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Abort Guidance System
Auxiliary Power Unit
Abort to Orbit
Russian Micropurification Unit (Russian)
Carbon Dioxide Removal System
Colony Forming Unit
Control Moment Gyroscope
Cell Performance Monitor
Compound Specific Analyzer-Combustible Products
Extravehicular Mobility Unit
Electrical Power System
Fuel Cell Monitoring System
Functional Cargo Block (Russian)
Flight Safety Office
Galley Iodine Removal Assembly
Guidance, Navigation, and Control
General Purpose Computer
Global Positioning System
Inertial Measurement Unit
International Space Station
Internal Thermal Control System
Launch Control Officer
Low Iodine Residual System
Loss of Crew
Loss of Vehicle
Minimum Duration Flight
Master Events Controller
Main Landing Gear
Micro-Meteoroid Orbital Debris
Marshall Space Flight Center
NASA Standard Initiator
Office of Safety & Mission Assurance (NASA HQ)
Protuberance Air Load
Precision Approach Path Indicator
Primary Avionics Software System
Pyrotechnic Initiator Controller
Partial Pressure of CO2
Reaction Control System/Subsystem
Remote Manipulator System
Russia or Russian
Return to Launch Site
Safety & Mission Assurance
Solid Fuel Oxygen Generator
Solid Rocket Booster
Condensate Water Processor Unit (Russian)
Space Shuttle Main Engine
Space Shuttle Program
Thermal Protection System
Loss of Crew
Gemini 5 8/29/1965
Skylab 4 2/8/1974
Soyuz TM-5 9/6/1988
Mercury MA-7 5/24/1962
Soyuz TM-25 8/17/1997
O2 Fire - Soviet
Apollo 10 5/22/1969
STS-3 | 3/30/1982 | Crew: 2
Pilot induced oscillation during derotation. Stronger than predicted winds contributed.
On March 30, 1982 during orbiter derotation on rollout, the vehicle pitched up to approximately six degrees after having been down to -3 degrees pitch. This pitch up occurred because the pilot was preventing premature nose wheel contact. The planned late transition from autoland to manual control did not provide sufficient time for the pilot to feel the vehicle response, and attempts by the pilot to make minor trajectory adjustments resulted in a touchdown sooner than intended and at a higher than planned airspeed (225 Keas vs. 195 Keas). Subsequently, the derotation after main landing gear touchdown started at too high an airspeed and required the pilot to try and stop it at too low a pitch angle. The rapidly changing elevator trim requirements made it difficult to avoid over-controlling in this situation.
On all future missions, manual takeover from autoland was not planned to occur between the start of the preflare maneuver and touchdown. Flight procedures and crew training were also revised to be more explicit about keeping the nose up until the vehicle slows to 180 knots.
STS-9 | 12/8/1983 | Crew: 6
A. Two APUs caught fire during rollout.
B. GPC failed on touchdown.
C. Incorrect flight control rechannelization on rollout.
A) During rollout on December 8, 1983 two Auxiliary Power Units (APUs) caught fire. Six minutes and fifty seconds after the orbiter landed, APU-1 shut down automatically due to a turbine-underspeed condition. Four minutes and twenty-four seconds later, a detonation occurred in APU-1, along with simultaneous automatic shutdown of APU-2, also the result of a turbine-underspeed condition. Fourteen minutes and forty-two seconds after APU-2 shutdown, a detonation occurred on APU-2. Post-flight examination of the orbiter aft compartment revealed fire damage to both APUs and minor shrapnel damage. Post-flight analysis indicated that both APU failures were the result of stress-corrosion cracking in the injector stems of both APUs, which resulted in leakage of hydrazine and subsequent fire/explosion events. The injector stems were subsequently redesigned to reduce susceptibility to corrosion by chromizing the stem, and to reduce material stresses by making changes in the installation processes.
