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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
On-pad Abort Events (1984-93)
Gemini 6 12/12/1965
Apollo 1 (AS-204) 1/27/1967
Soyuz T-10-1(T-10a) 9/26/1983
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.
Soyuz T-10-1 (T-10a) | 9/26/1983 | Crew: 2 | Loss of Vehicle/Mission | Related or Recurring event
Pad booster fire/explosion. Capsule Escape System used.
Shortly before liftoff on September 26, 1983 fuel spilled around the base of the Soyuz launch vehicle and ignited the vehicle. Launch control activated the escape system, but the control cables were burnt.
Twenty seconds later ground control activated the escape system by radio command. By this time the booster was engulfed in flames.
Explosive bolts fired to separate the descent module from the service module. Explosive bolts also fired to separate of the upper shroud from the lower shroud. The escape system motor pulled the orbit module and descent module, still encased within the upper shroud, away from the booster at 14 to 17g of acceleration.
Seconds after the escape system activated, the booster exploded, destroying the launch complex.
The descent module separated from the orbital module and dropped free from the shroud. The descent module heat shield was discarded to expose the solid-fueled landing rockets. A fast-opening emergency parachute was deployed and landing occurred about four km from the launch pad.
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.
Related or Recurring event
LANDING & POSTLANDING