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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
Crew Injury/Illness and/or Loss of Vehicle or Mission
Other significant STS TPS anomalies:
STS-6, 41B, 51G, 27*, 28, 40, 42, 45
*Most severe tile damage to date.
STS-107 (Columbia) 2/1/2003
Apollo 12 11/24/1969
Apollo 15 8/7/1971
Mercury MA-6 2/20/1962
Mercury MA-7 5/24/1962
Gemini 4 6/7/1965
Gemini 5 8/29/1965
Skylab 4 2/8/1974
Mercury MA-7 5/24/1962
Apollo ASTP 7/24/1975
SpaceShipTwo, PF04 10/31/2014
STS-40, 6/91, Crew: 7*
STS-35, 12/90, Crew: 7*
STS-28, 8/89, Crew: 5*
STS-6, 4/83, Crew: 4*
*toxic byproducts released
ISS, 10/10/08, Crew: 3
ISS, 9/18/06, Crew: 3*
ISS, 3/05, Crew: 2
Skylab 2 5/26/1973
Mercury MA-9 5/16/1963
Gemini 8 3/16-3/17/1966
ISS, Flight 2A.1 5/1999
ISS, Increment 2 4/24/2001
ISS, Increment 4 2/2002
ISS, Increment 2-4 4/2001-3/2002
ISS, Increment 5&6 mid 2002-2/03
ISS, Increment 10 2/2005
ISS, Increment 13 8/2006
ISS, Increment 15 6/10-6/18/2007
ISS, Increment 17 4/30/2008
ISS Increment 38 12/1/2013
Apollo 11 7/21/1969
Apollo 10 5/22/1969
STS-114, 115, 118, 119, 124, 126
STS-116 and STS-125
Gemini 6 12/12/1965
Apollo 1 (AS-204) 1/27/1967
Other On-pad Abort Events:
STS-51F, STS-55, STS-51, STS-68
Gemini 10 7/18/1966
Apollo 12 11/14/1969
Apollo 13 4/11/1970
Other SRB gas seal anomalies:
STS-2, 6, 41B, 41C, 41D, 51C, 51D, 51B, 51G, 51F, 51I, 51J, 61A, 61B, 61C, 42, 70, 71, 78
STS-51L (Challenger) 1/28/1986
X-15 3-65-97 | 11/15/1967 | Crew: 1 | Loss of Crew
Electrical short and crew error led to loss of control at 230,000 feet. First U.S. spaceflight fatality.
On November 15, 1967 an electrical short and crew error led to loss of control of the X-15 at 230,000 feet. During re-entry of the vehicle, the aircraft deviated off course due to a combination of the pilot's distraction, misinterpretation of instrumentation display, and possible vertigo. An electrical disturbance that occurred early in the flight had degraded the overall effectiveness of the aircraft's control system and further added to pilot workload. The aircraft entered into a high Mach spin.
The pilot was able to break free from the spin, but the aircraft was in a high-speed inverted dive. While the aircraft was still at sufficient altitude to recover from the dive, the hand controller began forcing the horizontal stabilizers to oscillate. Because of the buffeting in the spin and dive, the pilot likely lost consciousness and the aircraft broke apart.
This was the first United States spaceflight fatality.
Related or Recurring event
Other Thermal Protection System Damage Events
In addition to the Thermal Protection System (TPS) damage on STS-1, STS-51D, and STS-107, the following Space Shuttle flights experienced TPS damage:
STS-6 (April 1983)
STS-41B (February 1984)
STS-51G (June 1985)
STS-27 (December 1988)
STS-28 (August 1989)
STS-40 (June 1991)
STS-42 (January 1992)
STS-45 (March 1992)
Additional information can be found in the reports linked below.
STS-6 Mission Report STS-41B MER Report STS-41B Mission Report STS-51G MER Report STS-51G Mission Report STS-27 MER Report STS-27 Mission Report STS-27 Close Call STS-28 Mission Safety Eval Record STS-28 MER Report STS-28 Mission Report STS-40 Debris, Ice, TPS Assessment STS-40 Mission Safety Evaluation STS-40 Mission Report STS-42 Debris, Ice, TPS Assessment STS-42 Mission Report
STS-1 | 4/14/1981 | Crew: 2 | Related or Recurring event
Right-hand main landing gear door warped due to entry heating.
On April 14, 1981 the right-hand main landing gear door warped due to entry heating. A forward facing step, a tile gap, a tile-to-filler bar gap and an inadequate flow restrictor resulted in excessive gap heating on the forward portion of the right main landing gear door. This excessive heating resulted in severe tile sidewall shrinkage (on four tiles), a charred filler bar, and a localized buckle in the door structure. The structure and Thermal Protection System on the door was refurbished, and the flow restrictor was modified to increase the effectiveness of the Thermal Protection System in the area of the main landing gear doors.
STS-51D | 4/19/1985 | Crew: 7 | Related or Recurring event
TPS burn-through on left outboard elevon.
The post-flight inspection of the Thermal Protection System (TPS) revealed that significant damage occurred during landing on April 19, 1985.
The outboard end of the left-hand lower outboard elevon had received significant heat damage, specifically the outboard forward corner of the elevon lower-honeycomb outer-face-sheet. This area was buckled and delaminated and had two small burn-through holes. The outboard elevon lower-leading-edge tile-carrier panel was completely melted under the outboard tile, and a hole was melted in the elevon-cove primary-seal support plate. Because of the damage the lower-outboard carrier-panel outermost tile fell onto the runway when the elevon was deflected upward after landing.
Evidence indicates that the entry plasma flow entered the inboard gap of the outboard tile, then progressed under the tile flowing outboard, where eventually the tile-attachment strain isolation pad was burned, allowing the tile to become loose. This allowed more plasma flow under the tile, resulting in the melting of the aluminum carrier panel, primary seal panel structure, and elevon honeycomb outer face sheet, as well as the melting of two tiles aft of the plasma entry point and two elevon sidewall tiles. The cause of the TPS and structural damage that occurred during descent has not been positively identified. The most probable cause is an out-of-spec step or gap in the lower wing surface forward of the elevon leading edge. It is believed that this flow path may have existed for the two previous flights, with progressive deterioration of the bond, but was not evident from outside inspection of this area during post-flight inspections.
A requirement was established to remove the outboard leading-edge carrier-panel on each side of all orbiters for detailed inspection after the next several flights. In addition, a more comprehensive detailed inspection of each outboard elevon/wing area was accomplished during the normal TPS post-flight inspections.
STS-107 (Columbia) | 2/1/2003 | Crew: 1 | Loss of Crew | Related or Recurring event
TPS damage from ascent debris strike resulted in loss of crew and vehicle on entry. Similar bipod ramp foam loss occurred on STS-7, STS-32, STS-50, STS-52, STS-62, and STS-112.
Damage to the Thermal Protection System from a debris strike on ascent resulted in the loss of crew and vehicle on entry on February 1, 2003.
At 81.7 seconds Mission Elapsed Time a piece of foam insulation from the External Tank (ET) left bipod ramp separated from the ET and struck the orbiter left wing leading edge in the vicinity of the lower half of reinforced carbon-carbon (RCC) panel #8, causing a breach in the RCC. During re-entry this breach allowed super-heated air to penetrate through the leading edge insulation and progressively melt the aluminum structure of the left wing, resulting in a weakening of the structure until increasing aerodynamic forces caused loss of control, failure of the wing, and break-up of the orbiter. This breakup occurred in a flight regime in which, given the design of the orbiter, there was no possibility for the crew to survive. (Similar bipod ramp foam releases prior to STS-107 occurred on STS-7, STS-32, STS-50, STS-52, STS-62, and STS-112.
Seven crew members were lost.
Mercury MR-4 | 7/21/1969 | Crew: 1 | Loss of Capsule
Inadvertent hatch pyrotechnic firing. Capsule sunk. Astronaut nearly drowned.
After landing on July 21, 1961 the spacecraft hatch pyrotechnic charges prematurely fired. The crew member was able to escape from the emergency situation, but because of waves flooding the capsule, the capsule sunk. The crew member was nearly drowned when the flight suit took on water from an unsealed neckdam. The crew member was rescued after three to four minutes in the water.
