COLUMBIA ACCIDENT INVESTIGATION BOARD (CAIB)/ 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION (NASA) 

ACCIDENT INVESTIGATION TEAM (NAIT) 

WORKING SCENARIO 

FINAL VERSION: JULY 8, 2003 
PREFACE 

This Working Scenario report was written to document the collection of known facts, 

events, timelines, and historical information of particular interest to the final flight of 

Columbia. The report was written with the understanding that it could be published, 

either in part or in its entirety, as part of the official Columbia Accident Investigation 

Board (CAIB) report. The report includes information and results from numerous 

analyses, tests, and simulations related to the Columbia investigation that have been 

completed, or were ongoing at the time that this report was completed. It is anticipated 

that additional analytical and test results will emerge from ongoing work, as well as from 

future activities associated with the Columbia investigation and efforts related to the 

Return-To-Flight work. This Working Scenario includes information and results as they 

existed up to and including July 8, 2003. 
CONTENTS 

Section Page 

1.0 INTRODUCTION ........................................................................................ 1-1 

1.1 SCOPE....................................................................................... 1-1 

1.2 MISSION BACKGROUND.......................................................... 1-2 

2.0 LAUNCH COUNTDOWN............................................................................ 2-1 

3.0 LAUNCH..................................................................................................... 3-1 

3.1 INTRODUCTION........................................................................ 3-1 

3.2 LAUNCH DEBRIS IMPACT OBSERVATION............................. 3-2 

3.2.1 Launch/Ascent Conditions.......................................................... 3-2 

3.2.2 Launch Debris Impact Area........................................................ 3-4 

3.2.3 Launch Photo and Transport Analysis........................................ 3-6 

3.2.4 Debris Velocity and Size Assessment ........................................ 3-10 

3.2.5 Impact Damage Testing and Analysis ........................................ 3-12 

3.3 LAUNCH MADS DATA............................................................... 3-14 

3.4 LAUNCH AREA RADAR ANALYSIS.......................................... 3-19 

3.5 LAUNCH GUIDANCE NAVIGATION AND CONTROL............... 3-25 

3.5.1 Wind Shear, Day of Launch Wind Effects .................................. 3-25 

3.5.2 Predicted/Actual Loads .............................................................. 3-29 

3.5.3 ET Liquid Oxygen Slosh ............................................................. 3-33 

3.5.4 Nozzle Positions ......................................................................... 3-35 

3.5.5 ET Separation Yaw Rate............................................................ 3-40 

3.5.6 Data Correlation of Flights that Used a LWT and PEs ............... 3-41 

3.5.7 Data Correlation of Flights with ET Bipod Foam Liberation........ 3-42 

4.0 ORBIT......................................................................................................... 4-1 

4.1 INTRODUCTION........................................................................ 4-1 

4.2 ORBITAL DEBRIS...................................................................... 4-1 

4.2.1 Orbital Debris Risk Assessment ................................................. 4-1 

4.2.2 Micrometeoroid or Orbital Debris Detection................................ 4-1 

4.3 FLIGHT DAY 2 EVENT .............................................................. 4-7 

4.3.1 Radar Tracking of Flight Day 2 Object ....................................... 4-7 

4.3.2 Analysis of Mechanisms for Object Release .............................. 4-9 

4.3.3 Radar Cross Section and Ballistics Testing................................ 4-11 

4.3.4 KSC Lost and Found Items ........................................................ 4-13 

4.4 ORBIT SUMMARY ..................................................................... 4-13 

i 
Section Page 

5.0 DEORBIT/ENTRY....................................................................................... 5-1 

5.1 INTRODUCTION........................................................................ 5-1 

5.2 WEATHER ................................................................................. 5-1 

5.2.1 Upper Atmosphere Weather....................................................... 5-1 

5.2.2 Landing Weather........................................................................ 5-1 

5.3 HARDWARE FORENSICS......................................................... 5-4 

5.4 ENTRY EVENTS TIMELINE ...................................................... 5-12 

5.4.1 Early Entry Heating Events......................................................... 5-12 

5.4.2 First Roll Maneuver Through Wing Spar Breach........................ 5-16 

5.4.3 Wing Breach and Wire Failures.................................................. 5-28 

5.4.4 Aerodynamic Events................................................................... 5-37 

5.4.5 Wheel Well Gas Penetration and Final Aerodynamic Events .... 5-47 

5.5 AERODYNAMIC RECONSTRUCTION...................................... 5-53 

6.0 RE-USABLE SOLID ROCKET MOTOR..................................................... 6-1 

7.0 SOLID ROCKET BOOSTER...................................................................... 7-1 

8.0 SPACE SHUTTLE MAIN ENGINE ............................................................. 8-1 

9.0 ENVIRONMENTAL FACTORS................................................................... 9-1 

9.1 INTRODUCTION........................................................................ 9-1 

9.2 AGE AND EXPOSURE .............................................................. 9-1 

9.3 WEATHER FACTORS ............................................................... 9-5 

10.0 LEFT WING PROCESSING AND RCC DESIGN ....................................... 10-1 

10.1 INTRODUCTION........................................................................ 10-1 

10.2 LEFT WING PROCESSING (PALMDALE, J3-OMM)................. 10-1 

10.3 LEFT WING PROCESSING (STS-109) ..................................... 10-4 

10.4 LEFT WING PROCESSING (STS-107) ..................................... 10-5 

10.5 RCC DESIGN............................................................................. 10-7 

10.6 RCC IMPACT RESISTANCE ..................................................... 10-9 

