SUBMITTED BY:	- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -
	R. J. SCHWINGHAMER, CHAIRMAN, SOLID ROCKET MOTOR WORKING GROUP
APPROVED BY:	- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -
	JOHN W. THOMAS, DEPUTY, ACCIDENT ANALYSIS TEAM
	- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -
	J. R. THOMPSON, JR., LEAD, ACCIDENT ANALYSIS TEAM

 

 

TABLE OF CONTENTS

 

I. Organization/Responsibilities.

II. Method of Investigation.

III. Definition of Terms and Acronyms.

IV. Technical Description.

A. Hot Gas Leak Failure Scenarios.

1. Internal Heating/Inhibitor/Insulation Anomalies.

a. Scenario Description.

b. Investigation.

c. Findings.

2. LH₂ Leak Resulting in SRM Case Membrane or Joint Failure.

3. Assembly Conditions/Anomalies.

a. Scenario Description.

b. Investigation.

c. Findings.

4. Scenario 4: Primary O-Ring Blow-by or Low Temperature Effects.

a. Scenario 4a: Secondary O-Ring Defect.

b. Scenario 4b: Ice in Joint.

c. Scenario 4c: Leak Check Port Leaks.

d. Scenario 4d: O-Ring Actuation Time Delayed by Low Temperature.

5. Loads Exceedance.

6. Scenario 6: Putty Holds Pressure Off Primary O-Ring.

7. Case Rupture.

B. Linear Shaped Charge Failure (Motor).

C. Linear Shaped Charge Failure (Nozzle).

D. Propellant Detonation (External Shock Wave).

E. Safe and Arm Device Failure.

F. Igniter Failure.

G. Flex Bearing Failure.

H. System Tunnel Failure.

I. Propellant/Liner Debound Failure.

J. Insulation Failure.

K. Nozzle Failure.

L. Pressure Transducer/Port Failure.

M. Propellant Grain Failure.

N. Combustion Instability.

O. External Thermal Protection System Failure.

P. Electrostatic Discharge.

V. Summary.

VI. Findings.

 

APPENDIX A. Recovery Support Team Report

APPENDIX B. SRM Team Analyses

APPENDIX C. SRM Team Tests

 

 

[L51] LIST OF FIGURES

 

1. SRM Hot Gas Leak Failure Scenarios.

2. Solid Rocket Motor Elements.

3. Aft Segment/Aft Center Segment Field Joint Configuration.

4. Assumed Insulation Flaw.

5. Primary O-Ring Temperature for Assumed Insulation Flaw of Figure 4.

6. Secondary O-Ring Temperatures for Assumed Insulation Flaw of Figure 4.

7. Assumed Inhibitor Flaw.

8. Scenario 3 Logic Chart.

9. Handling.

10. Preassembly Ovality Assessment.

11. Rounding Tool.

12. Case Dimensional Reassessment at ROHR.

13. 51-L RH Aft Field Joint.

14. Scenario 4a Logic Chart.

15. Ice in Joint Scenario Evaluation.

16. Viewing Angles from Cameras E-60 and E-63.

17. Predicted Location of Smoke Plume from a Leaking Leak Check Port.

18. Location of Observed Smoke Plume at 0.854 Seconds.

19. Location of Observed Smoke Plume at 0.954 Seconds.

20. Location of Observed Smoke Plume at 0.731 Seconds.

21. Location of Observed Smoke Plume at 0.779 Seconds.

22. Location of Observed Smoke Plume at 0.829 Seconds.

23. Location of Observed Smoke Plume at 1.055 Seconds.

24. Pressurized Joint Deflection (Exaggerated).

25. Scenario 4-Primary O-Ring Blow-by or Low Temperature.

26. Scenario 4.1 Disposition.

27. Recovery vs. Time.

28. Comparison of MTI and MSFC Dynamic Test Reports.

29. Test Results Summary.

30. STS 51-L Right Hand SRM Aft Field Joint.

 

 

LIST OF TABLES

 

1. MTI O-Ring Inspection Plan-IU50228-24.

2. Parker Formulation/Process Combinations.

 

 

 

[L52] I. Organization/Responsibilities

The organization and responsibilities of the SRM Working Group were in accordance with the Shuttle Contingency Plan. The following is a listing of the SRM Group along with the area of responsibility for each member. A listing of other key participants, both NASA and contractor, is also given.

Key Morton Thiokol, Inc., personnel participating throughout the investigatory period in the Huntsville Operations Support Center are as follows:

CS-SAFETY OFFICE

ED-SYSTEMS DYNAMICS LABORATORY

EE-ASSOCIATE DIRECTOR OF ENGINEERING OFFICE

EG-RELIABILITY AND QUALITY ASSURANCE OFFICE

EH-MATERIALS AND PROCESSES LABORATORY

EP-STRUCTURES AND PROPULSION

ET-TEST LABORATORY

SA-SHUTTLE PROJECTS OFFICE

II. Method of Investigation

The SRM Working Group was activated in accordance with the Space Shuttle Contingency Plan, MSFC-SSCP-5-77, Rev. Q, July 8, 1985, immediately following the Challenger incident. Following the establishment of the NASA Task Force, the SRM Working Group activities were directed in support of the Accident Analysis Team of the Data and Design Analysis Task Force.

The investigation activities of the group were conducted in general accordance with MMI 1711.2d, Mishap Reporting and Investigation, April 16, 1984.

Build Records and Discrepancy Report Reviews

Build records and discrepancy report reviews were initiated early in the investigation. The object was to determine whether errors or omissions in the planning and discrepancy reporting documents existed and if so, to determine whether they could have contributed to the anomaly. The build records and discrepancy documents were obtained from MTI, Wasatch, and KSC and were reviewed by Senior personnel in Materials and Processes and Reliability and Quality Assurance. Discrepancies requiring special analysis to disposition were further reviewed by specialists in stress, O-ring application, flex bearing engineering, and metallurgy.

Build records and discrepancy documents assessed by the SRM Accident Investigation Document Review Team included: all discrepancy documents written against parts and materials used in the 51-L right booster, all shop travelers, all KSC stacking documents, and all KSC problem reports.

All discrepancy reports, process departures, and parts rejection reports reported against the 51-L right motor components by Morton Thiokol were reviewed. The majority of discrepancies can be categorized as minor in nature and many are associated with reuse of the hardware. The disposition given all the discrepancies, except one, were upheld by the Review Team.

