Just after liftoff at .678 seconds into the flight, photographic data show a strong puff of gray smoke was spurting from the vicinity of the aft field joint on the right Solid Rocket Booster. The two pad 39B cameras that would have recorded the precise location of the puff were inoperative. Computer graphic analysis of film from other cameras indicated the initial smoke came from the 270 to 310-degree sector of the circumference of the aft field joint of the right Solid Rocket Booster. This area of the solid booster faces the External Tank. The vaporized material streaming from the joint indicated there was not complete sealing action within the joint. Eight more distinctive puffs of increasingly blacker smoke were recorded between .836 and 2.500 seconds. The smoke appeared to puff upwards from the joint. While each smoke puff was being left behind by the upward flight of the Shuttle, the next fresh puff could be seen near the level of the joint. The multiple smoke puffs in this sequence occurred at about four times per second, approximating the frequency of the structural load dynamics and resultant joint flexing. Computer graphics applied to NASA photos from a variety of cameras in this sequence again placed the smoke puffs' origin in the 270- to 310-degree sector of the original smoke spurt. As the Shuttle increased its upward velocity, it flew past the emerging and expanding smoke puffs. The last smoke was seen above the field joint at 2.733 seconds. The black color and dense composition of the smoke puffs suggest that the grease, joint insulation and rubber O-rings in the joint seal were being burned and eroded by the hot propellant gases. At approximately 37 seconds, Challenger encountered the first of several high-altitude wind shear conditions, which lasted until about 64 seconds. The wind shear created forces on the vehicle with relatively large fluctuations. These were immediately sensed and countered by the guidance, navigation and control system. The steering system (thrust vector control) of the Solid Rocket Booster responded to all commands and wind shear effects. The wind shear caused the steering system to be more active than on any previous flight. Both the Shuttle main engines and the solid rockets operated at reduced thrust approaching and passing through the area of maximum dynamic pressure of 720 pounds per square foot. Main engines had been throttled up to 104 percent thrust and the Solid Rocket Boosters were increasing their thrust when the first flickering flame appeared on the right Solid Rocket Booster in the area of the aft field joint. This first very small flame was detected on image enhanced film at 58.788 seconds into the flight. It appeared to originate at about 305 degrees around the booster circumference at or near the aft field joint. One film frame later from the same camera, the flame was visible without image enhancement. It grew into a continuous, well-defined plume at 59.262 seconds. At about the same time (60 seconds), telemetry showed a pressure differential between the chamber pressures in the right and left boosters. The right booster chamber pressure was lower, confirming the growing leak in the area of the field joint. As the flame plume increased in size, it was deflected rearward by the aerodynamic slipstream and circumferentially by the protruding structure of the upper ring attaching the booster to the External Tank. These deflections directed the flame plume onto the surface of the External Tank. This sequence of flame spreading is confirmed by analysis of the recovered wreckage. The growing flame also impinged on the strut attaching the Solid Rocket Booster to the External Tank. The first visual indication that swirling flame from the right Solid Rocket Booster breached the External Tank was at 64.660 seconds when there was an abrupt change in the shape and color of the plume. This indicated that it was mixing with leaking hydrogen from the External Tank. Telemetered changes in the hydrogen tank pressurization confirmed the leak. Within 45 milliseconds of the breach of the External Tank, a bright sustained glow developed on the black-tiled underside of the Challenger between it and the External Tank. Beginning at about 72 seconds, a series of events occurred extremely rapidly that terminated the flight. Telemetered data indicate a wide variety of flight system actions that support the visual evidence of the photos as the Shuttle struggled futilely against the forces that were destroying it. At about 72.20 seconds the lower strut linking the Solid Rocket Booster and the External Tank was severed or pulled away from the weakened hydrogen tank permitting the right Solid Rocket Booster to rotate around the