10.7 RCC CORROSION..................................................................... 10-10 

11.0 EXTERNAL TANK...................................................................................... 11-1 

11.1 INTRODUCTION........................................................................ 11-1 

11.2 TPS REQUIREMENTS............................................................... 11-1 

11.3 HISTORY OF FOAM CHANGES AND DEBRIS EVENTS ......... 11-2 

ii 
Section Page 

11.4 STS-107/ET-93 CHRONOLOGY................................................ 11-6 

11.4.1 Bipod Ramp TPS Configuration ................................................. 11-6 

11.4.2 Bipod Ramp Certification............................................................ 11-8 

11.4.3 Bipod Ramp Build Process......................................................... 11-8 

11.4.4 Bipod Ramp Foam Acceptance/Non-Destructive Evaluation ..... 11-9 

11.4.5 ET Shipping and Handling.......................................................... 11-9 

11.4.6 KSC Processing Activities .......................................................... 11-9 

11.4.7 ET Pre-launch Operations.......................................................... 11-11 

11.4.8 Launch/Ascent............................................................................ 11-11 

11.4.9 Possible Contributors to Strain Energy at ET Separation........... 11-12 

11.5 STS-107/ET-93 TPS BIPOD DEBRIS........................................ 11-14 

11.5.1 Bipod Foam Failure Modes and Contributors............................. 11-14 

11.5.2 Test Results for Debris Assessment .......................................... 11-16 

11.5.3 Max Bipod SLA Temperatures (80 seconds MET) ..................... 11-17 

11.5.4 Bipod Ramp As-Built Hardware Assessment ............................. 11-18 

11.5.5 Multi-Failure Mode TPS Bipod Debris ........................................ 11-19 

12.0 SUMMARY ................................................................................................. 12-1 

Appendix 

A ACRONYMS AND ABBREVIATIONS ........................................................ A-1 

iii 
 

TABLES 

Table Page 

3-1 Wing damage analysis methods and results......................................... 3-5 

3-2 Transport analysis and ET Working Group estimates of ET bipod 

debris size, weight, and volume ............................................................ 3-12 

3-3 STS-107 ascent radar events ............................................................... 3-20 

3-4 Material samples from post-STS-27 radar calibration tests .................. 3-23 

3-5 LWT and PE flights ............................................................................... 3-41 

3-6 STS flights with ET left bipod foam liberation........................................ 3-43 

4-1 Summary of analysis of 13 rate events ................................................. 4-3 

4-2 Summary of analysis of the lower bound of MMOD (based on body 

rate data)............................................................................................... 4-4 

4-3 Chronology of events related to flight day 2 object ............................... 4-10 

4-4 Summary of nominal launch day events ............................................... 4-11 

4-5 Lost tools in Columbia processing for STS-107, STS-109, 

and OMM J3.......................................................................................... 4-13 

10-1 RCC refurbishment limits ...................................................................... 10-10 

11-1 STS-Orbiter-ET configuration, age, and exposure................................ 11-3 

FIGURES 

Figure 

1-1 STS-107 payload bay configuration ...................................................... 1-2 

2-1 STS-107 Launch Countdown (LCD) overview flowchart ....................... 2-2 

2-2 Shuttle vehicle coordinate system......................................................... 2-4 

3-1 Launch of STS-107 at pad 39A at Kennedy Space Center................... 3-2 

3-2 STS-107 reconstructed altitude during first stage (prior to SRB 

separation) ............................................................................................ 3-3 

3-3 STS-107 flight reconstruction data for mach number and dynamic 

pressure (Q-bar) prior to SRB separation. Note that Q-bar is highest 

during first stage (prior to SRB separation), and reduces to a very 

small number after SRB separation ...................................................... 3-3 

3-4 Area of most likely wing damage........................................................... 3-4 

iv 
 

 

 

Figure Page 

3-5 Multiple analyses determine foam impacted lower RCC panels 6 

through 8 area....................................................................................... 3-6 

3-6 Photographic analysis techniques determined foam size: debris 

appears almost circular in frame 4914 and elongated in frame 4919 ... 3-7 

3-7 Camera geometry for ascent video analysis; note that video camera 

ET208 is at same location as film camera E208 ................................... 3-7 

3-8 Orbiter view from Cameras E212 and ET208 ....................................... 3-8 

3-9 Multiple analyses indicate foam is from ET left bipod area ................... 

Red line depicts the estimated foam trajectory as it moved from the 

bipod ramp area toward the left wing. ................................................... 3-9 

3-10 Video analysis shows impact is below wing leading edge stagnation 

line. Trajectories of particles are depicted after the impact.................. 3-9 

3-11 Pre-impact vs. post-impact shows no observable damage within the 

resolution limits ..................................................................................... 3-10 

3-12 Sample CFD flow field with debris modeling ......................................... 3-11 

3-13 CFD surface flow with lower left wing pressure sensors ....................... 3-14 

3-14 Unusual behavior of pressure sensor V07P8074A ............................... 3-15 

3-15 Close-out photo shows RCC panel 9 wing leading edge temperature 

measurement ........................................................................................ 3-16 

 3-16 Three-bit rise (7.5 degrees F) on MADS wing leading edge spar 

temperature measurement (V09T9895A) during ascent....................... 3-17 

3-17 Correlation between simplified thermal math model and STS-107 

ascent and entry flight data ................................................................... 3-17 