The single exception, DR 110487-04, was considered improperly dispositioned in that pitting on the nozzle throat housing sealing surface was accepted for "Use As Is" without rework. Sealing surfaces typically are considered critical to part functioning and should not be compromised. But after careful analysis it was concluded that this discrepancy did not have any influence on the 51-L incident. Photographic evidence shows the nozzle throat housing was not involved.

Two Problem Reports (PRs) were issued at KSC against the 51-L, right motor, aft center segment to aft segment field joint. These were minor in nature and considered properly dispositioned.

A total of 287 Common Planning Indices (CPIs) used to construct the 51-L right motor were reviewed by the Group. The CPIs were reviewed for clarity of instructions and for completeness of buyoff; i.e., all steps and logs were properly stamped and approved by MTI manufacturing, engineering, and Quality and Air Force Quality.

The CPIs were determined to be well written with extensive detail including numerous inspection and verification points.

All CPIs were properly dispositioned. All manufacturing quality and Air Force mandatory verifications were properly accomplished. The CPIs were mature planning documents; some had been used more than 100 times to fabricate major components for qualification, test and flight. The majority had been used between 14 and 60 times.

It was determined that there had been no Class 1 changes to the planning system or individual documents between the previous use and 51-L. It was further determined that there was no limited use hardware employed in any field joint for the right SBM. Some limited use restrictions were in place but they were confined to factory joints and were well documented.

Finally, the extensive review of SRM build records did not reveal any irregularities that could have contributed to failure of the 51-L right Solid Rocket Motor.

[L54] Possible SRM Failure Modes

Failure modes were identified and each assessed against the failure investigation data base. This data base consisted of Mission 51-L photographic, telemetry, prelaunch, design, and manufacturing data. This database was progressively expanded during the course of the investigation through a review of previous flight and test data and with tests and analyses performed in direct support of the investigation.

During the investigation the potential failure modes listed below were identified and evaluated:

A. SRM Hot Gas Leak

1. Internal Heating/Inhibitor/Insulation Anomalies

2. LH₂ leak impingement on SRM

3. Assembly Conditions/Anomalies

4. Primary O-ring blow-by or low temperature effects coupled with either:

• secondary O-ring defect or
• ice in joint or
• leak check port leaks or
• O-ring actuation time delayed by lower temperature

5. Load exceedance effects

6. Putty delaying pressure to primary O-ring

7. Case rupture

Other Possible Modes

B. Linear shaped charge (LSC) failure (motor)

C. Linear shaped charge failure (nozzle)

D. Propellant detonation from external shock wave

E. Safe and arm (S&A) failure

F. Igniter failure

G. Flex bearing failure

H. Systems tunnel failure

1. Propellant/liner debond failure

J. Insulation failure

K. Nozzle failure

L. Pressure transducer/port failure

M. Propellant grain failure

N. Combustion instability

0. External thermal protection system (TPS) failure

P. Electrostatic discharge (ESD)

Much of the initial photographic data review resulted in focusing on investigation of failure modes that could result in a hot gas leak from the right SRM. The primary emphasis of the investigation was placed on failure modes A.1 thru A.7. Subsequent hardware recovery has confirmed this to be the proper focus for the failure investigation. Failure modes B thru P were discounted early based on empirical data and analyses. Supporting rationale for conclusions of all the failure modes will be presented in Section IV of this report.

A failure scenario diagram (Figure 1) was devised to assist in a systematic evaluation of each hot gas leak failure scenario mechanism. Each scenario will be treated in subsequent sections of this report, except Scenarios A.2 and A.5. Scenario A.2 is covered in the External Tank Final Report and Scenario A.5 is covered in the Systems Final Report.

III. Definition of Terms and Acronyms

APU	AUXILIARY POWER UNIT
BET	BEST ESTIMATED TRAJECTORY
BSM	BOOSTER SEPARATION MOTOR
BTU	BRITISH THERMAL UNIT
CDF	CONFINED DETONATING FUSE
CDR	CRITICAL DESIGN REVIEW
CEI	CONTRACT END ITEM
C/O	CHECKOUT
CPI	COMMON PLANNING INDICES
CTPB	CARBOXYL TERMINATED POLYBUTADIENE
DD	DEEP DRONE (UNMANNED SUBMARINE)
DFI	DEVELOPMENT FLIGHT INSTRUMENTATION
DM	DEVELOPMENT MOTOR
DR	DISCREPANCY REPORT
E&I	ELECTRICAL AND INSTRUMENTATION
EPDM	ETHYLENE PROPYLENE TER POLYMER
ESD	ELECTROSTATIC DISCHARGE
ET	EXTERNAL TANK
ETA	EXTERNAL TANK ATTACH
ETR	EASTERN TEST RANGE
EWAT	END WEB ACTION TIME
EWT	END WEB TIME
FLT	FLIGHT
FSM	FUEL SUPPLY MODULE
FSS	FIXED SERVICE STRUCTURE
FTU	ULTIMATE TENSILE STRESS
FTY	YIELD TENSILE STRESS
FWD	FORWARD
GSE	GROUND SUPPORT EQUIPMENT
HMX	CYCLOTETRAMETHYLENE TETRA
HNS	NITRAMINE
HP	HEXA NITRO STILBENE
HPM	HIGH PRESSURE
HPU	HIGH PERFORMANCE MOTOR
IMU	HYDRAULIC POWER UNIT
Lus	INERTIAL MEASUREMENT UNIT INERTIAL UPPER STAGE
JSL	JOHNSON SEA LINK (FOUR-MAN SUBMARINE)
KSI	THOUSANDS OF POUNDS PER INCH
LANL	LOS ALAMOS NATIONAL LABORATORY
LBM	POUNDS MASS
LG	LARGE
LH	LEFT HAND
LSC	LINEAR SHAPED CHARGE
LT	LIGHT
LVDT	LINEAR VARIABLE DIFFERENTIAL TRANSFORMER
MAX Q	MAXIMUM DYNAMIC PRESSURE
MEOP	MAXIMUM EXPECTED OPERATING PRESSURE
MET	MISSION ELAPSED TIME
MiLS	THOUSANDTHS OF AN INCH
MIN	MINIMUM