Apollo 12 | 11/24/1969 | Crew: 3
Harder than normal splashdown knocked loose a camera. The camera knocked lunar module pilot unconscious.
Due to a harder than normal splashdown on November 24, 1969, a camera broke free from the window bracket and struck the lunar module pilot on the forehead. The crew member was unconscious for five seconds after the injury and required sutures following retrieval.
Apollo 15 | 8/7/1971 | Crew: 3
Landed with only 2 of 3 parachutes.
On August 7, 1971 the Apollo capsule, Endeavour, dropped into the Pacific Ocean about 320 miles (515 kilometers) north of Hawaii. During the Earth landing phase, after the main parachutes were deployed and shortly after Reaction Control System (RCS) propellant dumping, one of the main parachutes was observed to be deflated when exiting the clouds (3 of 6 fabric risers failed and two-thirds of the suspension lines were missing). One of the main parachutes was recovered after landing, but the failed parachute was not recovered.
The investigation was divided into three areas which were likely causes of the parachute failure.
The forward heat shield was suspected because of the close proximity to the spacecraft flight path during the period when the failure occurred.
A broken riser/suspension-line connector link was found on the recovered parachute, indicating the possibility of broken links in the failed chute.
The Command Module RCS propellant depletion firing had just been completed, and fuel (monomethyl hydrazine) expulsion was in progress at the time of the failure, indicating the possibility of damage from propellants.
Analysis and testing ruled out possible causes one and two, but a test of raw fuel expulsion after RCS firing produced burning outside of the engine. The flame front extended up to eight feet from the engine exit and unburned fuel was sprayed up to 10 feet from the engine and ignited by burning droplets. This was considered the most likely cause of the parachute failure.
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.
STS-51D | 4/19/1985 | Crew: 7
Right brake failed (locked up) causing blowout of inboard tire and significant damage to outboard tire.
On April 19, 1985 the right brake on the orbiter failed, causing the blowout of an inboard tire and significant damage to an outboard tire. A crosswind of about 8 knots, gusting up to 12, resulted in extra brake energy on the right brake while returning to and holding the runway centerline during rollout. The number 3 stator on both the inboard and outboard right main landing gear (MLG) brakes broke into several pieces, causing both brakes to lock during rollout. The inboard right brake locked at 20.6 knots and 113 feet, before the orbiter stopped and the outboard right brake locked for the last 5 feet of rollout. The right MLG inboard tire burst 33 feet after the inboard brake locked. Eleven of 16 cord layers were worn through before the tire burst. The right MLG brakes failed and locked due to thermal soak-back when the number 3 stators broke.
The corrective action includes the following standard procedures to prevent heat soak-back:
1. Brake-on velocity between 140 and 120 knots.
2. Deceleration rate between 8 and 10 ft/sec2.
3. Deceleration rate reduced to 6 ft/sec2 at 40 knots. If brake-on velocity exceeds 140 knots, continue 8 to 10 ft/sec2 deceleration.
STS-37 | 4/11/1991 | Crew: 5
Several factors contributed to a low-energy landing 623 feet prior to the threshold of the runway at the backup landing location.
Low Energy Landing
On April 11, 1991 the first landing opportunity at Kennedy Space Center was waived due to fog, and a decision was made to land at the alternate landing site at Edwards Air Force Base. The entry wind profile included a large wind shear (from 90 kts at 13,000 feet to 10 kts at 8,000 feet). These conditions fell outside the Edwards Air Force Base 99 percent wind profile from 20,000 feet to 10,000 feet and significantly outside the shuttle experience base from 30,000 feet to 10,000 feet.
Around the Heading Alignment Circle (HAC), a significant amount of energy was lost due to a consistent negative pitch attitude error, and being outside of the HAC reference. The HAC is designed to provide an “energy pad” for use while making the approach. If guidance senses that the vehicle's energy state is getting very low (it uses altitude below a reference altitude and degrees of turn remaining to make the judgment), the HAC radius is decreased to help make up for the lower energy state. At approximately 23,000 feet the low-energy state triggered HAC shrink (when the range-to-go “distance from the vehicle to the runway” falls below the max lift/drag line). The HAC shrink was triggered due to the altitude being approximately 2-3,000 feet low, and increased the error from the HAC reference. A slow convergence back to the altitude reference was seen though the energy state remained low. At 13,500 feet the vehicle encountered a sharp wind shear reducing the vehicle's airspeed driving the energy state even lower. The vehicle again encountered wind shear at 8,600 feet in altitude. Touchdown occurred 623 feet prior to the threshold at 168 kts.
It is believed that the loss of energy on the HAC combined with the inadequate correction back to the altitude profile, both coming off the HAC and through the wind shear, resulted in the low-energy touchdown.
STS-90 | 5/3/1998 | Crew: 7
Hard, fast landing due to human factors and rogue wind gust. Hardest shuttle landing.
Following the landing on May 3, 1998 the post-mission report indicated a harder than normal landing. The main gear touchdown speed was 196.2 Keas with a sink rate of -3.18 feet/second. Brake energies were all below 15 million ft-lbs. The rollout distance was 9769.3 feet. Imagery analysis indicated a main landing gear sink rate of -6.7 ft/sec and a "harder than normal" landing. The Mission Operations Directorate Space Shuttle Summary reported touchdown of the main landing gear at 218 Keas, a -6.0 ft/sec sink rate, and a rollout distance of 9998 feet.
STS-108 | 12/7/2001 | Crew: 7
Violation of minimum landing weather requirements.
On December 17, 2001 the Shuttle Meteorology Group forecasted a “no go” for the de-orbit burn due to a weather forecast predicting the creation of a cloud ceiling at landing time. The Shuttle Training Aircraft reported a “go” based on observed conditions. Several positive factors provided the Flight Director with confidence to give a GO for landing on orbit 186, despite a weather forecast which could result in the crew being unable to see the Precision Approach Path Indicators (PAPIs) or runway environment until 3,000 feet or below. The GO was given with the belief that the cloud layer at 3,000 feet would break and that the PAPIs and runway environment would be visible by 6,500 feet. However, the cloud ceiling did not break and a flight rule was violated, but waived following the flight.
STS-137 | 6/1/2011 | Crew: 7
Brief fire observed between the left main landing gear tires during runway rollout.
STS-134 landed on June 1, 2011. During analysis of the post-landing imagery, a fire was briefly observed during the rollout period, located between the left Main Landing Gear tires shortly after the drag chute was jettisoned. Detailed visual inspections, material analysis, and landing gear systems tests were performed in an effort to determine the root cause of the fire. However, no definitive root cause could be determined. The fire may have been operational, because the commander applied the brakes in excess of recommended deceleration rates, especially at the lower speeds, resulting in the shortest landing rollout on a concrete runway surface since the drag chute had been used. The excessive braking may have generated higher than normal temperatures within the brake.
Mercury MA-6 | 2/20/1962 | Crew: 1
False landing-bag indicator light led to entry with retropack in place as a precaution.
On February 20, 1962 a sensor indicated the heatshield was in an unlatched condition while still in orbit. If the sensor's reading were true, the heatshield could have been lost during entry, resulting in the loss of the vehicle and crew. Because the indictor said the heatshield had been dropped to the landing position, entry procedures were changed to eliminate the jettisoning of the retropack. The retropack was used as a redundant heatshield hold-down device to keep the heatshield in place. The straps holding the retropack burned through during entry, but it was thought that the aerodynamic pressure would hold the heatshield in place. After landing it was discovered that the indicator was incorrect and that the heatshield had not been dropped to the landing position.
Mercury MA-7 | 5/24/1962 | Crew: 1
Pitch horizon scanner failed, resulting in manual entry and off-target landing. Delayed crew recovery.
On May 24, 1962 the failure of the spacecraft pitch horizon scanner required the pilot to assume manual control of the spacecraft for retrofire. As a result, the spacecraft attitude was outside of the recommended range for automatic initiation of the retrofire signal. Manual initiation of the retrofire signal occurred several seconds later than scheduled.