upper attachment strut. This rotation is indicated by divergent yaw and pitch rates between the left and right Solid Rocket Boosters. At 73.124 seconds,. a circumferential white vapor pattern was observed blooming from the side of the External Tank bottom dome. This was the beginning of the structural failure of hydrogen tank that culminated in the entire aft dome dropping away. This released massive amounts of liquid hydrogen from the tank and created a sudden forward thrust of about 2.8 million pounds, pushing the hydrogen tank upward into the intertank structure. At about the same time, the rotating right Solid Rocket Booster impacted the intertank structure and the lower part of the liquid oxygen tank. These structures failed at 73.137 seconds as evidenced by the white vapors appearing in the intertank region. Within milliseconds there was massive, almost explosive, burning of the hydrogen streaming from the failed tank bottom and liquid oxygen breach in the area of the intertank. At this point in its trajectory, while traveling at a Mach number of 1.92 at an altitude of 46,000 feet, the Challenger was totally enveloped in the explosive burn. The Challenger's reaction control system ruptured and a hypergolic burn of its propellants occurred as it exited the oxygen-hydrogen flames. The reddish brown colors of the hypergolic fuel burn are visible on the edge of the main fireball. The Orbiter, under severe aerodynamic loads, broke into several large sections which emerged from the fireball. Separate sections that can be identified on film include the main engine/tail section with the engines still burning, one wing of the Orbiter, and the forward fuselage trailing a mass of umbilical lines pulled loose from the payload bay. STS 51-L SEQUENCE OF MAJOR EVENTS Mission Time Elapsed GMT (hr:min:sec) Event Time (secs.) Source 16:37:53.444 ME-3 Ignition Command -6.566 GPC 37:53.564 ME-2 Ignition Command -6.446 GPC 37:53.684 ME-1 Ignition Command -6.326 GPC 38:00.010 SRM Ignition Command (T=0) 0.000 GPC 38:00.018 Holddown Post 2 PIC firing 0.008 E8 Camera 38:00.260 First Continuous Vertical Motion 0.250 E9 Camera 38:00.688 Confirmed smoke above field joint on RH SRM 0.678 E60 Camera 38:00.846 Eight puffs of smoke (from 0.836 thru 2.500 sec MET) 0.836 E63 Camera 38:02.743 Last positive evidence of smoke above right aft SRB/ET attach ring 2.733 CZR-1 Camera 38:03.385 Last positive visual indication of smoke 3.375 E60 Camera 38:04.349 SSME 104% Command 4.339 E41M2076D 38:05.684 RH SRM pressure 11.8 psi above nominal 5.674 B47P2302C 38:07.734 Roll maneuver initiated 7.724 V90R5301C 38:19.869 SSME 94% Command 19.859 E41M2076D 38:21.134 Roll maneuver completed 21.124 VP0R5301C 38:35.389 SSME 65% Command 35.379 E41M2076D 38:37.000 Roll and Yaw Attitude Response to Wind (36.990 to 62.990 sec) 36.990 V95H352nC 38:51.870 SSME 104% Command 51.860 E41M2076D 38:58.798 First evidence of flame on RH SRM 58.788 E207 Camera 38:59.010 Reconstructed Max Q (720 psf) 59.000 BET 38:59.272 Continuous well defined plume on RH SRM 59.262 E207 Camera 38:59.763 Flame from RH SRM in +Z direction (seen from south side of vehicle) 59.753 E204 Camera 39:00.014 SRM pressure divergence (RH vs. LH) 60.004 B47P2302 39:00.248 First evidence of plume deflection, intermittent 60.238 E207 Camera 39:00.258 First evidence of SRB plume attaching to ET ring frame 60.248 E203 Camera 39:00.998 First evidence of plume deflection, continuous 60.988 E207 Camera 39:01.734 Peak roll rate response to wind 61.724 V90R5301C 39:02.094 Peak TVC response to wind 62.084 B58H1150C 39:02.414 Peak yaw response to wind 62.404 V90R5341C 39:02.494 RH outboard elevon actuator hinge moment spike 62.484 V58P0966C 39:03.934 RH outboard elevon actuator delta pressure change 63.924 V58P0966C 39:03.974 Start of planned pitch rate maneuver 63.964 V90R5321C 39:04.670 Change in anomalous plume shape (LH2 tank leak near 2058 ring frame) 64.660 E204 Camera 39:04.715 Bright sustained glow on sides of ET 64.705 E204 Camera 39:04.947 Start SSME gimbal angle large pitch variations 64.937 V58H1100A 39:05.174 Beginning of transient motion due to changes in aero forces due to plume 65.164 V90R5321C 39:06.774 Start ET LH2 ullage pressure deviations 66.764 T41P1700C 39:12.214 Start divergent yaw rates (RH vs. LH SRB) 72.204 V90R2528C 39:12.294 Start divergent pitch rates (RH vs. LH SRB) 72.284 V90R2525C 39:12.488 SRB major high-rate actuator command 72.478 V79H2111A 39:12.507 SSME roll gimball rates 5 deg/sec 72.497 V58H1100A 39:12.535 Vehicle max +Y lateral acceleration (+.227 g) 72.525 V98A1581C 39:12.574 SRB major high-rate actuator motion 72.564 B58H1151C 39:12.574 Start of H2 tank pressure decrease with 2 flow control valves open 72.564 T41P1700C 