MOF	MANNED ORBITAL FLIGHT
MLP	MOBILE LAUNCH PLATFORM
MMI	MARSHALL MANAGEMENT INSTRUCTION
M&P	MATERIALS AND PROCESSING
MTI	MORTON THIOKOL, INCORPORATED
NBR	NITRILE BUTADIENE RUBBER
NDI	NONDESTRUCTIVE INSPECTION
NDT	NONDESTRUCTIVE TESTING
NPIC	NATIONAL PHOTOGRAPHIC INTERPRETATION CENTER
NRA	NUCLEAR RESEARCH SUBMARINE
NSI	NASA STANDARD INITIATOR
NSTL	NATIONAL SPACE TECHNOLOGY LABORATORY
OF	OPERATIONAL FLIGHT
OFI	OPERATIONAL FLIGHT INSTRUMENTATION'
OPT	OPERATIONAL FLIGHT TRANSDUCERS
ORB	ORBITER
OV	ORBITER VEHICLE

[L56] PBAN	POLYBUTADIENE ACRYLONITRILE (PROPELLANT)
PETN	PENTA ERYTHRITE TETRA NITRATE
PDR	PRELIMINARY DESIGN REVIEW
PIC	PYROTECHNIC INITIATOR CONTROLLER
PR	PROBLEM REPORT
PRI	PRIMARY
PSI	POUNDS PER SQUARE INCH
PSIA	POUNDS PER SQUARE INCH -ABSOLUTE
PSID	POUNDS PER SQUARE INCH-DIFFERENTIAL
PSIG	POUNDS PER SQUARE INCH GAGE
Q	DYNAMIC PRESSURE (POUNDS PER SQUARE FOOT)
Q MAX	MAXIMUM DYNAMIC PRESSURE DURING ASCENT
RDX	CYCLO TRIMETHYLENE TRINITRAMINE
RH	RIGHT HAND, or RELATIVE HUMIDITY
RPC	REMOTE POWER CONTROLLER
R&QA	RELIABILITY AND QUALITY ASSURANCE
RSS	RANGE SAFETY SYSTEM or ROOT SUM SQUARED
S&A	SAFE AND ARM DEVICE
SBPE	SPACE BOOSTER PROJECT ENGINEERING
SEG	SEGMENT
SRB	SOLID ROCKET BOOSTER
SRM	SOLID ROCKET MOTOR
SRMPO	SOLID ROCKET MOTOR PROJECT OFFICE
SSME	SPACE SHUTTLE MAIN ENGINE
STA	STATION
STS	SPACE TRANSPORTATION SYSTEM
TBD	TO BE DETERMINED
T/C	THERMOCOUPLE
TDRS	TRACKING AND DATA RELAY SATELLITE
TPS	THERMAL PROTECTION SYSTEM
TVC	THRUST VECTOR CONTROL
USBI	UNITED SPACE BOOSTER, INCORPORATED
VAB	VERTICAL ASSEMBLY BUILDING
VAFB	VANDENBERG AIR FORCE BASE
WTR	WESTERN TEST RANGE

IV. Technical Description

SRM Description

The total SRB element of the Space Shuttle is composed of six subsystems: The Solid Rocket Motor (SRM), the structural subsystem, the Thrust Vector Control (TVC) subsystem, the mechanical and ordnance equipment subsystem, the recovery subsystem containing the mechanical and parachute equipment, and the electrical subsystem including the Range Safety System (RSS). All elements except the nose cap and separation motors are intended to be reusable and are recovered via drogue and main parachutes.

The SRM is the primary propulsive element providing impulse and TVC from ignition to SRB staging. The SRM consists essentially of lined, insulated, segmented, weldless D6AC steel rocket motor cases loaded with TP-H1148 propellant. TP-H1148 is a composite type solid propellant formulated of polybutadiene acrylonitrile (PBAN), terpolymer binder, ammonium perchlorate, and aluminum powder, with a small amount of iron oxide burning rate catalyst. An ignition system, initiators, igniter, movable nozzle, systems tunnel, and instrumentation are other essential parts of the SRM.

Performance interchangeability and replaceability between a flight set of SRBs are maintained by matching the burning rates of motor segments cast in matched pairs from the same propellant lot. The sea level thrust of the SRM is 2.65 million pounds. The propellant grain design is performance tailored, consisting of a forward segment with an 11-point star and transitioning into a cylindrical perforated configuration in the cylindrical portion of the segment, two identically configured center segments that are tapered cylindrical perforated and an aft segment with a dual taper cylindrical perforated configuration. Figure 2 shows how the SRM is segmented indicating the field joints.

The 11-point star configuration in the forward SRM segment produces high level lift-off thrust until burnout of the star. Burning of the cylindrical perforated configuration continues until thrust decay due to burnout. A linear, 10-second thrust decay is achieved by the programmed burnout of slivers in all segments.

The insulation used in the chamber, propellant relief flaps, and forward inhibitors is asbestos- silica-filled nitrile butadiene rubber (NBR). The aft inhibitor insulation material is an asbestos filled carboxyl terminated polybutadiene (CTPB) polymer. The inhibitors typically are designed to prevent ignition and burning of the propellant grain in a direction perpendicular to the inhibitor surface. The case liner material is an asbestos-filled CTPB polymer (UF-2137).

The SRM nozzle is a convergent divergent movable design containing an aft pivot point, flexible bearing as the gimbal mechanism. This type of bearing has been used previously in the Poseidon missile; however, the SRM nozzle and flexible bearing are larger than any others in current use.

STS 51-L Incident Timeline Summary

The STS 51-L Mission Space Shuttle, Challenger, received the SRM ignition command at 11:38:01 a.m. EST on January 28, 1986, at the Kennedy Space Center. At 0.678 seconds, mission elapsed time (MET), an anomalous cloud of smoke appeared (according to camera E60 film) in the 270° to 0° quadrant of the right SRM at (or near) the aft field joint. This smoke persisted at least until 3.35 seconds MET. Subsequent analysis of the right chamber pressure trace showed that the motor performance at 5.674 seconds MET peaked at about 1.4 standard deviations high, so that the right booster chamber pressure was neither deficient nor excessive at this time. At 58.788 seconds MET, camera E207 recorded flickering flame visible between the right SRM and the External Tank (ET). At 60.004 seconds MET, the chamber pressure of the right booster began to diverge from that of the left booster. By 73.124 seconds MET, this chamber pressure disparity had progressed to produce a 19 psi differential between left and right motors. This 19 psi differential constituted a 4.9 standard deviation low departure from the performance of the left booster, which was responding normally at that time.

