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.
(NASA PICTURE -EDITED BY TSE)
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.
(NASA PICTURE -EDITED BY TSE)
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.
(NASA PICTURE -EDITED BY TSE)
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.