The delay in retrofire initiation and the less-than-ideal spacecraft attitude contributed to the spacecraft landing 250 nautical miles downrange of the intended landing point which delayed crew recovery.
Gemini 4 | 6/7/1965 | Crew: 2
Erroneous entry data uplinked; crew manually corrected entry flight profile.
On June 7, 1965 the computer could not be updated for entry, could not be turned off, and then stopped working entirely. The crew resorted to a rolling Mercury-type entry, rather than the lifting bank angle the computer was supposed to help them achieve.
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.
SpaceShipOne Flight 11P | 10/31/2014 | Crew: 1
Left main landing gear collapsed.
A nominal landing pattern was flown on December 17, 2003. However, touchdown caused the left main gear to collapse, and the vehicle rolled to a stop off the runway in the soft sand.
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
Apollo ASTP | 7/24/1975 | Crew: 3 | Crew Injury
N2O4 in crew cabin. Crew hospitalized for 2 weeks.
On July 24, 1975 as the spacecraft descended, the commander, who was reading the checklist, failed to tell the command module pilot to move the Earth Landing System auto/manual switch to auto. The crew saw that the spacecraft was well below the deployment altitude and proceeded to manually deploy the chutes. Drogue chutes were deployed manually at 18,550 feet instead of 23,500 feet as the automatic system would have done. At 10,000 feet the commander realized that ELS was not in AUTO and quickly switched ELS Logic and AUTO, deploying the main parachutes at 7,150 feet and disabling the RCS instead of 10,500 feet.. The Reaction Control System (RCS) was not disabled manually (RCS command switch turned to “off”) at this time. It was disabled manually at 16,000 feet instead of when the checklist indicated at 24,000 feet. The cabin pressure relief valve opened automatically at 24,500 feet.
During a 30-second period of high thruster activity after drogue parachute deployment, a mixture of air and propellant combustion products followed by a mixture of air and nitrogen tetroxide oxidizer (N2O4) vapors were sucked into the cabin. One of the positive roll thrusters is located only two feet away from the steam vent that pulls in outside air when the cabin relief valve is open. This exposed the crew to a high level of N2O4 since emergency oxygen masks were not available until landing. The pilot passed out, but the commander quickly put the oxygen mask on him and he was revived. The exposure resulted in a two-week hospital stay for the crew after landing.
SpaceShipTwo PF04 | 10/31/2014 | Crew: 2 | Loss of Crew (1)
Vehicle breakup during powered flight.
On October 31, 2014 shortly after separating from the WhiteKnightTwo carrier aircraft, the SpaceShipTwo vehicle broke apart resulting in the loss of one crew member. A National Transportation Safety Board investigation into the accident is ongoing.
Related or Recurring event
In addition to the three overheating/fire events on the ISS and the two significant events on Mir in 1997 and 1998, other overheating/fire events also occurred on:
Mir (October 1994) (A)
STS-40 (June 1991) (B)
STS-35 (December 1990) (C)
STS-28 (August 1989) (D)
STS-6 (April 1983) (E)
Salyut 7 (September 1982) (F)
Salyut 6 (1979) (G)
Salyut 1 (June 1971) (H)
Information on these events is contained in the reports below.
Related or Recurring event
On October 10, 2008 the crew reported smoke and odor emitting from the Russian condensate water processor unit [SRV-K]. The equipment housing was hot. When the air quality was tested using the Compound Specific Analyzer-Combustible Products (CSA-CP), carbon dioxide was found at five parts per million and acid gases, hydrogen chloride and hydrogen cyanide, were zero. The [SRV-K] was powered off and replaced, which resolved the issue.
On September 18, 2006 the crew reported smoke and a solvent smell. The Elektron was found to have released toxic byproducts. The CSA-CP registered carbon dioxide at seven parts per million and hydrogen chloride and hydrogen cyanide above one part per million.
In March 2005 an electrical odor was traced to a lamp on the Service Module.
Skylab 2 | 5/26/1973 | Crew: 3 | Related or Recurring event
Multiple failed automatic docking attempts resulted in manual docking to Skylab.
On May 26, 1973 numerous failed docking attempts resulted in the use of contingency in-flight procedures to bypass the automated docking system. Successful docking to the Skylab station ultimately relied on manual control and crew piloting skills.
The contingency procedure required the Skylab 2 crew members to don pressure suits, depressurize the command module cabin, open the tunnel hatch, cut wires in the probe, and connect the emergency probe-retract cable using a utility power outlet. The crew members were able to fire the probe-retract pyrotechnic and complete docking manually.
The failure to dock would have resulted in the loss of Skylab due to the inability to perform critical repairs.
STS-130 | 2/10/2010 | Crew: 6 | Related or Recurring event
Experienced significant misalignment between orbiter and ISS during post-capture free drift due to gravity-gradient-induced motion.
On February 10, 2010 there were significant oscillations between the orbiter and ISS on STS-130 during final ring retraction, requiring an additional 34 minutes in free drift to complete docking. Similar oscillations were observed during STS-133 docking.
Post-flight analysis from the STS-130 event indicated that the oscillations were caused by gravity gradient effects on the integrated vehicle stack (ISS/Shuttle) resulting in misalignment of the final docking ring and loss of the “RING ALIGN” indication. When the “RING ALIGN” indication is lost, the fixers are released, resulting in large misalignments.
Concern for ISS longeron shadowing and lack of data resulted in the Missions Operations Directorate not accepting the recommendation to rely on the fixers to maintain the alignment even after loss of “RING ALIGN” until after STS-133.
STS-133 | 2/26/2011 | Crew: 6 | Related or Recurring event
Experienced significant misalignment between orbiter and ISS during post-capture free drift due to gravity-gradient-induced motion.
On February 26, 2011 STS-133 experienced significant misalignment between the orbiter and the ISS during post-capture free drift due to gravity-gradient-induced motion.
There were significant oscillations between the orbiter and ISS on STS-133 during final ring retraction. The orbiter tipped approximately 10 degrees in pitch and four degrees in roll while ring retraction was paused. The time from contact to hardmate took 50 minutes.
Post-flight analysis of STS-133 docking operations raised several concerns, including vehicle-to-vehicle clearance, mechanism-to-mechanism contact, ISS free drift risks (longeron shadowing risk), timeline impacts, and the lack of a good integrated analysis tool.
At the Generic Joint Operations Panel on April, 6 2011 it was stated that the docking mechanism fixers will be used to maintain alignment during retraction. Engineering and safety agreed with the recommendation. The docking procedures were updated prior to STS-134/ULF6.
SR-71 | 7/30/1966 | Crew: 2 | Loss of Crew (1)
Loss of control at high speed and altitude.
On January 25, 1966 the SR-71 aircraft disintegrated during a high-speed, high-altitude test flight when the breakdown of super sonic airflow resulted in engine cutoff (also known as engine un-start). This occurred during a turn at speeds exceeding Mach 3.17 and a bank of 35 degrees. The bank immediately increased to 60 degrees. The nose pitched up and the aircraft broke apart. The pilot was thrown clear (his ejection seat never left the plane). He blacked out during the accident, but recovered and landed on the ground safely. His Reconnaissance System Officer did not survive the high-g bailout.
M21-D21 | 7/30/1966 | Crew: 2 | Loss of Crew (1)
D21 drone collided with M21 during launch, causing M21 breakup. Crew survived breakup but one was lost after water landing.
On July 30, 1966 as the M-21 mothership was performing a flight test for launching the D-21 drone, while traveling at high Mach speeds the drone was not able to penetrate the shock wave coming off the mothership. The D-21 had almost cleared the rudders of the M-21 when the drone bounced off the shockwave and pitched down, striking the M-21 and breaking it in half. The Pilot and Launch Control Officer (LCO) stayed with the tumbling wreckage of the plane a short time until a lower altitude was reached, then ejected over the Pacific Ocean.
Both crew members made safe ejections and landings, but after landing the LCO opened his helmet visor by mistake and his suit filled with water, causing him to drown. All subsequent flights of the D-21 were as D-21Bs, which were reconfigured to launch the drone from an under wing pylon of a B-52 (much like the X-15 had been), boosted to Mach 3 by a rocket motor that was jettisoned after the D-21Bs Marquardt ramjet was started.