39:12.634 Last state vector downlinked 72.624 Data reduction 39:12.974 Start of sharp MPS LOX inlet pressure drop 72.964 V41P1330C 39:13.020 Last full computer frame of TDRS data 73.010 Data reduction 39:13.054 Start of sharp MPS LH2 inlet pressure drop 73.044 V41P1100C 39:13.055 Vehicle max -Y lateral accelerarion (-.254 g) 73.045 V98A1581C 39:13.134 Circumferential white pattern on ET aft dome (LH2 tank failure) 73.124 E204 Camera 39:13.134 RH SRM pressure 19 psi lower than LH SRM 73.124 B47P2302C 39:13.147 First hint of vapor at intertank E207 Camera 39:13.153 All engine systems start responding to loss of fuel and LOX inlet pressure 73.143 SSME team 39:13.172 Sudden cloud along ET between intertank and aft dome 73.162 E207 Camera 39:13.201 Flash between Orbiter & LH2 tank 73.191 E204 Camera 39:13.221 SSME telemetry data interference from 73.211 to 73.303 73.211 39:13.223 Flash near SRB fwd attach and brightening of flash between Orbiter and ET 73.213 E204 Camera 39:13.292 First indication intense white flash at SRB fwd attach point 73.282 E204 Camera 39:13.337 Greatly increased intensity of white flash 73.327 E204 Camera 39:13.387 Start RCS jet chamber pressure fluctuations 73.377 V42P1552A 39:13.393 All engines approaching HPFT discharge temp redline limits 73.383 E41Tn010D 39:13.492 ME-2 HPFT disch. temp Chan. A vote for shutdown; 2 strikes on Chan. B 73.482 MEC data 39:13.492 ME-2 controller last time word update 73.482 MEC data 39:13.513 ME-3 in shutdown due to HPFT discharge temperature redline exceedance 73.503 MEC data 39:13.513 ME-3 controller last time word update 73.503 MEC data 39:13.533 ME-1 in shutdown due to HPFT discharge temperature redline exceedance 73.523 Calculation 39:13.553 ME-1 last telemetered data point 73.543 Calculation 39:13.628 Last validated Orbiter telemetry measurement 73.618 V46P0120A 39:13.641 End of last reconstructured data frame with valid synchronization and frame count 73.631 Data reduction 39:14.140 Last radio frequency signal from Orbiter 74.130 Data reduction 39:14.597 Bright flash in vicinity of Orbiter nose 74.587 E204 Camera 39:16.447 RH SRB nose cap sep/chute deployment 76.437 E207 Camera 39:50.260 RH SRB RSS destruct 110.250 E202 Camera 39:50.262 LH SRB RSS destruct 110.252 E230 Camera ACT POS -- Actuator Position APU -- Auxilixary Power Unit BET -- Best Estimated Trajectory CH -- Channel DISC -- Discharge ET -- External Tank GG -- Gas Generator GPC -- General Purpose Computer GMT -- Greenwich Mean Time HPFT -- High Pressure Fuel Turbopump LH -- Lefthand LH2 -- Liquid Hydrogen LO2 -- Liquid Oxygen (same as LOX) MAX Q -- Maximum Dynamic Pressure ME -- Main Engine (same as SSME) MEC -- Main Engine Controller MET -- Mission Elapsed Time MPS -- Main Propulsion System PC -- Chamber Pressure PIC -- Pyrotechnics Initiator Controller psf -- Pounds per square foot RCS -- Reaction Control System RGA -- Rate Gyro Assembly RH -- Righthand RSS -- Range Safety System SRM -- Solid Rocket Motor SSME -- Space Shuttle Main Engine TEMP -- Temperature TVC -- Thrust Vector Control NOTE: The Shuttle coordinate system used is relative to the Orbiter, as follows: +X direction = forward (tail to nose) -X direction = rearward (nose to tail) +Y direction = right (toward the right wing tip) -Y direction = left (toward the left wing tip) +Z direction = down -Z direction = up
THE CAUSE OF THE ACCIDENT (Source: The Presidential Commission on the Space Shuttle Challenger Accident Report, June 6, 1986) THE CAUSE OF THE ACCIDENT The consensus of the Commission and participating investigative agencies is that the loss of the Space Shuttle Challenger was caused by a failure in the joint between the two lower segments of the right Solid Rocket Motor. The specific failure was the destruction of the seals that are intended to prevent hot gases from leaking through the joint during the propellant burn of the rocket motor. The evidence assembled by the Commission indicates that no other element of the Space Shuttle system contributed to this failure. In arriving at this conclusion, the Commission reviewed in detail all available data, reports and records; directed and supervised numerous tests, analyses, and experiments by NASA, civilian contractors and various government agencies; and then developed specific scenarios and the range of most probable causative factors. FINDINGS 1. A combustion gas leak through the right Solid Rocket Motor aft field joint initiated at or shortly after ignition eventually weakened and/or penetrated the External Tank initiating vehicle structural breakup and loss of the Space Shuttle Challenger during STS Mission 51-L. 2. The evidence shows that no other STS 51-L Shuttle element or the payload contributed to the causes of the right Solid Rocket Motor aft field joint combustion gas leak. Sabotage was not a factor. 