3-18 STS-107 ascent and entry heating environments on RCC panel 9....... 3-18 

3-19 Limits of dimensional detectability for three simple shapes .................. 3-22 

3-20 Limits of radar cross section (RCS) detectability and measured 

STS-107 debris for three radar source sites ......................................... 3-22 

3-21 Out-of-plane wind velocity..................................................................... 3-26 

3-22 Side-slip angle....................................................................................... 3-27 

3-23 Wing loads during wind shear and side-slip angle ................................ 3-29 

3-24 ET interface loads at forward attachment during wind shear and 

side-slip angle. Q-beta is side-slip angle multiplied by the dynamic 

pressure and represents the side-slip angle contribution of the 

interface load......................................................................................... 3-30 

3-25 Slosh effect on ET interface loads ........................................................ 3-30 

3-26 ET bipod axial aerodynamic loads ........................................................ 3-31 

3-27 ET bipod side-force aerodynamic loads................................................ 3-32 

3-28 ET bipod radial aerodynamic loads....................................................... 3-32 

3-29 STS-107 SRB tilt actuators experienced more than typical 0.6 Hz 

content .................................................................................................. 3-33 

3-30 STS-107 SRB gimbal responses at 0.6 Hz frequency correlated to 

wind....................................................................................................... 3-34 

3-31 Center SSME yaw position.................................................................... 3-36 

v 
 

 

Figure Page 

3-32 Right SSME yaw position...................................................................... 3-36 

3-33 RSRM burn rate at propellant mean bulk temperature (PMBT) ............ 3-37 

3-34 SRB thrust mismatch ............................................................................ 3-38 

3-35 SRB nozzle position for PE flights versus non-PE flights...................... 3-39 

3-36 ET separation yaw rate ......................................................................... 3-40 

4-1 Jet firing example for vehicle rates........................................................ 4-5 

4-2 Sample data from SAMS and ODRC .................................................... 4-6 

4-3 SAMS data frequency content .............................................................. 4-7 

4-4 Tracking of flight day 2 object through various sensor passes.............. 4-8 

4-5 On-orbit RCS shows increased tumble/rotation rate over time ............. 4-9 

4-6 Leading edge structural subsystem components matching RCS 

and ballistics.......................................................................................... 4-12 

5-1 Wind profile developed by DAO as part of the STS-107 investigation 

(time referenced to 8:min:sec EST) ...................................................... 5-2 

5-2 Slag deposition in the RCC panel 8/9 area relative to the other parts 

of the left wing leading edge ................................................................. 5-4 

5-3 Samples of severe slag deposition on the panel 8 rib........................... 5-5 

5-4 Example of rib erosion .......................................................................... 5-6 

5-5 Flow on the lower carrier panel 9 tiles................................................... 5-7 

5-6 CAD drawing of the recovered debris showing overall slag deposition 

and erosion patterns ............................................................................. 5-7 

5-7 Analysis results show possible flow direction and deposition of 

metals.................................................................................................... 5-9 

5-8 RCC panel debris location .................................................................... 5-10 

5-9 Three possible orbiter locations of the Littlefield tile on left wing .......... 5-11 

5-10 STS-107 stagnation heat flux and dynamic pressure. Note that EI 

was at 8:44:09 EST............................................................................... 5-13 

5-11 Left wing RCC panel 9 strain gauge is first measurement to indicate 

an off-nominal event. Note that EI was at 8:44:09 EST ....................... 5-13 

5-12 MADS sensors inside left wing.............................................................. 5-14 

5-13 Left wing RCC panel 9/10 clevis temp sensor is second 

measurement to indicate an off-nominal event ..................................... 5-15 

5-14 Typical off-nominal OMS pod thermocouple (V07T9220A)................... 5-16 

5-15 Location of OMS pod thermocouples off-nominal low........................... 5-17 

5-16 Postulated orbiter leeside flow field associated with wing leading 

edge damage ........................................................................................ 5-18 

5-17 Orbiter wind tunnel model with vent gap along wing leading edge........ 5-19 

5-18 Wind tunnel model results for sensitivity of orbiter side fuselage 

and OMS pod heating patterns to mass addition along WLE leeside 

vent gap ................................................................................................ 5-19 

5-19 Location of left sidewall temperature sensor......................................... 5-20 

5-20 Off-nominal temperature indication on the left sidewall ........................ 5-21 

vi 
 

25 

30 

35 

40 

45 

50 

Figure Page 

5-21 Temperature rise on tile surfaces aft of RCC panel 9 ........................... 5-22 

5-22 Left wing MADS sensors, including Measurement Stimulation 

Identification (MSID) number, and start time of loss of signal 

(EI + sec.).............................................................................................. 5-23 

5-23 RCC panel 9 MADS strain and temperature measurements, 

STS-107................................................................................................ 5-24 

5-24 Off-nominal low OMS pod thermocouple (V07T9972A)........................ 5-25 

5- Left side fuselage/OMS pod off-nominal responses indicate 

increased heating.................................................................................. 5-26 

5-26 Wing tunnel test results for RCC panel 9 missing and resulting in 

increased heating to OMS pod.............................................................. 5-27 

5-27 CFD results for no damage, partial damage, and full damage to RCC 

panel 9 show increased heating on side fuselage and OMS pod ......... 5-27 

5-28 STS-107 entry heating rate profile ........................................................ 5-28 