Navy Chamber | 11/17/1962 | Crew: 4 | Crew Injury (4)
Fire started in a 100% oxygen environment at 5 psi. Four officers injured.
On November 17, 1962 four Navy officers were injured, two seriously, when a fire started in the altitude chamber they were occupying in a 20 day experiment at the U.S. Navy Air Crew Equipment Laboratory as part of a NASA atmosphere validation program.
The chamber contained 100% oxygen at 5 psi. The fire started when one officer changed a light bulb in an energized 24 volt DC light fixture. One wire in the fixture became disconnected resulting in arcing. A cotton towel was used in an attempt to smoother the fire. The towel caught fire, and the flames spread to the officers' clothes.
Mercury MA-9 | 5/16/1963 | Crew: 1 | Manual Entry
Electrical faults caused loss of some systems and need to perform manual entry. Also experienced high PPCO2 levels in suit during entry operations.
During the May 16, 1963 flight electrical faults caused the loss of some systems and the need to perform manual entry. The alternating current power supply for the control system failed to operate, and it was determined that the pilot would have to make a manual retrofire and re-entry. He performed these maneuvers with close precision and landed a short distance from the prime recovery ship in the Pacific Ocean.
The malfunction during re-entry on MA-9 was traced to two connectors in an electrical amplifier.
Gemini 8 | 3/16-3/17/1966 | Crew: 2 | Emergency De-orbit | Loss of Mission
Stuck thruster caused loss of control and led to 1st U.S. emergency
During the Gemini 8 flight from March 16 – 17, 1966 a stuck thruster, number 8, which controls roll, caused a loss of control and rapid spin rate of the capsule that could have led to the crew losing consciousness. To counter the effects the stuck thruster was turned off and the re-entry control system had to be used to stabilize the capsule. Use of the re-entry control system led the Gemini safety group to declare an end to the mission, which led to the first United States emergency de-orbit. The thruster apparently short circuited while attached to the Agena target vehicle.
STS-2 | 11/21/1981 | Crew: 2 | Minimum Duration Flight | Crew Injury
Failure of fuel cell resulted in a MDF being declared. The fuel cell failure also resulted in hydrogen in the drinking water leading to crew dehydration.
During the flight of STS-2, which spanned from November 12 – 14, 1981, fuel cell failure led to the declaration of a minimum duration flight (MDF). In addition to the MDF, the failure of the fuel cell also led to high hydrogen levels in the drinking water. The fuel cells used produce drinking water as a by product. When the crew drank this water it provoked a need to belch. Belching in zero g leads to regurgitation. The crew avoiding drinking the water in order to avoid belching, which caused crew dehydration. Prior to entry, crew members fluid load to offset fluid shift when returning from orbit. The crew dehydration increased the effects of the fluid shift and could have posed a risk during high g entry procedures due to a higher chance for loss of consciousness.
STS-9 | 12/8/1983 | Crew: 3
Two GPCs failed during reconfiguration for entry. One GPC could not be recovered.
On December 8, 1983 about five hours prior to the planned landing time, the orbiter's General Purpose Computer (GPC) 1 failed when the primary Reaction Control System jets were fired. About six minutes later GPC 2 also failed, leaving the orbiter in free drift for approximately five minutes before GPC 3 was brought online in OPS 3 entry mode (GPC 3 had been freeze dried for on-orbit operations). Attempts to bring GPC 1 back online were unsuccessful, and it was powered down.
Although problems had occurred, GPC 2 was reinitialized and placed back online, and GPCs 2, 3, 4, and 5 were configured for entry. This off-nominal configuration led to further problems, and delayed the landing time by about eight hours. Entry was set up without GPC 1, and upon landing GPC 2 failed again. Particle Impact Noise Detection testing was instituted to screen out any contamination of the GPC boards, and a spare GPC was flown for several flights after STS-9, but was later dropped as a requirement.
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-44 | 11/24/1991 | Crew: 6 | Minimum Duration Flight
Failure of IMU 2 caused MDF to be declared. 10-day mission shortened to 7 days.
Failure of Inertial Measurement Unit (IMU) number 2 on November 24, 1991 caused minimum duration flight to be declared. The 10-day mission was shortened to seven days. In an attempt to recover normal operation of the IMU, it was placed in standby, operate, and then power cycled. These actions were not successful. Failure of this IMU invoked a flight rule requiring minimum duration flight for loss of one IMU.
Post-flight troubleshooting in the Inertial Systems Laboratory at Johnson Space Center isolated the problem to a failed computer interface card. This card converts analog acceleration signals into digital signals. The failed card was sent to the manufacturer for further analysis which revealed that a filter capacitor (C14), located within a chopper-stabilized amplifier hybrid component (U12) in the Z-accelerator channel, had shorted. This short circuit caused a bond wire from U12 pin 9 to the card case to fuse open.
STS-51 | 9/12/1993 | Crew: 5 |
Both port-side primary and secondary SUPER*ZIP explosive cords fired, resulting in containment tube failure and damage in the payload bay.
On September 12, 1993 the STS-51 mission commands intended to initiate the primary SUPER*ZIP explosive cord resulted in the simultaneous firing of both the primary explosive cord and back-up explosive cord. This simultaneous explosive cord firing resulted in the rupture of a SUPER*ZIP containment tube and release of contaminants and high-energy debris into the orbiter cargo bay. The orbiter sustained damage to blankets, wire tray covers, the 1307 bulkhead, and Thermal Protection System tiles. If debris had hit critical items it could have resulted in a loss of the orbiter and crew.
STS-83 | 4/6/1997 | Crew: 7 | Loss of Mission
Failure of fuel cell number 2 resulted in MDF being declared. The 15-day mission was shortened to 3 days.
A failure of fuel cell (FC) number two resulted in a minimum duration flight being declared on April 6,1997. The 16-day mission was reduced to four days due to FC problems encountered on flight day two. During prelaunch operations the differential voltage on FC 2, substack 3 remained above the 150mV limit (defined in the Operations and Maintenance Requirements Specification Document) for an unusually long period of time before dropping below 150 mV. The substack delta voltage began to trend upward shortly after on-orbit operations began at approximately two hours Mission Elapsed Time. FC purges were ineffective at stopping the trend. FC 2 was subsequently shut down and safed to prevent the possibility of a crossover condition, and multiple payloads had to be powered down. FCs 1 and 3 continued to carry the total orbiter load and performed nominally.
Post-flight failure analysis of FC 2 did not identify a root cause for the on-orbit anomaly experienced, but did identify degraded cells and verified the cell performance monitor (CPM) was functioning properly. No foreign material/contaminant was found and the most credible scenario implies an abnormal external event affected a group of cells prior to start–up. It has been hypothesized that this external event was the presence of oxygen in the oxygen side of substack 3 of the FC at a time when the FC was supposed to be inerted with helium. This event over time would cause oxidation of the nickel Electrolyte Reservoir Plate and dissolution of palladium and platinum in the anodes. Migration and plating of the palladium onto the cathode catalyst would cause high open circuit voltage once full reactants are applied to the FC.
As a result of this anomaly and failure analysis, the Launch Commit Criteria was revised to not allow launching with an FC showing similar prelaunch CPM readings. Kennedy Space Center FC purge procedures were also modified to preclude the potential for the presence of oxygen in inerted fuel cells. The program also designed and tested a fuel-cell monitoring system (FCMS) which finally provided individual cell-health monitoring capability (STS-87 first flight). If the FCMS had been available for STS-83, it may have precluded the shutdown of the FC and may have allowed the mission to complete its planned duration.
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.
STS-95 | 10/29/1998 | Crew: 7
Preflight sterilization process chemically altered the Low Iodine Residual System resulting in contaminated drinking water.
During STS-95 on October 29, 1998 the preflight sterilization process chemically altered the Low Iodine Residual System (LIRS) resulting in contaminated drinking water.
The crew reported the water from the galley with the LIRS installed had a bad taste. Samples of the water were taken post-flight for analysis. The LIRS was removed. A purge of the LIRS water from the galley plumbing was conducted. After the purge the Galley Iodine Removal Assembly was reinstalled and the crew reported that the water taste was normal after the change-out of hardware.