3. Evidence examined in the review of Space Shuttle material, manufacturing, assembly, quality control, and processing on non-conformance reports found no flight hardware shipped to the launch site that fell outside the limits of Shuttle design specifications. 4. Launch site activities, including assembly and preparation, from receipt of the flight hardware to launch were generally in accord with established procedures and were not considered a factor in the accident. 5. Launch site records show that the right Solid Rocket Motor segments were assembled using approved procedures. However, significant out-of-round conditions existed between the two segments joined at the right Solid Rocket Motor aft field joint (the joint that failed). a. While the assembly conditions had the potential of generating debris or damage that could cause O-ring seal failure, these were not considered factors in this accident. b. The diameters of the two Solid Rocket Motor segments had grown as a result of prior use. c. The growth resulted in a condition at time of launch wherein the maximum gap between the tang and clevis in the region of the joint's O-rings was no more than .008 inches and the average gap would have been .004 inches. d. With a tang-to-clevis gap of .004 inches, the O-ring in the joint would be compressed to the extent that it pressed against all three walls of the O-ring retaining channel. e. The lack of roundness of the segments was such that the smallest tang-to-clevis clearance occurred at the initiation of the assembly operation at positions of 120 degrees and 300 degrees around the circumference of the aft field joint. It is uncertain if this tight condition and the resultant greater compression of the O-rings at these points persisted to the time of launch. 6. The ambient temperature at time of launch was 36 degrees Fahrenheit, or 15 degrees lower than the next coldest previous launch. a. The temperature at the 300 degree position on the right aft field joint circumference was estimated to be 28 degrees plus or minus 5 degrees Fahrenheit. This was the coldest point on the joint. b. Temperature on the opposite side of the right Solid Rocket Booster facing the sun was estimated to be about 50 degrees Fahrenheit. 7. Other joints on the left and right Solid Rocket Boosters experienced similar combinations of tang-to-clevis gap clearance and temperature. It is not known whether these joints experienced distress during the flight of 51-L. 8. Experimental evidence indicates that due to several effects associated with the Solid Rocket Booster's ignition and combustion pressures and associated vehicle motions, the gap between the tang and the clevis will open as much as .017 and .029 inches at the secondary and primary O-rings, respectively. a. This opening begins upon ignition, reaches its maximum rate of opening at about 200-300 milliseconds, and is essentially complete at 600 milliseconds when the Solid Rocket Booster reaches its operating pressure. b. The External Tank and right Solid Rocket Booster are connected by several struts, including one at 310 degrees near the aft field joint that failed. This strut's effect on the joint dynamics is to enhance the opening of the gap between the tang and clevis by about 10-20 percent in the region of 300-320 degrees. 9. O-ring resiliency is directly related to its temperature. a. A warm O-ring that has been compressed will return to its original shape much quicker than will a cold O-ring when compression is relieved. Thus, a warm O-ring will follow the opening of the tang-to-clevis gap. A cold O-ring may not. b. A compressed O-ring at 75 degrees Fahrenheit is five times more responsive in returning to its uncompressed shape than a cold O-ring at 30 degrees Fahrenheit. c. As a result it is probable that the O-rings in the right solid booster aft field joint were not following the opening of the gap between the tang and cleavis at time of ignition. 10. Experiments indicate that the primary mechanism that actuates O-ring sealing is the application of gas pressure to the upstream (high-pressure) side of the O-ring as it sits in its groove or channel. a. For this pressure actuation to work most effectively, a space between the O-ring and its upstream channel wall should exist during pressurization. b. A tang-to-clevis gap of .004 inches, as probably existed in the failed joint, would have initially compressed the O-ring to the degree that no clearance existed between the O-ring and its upstream channel wall and the other two surfaces of the channel. c. At the cold launch temperature experienced, the O-ring would be very slow in returning to its normal rounded shape. It would not follow the opening of the tang-to-clevis gap. It would remain in its compressed position in the O-ring channel and not provide a space between itself and the upstream channel wall. Thus, it is probable the O-ring would not be pressure actuated to seal the gap in time to preclude joint failure due to blow-by and erosion from hot combustion gases. 