5-29 Cable routing on wing leading edge and wheel well wall ...................... 5-29 

5- Thermal model prediction of wing spar burn through............................ 5-30 

5-31 Hot gas begins to fill left wing................................................................ 5-31 

5-32 Columbia LH wing and wheel well geometry......................................... 5-32 

5-33 Columbia LH wing and wheel well vent model (wheel well leak paths 

based on Atlantis test comparison) ....................................................... 5-33 

5-34 MADS data failure due to wire burning.................................................. 5-34 

5- View of cables running along outside of wheel wall cavity bulkhead..... 5-35 

5-36 Strain measurements on 1040 spar...................................................... 5-36 

5-37 Off-nominal temperature rise rate in nose cap RCC attach clevis ........ 5-37 

5-38 Location of sensors in the LH wing wheel well...................................... 5-38 

5-39 Strain rise in lower 1040 spar cap......................................................... 5-39 

5- Outboard elevon accelerometer responses at 8:52:25 

and 8:52:31 EST (EI + 496 and 502 sec.)............................................. 5-40 

5-41 Location of supply dump and vacuum vent nozzles.............................. 5-41 

5-42 Off-nominal temperature for supply nozzle and vacuum vent nozzle.... 5-41 

5-43 First noted off-nominal aero event (Greenwich Mean Time, GMT, is 

EST + 5 hours)...................................................................................... 5-42 

5-44 Sharp change in rolling moment (GMT is EST + 5 hours)..................... 5-45 

5- Modeling results show potential area of damage and that significant 

deformation of the intermediate wing area and/or a recession in the 

lower surface are possible .................................................................... 5-46 

5-46 Temperature data in left wheel well trends up (GMT is EST + 5 

hours).................................................................................................... 5-47 

5-47 Hot gas breaches the wheel well........................................................... 5-48 

5-48 Kirtland photo........................................................................................ 5-49 

5-49 Increased wing deformation and wing recession leads to significant 

vehicle aerodynamic changes (GMT is EST + 5 hours) ........................ 5-51 

5- Wind tunnel testing configurations that match delta roll moment data.. 5-53 

vii 
Figure Page 

5-51 Wind tunnel configurations that match delta yaw moment data............ 5-54 

5-52 Wind tunnel configurations that match delta pitch moment data........... 5-54 

5-53 Delta roll for lower half and full panel RCC panel missing ................... 5-55 

5-54 Delta yaw for lower half and full panel RCC panel missing................... 5-56 

5-55 Wind tunnel testing results for missing lower carrier panel 8 

and a slot and hole through wing .......................................................... 5-57 

5-56 CFD analysis of wing deformation......................................................... 5-58 

5-57 LaRC wind tunnel testing of lower surface depressions........................ 5-59 

7-1 Details of SRB/ET forward separation bolt catcher assembly............... 7-2 

9-1 ET age for all STS missions.................................................................. 9-1 

9-2 ET age for STS-107 compared to ET age for missions with 

and without bipod foam loss.................................................................. 9-2 

9-3 ET exposure time (to weather) prelaunch for all STS missions ............ 9-3 

9-4 ET exposure time (to weather) for STS-107 compared to ET 

exposure time for missions with and without bipod foam loss............... 9-4 

9-5 Total prelaunch rainfall for all STS missions ......................................... 9-5 

9-6 Total rainfall for STS-107 compared to total rainfall for missions with 

and without bipod foam loss.................................................................. 9-6 

9-7 Average daily rainfall prelaunch for all STS missions ........................... 9-7 

9-8 Average daily rainfall prelaunch for STS-107 compared to average 

daily rainfall for missions with and without bipod foam loss .................. 9-7 

9-9 Day-of-launch average temperature for all STS missions..................... 9-8 

9-10 Prelaunch average dewpoint for all STS missions ................................ 9-9 

9-11 Prelaunch average humidity for all STS missions ................................. 9-9 

10-1 RCC components.................................................................................. 10-2 

10-2 RCC panel assembly ............................................................................ 10-3 

10-3 Typical tile installation ........................................................................... 10-4 

10-4 Wing leading edge RCC........................................................................ 10-6 

10-5 RCC cross section ................................................................................ 10-8 

10-6 RCC impact resistance ......................................................................... 10-9 

10-7 RCC corrosion process ......................................................................... 10-11 

10-8 Tee seal crack location ......................................................................... 10-12 

10-9 Tee seal cracking.................................................................................. 10-13 

10-10 RCC pinholes ........................................................................................ 10-14 

10-11 RCC impact damage............................................................................. 10-15 

11-1 History of foam changes. Blowing agent shown in parentheses, 

no changes to SLA................................................................................ 11-2 

11-2 Right (+Y) bipod ramp ........................................................................... 11-7 

11-3 Left and right bipod ramp flow differences, CFD results ....................... 11-7 

viii 
 

 

 

 

 

 

 

 

Figure Page 

11-4 ET-93 processing timeline at KSC ........................................................ 11-10 

11-5 ET LO2 slosh baffle changes  ET-14 .................................................. 11-13 

11-6 ET LO2 slosh baffle changes  ET-87 .................................................. 11-13 

11-7 Bipod foam failure modes ..................................................................... 11-14 

11-8 Schematic of bipod ramp - potential cryopumping ................................ 11-15 