The bad tasting water could have led to possible crew dehydration due to the crew drinking less water.
ISS, Flight 2A.1 | 5/1999 | Crew: 7
Crew sickened in FGB; likely a result of high localized CO2 levels due to poor ventilation.
During ISS Flight 2A.1 in May 1999, the crew was sickened in the Functional Cargo Block [FGB], likely as result of high localized carbon dioxide levels due to poor ventilation. The evidence suggests that human metabolic products (carbon dioxide, water vapor, heat) were not being effectively removed from the crew member work area, and therefore caused the symptoms.
The number of crew members working in the [FGB] may also have contributed to the air quality issues.
Additionally, the flexible air duct running between the orbiter's Pressurized Mating Adapter 1 and the [FGB] may have contributed to poor air quality. The flexible air duct has a tendency to collapse with only minor incidental contact.
This duct was later redesigned to minimize the potential for collapse and restricted air flow.
STS-99 | 2/2000 | Crew: 6
High bacterial count in postflight sample after GIRA installed to removed iodine.
During the February 2000 flight, a high bacterial count of 160 colony forming units (CFUs)per 100 ml was discovered in a post-flight sample after the Galley Iodine Removal Assembly was installed to remove iodine. The level should have been less than 100 CFUs/100 ml.
ISS, Increment 2 | 4/24/2001 | Crew: 10
Failure of all U.S. command and control computers on ISS.
On April 24, 2001 the ISS Command and Control (C&C) Multiplexer/Demultiplexer (MDM)-1 suffered hard drive errors that resulted in C&C-1 going offline.C&C-2 automatically switched from backup to primary mode, but suffered hard drive errors. C&C-3 was brought online but also failed. This resulted in complete loss of command and control to the United States orbital segment. C&C-2 was restored and placed into operation in primary mode. Flight controllers were able to uplink critical C&C software into the dynamic random access memory of C&C-3. C&C-3 was declared operational except the hard drive. C&C-1 was replaced with an identical payload computer.
If the MDMs were unrecoverable, the failure could have resulted in the loss of the United States orbital segment.
STS-104 | 7/2001 | Crew: 5
EMU battery leaked hazardous KOH. Discovered during EMU checkout.
During the first pre-extravehicular activity checkout of the July 2001 flight, an increased capacity extravehicular mobility unit (space suit) battery was discovered to have leaked hazardous potassium hydroxide. The leakage resulted in potassium hydroxide deposits on the contamination control cartridge, the water tank structure, and other locations on the primary life support system of the space suit.
ISS | 8/2001 | Crew: 3
Extremely high methanol levels in FGB air sample.
During August 2001 Functional Cargo Block [FGB] air samples contained extremely high methanol levels. The source of the methanol was never identified.
ISS, Increment 4 | 2/2002 | Crew: 3
MetOx regeneration caused noxious air.
During ISS Increment 4, February 2002, the United States Orbital Segment metal oxide canister regeneration caused the release of pollutants into the air.
ISS, Increment 2-4 | 4/2001-3/2002 | Crew: 3
Freon 218 leaked from SM AC.
During ISS Increments 2-4, which spanned from April 2001 until March 2002, Freon 218 leaked from the Service Module air conditioner.
Freon 218, a coolant fluid used in the air conditioning system, was found in cabin air quality samples after Zvezda was activated. The concentration of the Freon increased slowly once detected, but then rapidly as time passed. Peak concentrations reached over 600 mg/m3.
ISS, Increment 5&6 | mid 2002-2/03 | Crew: 3-10
Formaldehyde periodically exceeded long-term limits.
During ISS Increments 5 and 6, which spanned from mid-2002 until February 2003, formaldehyde levels onboard the station periodically exceeded the long-term limits.
ISS, Increment 10 | 2/2005 | Crew: 2
Potential acid preservative aerosol escape from Russian urinal.
In February 2005 during ISS Increment 10 an acid preservative aerosol escaped from the Russian urinal.
ISS, Increment 13 | 8/2006 | Crew: 3
Triol coolant leak in SM.
In August 2006 during ISS Increment 13, about 150 grams of Triol coolant leaked from the Docking Compartment-1 Hydraulic Cooling Loop Connector in the Service Module. The crew cleaned and monitored the area throughout the day. No additional leaking was reported.
ISS, Increment 15 | 6/10-6/18/2007 | Crew: 10
Power switch failures caused loss of ISS propulsive attitude control capability.
On June 10-18 2007 Russian computers that provide ISS propulsive attitude control [ТВМ], and Russian segment command and control capability [ЦВМ], experienced multiple automatic and manual restarts. ISS attitude control was maintained by the docked shuttle (Atlantis STS-117/13A) while Russian specialists and US teams worked to restore consistent power to the computers. The Russian cosmonauts were able to re-establish two of three computers on both systems ([ТВМ], [ЦВМ]) by June 18 after bypassing the secondary power circuitry to provide a continuous “ON” command.
Troubleshooting later identified the root cause to be an electrical short in the line resulting from corrosion of cabling within the Command Acquisition (Processing) Unit [БОК3] which monitors power. The short caused a power-off command to be passed to all six computers. The corrosion was presumed to be caused by increased humidity resulting from the close proximity of an air separator to the [БОК3]. The [БОК3] was subsequently relocated to a separate compartment.
If the Russian computers were unrecoverable, the failure could have resulted in the loss of ISS attitude control and loss of ISS.
ISS, Increment 17 | 4/30/2008 | Crew: 3
Freon 218 leaked from SM AC.
On April 30, 2008 Freon 218 leaked from the Service Module air conditioner. The Russian [BMP] (Russian Microimpurities Removal System) was modified to allow for faster removal of the Freon 218.
ISS Increment 38 | 12/1/2013 | Crew: 6
ITCS configuration errors resulted in near freezing and potential rupture of water-to-ammonia heat exchanger.
On December 11, 2013 the failure of a flow control valve in the pump module of the External Thermal Control System (ETCS) and subsequent Internal Thermal Control System (ITCS) reconfiguration led to a drop in water temperature to nearly freezing in the Columbus module's Moderate Temperature Heat Exchanger (MTHX). If the water in the Interface Heat Exchanger (IFHX) had frozen, the expansion could have ruptured the barrier between the ITCS and the ETCS. A rupture of this barrier could allow ammonia to enter the interior crew portions of the ISS, causing a potential loss of crew/loss of vehicle.
SpaceShipOne 14P | 5/13/2004 | Crew: 1
Flight computer unresponsive. Recovered by rebooting.
On May 13, 2004 the flight computer on SpaceShipOne became unresponsive. During the boost following the vertical part of the trajectory, the avionics display flickered and went blank. The ground displays did not show an error. The avionics display on SpaceShipOne came back on as soon as the motor shut down.
Due to the loss of avionics during the boost, the trajectory was not precise. The avionics malfunction was traced to a dimmer, a small electrical component.
SpaceShipOne 16P | 9/29/2004 | Crew: 1
Uncommanded vehicle roll. Control regained prior to apogee.
On September 29, 2004 SpaceShipOne performed a series of 60 rolls during last stage of engine burn. SpaceShipOne coasted to 103 km of altitude and successfully completed the first of two X-Prize flights. The motor was shut down when the pilot noted that his altitude predictor exceeded the required 100 km mark. During the motor burn the spacecraft began to roll uncontrollably, but the pilot continued despite advice from the ground to shut the motor down and abort the attempt.
The thin air at that altitude meant that the control surfaces didn't have enough air flowing over them, so they lost effectiveness to compensate for the roll as the spacecraft pointed nearly straight up. The pilot needed to correct the rolling that occurred because of asymmetric thrust coming from the engine.
To correct the issue for the 17P flight, the amount of allowable “down pitch trim” was limited, to avoid the negative-lift condition. The solution was to more gently turn the corner, such that a forward correction later would not be needed. Pointing straight up at burnout was determined to be acceptable, as long as negative lift was not created. This problem was corrected on SpaceShipTwo.
Apollo 11 | 7/21/1969 | Crew: 2
Engine arm circuit breaker knob broke off. Circuit breaker successfully reset allowing ascent.