Scenarios

The initial identification of possible causes resulted in 22 possible failure modes (seven hot gas leak scenarios "subsection A" and 15 others- "subsections B through P"). As data subsequently became available, this list was systematically reduced to the seven hot gas leak scenarios (see Figure 1). Of these seven, continued effort has resulted in scenarios 1, 2, 4a, 4c, 5, and 7 being categorized as improbable. Data accumulated during the investigation indicate that the mechanisms or combinations of mechanisms contained in scenarios 3, 4b, 4d, and 6 could have resulted in or contributed to the failure. However, it must be noted that all six of the field joints of the two STS 51-L SRMs were subject to several of these mechanisms. To define the STS 51-L failure in the context of the failure mechanisms of these scenarios, it is desirable, if possible, to determine either a unique feature or a unique application of each or combinations as they may....

[L58] ....affect the STS 51-L right SRM aft field joint. This will be treated in each of the scenario to assessments and in the Accident Analysis Team Report.

Following in this section are descriptions and dispositions of all 22 scenarios including a narrative description for each with supporting or refuting data. It should be noted that the SRM Working Group does not carry any of the scenarios beyond 58.788 seconds MET. The Systems Working Group was charged with the responsibility to address events past 58.788 seconds MET.

A. Hot Gas Leak Failure Scenarios

1. Internal Heating/Inhibitor/lnsulation Anomalies

a. Scenario Description

SRM hot gas leak failure scenario I considers the possibility that a case membrane or joint burn through could have occurred at approximately 58 seconds due to increased internal heat transfer or insulation or inhibitor flaws. This scenario requires that the initial "puff of smoke" observed at 0.678 seconds was an independent event.

b. Investigation

The investigation consisted of: (1) definition of possible mechanisms by which internal convective heat transfer could cause abnormal hearing to the joint or membrane, (2) definition of insulation or inhibitor flaws which would result in abnormal heating to the metal with nominal convective and radioactive heat transfer, and (3) evaluation of the identified heat transfer or failure mechanisms by flow and thermal analysis, and flight and experimental data. All failures were evaluated against the following criteria as dictated by flight data or visual observations: (1) enough heating could be generated to either damage the O-ring to the point of failure or burn through the metal in 58 seconds, (2) the burn through would occur at or near the aft field joint, and (3) no effect on motor internal pressure would be observed (no deviation observed prior to approximately 60 seconds).

The nominal convective and radioactive heat transfer rates were obtained by comparing the measured material loss rate in the SRM aft field joint slot with the correlation of erosion rate to convective heat transfer coefficient from char-motor tests (Pendleton, S. B.; Evaluation of Elastomeric Insulation Materials at High Chamber Pressures, Morton Thiokol TWR-3896). The effect of nominal radioactive and convective heating on the aft field joint O-rings and metal was evaluated using a 2-dimensional thermal model of the joint. A simplified ballistics analysis was performed to estimate the effect of hypothesized failures on motor internal pressure. A simplified flow analysis was performed to evaluate the possibility that an inhibitor failure or uneven burning of an inhibitor could create circumferential flow of hot gas in the O-rings groove, causing erosion of the O-rings in approximately 58 seconds.

The SRM aft field joint configuration is illustrated in Figure 3. Flow disturbances were investigated as a possible cause of abnormal heat transfer. The following is a discussion of the effects of flow disturbances on convective heat transfer in the SRM.

(1) Axial Vortices

Under normal conditions, axial swirl and axial vortex pairs do not exist for reason of axial symmetry. Swirl can conceivably originate from uneven burning in the forward part of the forward motor segment. It would be restricted to a stream tube near the motor axis, causing a slight helical twist of the streamlines. It would have no effect on structural heating. Vortex pairs could originate from uneven burning of axial strips of the burning surface. They should be weak, have no effect on the combustion process, be restricted to the central motor cavity, and have negligible effects on structural heating.

(2) Circumferential Vortex Rings

Circumferential vortex rings exist in the space (or slots) between the propellant of the motor segments of the SRM from about t - 47 seconds to the end of burn; i.e., when the exit area of the slots exceeds the surface area of the axial burning faces. They are a. normal feature of all SRMs and drive the convective heat transfer on the surfaces they touch. No reason can be identified why they should have been different in the right motor on STS 51-L.

(3) Internal Acoustics

The noise level in the SRM cavity is of the order of 175 db, dominated by cavity noise, and decreasing with increasing burn time. Plane, coherent sound waves of this sound pressure level produce wave-induced velocities of about 1 percent or less of the mean gas velocity; a standing wave would only double this value. The actual random waves produce velocities of a much smaller magnitude. The convective heat transfer is therefore not affected by the acoustic noise level.

A failure due to increased convective heating on the joint is improbable. Therefore, assumed types of insulation flaws were examined to determine the probability of overheating the O-ring or joint with nominal convective and radioactive heat transfer. Figure 4 gives an example of the magnitude of an insulation flaw which would be required in order to result in a failure prior to 60 seconds. A convective heating rate (based on "char-motor" data and an observed erosion rate in the SRM aft field joint slot of 5.4 mils per second) of 380 Btu/ft² -sec and a radiation heating rate of 480 Btu/ ft² -sec were assumed and incorporated in a two-dimensional thermal model of the joint. The insulation thickness which would erode away at 5.4 mils per second to produce a failure at 60 seconds is shown in Figure 4. A flaw of this magnitude is unlikely and would be easily detected. Testing conducted by Morton Thiokol as part of this failure investigation indicated O-ring failure near 1,000°F. The primary and secondary O-ring temperature transients are shown in Figures 5 and 6 for the insulation flaw in Figure 4.

Insulation flaws at the field joint must be large to create a failure in 60 seconds. However, a relatively small flaw-in the forward-facing inhibitor could cause ignition of the propellant behind the inhibitor which would burn back prematurely. It is not probable that a failure of the castable inhibitor would cause a flight failure since this inhibitor normally burns away in approximately 13 seconds.