11-9 Critical test results in debris assessment .............................................. 11-16 

11-10 Max bipod SLA temperatures (80 seconds MET), o

F............................ 11-17 

11-11 Defects found at critical locations.......................................................... 11-18 

11-12 Weakened plane defect found .............................................................. 11-19 

11-13 Multi-failure mode bipod TPS debris estimated by the ET Working 

Group. Note that this size and weight were not used in the 

RCC impact testing as part of the STS-107 investigation. .................... 11-20 

ix 
1-1 

1.0 INTRODUCTION 

1.1 SCOPE 

The Working Scenario is the result of a joint effort between the Columbia Accident 

Investigation Board (CAIB) and the NASA Accident Investigation Team (NAIT). This 

effort collates and documents the principal facts related to specific vehicle element 

events, timelines, and data. It also includes pertinent historical data surrounding some 

of the key vehicle element considerations in the investigation. The scenario addresses 

the chronology of vehicle events from prelaunch, Launch Countdown (LCD), 

launch/ascent, orbit, and entry as well as specific information for the External Tank (ET) 

and the Left Hand (LH) wing, including aspects of the Reinforced Carbon-Carbon 

(RCC) and attachment hardware. Vehicle processing and significant preflight events 

and milestones are also discussed. The scenario addresses technical aspects only, 

and does not address management practices or philosophies, or other organizational 

considerations. 

The chronological portion of the scenario is contained in Sections 2 through 5 of this 

report. These sections discuss the prelaunch, launch, orbit, and deorbit/entry portions 

of the Space Transportation System 107 mission (STS-107). Sections 6 through 8 

address the facts related to the Reusable Solid Rocket Motor (RSRM), Solid Rocket 

Booster (SRB), and Space Shuttle Main Engine (SSME) elements. Section 9 

addresses relevant environmental factors such as weather and age of the ET. 

Section 10 addresses the details of Columbia vehicle processing, specifically as it 

pertains to the LH wing, from the most recent Orbiter Major Maintenance (OMM) at 

Palmdale, California, through the processing for STS-107. This section also includes a 

number of design and historical considerations for the LH wing and for the RCC in 

general. Section 11 addresses several aspects of the ET, including manufacturing, 

Kennedy Space Center (KSC) processing, Thermal Protection System (TPS) 

requirements, and numerous aspects of the foam insulation. These discussions 

provide the history of the bipod foam ramp design, fabrication, testing, and address the 

details of bipod foam ramp debris failure modes, testing, and analyses. Section 12 

briefly summarizes the discussion of the working scenario. 

The data sources and types include, but are not limited to, telemetry from all flight 

phases, Modular Auxiliary Data System (MADS) data from ascent and entry, video and 

imagery from launch/ascent and entry, and launch/ascent radar. It also includes 

reconstructed aerodynamic and vehicle loads, Radar Cross Section (RCS) and 

ballistics, aero/thermal, structural, debris forensics, post-flight test data (TPS impact, 

ET/SRB bolt catcher, wind tunnel, etc.), and prelaunch processing. 
1-2 

1.2 MISSION BACKGROUND 

STS-107 was the 113th mission in the Space Shuttle program and Columbias 28th trip 

into space. These 28 missions spanned 22 years with the first being STS-1, launched 

on April 12, 1981. The STS-107 mission was a science research mission and the 

payload complement consisted of the Spacehab Double Research Module and the Fast 

Reaction Enabling Science, Technology, and Research (FREESTAR). The mission 

altitude was approximately 150 nautical miles with an inclination of 39 degrees. Figure 

1-1 depicts the STS-107 payload bay configuration. 

Figure 1-1. STS-107 payload bay configuration 

STS-107 was a Shuttle mission dedicated to investigating human physiology, fire 

suppression, and other areas of research, with 80-plus experiments representing the 

latest application of micro gravity research. The seven-member crew devoted 16 days 

on-orbit to a mixed complement of research in the space, life, and physical sciences 

including biology, physics, and chemistry. Other investigations studied factors that 

control our terrestrial climate. Participants included several NASA centers, universities, 

and education and research organizations throughout the United States, along with the 

European Space Agency (ESA), the Canadian Space Agency (CSA), the Japanese 

National Space Development Agency (NASDA), the German Aerospace Research 

Establishment (DLR), and the Israeli Space Agency. 

The primary payload carrier on STS-107 was the new SPACEHAB Research Double 

Module (RDM), doubling the volume available for, and significantly increasing the 

amount and complexity of, micro-gravity research. The RDM was a pressurized 

environment carried in Columbia's payload bay and accessible to the crew via a tunnel 

from the Shuttle's middeck. 

SPACEHAB Inc., via commercial contracts, enabled many universities, companies, and 

other government agencies to conduct important research in space on STS-107. As an 

example, the CSA conducted three bone-growth experiments and the DLR measured 

the development of the gravity-sensing organs of fish in the absence of gravity's effects. 
1-3 

One university grew ultra-pure protein crystals for drug research while another 

university tested a navigation system for future satellites. The U.S. Air Force conducted 

communications experiments. Elementary school students in Australia, China, Israel, 

Japan, Liechtenstein, and the United States studied the effects of space flight on fish, 

spiders, ants, silkworms, bees, and even inorganic crystals. 