On July 21, 1969 while preparing for extravehicular activities, the engine arm circuit breaker broke, probably due to an impact from the oxygen purge system. A felt tipped pen was used to successfully depress the circuit breaker when needed. Circuit breaker guards were installed on Apollo 12 and subsequent vehicles to prevent the oxygen purge system from impacting the circuit breakers.
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.
Apollo 13 | 4/13/1970 | Crew: 3 | Loss of Mission
Explosion due to electrical short. Loss of O2 and EPS.
Apollo 13 launched on April 11, 1970. On April 13, 1970 during trans-lunar flight at approximately 56 hours, one of the two Service Module oxygen tanks over-pressurized and exploded. This caused the loss of oxygen in that tank and a leak of oxygen out of the remaining tank. This resulted in the loss of all three fuel cells, loss of the primary oxygen source, and the loss of electrical power to the Command Module (except for the entry batteries). The mission was able to continue with the use of the Lunar Module, and the crew safely returned.
Prior to launch, the following conditions resulted in the oxygen tank failing during the mission: By design the cryogenic oxygen tank required both electrical heaters to maintain pressure, and fans to prevent stratification. The tank was a complex assembly with blind installation of the quantity probe, heater/fan assembly, and fill tube. This design leaves wiring insulation vulnerable to damage during assembly with no way to inspect after installation. The Teflon insulated wiring, which is a combustible material in the oxygen tank, was in close proximity to the heater elements and fan.
The Apollo 13 tanks were originally installed on Apollo 11, but a change required the tanks to be removed. During removal of the oxygen shelf, one bolt was left in place causing the fixture to break and resulting in a two-inch drop of the shelf and tanks. Although a loosely fitting (due to loose specification tolerances) fill tube could have been displaced by this, all testing was passed. No cryogenic tests were performed which would have revealed the problem. During the Count Down Demonstration Test the oxygen tank could not be emptied by the normal means of pressurized oxygen gas due to a leak at the fill tube. Instead, the tank heaters were turned on to boil off the oxygen in the tank. The thermostatic switches were rated for 30 volts direct current, but several years earlier the heater ground power supply voltage was raised to 65 volts to reduce the pressurization time. As the temperature increased the thermostatic switch opened and the higher voltage caused the contacts to weld closed. With the heaters continuously on, the temperature approached 1000 degrees and damaged the wire insulation, setting up the conditions for a short and ignition inside the tank. Ground personnel did not notice the continuous heater operation. During the prelaunch problem solving neither the Apollo Spacecraft Program Manager nor the Kennedy Director of Launch Operations knew the tank had previously been dropped or that the heaters had been on for eight hours.
Apollo 14 | 1/31/1971 | Crew: 3
Multiple failed docking attempts. Contingency procedures developed to mitigate risk of recurring docking anomaly. Docking successful.
On January 31, 1971 six docking attempts were unsuccessful following translunar injection. On the seventh try the command module pilot was told to fire thrusters to hold the command module to the lunar module while the docking probe was retracted. The docking capture latches were triggered to fire the probe-retract pyrotechnic, and docking was successful. After docking, the drogue and probe were examined by the crew and appeared normal. No other issues arose with the docking mechanisms.
The most likely cause of the docking issue was a piece of debris or ice on the docking probe from rain water entering the boost shroud the day before launch. The mission would have been No-Go for lunar module separation and landing if a backup procedure for docking and retracting the docking probe in an emergency had not been developed.
Late Release Orbiter Tyvek Covers
During the ascent phase of:
STS-114 | 7/26/2005
STS-115 | 9/9/2006
STS-118 | 8/8/2007
STS-119 | 3/15/2009
STS-124 | 7/31/2008
STS-126 | 11/14/2008
The Tyvek covers for the reaction control system jets released later than intended. The late release resulted in the Tyvek reaching a higher velocity than designed, thus posing a higher risk of damage to the orbiter.
Tyvek covers were used to keep rain water and other debris out of the reaction control system jets while the shuttle was on the launch pad.
STS-116 | 12/10/2006 | Related or Recurring event
STS-125 | 5/11/2009 | Related or Recurring event
Debris from the flight vehicle or the launch facility poses a risk to all crewed space launches.
Other significant ascent debris events have occurred on:
On December 10, 2006, during ascent, booster trowelable ablative (BTA) around the solid rocket booster (SRB) left hand aft booster separation motor (BSM) nozzle liberated and was seen striking the bottom of the orbiter shortly after SRB separation began. The root cause of the observed debris liberation was determined to be stress buildup in the thermal protection system (TPS) and BTA closeout configuration at the aft BSM exit cone. The failure mechanism that produced the forward transport of the BTA was a hinging effect on the liberated BTA coupled with plume expansion at altitude, which projected the debris forward to where it struck the orbiter.
The BTA closeouts around the aft BSM nozzles were modified to eliminate the failure mechanism that caused the forward debris transport as well as limit the size and release angle of BTA liberated at SRB separation.
On May 11, 2009 during ascent the Wing Leading Edge Impact Detection System (WLEIDS) recorded two impacts on the starboard chine, which was confirmed by video imagery analysis.
Damage detected during on-orbit inspection was below the damage tolerance threshold.
STS-95 | 10/29/1998 | Crew: 7 | Related or Recurring event
Drag chute door separated during launch and impacted main engine bell.
On October 29, 1998 during the main engine ignition sequence, the drag chute panel fell away from the vehicle. Video of the launch confirmed the drag chute door detached three seconds prior to liftoff and hit the engine nozzle of Space Shuttle Main Engine (SSME) 1.
The remains of the door were found during a post-launch pad inspection, revealing that at least one aluminum shear pin used to attach the door sheared. The root cause of the shearing was a combination of the high pressure environment caused by SSME ignition and a low margin hinge pin. The drag chute door struck the nozzle of the center main engine but did not do any appreciable damage to the engine or vehicle. Concerns about the status and condition of the chute contained in the drag chute compartment resulted in the decision not to deploy the chute during landing.
As a precautionary measure, two subsequent missions used a solid closeout panel bolted over the drag chute compartment and did not utilize the drag chute. All flights following these missions used Inconel instead of aluminum as shear pin material.
STS-124 | 5/31/2008 | Crew: 7 | Related or Recurring event
Pad 39-A flame trench suffered significant damage causing about 3,500 refractory bricks to be blown away from the flame trench wall.
On May 31, 2008 during the launch of STS-124, the Pad 39-A flame trench suffered significant damage causing about 3,500 refractory bricks to be blown away from the flame trench wall. The bricks scattered away from the flame trench beyond the pad perimeter fence. This debris could have damaged the vehicle or the launch complex.
Pad generated debris is a concern for any spaceflight, due to the potential for debris to travel in a manner that can damage the vehicle or the launch complex.
STS-93 | 7/23/1999 | Crew: 5
At T+5 a short on AC1 Phase A resulted in loss of SSME1 Controller A and SSME3 Controller B.
SSME3 H2 leak: early LOX depletion and shutdown.
STS-93 encountered two close-call events.
STS-114 | 5/26/2006 | Crew: 7
Bird strike on External Tank.
Loss of foam from External Tank PAL ramp.
TPS gap filers protruding. Removed during third mission EVA.
Missing O-ring resulted in ejection of one of two NSIs, compromising the ET forward
separation bolt function and damaging secondary structure and a thermal blanket.
STS-114 encountered four close-call events.
STS-117 | 6/8/2007 | Crew: 7
Thermal blanket damage. EVA performed to repair damage.
On June 8, 2007 during ascent, a thermal blanket covering the port orbital maneuvering system (OMS) pod was damaged.
An unplanned extravehicular activity, a high risk operation, was performed to repair the damaged blanket, so the blanket could effectively prevent potential damage to the vehicle from heating during entry. Failure of the thermal protection during entry could have resulted in overheating of the OMS and catastrophic structural failure of the vehicle and loss of crew.
The blanket was repaired by inserting pins between the thermal blanket and the surrounding shuttle tiles. A surgical stapler was also used in fastening the two thermal blankets together.