The effects of a forward-facing inhibitor failure was evaluated by assuming a 1-inch-diameter hole at the base of the inhibitor as shown in Figure 7. Analysis indicated that the change in motor internal pressure resulting from this failure would probably not be detected. However, an extremely high erosion rate would be required (as compared to observed values) to produce a burn through of the membrane by 58 seconds. In addition, the assumed flaw is unlikely since the inhibitor is constructed by vulcanizing eight individual plies and subsequent damage of the magnitude required is improbable and would be easily detected.

Other analyses conducted in support of assessments of this scenario are reported in Appendix B.

c. Findings

No credible failure mode based on inhibitor or insulation flaws or increased heating has been identified which results in a case membrane or joint burn through in 58 seconds and also reproduces flight data and visual observations. In addition, this failure scenario requires an independent and unrelated event to explain the smoke observed at 0.678 seconds. This scenario is judged to be improbable.

Figure 4. Assumed Insulation Flaw [top].

Figure 5. Primary O-Ring Temperature for Assumed Insulation Flaw of Figure 4 [bottom].

---

Figure 6. Secondary O-Ring Temperatures for Assumed Insulation Flaw of Figure 4 [top].

Figure 7. Assumed Inhibitor Flaw [bottom].

---

 

[L62] 2. LH₂ Leak Resulting in SRM Case Membrane or joint Failure

This hot gas leak failure scenario is discussed in the External Tank Report, Section IV.I.2. Based on the rationale presented therein, which draws on build data, prelaunch data, flight data, and analysis, this scenario was judged to be improbable.

3. Assembly Conditions/Anomalies

a. Scenario Description

This hot gas leak failure scenario hypothesizes damage of the right SRM aft field joint and/or O-rings during mating, leading to a hot gas leak during flight or other assembly anomalies which could have contributed to the failure. Possible event paths leading to joint/O-ring failure are shown in Figure 8.

Several events and photographic coverage of the STS 51-L mission, which depicted a hot gas leak in the region of the aft center segment to aft segment field joint of the right SRM, provided the initial rationale for this scenario. A review of the right aft SRM field joint dimensions at mating revealed that an interference fit resulted in the quadrant where the leak was observed. This condition could have resulted in generating contaminants (metal shavings), physical damage of the O-ring, or stress-related failure of the clevis leg. It also created a condition of near metal-to-metal at the clevis and tang sealing surfaces which could cause essentially complete compression of the O-rings into their grooves. The effect of this condition (max O-ring squeeze) is also involved in other scenarios and is discussed in subsequent sections of this report. Additional data were provided from an investigative activity to assess dimensional changes associated with case use. This activity revealed the possibility that growth was occurring in the critical sealing diameters of both clevis and tang with the greater growth occurring in the inner clevis leg. This could further increase the degree of interference known to exist as well as maximum squeeze on the O-rings.

The related events supporting this scenario included SRM stacking reports and procedures indicating difficulty in mating as a result of shipping-induced case segment ovality, postflight inspection reports indicating mechanical damage or pinch marks on field joint O-rings from SRM 14A, one ground test assembly incident of a damaged primary O-ring, and a reported incident of metal shavings in the sealing area after case segment disassembly. The incident involving metal shavings, although not directly relatable to flight assembly, was reported in TWR 13-515, covering the results of an investigation of DM-5 O-ring leak check failure.

Six major activities were initiated to assess the validity of Scenario 3. A detailed technical discussion and relevant findings relating to each activity are provided. A review of build records, including discrepancy reports, was also initiated to determine if errors or omissions in these documents existed. The review of SRM build records did not reveal any irregularities that could have contributed to the failure of 51-L, right SRM.

b. Investigation

(1) Stacking Conditions for 51- L and Other Flights (Shipping, Handling and Assembly)

There was difficulty in mating the right SRM aft field joint on 51-L. As a result of this report, a review of stacking procedures was initiated to determine the conditions surrounding the assembly of this joint and others. Data for the review consisted of KSC stacking records and pre-mate ovality measurements for the 51-L right SRM field joint and most other field joints assembled to date.

The SRB case segments are transported to KSC in the horizontal position resting in a shipping fixture with the tunnel (located at 90°) in the twelve o'clock position. Because of the flexibility of the segments, some ovality is induced with the major axis of the ellipse in the 0°/180° direction. Also, because of nonsymmetrical placement of the aft skirt holddown posts, ovality is induced in the aft casting segments, with the major axis in the 90°/270° direction. Thus two ovals must be assembled with their major axis out of phase by 90°.

Experience and analysis have shown that the segments can be mated and various techniques have been developed. The four-point handling beam was developed to lift and recontour the cases (Figure 9). This is accomplished by lifting the case segment and comparing tang end ovality measurements with those of the clevis to which it is being mated. The handling beam has provision to release two of the four pickup points thus permitting the weight of the loaded case to induce slumping and tend to round according to which two pickup points are released. Over time, while hanging from the crane, the case will tend to release some of the ovality induced by the shipping fixture. There is a requirement that prohibits mating flat-to-flat and the mating procedures state that ovality cannot exceed 0.250 inch in the positive direction (tang outer diameter greater than clevis outer leg inner diameter) to prevent flat-on-flat while attempting to lower the tang into the clevis. However, no requirement exists for controlling the situation which would exist for the tang inner diameter being less than the inner clevis leg (negative direction. Figure 10). This situation existed for the right SRM aft field joint and to a lesser extent for other STS 51-L joints.

Should the handling beam be unable to achieve conditions suitable for mating in a reasonable time (hours), a further attempt to round can be made using a rounding tool. (Reference Figure 11.) The rounding tool is essentially two outside fitting chocks with a connecting rod through two opposing tang pin holes. The rod is tensioned using a small hydraulic pump until the segment is brought into the approximately correct diameter for mating. The rounding tool is removed just before mating.

Use of the crane hoist fixture and the rounding fixture as currently designed do not always correct all ovality sufficiently for mating. Some difficulty in assembly is still experienced, particular1y with the aft field joint. The 51-L right SRM aft and 61-G left SRM aft field joints were among the most difficult to assemble of any attempted to date because of excessive ovality.