Columbia's payload bay also housed six science payloads known as FREESTAR, 

which were mounted on a Multi-Purpose Experiment Support Structure bridge spanning 

the width of the Payload Bay (PLB). These experiments performed solar observations, 

earth science and atmospheric observations, fluid physics, and demonstrated new 

communications technology for future spacecraft. Columbia was also outfitted with an 

Extended Duration Orbiter (EDO) cryogenic pallet, which provided the required 

consumables for the long duration of the mission. 

The Mediterranean Israeli Dust Experiment (MEIDEX), managed by the Israeli Space 

Agency and Tel-Aviv University, was one of the key FREESTAR experiments. The 

primary objective of MEIDEX was to observe dust storms in the Mediterranean and the 

Atlantic coast of Africa using a radiometric camera mounted in the payload bay, which 

was remotely controlled by the ground or astronauts in the crew cabin. Secondary 

objectives of MEIDEX included observations of slant visibility, sea-surface and desertsurface reflectivity, and Transient Luminous Events, such as sprites. 
2-1 

2.0 LAUNCH COUNTDOWN 

The STS-107 Launch Countdown (LCD) was approximately 24 hours longer than a 

typical International Space Station (ISS) countdown, but within the experience base of 

other SPACELAB or SPACEHAB-type missions. There were some differences in this 

countdown as compared to most LCDs, primarily because this was not an ISS mission. 

Some of the more significant differences were due to the Extended Duration Orbiter 

(EDO) pallet that provided additional electrical power generation capability for this 

16-day science mission, and the fact that the SPACEHAB module had to receive final 

stowage late in the countdown to accommodate the live animals and other unique 

science payloads. Figure 2-1 details the STS-107 LCD overview flowchart. 

There were no significant issues during the LCD including the Power Reactants Storage 

Device (PRSD) cryogenic load or EDO planned offload operations. The crew module 

activities were in the critical path from L-48 hours (post-PRSD) through the start of 

External Tank (ET) loading due to the amount of SPACEHAB and middeck stowage 

items. The SPACEHAB stowage activities were completed approximately 90 minutes 

late due to configuration issues and the significant amount of equipment to stow. 

However, the LCD team was back on the critical path timeline by the completion of the 

communication system activation (~ L-24 hours). 

ET propellant loading was delayed by approximately 70 minutes (started at L-7 hours, 

20 minutes) due to several factors. These factors included the fuel cell 

activation/calibration running longer than planned because the time allocated for this 

activity was not adequate for the additional cryogenic tanks on the EDO pallet. Also, 

the work to resolve Interim Problem Report 110 (IPR 107V-0110), which was written to 

document a Liquid Oxygen (LO2) replenish valve problem, required access to the 

Mobile Launch Platform (MLP) and delayed preparation for ET LO2 filling operations. 

As a result of troubleshooting for another IPR (IPR 107V-0108, Front-End Processor 

(FEP) 661 Unplanned Swap), a Launch Processing System (LPS) reconfiguration of the 

active/standby launch data bus FEP power supplies was required to provide power 

redundancy for ET loading. 

LO2 and LH2 tank loading were both normal, and all loading cycles were within 

previous experience. According to postflight analysis, at the end of propellant loading 

(end of replenish), the LH2 tank load was 231,035 pounds mass (lbm), and the LO2 

tank load was 1,382,980 lbm. The postflight analysis includes corrections for the 

specific ET volume for both tanks and helium injection density corrections for the LO2 

tank. 
2-2 

Final

Load Prep

SpaceHab F

And Clo

MVAK

Setups

S/H

Pwr Up

Avionics

Checkout

Final

Load Prep

PRSD EDO Cryo Load

SSME Final Preps/TSM Secur

OMBU

PRSD Load

Preps

4 Hr

BIH

8 Hr

BIH

STS-107 LCD Timeline Summary

KSC Launch & Landing

PR K60-0055 Facility Pneumatics Dome Regulator (GSE)PR K61-4318-01 LH2 Disconnect Tower Circuit Breaker (GSE)IPR 0097 MADS FDM Control Switch Switch Out of Config

PR APU 0612 APU3 Flt Cap Dash Number (MRB)

( Defer APU 2-A0037) TVAS PIC

Failed Cal (

ress Valv espond (GSE

P H2 Detect SE)

IPR fold De

ft Booster C/

aken Docu

IPR

2 Anti Ic

(GSE)

PRSD EDO Offload

i

k 

Final

Load Prep

SpaceHab F

And Clo

MVAK

Setups

S/H

Pwr Up

Avionics

Checkout

Final

Load Prep

PRSD EDO Cryo Load

SSME Final Preps/TSM Secur

OMBU

PRSD Load

Preps

4 Hr

BIH

8 Hr

BIH

STS-107 LCD Timeline Summary

KSC Launch & Landing

PR K60-0055 Facility Pneumatics Dome Regulator (GSE)PR K61-4318-01 LH2 Disconnect Tower Circuit Breaker (GSE)IPR 0097 MADS FDM Control Switch Switch Out of Config

PR APU 0612 APU3 Flt Cap Dash Number (MRB)

( Defer APU 2-A0037) TVAS PIC

Failed Cal (

ress Valv espond (GSE

P H2 Detect SE)

IPR fold De

ft Booster C/

aken Docu

IPR

2 Anti Ic

(GSE)