Gemini 6 | 12/12/1965 | Crew: 2
Main engine shutdown. Booster left unsecured on pad. Crew elected not to eject. Launched 3 days later. After the failed launch attempt, review of engine data and an inspection of the number 2 engine revealed that a plastic dust cover had been inadvertently left on the oxidizer gas generator inlet port causing blockage of oxidizer to the gas generator. Ground procedures were modified to ensure removal of dust covers during engine assembly.
There was a main engine shutdown during the attempted launch on December 12,1965.
About 1.5 seconds after main engine ignition, an electrical plug fell from the vehicle and accidentally started a clock that normally starts during vehicle liftoff.
The rocket malfunction detection system sensed an anomaly since there was no upward motion associated with the start of the clock and triggered engine stop. The booster was left unsecured on the pad with the crew inside. The crew members elected to remain in the capsule until the gantry was returned.
A successful launch occurred three days later.
Apollo I (AS-204) | 1/27/1967 | Crew: 3 | Loss of Crew
Crew cabin fire (electrical short + high pressure O2 atmosphere).
On January 27, 1967 the crew cabin of Apollo 1 caught fire during a test with three crew members inside. The cabin was filled with a pure oxygen atmosphere and pressurized greater than ambient pressure (16.7 psi). Over the course of several hours, the oxygen permeated all materials in the cabin, which had been tested to the normal flight pressure of pure oxygen (5 psi). When the fire began it spread rapidly. Due to the pressure in the cabin, the crew members could not open the hatch to escape. Technicians in the room outside the capsule attempted to open the hatch but were driven back by the heat and smoke. Some technicians donned the available gas masks, but the masks were designed to protect against hypergolic propellant fumes, not smoke. Consequently, these technicians lost consciousness after a short time in the smoke-filled room.
All three crew members were lost.
The fire was caused by an electrical short from an unprotected wire. A subsequent review of all wiring dioded to both Main Bus A and B identified a problem with an environmental control system instrumentation wire powered from Main Bus A and B. The wire was routed over plumbing lines on the crew compartment floor, located below the left-hand crew seat, going into the left-hand equipment bay, between the environmental control unit and the oxygen panel. This Teflon-insulated wire should have had a protective Teflon overwrap, but closeout photos showed that the overwrap had slipped down, no longer providing protection. The commander likely contacted this wire with his foot when he turned to change his communications cable. The most probable initiator of the fire is an electrical arc from this wire, which was unprotected from external damage.
Factors contributing to this accident include:
STS-1 | 4/12/1981 | Crew: 2
SRB ignition pressure wave caused TPS and structural damage.
During the April 12, 1981 launch of STS-1, a higher than expected solid rocket booster ignition pressure wave caused damage to both the thermal protection system and structure.
Other On Pad Aborts | Related or Recurring events
On pad aborts have also occurred on:
STS-51F: On July 12, 1985 the first launch attempt of STS-51F was aborted at T-4.2 seconds. The chamber coolant valve on Space Shuttle Main Engine 2 needed to move from the 100% open setting to the 70% open setting required for startup, but the valve responded slowly. The launch was aborted and the vehicle safely shut down.
STS-55: On March 22, 1993 the third launch attempt of STS-55 was aborted at T-3 seconds when the Space Shuttle Main Engine (SSME) 3 oxidizer preburner purge pressure exceeded the maximum pressure. The monitor detected that the combustion product pressure exceeded the 50 psi redline due to one of the five check valves in the purge system leaking. The launch was aborted and the vehicle safely shut down. All three SSMEs were replaced before the next launch attempt.
STS-51: On August 12, 1993 the first launch attempt of STS-51 was aborted at T-3 seconds because of a disagreement in the turbine fuel flow sensors of Space Shuttle Main Engine (SSME) 2. At 0.6 seconds after ignition SSME 2 experienced a failure of the fuel flow meter channel A2 speed pickup coil sensor (loss of redundancy) to respond to the start transient. The main engine controller used the A and B flow rate measurement data for closed-loop thrust/mixture ratio control, and the failure of either the A or B measurements resulted in lockup of the preset engine mixture ratio. As a result of the failure, the engines were safely shut down. All the main engines were replaced before the next launch attempt.
STS-68: On August 18, 1994 the first launch attempt of STS-68 was aborted at T-1.9 seconds. The abort was triggered after the discharge temperature of the high pressure oxidizer turbopump in Space Shuttle Main Engine (SSME) 3 exceeded the redline temperature, causing the main engine controller to issue a shutdown of SSME 3, followed shortly by the General Purpose Computers issuing a shutdown of the other two engines.
STS-41D | 6/26/1984 | Crew: 6 | Related or Recurring event
During the launch attempt on June 26, 1984 all aspects of launch countdown were nominal until T-4 seconds when an irregular operation of the Space Shuttle Main Engine (SSME) 3 main fuel valve resulted in an engine shutdown and pad abort condition.
The SSME 3 main fuel valve failed to open when commanded and leaked hydrogen for approximately 21 minutes following the shutdown. After the abort, fire was seen on the starboard side of the body flap for approximately 12 minutes. The aft base heat shield water deluge system was activated and was able to extinguish the fire. Damage to the orbiter was mostly confined to the body flap despite invisible flames of burning hydrogen reaching 190 feet.
The Mobile Launcher Platform sustained minor scorching damage to a few purge ducts and one burnt ground wire. The cause of the main fuel valve anomaly was attributed to contamination in the hydraulic actuator that was likely present at time of installation.
Additional inspection and ground tests were implemented to detect valve anomalies prior to installation into flight engines. Real-time monitoring of valve operation prior to engine start was also enhanced, and ultraviolet fire sensors were added.
STS-61C | 1/6/1986 | Crew: 7
System configuration errors resulted in inadvertent drain back of 14,000 lbs of LOX prelaunch, which would have resulted in a Trans-Atlantic Abort Landing.
On January 6, 1986 during the second launch attempt of STS-61C, the MPS liquid-oxygen inboard fill-and-drain valve was not commanded closed because the liquid-oxygen (LOX) loading automatic sequencer (terminal countdown sequencer / control software) did not receive the closed-switch indication from the replenish valve as required by the prerequisite control logic. This resulted in the automatic sequencer initiating a hold at launch minus 4 minutes 20 seconds. The ground operator verified replenish-valve closure using flowrate and other parameters, but did not close the inboard fill-and-drain valve prior to issuing the resume command to the automatic sequencer at launch minus 2 minutes 55 seconds. This allowed LOX to drain back out of the external tank through the tail service mast vent-and-drain valves until the ground operators noticed the inboard fill-and-drain valve was still open and manually closed the valve. Although unknown at the time, approximately 14,000 to 18,000 lbm of LOX had been inadvertently drained out of the external tank.
Another hold was initiated by ground personnel at launch minus 31 seconds to review the previous out-of-sequence loading termination and obtain a 5-minute liquid-oxygen drain through the main engines. During the hold, the liquid-oxygen main engine temperature dropped below the engine start requirement of 168.3 degrees Rankine by approximately 3 degrees. The engine limit was exceeded because the amount of LOX lost overboard through the fill-and-drain valve caused the colder, more-dense LOX to be drawn in from the external tank. The countdown was recycled to launch minus 20 minutes and oxygen replenish flow was reestablished. The launch was scrubbed when it was determined that the vehicle could not be recycled within the allowable launch window. If the launch had occurred, the reduced LOX quantity in the external tank would have caused early SSME shutdown due to LOX depletion resulting in a Trans-Atlantic Abort Landing (TAL).
Corrective action incorporated in response to this close call included modifications to automatic sequencer software to prevent the prerequisite control logic from blocking LOX inboard fill-and-drain valve close commands, updates to countdown procedures and launch constraints to verify closure of the inboard fill-and-drain valve after replenish valve closure and prior to tail service mast vent-and-drain valve opening, monitoring and initiation of holds if the fill-and-drain valve closed indication is lost, implementation of helium repressurization “pulse purge” if ET ullage pressure drops below 0.25 psi, and verification of minimum ET ullage pressure rise rate at T-120 seconds.