The right SRM aft field joint of 51-L required numerous attempts to match the tang ovality with that of the clevis. Following the initial four-point lift, three two-point lifts were employed and then the rounding tool was used. The last measurement before mating indicated a positive ovality measurement of 0.216 inch, only 0.034 inch from the maximum allowed by procedure. Negative ovality measurements were 0.334 and 0.393 inches at the 90°/270° and 120°/300° positions respectively, placing the tang end in contact with the inner clevis leg and O-rings at these locations. Because of the extreme ovality at mating, the possibility of the flat end of the tang resting on the flat top of the inner clevis leg before sliding down the taper into the slot, cannot be discounted. Negative measurements in the range of 0.340 to 0.355 inches can result in flat-on-flat if the segment center lines are not co-aligned within approximately 0.200 inch. Testing has demonstrated that this mating mode can generate metal shavings which in turn could compromise performance of the O-ring seals.

Although the right SRM aft field joint was the only one on 51-L requiring use of the rounding tool to achieve mating, other field joints also exhibited ovality with varying degrees of negative dimensions. The maximum negative ovality at mating for each of the remaining five field joints ranged from 0.178 to 0. 374 inches. Some of these conditions place the end of the tang in direct contact with the O-rings on insertion and result in near metal-to-metal contact at the sealing surfaces after assembly.

(2) Dimensional Assessment of Refurbished Cases

Case dimensions in the region of the tang and clevis are critical to calculations of O-ring squeeze and subsequent seal performance. It was determined in the early phases of the investigation that tang and clevis sealing surface diameters are not reestablished after case use, prompting the concern that changes in these critical.....

[L65] ....dimensions may have occurred.

To address this concern, two case segments (SN0000056 with four uses and SN00022 with five uses) were returned to Rohr for dimensional assessment. Measurements including Pi tape of the sealing surface diameters were made with the segments mounted in the machining lathe. The same procedures and equipment originally employed by Rohr prior to initial shipment of these segments to MTI, were used. These same measurements were also recorded with the segment in the unrestrained condition using the same equipment and procedures employed in initial new case machining buyoff. Case SN0000022 exhibited a 0.028-inch increase in diameter at the tang sealing surface and an increase of 0.034 inch in diameter at the clevis sealing surface. Case SNO000056 exhibited an increase of 0.022 inch and 0.032 inch at the tang and clevis sealing surfaces respectively. (See Figure 12 for tabulation of these dimensions and deltas between new case measurements and measurements taken on used cases.) Figure 13 presents a ratioed extrapolation of the growth experienced on Case Segment SNOO00056 with four prior uses to the 51-L aft field joint containing a tang end with one prior use and clevis end with two prior uses. This analysis disclosed a potential of a 0.002inch larger diameter clevis than tang resulting in an interference fit.

A similar analysis performed on the other 51-L field joints revealed a potential interference fit on the right SRM centerjoint of 0.005 inch and a near interference fit (0.0005-inch clearance) on the right SRM forward field joint. The joints on the left motor all exhibited a clearance greater than 0.005 inch, assuming concentricity. The greatest increase in clearance was noted on the left SRM forward field joint. Again, assuming concentricity, the clearance increased from 0.008 inch, as calculated from the original machined dimensions, to 0.013 inch using the ratioed extrapolation of growth based on case segment reuse.

(3) Sructural Analysis of 51-L Right SRM Aft Field joint Mating

Analysis of the 51-L right SRM aft field joint mating conditions was performed because of the difficulty experienced in assembling this hardware. Specifically, the study was structured to determine the deflections and stresses experienced during mating the right SRM aft center segment to the aft segment and an analysis of the fracture mechanics of the O-ring groove to determine the criticality of a flaw in the highest stressed area; i.e., the maximum size of a flaw that would not fail under the life cycle history of the 51-L case segment. Included in this analysis was the point load needed to deflect a suspended segment to the side by 0.200 inch and the maximum stress on the case clevis this causes. Additionally, the study was structured to determine the load/deflection relationship required to assemble an ovalized tang to an ovalized clevis. Finally, the study addressed an occasionally encountered condition where the tang rests on top of the inner leg of the clevis on one side and slips down into the clevis groove 180° opposite. The analysis determined the loads involved and if damage to the assemblies could occur.

The fracture mechanics analysis showed that flaws 0.21-inch deep x 2.1-inch long in the secondary O-ring groove will survive over 100 uses without any substantial growth. Flaws of this size would be readily detectable by the magnetic particle inspection performed on all case segments. The detectability limit specified for this inspection is a flaw length of 0.1 inch. The load to deflect a suspended segment over 0.2 inch in the horizontal direction was calculated to be 75 pounds and causes a maximum stress in the secondary O-ring groove of only + 966 psi. Resting a segment on the top of the inner leg of the clevis and allowing the segment to rotate down 0.7 inch to the outer leg of the clevis on the opposite side results in an 8,000-pound vertical load on the inner leg. This produces a maximum compression stress of 62,000 psi in the primary O-ring groove. Applying a horizontal point load of 1,000 pounds at the top of the inner leg of the clevis produces a maximum stress of + 12,871 psi at the secondary O-ring groove.

The results of the analysis indicate that the stresses induced during the mating operation were low and would not have resulted in hardware damage. Also the stresses would not have resulted in significant growth of an undetected flaw of a length substantially greater than that which is detectable by the inspection techniques employed.

(4) Testing to Simulate 51-L Right SRM Aft Field joint Mating

This test program was designed to simulate probable conditions during the assembly of the right SRM aft to aft center segment field joint of STS 51-L. Careful attention was focused on the efforts of controlled assembly conditions on the O-rings and the mating metal surfaces of the tang and clevis. Testing included laboratory studies with a 1-foot section of tang and clevis and full-scale studies using fixtures at MTI.

A section of a field joint consisting of a tang and clevis (each approximately 12 inches x 13 inches) was secured in a tensile test machine (50,000-pound capacity) with two rigid fixtures. Two lengths of O-ring material (taken from flight quality sections of SN0001669 and SN0001777) were held in place with fasteners at point's tangent to the arc of the clevis and were tensioned within a predetermined range. The tang was positioned in its fixture above the clevis in a specific position for each test to simulate several different probable combinations of mating surface interference and relative angles. The clevis was then raised into the tang by the tensile machine drive mechanism. Various measurements and observations were made before during, and after the test. Any damage to O-ring or metal surfaces or any occurrence that would tend to increase the probability of damage was noted.