PRSD EDO Offload

i

k 

Final

Load Prep

SpaceHab F

And Clo

MVAK

Setups

S/H

Pwr Up

Avionics

Checkout

Final

Load Prep

PRSD EDO Cryo Load

SSME Final Preps/TSM Secur

OMBU

PRSD Load

Preps

4 Hr

BIH

8 Hr

BIH

STS-107 LCD Timeline Summary

KSC Launch & Landing

PR K60-0055 Facility Pneumatics Dome Regulator (GSE)PR K61-4318-01 LH2 Disconnect Tower Circuit Breaker (GSE)IPR 0097 MADS FDM Control Switch Switch Out of Config

PR APU 0612 APU3 Flt Cap Dash Number (MRB)

( Defer APU 2-A0037) TVAS PIC

Failed Cal (

ress Valv espond (GSE

P H2 Detect SE)

IPR fold De

ft Booster C/

aken Docu

IPR

2 Anti Ic

(GSE)

PRSD EDO Offload

i

k 

Final

Load Prep

1/13/03 1/13/03 1/13/03 1/14/03 

2300 0700 1500 2300 0700 1500 

1/13/03 1/14/03 

2300 0700 1500 2300 0700 1500 

LCD Timeline

SSV PIC Test

Shuttle Timeline

CTS T-43 Hrs T-27Hrs T-27 Hrs T-19 Hrs T-19 Hrs

L-83:39 L-59:39 

IPR Summary 

- 

-

-

IPR-0103 E 

G

IPR-0098GSE Accum LO2 P e Failed to R )

IPR-0099 FC or Failure (G

-0100 FCP O2 Mani cay 

IPR-0101 R/H A O 

Not T  m

-0104 

Va

IPR-0102 LH - e 

Payload & C/M T melines

IPR-0096 PCM FEP Loc (GSE)

LCD Timeline

SSV PIC Test

Shuttle Timeline

CTS T-43 Hrs T-27Hrs T-27 Hrs T-19 Hrs T-19 Hrs

L-83:39 L-59:39 

IPR Summary 

- 

-

-

IPR-0103 E 

G

IPR-0098GSE Accum LO2 P e Failed to R )

IPR-0099 FC or Failure (G

-0100 FCP O2 Mani cay 

IPR-0101 R/H A O 

Not T  m

-0104 

Va

IPR-0102 LH - e 

Payload & C/M T melines

IPR-0096 PCM FEP Loc (GSE)

LCD Timeline

SSV PIC Test

Shuttle Timeline

CTS T-43 Hrs T-27Hrs T-27 Hrs T-19 Hrs T-19 Hrs

L-83:39 L-59:39 

IPR Summary 

- 

-

-

IPR-0103 E 

G

IPR-0098GSE Accum LO2 P e Failed to R )

IPR-0099 FC or Failure (G

-0100 FCP O2 Mani cay 

IPR-0101 R/H A O 

Not T  m

-0104 

Va

IPR-0102 LH - e 

Payload & C/M T melines

IPR-0096 PCM FEP Loc (GSE)

LCD Timeline 

SpaceHab F

And Clo

MVAK 

Setups 

S/H 

Pwr Up 

SSV PIC Test 

Shuttle Timeline 

Avionics 

Checkout 

Final 

Load Prep 

PRSD EDO Cryo Load 

SSME Final Preps/TSM Secur

OMBU

PRSD Load 

Preps 

4 Hr 

BIH 

8 Hr 

BIH 

CTS T-43 Hrs T-27Hrs T-27 Hrs T-19 Hrs T-19 Hrs T-19 Hrs 

L-83:39 L-59:39 

STS-107 LCD Timeline Summary 

KSC Launch & Landing 

IPR Summary 

PR K60-0055 Facility Pneumatics Dome Regulator (GSE)PR K61-4318-01 LH2 Disconnect Tower Circuit Breaker (GSE)

neumatics Dome Regulator (GSE) 

PR K61-4318-01 LH2 Disconnect Tower Circuit Breaker (GSE) IPR-0097 MADS FDM Control Switch  Switch Out of Config IPR-0097 MADS FDM Control Switch  Switch Out of Config 

PR APU-0612 APU3 Flt Cap Dash Number (MRB)

( Defer APU-2-A0037)

PR APU-0612 APU3 Flt Cap Dash Number (MRB) 

( Defer APU-2-A0037) IPR-0103 ETVAS PIC 

Failed Cal (G

IPR-0103 ETVAS PIC 

Failed Cal (G

IPR-0098GSE Accum LO2 Press Valve Failed to Respond (GSE) IPR-0098GSE Accum LO2 Press Valve Failed to Respond (GSE) 

IPR-0099 FCP H2 Detector Failure (GSE) IPR-0099 FCP H2 Detector Failure (GSE) 

IPR-0100 FCP O2 Manifold Decay IPR-0100 FCP O2 Manifold Decay 

IPR-0101 R/H Aft Booster C/O 

Not Taken  Docum

IPR-0101 R/H Aft Booster C/O 

Not Taken  Docum

IPR-0104 

Va

IPR-0104 

Va

IPR-0102 LH2 Anti-Ice 

(GSE)

IPR-0102 LH2 Anti-Ice 

(GSE) 

PRSD EDO Offload 

Payload & C/M Timelines 

IPR-0096 PCM FEP Lock (GSE) IPR-0096 PCM FEP Lock (GSE) 

Figure 2-1. STS-107 Launch Countdown (LCD) overview flowchart 
2-3 

i

R i

e 

I

Fl o

m

e

Pad Closeouts and

Configure for Launch

Final ET Load

Preps

Term.