B) Also during landing on December 8, a General Purpose Computer (GPC) failed on touchdown and an incorrect flight control rechannelization occurred on rollout. Due to a failure on orbit, GPC 1 was powered down prior to entry (creating an off-nominal configuration), and the remaining GPCs (2, 3, 4, and 5) were configured for entry landing. During landing rollout, GPC 2, which had previously failed on orbit but was recovered prior to entry, failed again at nose-wheel slap down.
C) The crew reacted with procedures for computer loss in a nominal configuration with GPC 1 active and nominal Flight Control System channel assignments. The crew's execution of GPC 2 malfunction procedures in this off-nominal GPC string configuration resulted in the loss of the remaining two redundant flight control strings. This was not a problem on the runway, but could have resulted in loss of control in flight.
Gemini 5 | 8/29/1965 | Crew: 2
Erroneous entry data uplinked; crew manually corrected entry flight profile.
During entry on August 29, 1965 a crew member used attitude controls to correct the entry flight profile of the vehicle. The computer guiding the capsule was functioning as intended. However, the rotation rate of the Earth was incorrectly entered as 360 degrees per day, instead of the correct 360.98 degrees per day. The crew member recognized the error in the readings and was able to counter the effects. The landing fell 130 kilometers short of the target, but this short landing was closer to the U.S. Navy recovery ship than it would have been if the crew member had not taken action.
Skylab 4 | 2/8/1974 | Crew: 3
Incorrect circuit breakers opened, resulting in the loss of the automatic control.
On February 8, 1974 while preparing foar entry, the crew inadvertently opened the stabilization and control system (SCS) pitch and yaw circuit breakers instead of the service propulsion system pitch and yaw circuit breakers. The vehicle was in an apex forward configuration for service module jettison. The commander attempted to orient the vehicle to the proper heat shield forward attitude for entry. The control commands produced no effect due to the SCS being inadvertently unpowered, and the vehicle failed to change attitude. The crew switched to “manual reaction control system direct” and oriented the vehicle to the proper attitude. The circuit breakers being in close proximity and similarly labeled, increased the potential for human error.
The failure to orient the heat shield forward would have caused loss of crew.
Soyuz TM-5 | 9/6/1988 | Crew: 2
Two de-orbit attempts failed. Crew confined to DM due to OM being jettisoned prior to 1st de-orbit attempt. Crew prevented erroneous firing of SM separation pyrotechnics.
Two de-orbit burn attempts failed and nearly led to the loss of the crew. The crew was confined to the descent module due to the orbital module being jettisoned prior to the first deorbit attempt. The first deorbit burn was prevented by a sensor glitch which disappeared after seven minutes, and then the burn started. However, the crew manually shut down the burn after three seconds.
A second burn two revolutions later occurred on time for six seconds, then stopped, and the crew manually restarted the burn. However, after an additional 60 seconds it was cut off by the autopilot. The crew manually interrupted the command sequence shortly before the descent/equipment module separation pyros were to have been fired, preventing an erroneous firing. The main cause of the crew's problems was acknowledged to be a combination of incorrect actions of the crew commander and mission control personnel.
Mercury MA-7 | 5/24/1962 | Crew: 1
RCS depletion at 80,000 ft.
This incident on May 24, 1962 involved the use of double authority control and the accidental actuation of the fly-by-wire high thrust units during certain maneuvers. The manual-system fuel was depleted near the end of the retrofire maneuver, and the automatic-system fuel was depleted at about 80,000 and 70,000 feet. Because of the early depletion of automatic-system fuel, attitude control during re-entry was not available for the required duration. Attitude rates built up after the Automatic Stabilization Control System became inoperative because of the lack of fuel, and these rates were not sufficiently damped by aerodynamic forces. The pilot chose to deploy the drogue parachute manually at an altitude of approximately 25,000 feet to stabilize the spacecraft.
