+  "initializer_range": 0.02,
+  "layer_norm_epsilon": 1e-05,
+  "model_type": "gpt2",
+  "n_ctx": 1024,
+  "n_embd": 768,
+  "n_head": 12,
+  "n_inner": null,
+  "n_layer": 12,
+  "n_positions": 1024,
+  "reorder_and_upcast_attn": false,
+  "resid_pdrop": 0.1,
+  "scale_attn_by_inverse_layer_idx": false,
+  "scale_attn_weights": true,
+  "summary_activation": null,
+  "summary_first_dropout": 0.1,
+  "summary_proj_to_labels": true,
+  "summary_type": "cls_index",
+  "summary_use_proj": true,
+  "task_specific_params": {
+    "text-generation": {
+      "do_sample": true,
+      "max_length": 50
+    }
+  },
+  "torch_dtype": "float32",
+  "transformers_version": "4.30.1",
+  "use_cache": true,
+  "vocab_size": 50257
+}

generation_config.json

@@ -0,0 +1,6 @@
+{
+  "_from_model_config": true,
+  "bos_token_id": 50256,
+  "eos_token_id": 50256,
+  "transformers_version": "4.30.1"
+}

main.py

@@ -0,0 +1,124 @@
+import pandas as pd
+import numpy as np
+import os
+
+def read_txt(file_path):
+    text = ""
+    try:
+        with open(file_path, "r") as file:
+            text = file.read()
+    except:
+        text = ""
+    return text
+
+with open("train.txt", "w") as f:
+    f.write('')
+    
+data = ""
+for filename in os.listdir("./"):
+    file_path = os.path.join("./", filename)
+    if file_path.endswith(".txt") and (file_path != "train.txt"):
+        data += read_txt(file_path)
+        data =  ' '.join(data.split('\n'))
+            
+        with open("train.txt", "a") as f:
+            f.write(data)
+            
+from transformers import TextDataset, DataCollatorForLanguageModeling
+from transformers import GPT2Tokenizer, GPT2LMHeadModel
+from transformers import Trainer, TrainingArguments
+
+def load_dataset(file_path, tokenizer, block_size = 128):
+    dataset = TextDataset(
+        tokenizer = tokenizer,
+        file_path = file_path,
+        block_size = block_size,
+    )
+    return dataset
+
+def load_data_collator(tokenizer, mlm = False):
+    data_collator = DataCollatorForLanguageModeling(
+        tokenizer=tokenizer, 
+        mlm=mlm,
+    )
+    return data_collator
+
+def train(train_file_path,model_name,
+        output_dir,
+        overwrite_output_dir,
+        per_device_train_batch_size,
+        num_train_epochs,
+        save_steps):
+    
+    tokenizer = GPT2Tokenizer.from_pretrained(model_name)
+    train_dataset = load_dataset(train_file_path, tokenizer)
+    data_collator = load_data_collator(tokenizer)
+    tokenizer.save_pretrained(output_dir)
+    model = GPT2LMHeadModel.from_pretrained(model_name)
+    model.save_pretrained(output_dir)
+    
+    training_args = TrainingArguments(
+        output_dir=output_dir,
+        overwrite_output_dir=overwrite_output_dir,
+        per_device_train_batch_size=per_device_train_batch_size,
+        num_train_epochs=num_train_epochs,
+    )
+    
+    trainer = Trainer(
+        model=model,
+        args=training_args,
+        data_collator=data_collator,
+        train_dataset=train_dataset,
+    )
+      
+    trainer.train()
+    trainer.save_model()
+
+train_file_path = "train.txt"
+model_name = 'gpt2'
+output_dir = 'model'
+overwrite_output_dir = False
+per_device_train_batch_size = 8
+num_train_epochs = 50.0
+save_steps = 50000
+
+train(
+    train_file_path=train_file_path,
+    model_name=model_name,
+    output_dir=output_dir,
+    overwrite_output_dir=overwrite_output_dir,
+    per_device_train_batch_size=per_device_train_batch_size,
+    num_train_epochs=num_train_epochs,
+    save_steps=save_steps
+)
+
+from transformers import PreTrainedTokenizerFast, GPT2LMHeadModel, GPT2TokenizerFast, GPT2Tokenizer
+
+def load_model(model_path):
+    model = GPT2LMHeadModel.from_pretrained(model_path)
+    return model
+
+
+def load_tokenizer(tokenizer_path):
+    tokenizer = GPT2Tokenizer.from_pretrained(tokenizer_path)
+    return tokenizer
+
+def generate_text(model_path, sequence, max_length):
+    
+    model = load_model(model_path)
+    tokenizer = load_tokenizer(model_path)
+    ids = tokenizer.encode(f'{sequence}', return_tensors='pt')
+    final_outputs = model.generate(
+        ids,
+        do_sample=True,
+        max_length=max_length,
+        pad_token_id=model.config.eos_token_id,
+        top_k=50,
+        top_p=0.95,
+    )
+    print(tokenizer.decode(final_outputs[0], skip_special_tokens=True))
+
+model_path = "/model/"
+sequence = "Hello!"
+max_len = 50
+generate_text(model_path, sequence, max_len)

pytorch_model.bin

@@ -0,0 +1,3 @@
+version https://git-lfs.github.com/spec/v1
+oid sha256:8d4c87b447dd5a0cb5a4c98970fc1de035dc5548a34cb57b924e98453d10f097
+size 497805149

requirements.txt

@@ -0,0 +1,186 @@
+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==1.9.2

soyez.txt

@@ -0,0 +1,15 @@
+STS 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.

spacegpt.pem

@@ -0,0 +1,27 @@
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+-----END RSA PRIVATE KEY-----

special_tokens_map.json

@@ -0,0 +1,23 @@
+{
+  "bos_token": {
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  },
+  "eos_token": {
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  },
+  "unk_token": {
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  }
+}

tokenizer_config.json

@@ -0,0 +1,33 @@
+{
+  "add_bos_token": false,
+  "add_prefix_space": false,
+  "bos_token": {
+    "__type": "AddedToken",
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  },
+  "clean_up_tokenization_spaces": true,
+  "eos_token": {
+    "__type": "AddedToken",
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  },
+  "errors": "replace",
+  "model_max_length": 1024,
+  "pad_token": null,
+  "tokenizer_class": "GPT2Tokenizer",
+  "unk_token": {
+    "__type": "AddedToken",
+    "content": "<|endoftext|>",
+    "lstrip": false,
+    "normalized": true,
+    "rstrip": false,
+    "single_word": false
+  }
+}