Two significant findings were made. First, metal slivers were generated under mating conditions in which the flat end of the tang overlapped the flat end of the inner clevis leg by 0.005 to 0.010 inch before encountering the guiding taper. (Considering some change in ovality may have occurred between the time the last measurements were made on 51-L and final mating, such a condition could have existed for the right SRM aft field joint.) The metal slivers in turn were carried into the joint and deposited on and around the O-rings. The second significant finding, was that the length of the O-ring segments increased as the tang entered the clevis and compressed the O-ring diameter. This increase in length averaged 4.1 percent. The implication of this finding is that canted tang entry in a full diameter segment could "chase" the O-ring around the circumference, resulting in it bulging out of the groove on the opposite side, thus making it vulnerable to damage.

Full-scale tests using fixtures at MTI were conducted to further assess conditions of O-ring elongation noted in the laboratory. These tests were structured to determine if insertion of the tang into the clevis, when the tang is tipped relative to the clevis, will cause the O-ring to accumulate excessive slack such that it will locally extend beyond the groove and be damaged during assembly. Initial tests were conducted with a short stack fixture. This fixture is a full diameter representation of the joint area in the SRM. The results of these tests proved inconclusive in that the fixture was too flexible to induce any observable length change in the O-rings. Further tests employing a more rigid test fixture and/or SRM hardware are under consideration.

(5) Testing to Assess the Effects of Assembly-Induced Contamination on O-ring Sealing

These tests employed metal contaminants to simulate those generated in the segment mating tests. The tests were structured to determine if assembly-induced metal shavings, positioned between the O-ring and sealing surface, could pass a static leak check but fail under dynamic loading. Tests were designed to use the subscale static and dynamic blowby fixtures and the full-scale short....

Figure 11. Rounding Tool [top].

Figure 12. Case Dimensional Reassessment at ROHR [bottom].

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[L68] ....stack test fixture. The data show that all simulated metal contaminants tested produce essentially the same results within the variability of the data; the thickness of metal contaminant, which first passes the 50 psig level check, is in the 0.001 to 0.003 inch range. Tests to determine if this size contamination will compromise the seal under dynamic loading have not been completed at the time of this writing.

(6) Destacking STS 61-G

It was also determined that STS 61-G, the last vehicle to be stacked and still in the Vertical Assembly Building (VAB) at KSC. experienced similar mating problems with the left SRM aft field joint as did STS 51-L. STS 61-G afforded an unusual opportunity to look for assembly-related anomalies. Accordingly, a plan for destacking 61-G was prepared and forwarded to KSC.

The plan outlines specific actions to be taken during destacking to support an assessment of conditions of the motor and case field joints. The requirements identified were incorporated in specific destacking instructions issued by KSC. The plan includes a series of dimensional inspections, visual observations and assessments of hardware conditions with the operations sequenced to maximize data collection. The destacking operations are underway and this report will be augmented to incorporate the results of the destacking when completed.

c. Findings

Analysis indicates that the stresses induced during the mating operation would not have resulted in hardware damage nor would they have resulted in significant growth of an undetected flaw significantly larger than would have been detected by the inspection conducted.

Tests conducted in support of this specific scenario investigation did not result in damaged O-rings; however, such damage has been documented previously in ground test motor reports. The potential for O-ring damage due to "chasing" during assembly has not been adequately tested at this time. Such tests are under consideration. Metal shavings were generated during simulated mating tests and further testing demonstrated that small metal contaminants in the critical seal area could not be detected by the standard SRM leak check. Considering the conditions which existed during mating and reports of prior observations, the possibility that contamination was generated and could have contributed to a degraded seal cannot be dismissed. Although not totally unique to the right SRM aft field joint, the mating condition that could cause seal damage and/or contamination at assembly were more severe at this joint. Further, the location of the joint failure (burn through) was in the same circumferential location as the maximum interface at mating.

The dimensions taken during the mating operation indicated that a metal-to-metal fit existed in the quadrant where the leak was observed. This assembly condition could have been a contributing factor to the leak when combined with other conditions which existed. Such conditions will be treated in subsequent subsections.

4. Scenario 4: Primary O-Ring Blow-by or Low Temperature Effects

Hot Gas Leak Scenario 4 is a multi-branch scenario consisting of possible combinations of effects necessary to breach both primary and secondary O-rings. The common element of the scenario is primary O-ring blowby and/or low temperature effects on the primary O-ring. The four branches following the common element are 4a, Secondary O-ring Defect; 4b, Ice in joint: 4c, Leak Check Port Leaks; and 4d, O-ring Actuation Time Delayed by Low Temperature. Each branch will be discussed in subsequent subsections.

Following SRM ignition, internal pressure in the motor case causes structural deflection (rotation) of the joints in a direction that reduces the compression on the O-rings. This behavior is quantified in the analyses and tests reported in Appendix B, Part A. The sealing performance of the O-rings under this condition, as influenced by low temperature, is the subject of both the Scenario 4 common element and Branch 4d. Although joint geometry causes the primary O-ring to be more likely to experience blow-by than the secondary, the basic phenomena and method of disposition are common to both scenario blocks; therefore, a common discussion will be provided in subsection 4d.

Two primary mechanisms of this set of scenarios existed for all six SRM field joints on 51-L, specifical1y the low temperature of the O-rings and the joint rotation. The joint temperatures at launch were determined to be essentially the same (for all or a portion of the circumference) for all six joints. Joint rotation (seal gap opening due to motor pressurization) exists for all joints, although it is somewhat greater for the center and forward field joints as opposed to the aft field joints. Another important variable is the degree of O-ring squeeze. Tests and analyses indicate that squeeze conditions, which cause the O-ring to fill the groove to such an extent that the O-ring is compressed against the sides of the groove, are detrimental to sealing at low temperatures. This condition also existed to varying degrees on several joints.

a. Scenario 4a: Secondary O-Ring Defect

(1) Scenario Description

This scenario hypothesizes that there is blow-by of the primary O-ring and then leakage past the secondary O-ring due to an O-ring defect that existed prior to assembly.

(2) Investigation

This scenario was investigated when closeout photos of the right SRM aft field joint showed a suspect secondary O-ring. The investigation was conducted by an analysis of closeout photographs, review of records and review of results of special inspections and tests. (See Figure 14, Scenario 4a Logic Chart.)

Eleven closeout photographs were taken around the right SRM aft field joint showing the primary and secondary O-rings and other features of the clevis. These pictures were examined to determine if any damaged or defective areas existed in the O-rings or if any contamination existed in the joint or on the O-rings. Detailed examinations of the closeout photos were made using digital enhancement techniques. Laboratory experiments were performed to duplicate features found in the photographic analysis.