To avoid the same situation on later flights, Mercury MA-8 and subsequent spacecraft contained a switch which allowed the pilot to disable and reactivate the high-thrust units at his discretion. An automatic override reactivated these thrusters just prior to retrofire. Additionally, a revision of fuel management and control training procedures was instituted for subsequent missions
Soyuz TM-25 | 8/17/1997 | Crew: 3
Landing rockets fired at heat shield separation rather than at landing.
On August 17, 1997 the landing rockets on Soyuz TM-25 fired during heat shield separation rather than during landing. This failure resulted in a harder landing than normal.
The Mir with the docked Soyuz was experiencing high humidity levels in the atmosphere. Water condensing on the connectors in the Soyuz electrical box controlling the circuit probably caused a short circuit, which caused the rockets to fire when the system was armed at heat shield physical separation. Changes were made to either seal the connectors or separate the connectors to prevent a short from applying electrical power to the rockets when the system is armed.
The design has two primary types of inhibits. One inhibit consists of three mechanical switches that physically disconnect the firing circuit when the heat shield is attached. These switches are spring loaded and move approximately 20 mm as the heat shield is deployed to close the firing circuit. A minimum of two out of three switches must be closed for the initiating system to function. In addition to these mechanical inhibits, logic in the electronics prevents the ignition command from being sent until after the heat shield is deployed. Nominally the soft landing motors are ignited when the gamma ray altitude/velocity sensor detects proximity to the ground. The system automatically initiates either four or six motors depending on the velocity. Each motor has 1 kg of propellant and burns for approximately 0.1 to 0.14 seconds.
Mir | 8/30/1994 | Crew: Soyuz 2, Mir 3 | Related or Recurring event
Soyuz TM-17 collided twice with Mir during undocking.
On January 14, 1994 during the post-separation inspection fly-around of Mir, the crew lost manual translation control due to a configuration error. The loss of control led to the Soyuz colliding with Mir several times. The cause of the collision was traced to the hand controller in the orbital module which governed braking and acceleration being switched on, disabling the equivalent hand controller in the descent module.
Mir | 8/30/1994 | Mir Crew: 2 | Related or Recurring event
Progress M-24 collided with Mir during second docking attempt.
On August 30, 1994 during the second attempt of the Progress M-24 to dock with Mir, the Progress collided with Mir's forward docking unit two to four times, and then drifted away from the station. The docking problems with Progress M-24 have been variously attributed to software or Kurs electronics failures on Progress M-24, or the failure of control equipment in the Moscow Mission Control Center.
Mir | 6/25/1997 | Mir Crew: 3 | Loss of Mission | Related or Recurring event
Progress M-34 collided with Mir. Spektr pressure shell ruptured. Spektr module isolated. Cables through hatchway impeded hatch closing.
Mir Crew: 3
Loss of Element
On June 25, 1997 Progress M-34 collided with the Mir Spektr module rupturing the module. The crew of Mir had to cut through cables in the hatchway in order to seal off the leaking module from the rest of the station.
STS-32 | 1/9/1990 | Crew: 3 | Loss of Attitude Control
Erroneous state vector up-linked to flight control system, causing immediate and unpredictable attitude control problems.
An erroneous state vector up-linked to the flight control system on January 9, 1990 causing immediate and unpredictable attitude control problems.
At 17:23:46:51 Greenwich Mean Time, during a crew sleep period, a state vector update was commanded by the ground prior to the loss of signal. The uploaded state vector was erroneous, and the orbiter began to execute a multi-axis rotation at three degrees per second with a number of thruster firings. The rotation continued until the acquisition of signal period, about 10 minutes later, when the crew was awakened and instructed to switch to manual Digital Auto Pilot to arrest the unwanted rates. A good state vector was then uplinked.
STS-87 | 11/27/1997 | Crew: 6
Spartan satellite deployed without proper activation.
Recapture with RMS unsuccessful. Later captured by EVA crew.
Deployment of the SPARTAN satellite on November 21, 1997 occurred without proper activation.