A subsequent launch attempt on January 7, 1986 was scrubbed at the T minus 9 minute hold due to weather constraint violations at TAL sites. However, during post launch scrub operations a broken Ground Support Equipment (GSE) LOX temperature probe was found lodged in SSME #2 pre-valve post-detanking. This temperature probe had failed (off-scale high reading) during LOX loading but due to the absence of any mechanical failure history, the failure was assumed to be electrical in nature and the temperature data from this probe was not mandatory for pre-launch loading operations. The broken temperature probe would have prevented closure of pre-valve during flight, at MECO, resulting in an uncontained SSME failure and possible loss of crew, loss of vehicle.
As a result of the broken GSE LOX temperature probe, all GSE LOX temperature probes were inspected and screened for improper welds, monitored during pre-launch operations for any anomalies, and eventually replaced with redesigned probes. A coarse debris screen was also added upstream of the LOX prevalve to prevent large debris from entering into the prevalve.
STS-112 | 10/7/2002 | Crew: 6
T-0 umbilical issues resulted in none of the system A pyrotechnic charges firing.
The post-launch data review of the October 7, 2002 launch determined that none of the system A pyrotechnics (NASA Standard Initiators) for the Solid Rocket Booster hold-down posts nor the External Tank Vent Arm System discharged.
The Master Events Controller (MEC) provided the signal to the Pyrotechnic Initiator Controller (PIC) to discharge the pyrotechnics. Therefore the MEC common wiring, as well as the wiring between the MEC in the orbiter and the PIC rack on the ground, were suspected of not working properly.
All connectors and electrical circuits were inspected and tested, but no root cause was identified to explain the anomaly.
Gemini 10 | 7/18/1966 | Crew: 2
1st stage oxidizer tank exploded at staging. No discernible effects. Nominal ascent.
On July 18, 1966, 1.2 seconds after booster engine cutoff, tracking cameras displayed an amber cloud and an unusual amount of debris. The evidence indicated the first stage oxidizer tank ruptured after the normal staging sequence.
This event had no detectable effect on the operation of the second stage.
Apollo 12 | 11/14/1969 | Crew: 3
Lightning strike on ascent.
During the Apollo 12 launch on November 14, 1969 lightning struck the spacecraft.
Light rain was falling, but weather conditions did not indicate any thunderstorm activity. There were seven miles of visibility with cloud break estimated at 800 feet and overcast conditions at 10,000 feet.
At 11:22am, T+36 seconds, the crew saw a bright light.
At T+36.5 seconds many errors occurred: Fuel Cells 1, 2, and 3 disconnected; Main Buses A and B were under-voltage; Alternating Current (AC) Buses 1 and 2 overloaded. The warning lights and alarm came on in the cabin, indicating failure of the Inertial Stabilization System.
At T+52 seconds (13,000 feet) lightning struck the vehicle and the Inertial Measurement Unit platform tumbled.
The potential effect on the vehicle was induction into wiring, depending on the location and rate of change of potential and direct current flow in grounding. The high negative voltage spike (delta voltage/delta time) caused the Silicon Controlled Rectifiers to trip on the Fuel Cell and AC Inverter overload sensors. Failures occurred in four Service Module Reaction Control System helium tank quantity measurements, five thermocouples, and four pressure/temperature transducers.
Using power from the Battery Relay Bus, the crew reconnected the Fuel Cells to Main Bus A and B, and reconnected the inverters to AC Bus 1 and 2. The mission continued.
Apollo 13 | 4/11/1970 | Crew: 3
2nd stage center engine shutdown due to pogo oscillations.
During the April 11, 1970 launch of Apollo 13 severe pogo oscillations were experienced. Acceleration at the engine attachment reached an estimated 34 g (the accelerometer went out of recordable range) before the engine's combustion chamber low-level pressure sensor commanded an engine shutdown.
STS-51F | 7/29/1985 | Crew: 7 | Abort to Orbit
Temperature sensor problems resulted in SSME1 shutdown at T+5:45.
On July 29, 1985 at T+5:43, both temperature sensors for the Space Shuttle Main Engine (SSME) 1 high pressure fuel turbopump showed readings exceeding the redline limit. This resulted in a premature shutdown of SSME 1 and declaration of an Abort to Orbit condition, the first in program history. At T+8:13, one of the two temperature sensors on SSME 3 indicted a high reading, but auto-shutdown was inhibited to assure STS-51F achieved an acceptable orbit.
These events were a result of confirmed instrumentation failures. The temperature sensors were removed for engineering analysis following the flight, and a new configuration sensor was used on subsequent engines.
Other SRB Seal Events | Related or Recurring event
SRB gas sealing anomalies have also occurred on:
STS-2 November 12, 1981
STS-6 April 4, 1983
STS-11 (STS-41B) February 3,1984
STS-41C April 6,1984
STS-41D August 30, 1984
STS-51C January 24,1985
STS-51D April 12, 1985
STS-51B April 29, 1985
STS-51G June 17,1985
STS-51F July 29, 1985
STS-51I August 27, 1985
STS-51J October 3, 1985
STS-61A October 30, 1985
STS-61B November 26,1985
STS-61C January 12, 1986
STS-42 January 22, 1992
STS-71 June 27, 1995
STS-70 July 13, 1995
STS-78 June 20, 1996
STS-51L (Challenger) | 1/28/1986 | Crew: 7 | Loss of Crew | Related or Recurring event
SRB seal failure.
On January 28, 1986 a combustion gas leak developed in the right solid rocket motor aft field joint shortly after solid rocket booster ignition. The resulting hot gas plume exiting the joint impinged upon the SRB lower attachment strut and adjacent external tank structure weakening the structure to the point of failure.
Seventy-four seconds into flight, the Space Shuttle Challenger broke up.
All seven crew members were lost.
Soyuz TM-9 | 2/11/1990 | Crew: 2
DM insulation torn loose on ascent; contingency EVA repair.
During the docking of Soyuz TM-9 on February 11, 1990, the TM-8 crew aboard Mir noticed three of the eight descent module's thermal blankets had partially detached near the heat shield during ascent. This raised five concerns:
A rescue mission with a cosmonaut aboard Soyuz-TM 10 was considered, but not executed. The temperature of TM-9 was stabilized by the Mir directing it into alignment with the sun.
Four months later, the Kristall module (90-048A) arrived with the special tools needed to repair the decent module. Cosmonauts were able to secure the blankets out of the sensor's line of sight after an EVA longer than seven hours. The success of the EVA led to a nominal entry upon mission completion.
STS-91 | 6/2/1998 | Crew: 6
Main engine pressure chamber sensor failed. If it occurred later, logic error may have triggered at RTLS.
After the launch of STS-91 on June 2, 1998, a channel A main engine pressure chamber sensor froze. The sensor was disqualified by engine control software when the sensor exceeded allowable limits during the max Q engine throttle down sequence however the sensor remained qualified for engine redline monitoring since the reading was still within reasonableness limits / operational range for the sensor. The channel B sensor and the main engine performed nominally throughout ascent. If an engine problem had occurred, the channel B sensor would have displayed the correct information and indicated the proper corrective action. However, the channel B sensor would have been ignored due to the frozen channel A sensor.
The post-flight inspection revealed that contamination from a seal leak check caused the sensor to freeze. Marshall Space Flight Center (MSFC) project engineering suggested adding a V-seal leak check prior to the acceptance test, to ensure the newly added V-seal was installed properly. However, MSFC project engineering was unaware that performing the leak check would require the manufacturing personnel to plug the chamber pressure port with Viton. After the flight, the sensor was still plugged with a piece of Viton.
If the main engine pressure chamber sensor froze later in the flight, logic errors may have triggered a premature engine shutdown and a return to the launch site abort. Post-flight software changes were implemented to prevent this from occurring on subsequent flights. Additionally, a one-time flow check verified the reliability of the Lee-Jet for all engines in the fleet. A corrective action requires post-Lee-Jet installation flow check and borescope inspection for future engine builds. Lastly, the V-seal leak check was eliminated in future engine builds.
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.
Other significant ascent debris events have occurred on:
Late Release Orbiter Tyvek Covers
SRB Seal Events (1981-96)
On-pad Abort Events (1984-93)
Related or Recurring event
TPS Entry Events (1981-2003)
LANDING & POSTLANDING