11. The sealing characteristics of the Solid Rocket Booster O-rings are enhanced by timely application of motor pressure. a. Ideally, motor pressure should be applied to actuate the O-ring and seal the joint prior to significant opening of the tang-to-clevis gap (100 to 200 milliseconds after motor ignition). b. Experimental evidence indicates that temperature, humidity and other variables in the putty compound used to seal the joint can delay pressure application to the joint by 500 milliseconds or more. c. This delay in pressure could be a factor in initial joint failure. 12. Of 21 launches with ambient temperatures of 61 degrees Fahrenheit or greater, only four showed signs of O-ring thermal distress; i.e., erosion or blow-by and soot. Each of the launches below 61 degrees Fahrenheit resulted in one or more O-rings showing signs of thermal distress. a. Of these improper joint sealing actions, one-half occurred in the aft field joints, 20 percent in the center field joints, and 30 percent in the upper field joints. The division between left and right Solid Rocket Boosters was roughly equal. b. Each instance of thermal O-ring distress was accompanied by a leak path in the insulating putty. The leak path connects the rocket's combustion chamber with the O-ring region of the tang and clevis. Joints that actuated without incident may also have had these leak paths. 13. There is a possibility that there was water in the clevis of the STS 51-L joints since water was found in the STS-9 joints during a destack operation after exposure to less rainfall than STS 51-L. At time of launch, it was cold enough that water present in the joint would freeze. Tests show that ice in the joint can inhibit proper secondary seal performance. 14. A series of puffs of smoke were observed emanating from the 51-L aft field joint area of the right Solid Rocket Booster between 0.678 and 2.500 seconds after ignition of the Shuttle Solid Rocket Motors. a. The puffs appeared at a frequency of about three puffs per second. This roughly matches the natural structural frequency of the solids at lift off and is reflected in slight cyclic changes of the tang-to-clevis gap opening. b. The puffs were seen to be moving upward along the surface of the booster above the aft field joint. c. The smoke was estimated to originate at a circumferential position of between 270 degrees and 315 degrees on the booster aft field joint, emerging from the top of the joint. 15. This smoke from the aft field joint at Shuttle lift off was the first sign of the failure of the Solid Rocket Booster O-ring seals on STS 51-L. 16. The leak was again clearly evident as a flame at approximately 58 seconds into the flight. It is possible that the leak was continuous but unobservable or non-existent in portions of the intervening period. It is possible in either case that thrust vectoring and normal vehicle response to wind shear as well as planned maneuvers reinitiated or magnified the leakage from a degraded seal in the period preceding the observed flames. The estimated position of the flame, centered at a point 307 degrees around the circumference of the aft field joint, was confirmed by the recovery of two fragments of the right Solid Rocket Booster. a. A small leak could have been present that may have grown to breach the joint in flame at a time on the order of 58 to 60 seconds after lift off. b. Alternatively, the O-ring gap could have been resealed by deposition of a fragile buildup of aluminum oxide and other combustion debris. This resealed section of the joint could have been disturbed by thrust vectoring, Space Shuttle motion and flight loads inducted by changing winds aloft. c. The winds aloft caused control actions in the time interval of 32 seconds to 62 seconds into the flight that were typical of the largest values experienced on previous missions. CONCLUSION In view of the findings, the Commission concluded that the cause of the Challenger accident was the failure of the pressure seal in the aft field joint of the right Solid Rocket Booster. The failure was due to a faulty design unacceptably sensitive to a number of factors. These factors were the effects of temperature, physical dimensions, the character of materials, the effects of reusability, processing and the reaction of the joint to dynamic loading.