A crew input via the Payload and General Support Computer was not received by the spacecraft. Lack of telemetry and onboard verification procedures left this condition undetected by the Mission Control Center and flight crew. The SPARTAN was grappled with the Remote Manipulator System, removed from the Release/Engage Mechanism, and released per the flight plan.
The missed command step resulted in the failure of the SPARTAN to execute an expected preprogrammed maneuver ("pirouette") about 2.5 minutes after deploy. Attempts to re-grapple the SPARTAN after the deployment were unsuccessful. A previously scheduled extravehicular activity (space walk) had to be changed to manually recapture the satellite.
Altitude Chamber O2 Fire | 3/23/1961 | Crew: 1 | Loss of Crew
Alcohol wipe hit hot plate and started fire in oxygen-rich test chamber.
On March 23, 1961 a cosmonaut in an altitude chamber was removing the sensors that had been attached to him during an experiment. He cleaned the places where the sensors had been attached with cotton wool soaked in alcohol, and without looking threw away the cotton wool. The cotton wool landed on the ring of an electric hot plate in the oxygen-charged atmosphere of the chamber. In conditions of high oxygen concentration, normally non-flammable substances can burn vigorously. The cosmonaut's training suit caught fire. Unaccustomed to the vigor of high-oxygen fires, the cosmonaut would only have spread the flames further by attempting to smother them. The doctor on duty noticed the conflagration through a porthole and rushed to the hatch, which he could not open because the internal pressure kept the inward swinging hatch sealed. Releasing the pressure through bleed valves took several minutes and the cosmonaut later died in the hospital from the burns.
Apollo 10 | 5/22/1969 | Crew: 2
Switch misconfiguration resulted in lunar module control problems.
In May 22, 1969 a switch misconfiguration resulted in lunar lander control problems.
During the Lunar Module (LM) last pass, within eight miles of the moon and prior to the jettison of the LM Descent Stage, the Commander (while wearing a space suit) started to troubleshoot an electrical anomaly.
The Abort Guidance System (AGS) was inadvertently switched from HOLD ATTITUDE to AUTO, which caused the LM to look for the Command/Service Module (CSM) and flip end over end.
The attitude indicator was going to the red zone and in danger of tumbling the inertial platform. The Commander was able to grab the hand controller, switch to manual control, jettison the Descent Stage, control the LM Ascent Stage, and finally dock with the CSM.
STS-110 | 4/8/2002 | Crew: 7
STS-109 | 3/1/2002 | Crew: 7
STS-108 | 12/5/2001 | Crew: 7
Incorrect adjustments to the controller software resulted in SSME underperformance.
Prior to STS-108 a change had been made to the controller software coefficient for the Space Shuttle Main Engine (SSME) to compensate for an observed measurement bias in the SSME main combustion chamber pressure sensor, which controls the SSME fuel/oxidizer mixture ratio. The pressure chamber sensor was biased high causing the flight software to lower the chamber pressure by decreasing the liquid oxygen flow rate. To correct the high bias a coefficient in the equation was adjusted to compensate. Because of communication errors between ground systems engineers and deficiencies in the flight software verification and validation processes, the software coefficient was adjusted in the wrong direction, resulting in even larger dispersions in the mixture ratio and SSME performance.
The error in the coefficient was discovered during post-flight reconstruction of the data from STS-108. The cause of the error remained unknown until after STS-110. The erroneous coefficient was flown on three consecutive flights (STS-108, STS-109, and STS-110) resulting in a slight SSME underperformance on each flight, and was fixed with the proper coefficient and independent verification prior to STS-111. The error in software and resulting mixture ratio wasn't severe enough to cause any significant impacts to SSME performance, and all three flights achieved proper orbits. However, if the software error had been larger, more severe impacts to the missions and crew safety could have occurred, including a premature engine shutdown/failure resulting in on-pad or ascent abort, loss of mission.
Crew Injury/Illness and/or Loss of Vehicle or Mission
Related or Recurring event
Mir Collision Events (1994-1997)
LANDING & POSTLANDING