Assembly records were reviewed to establish the serial number of the O-rings used on STS 51-L. MTI receiving inspection records were reviewed to determine results of inspections and reject-history of all O-rings processed at MTI. Vendor process records were also reviewed to determine compliance with plans and specifications. The intent of this record review was to determine the probability of a defective O-ring not being detected prior to assembly.

All O-rings in inventory at MTI-Wasatch, VAFB, and KSC were reinspected to determine their compliance with specification requirements. Each O-ring was reinspected using a standard micrometer equipped with a 1/2 inch foot to measure the 0.280-inch diameter and a special tool to measure the 142-inch diameter. This same equipment had been used for the original acceptance inspection. During reinspection, the 0.280-inch diameter was measured every 9 inches in two planes. It was measured every 2 to 4 feet in only one plane during the original acceptance inspections. Each O-ring was also reinspected using a laser micrometer. This instrument provided continuous diameter measurements with a 0.007-inch measuring team width. Measurements were taken in four planes. All O-rings were also X-rayed to determine if voids and inclusions were present.

Color prints of the closeout photographs became available a few days after the incident. These photographs were not of engineering quality and little could be determined by direct inspection of the prints. However, an interesting thickening of the [L69] dark space between the upper surface of the secondary O-ring and the lower edge of the land was noted together with some notable streaks in the same area on both the secondary O-ring and the land. This print with supplemental engineering information was sent to the Los Alamos National Laboratory (LANL) for enhancement and mensuration. They, in turn, arranged for the mensuration to be accomplished by the National Photographic Interpretation Center (NPIC). Their report is attached as enclosure to MSFC Test Report 9. Mensuration data of the dark spaces indicated a gradual thickening of about 0.015 inch and a subsequent decrease to about 0.006 inch extending over about 0.5 inch of O-ring length. Note that these measurements were made after the image was rotated and the perspective corrected. These data suggested a possible O-ring defect. No other notable items were found in any of the closeout prints.

Since the mensuration and enhancement was made from a poor print of the closeout photo, fine details were not perceptible, having been lost in the reproduction process. The original negative was obtained and sent to LANL and NPIC for enhancement. The NPIC report of this enhancement and mensuration is contained in MSFC Test Report 9. There was a dramatic increase in quality of the print from the original negative. More details were available. The dark space gradual thickening was changed significantly. The upper edge of the dark space was found to contain two abrupt discontinuities which are faithfully mapped to the lower edge of the dark space, indicating that what was observed on the O-ring was a shadow of irregularities on the lower edge of the land caused by grease.

To reproduce the features seen in the photograph, a curved 3-foot-long section of the inner wall of the clevis was fabricated to nominal dimensions. This fixture was used for photographic investigations of the grease patterns and O-ring defects. Grease was applied to the metal plate using gloved hands. With relatively light coatings of grease, structures or ridges of about 0.005 inch height were easily and normally developed. Heavier coatings of grease produced larger structures. These data proved that small structures could be created on metal surfaces such as the O-ring land in the normal O-ring installation process without violating the process.

Examination of the remainder of the digitized and enhanced closeout photographs revealed another artifact, a circular structure about 0. 15 inches across. Tests proved that reasonably similar structures could be created in the grease without violation of the joint O-ring installation procedures. This artifact was concluded to be a grease structure and not contaminant.

A detailed study to understand and review the O-ring process flow was conducted. This study indicated that the 0.280-inch diameter of each O-ring is measured by MTI and MTI vendors at approximately 100 places prior to final acceptance. It is highly unlikely that a major defect would escape detection by this system. A review of O-ring receiving inspection history at MTI was conducted to determine the results of inspection of the O-rings used on 51-L and to determine the results of inspection of the O-rings received by MTI. Receiving inspection records indicate that the right SRM aft field joint secondary O-ring contained three splices (five allowed) and one repair (20 allowed). The primary O-ring contained three splices and five repairs. It was discovered that in March of 1985 a revision was made to the MTI receiving inspection plan that resulted in 2 of 24 MTI inspection steps being changed from direct buyoff to verification through vendor certification. Also, three of seven Air Force Quality Assurance inspection steps were deleted by mistake. The secondary O-ring in the right SRM aft field joint in 51-L plus six other field joint O-rings in 51-L were inspected to this less strenuous plan. Inspection steps associated with the 0.280-inch diameter were not affected by this change. Details of this inspection planning change is contained in Table 1.

A summary of O-ring receiving inspection history was reviewed to determine if O-rings had been rejected at MTI for diameter problems. It was found that 32 case O-rings had been rejected since 1978. None of these were rejected for the 0.280-inch diameter requirement. Seventy-five percent of these rejects were associated with dimensions of the scarf joint. Each scarf joint is measured at MTI receiving inspection. These data reflect a low total reject rate of 1.8 percent which indicates a good quality product being supplied by the vendor.

A review of records at Parker Seal Co. did reveal a records traceability problem associated with the O-ring batch containing the subject secondary O-ring. Parker Seal Co. has supplied O-ring material for the STS program since 1978. During this period, they have supplied material with 14 different combinations of formulation and processes. These combinations are shown on Table 2. The Parker Seal Co. was unable to produce records that would validate the batch of material used to fabricate the subject secondary O-ring. Although the exact formulation is unknown, it is known that it is one of the acceptable formulation used by the Parker Seal Co. Samples of O-ring material from this batch have been tested at MSFC for requirements of MILR-83248 and found to be in compliance.

All O-rings in inventory at MTI, KSC, and VAFB were reinspected using the same inspection equipment and acceptance criteria used when initially accepted. The 0.280-inch diameter was measured every 9 inches in two planes. Eight of 190 O-rings inspected were rejected. The worst out-of-tolerance condition found was a nine inch length that was 0.003 inches undersize. The average diameter obtained from the 18,000 measurements was 0.2810 inch with a standard deviation of 0.0013 inch. The average diameter ± 3a [standard deviation] is 0.2849 and 0.2771 which is within the specification requirements of 0,285 and 0.277. Statistical analysis indicated that the process will meet the specification requirements.

The inventory O-rings were also reinspected using a laser micrometer. This instrument has a 0.007-inch measuring beam arid will provide measurement to five places when calibrated with gage pins.