Count

ET

Load

r ti

S/H Final

Stow & C/O

Final

Load Prep

Middeck &

A/L Closeouts

Middeck Stow &

Switchlist

C/M 

Closeout

25 Hr

19 M BIH

6

rs

STS-107 LCD Timeline Summary

KSC Launch & Landing

Final

Load Prep

2300 0700 1500 2300 0700 1500 

1/15/03 1/15/03 1/16/03 1/16/03 

2300 0700 1500 2300 0700 1500 

LCD TimelineIPR Summary

IPR-0106 L/H SRB Bond Jumper

Sleeve (MRB)

IPR-0107 LCC UPS 2 Failure (GSE)

IPR-0105 Red Vinyl Tape on Aft

ET Dome (MRB)

IPR-0114 L/H SRB Fwd Pr mary Heate

(GSE)

IPR-0108 LDB FEP 661Unplanned Swap (GSE)

IP -0109 Fuel Cell 2 Pin 4 & 5 Shar ng (MRB, deferred FCP-2-A0052)

IPR-0110 LOX Replenish Valv Secondary Pressure Switch Dropout (GIPR-0111L/H SRB Sys A Fuel so Valve Position (OMRS Waiver  WK

IPR-0112 LH2 TSM 8 exline Leak Detector-No Fl w (GSE)

IPR-0113 LOX Replenish Valve Bias High (

IPR-0115 LOX ET V/V Actuation 

(GSE EK-10320)

Payload & C/M Timelines

Com Activation

Shuttle Timelin

RSS Ret act F/C Ac vation ET Ice Insp Flt Crew to Pad

T-11 Hrs T- H rs T-0 1039E

L-35:39 L-11:39

1 Hr BIH

T-6 Hrs

2 Hr BIH

T-3 H

LCD Timeline IPR Summary 

IPR-0106 L/H SRB Bond Jumper 

Sleeve (MRB) 

IPR-0107 LCC UPS 2 Failure (GSE) 

IPR-0105 Red Vinyl Tape on Aft 

ET Dome (MRB) 

IPR-0114 L/H SRB Fwd Primary Heate

(GSE) 

IPR-0108 LDB FEP 661Unplanned Swap (GSE) 

IPR-0109 Fuel Cell 2 Pin 4 & 5 Sharing (MRB, deferred FCP-2-A0052) 

IPR-0110 LOX Replenish Valve Secondary Pressure Switch Dropout (GIPR-0111L/H SRB Sys A Fuel Iso Valve Position (OMRS Waiver  WK

IPR-0112 LH2 TSM 8 Flexline Leak Detector-No Flow (GSE) 

IPR-0113 LOX Replenish Valve Bias High (

IPR-0115 LOX ET V/V Actuation 

(GSE EK-10320) 

Payload & C/M Timelines 

Comm Activation 

Shuttle Timeline 

Pad Closeouts and 

Configure for Launch 

Final ET Load 

Preps 

Term. 

Count 

ET Ice Insp 

ET 

Load 

RSS Retract F/C Activation Flt Crew to Pad 

S/H Final 

Stow & C/O 

Middeck & 

A/L Closeouts 

Middeck Stow & 

Switchlist 

C/M 

Closeout 

25 Hr 

19 M BIH 

T-11 Hrs T-6 Hrs T-0 1039E 

L-35:39 L-11:39 

1 Hr BIH 

T-6 Hrs 

2 Hr BIH 

T-3 Hrs 

STS-107 LCD Timeline Summary 

KSC Launch & Landing 

Figure 2-1. STS-107 Launch Countdown (LCD) overview flowchart (concluded
2-4 

The post-ET load Ice Team inspection was performed with no significant issues noted 

relative to previous inspections. The inspection began at 6:15 EST and finished at 

7:45 EST. The weather conditions at the start of inspection were as follows: 

temperature 48 degrees Fahrenheit, relative humidity 97 percent, winds from 

290 degrees at 5 knots. One item of interest was noted with respect to the -Y (left) 

bipod ramp closeout area (see Figure 2-2 for vehicle coordinate system orientation). 

The Liquid Hydrogen (LH2) section of the Ice Team report noted that there were visual 

indications of frost along the bondline of the ET -Y bipod, and that the frost dissipated 

by 7:15 EST, after sunrise. The ET bipod assembly is located at the forward ET/orbiter 

attach point, and indications of frost are not unusual in this area. 

Figure 2-2. Shuttle vehicle coordinate system 

The postlaunch debris walk down was performed at the launch pad per Operations and 

Maintenance Instruction (OMI) S6444, and no unusual debris or damage was noted. 

All IPRs and Problem Reports (PRs) recorded during the LCD were evaluated and 

three were noted as worthy of discussion since they involve possible ascent debris or 

the ET. The first, IPR 107V-0102, LH2 Anti-Ice Heater Failed Set Point, was written to 

document a Ground Support Equipment (GSE) heater that did not control to the 

required set point within the specified time. The Alternating Current (AC) phasing was 

found incorrectly wired due to a previous modification. The associated power leads 

were swapped and retested on the second day of the LCD without incident. This 

system performed nominally for the remainder of the LCD. 
2-5 

The second item, IPR 107V-0105, Red Vinyl Tape on Aft ET Dome, was written to 

document a small piece of red vinyl tape (1 in. by 1.5 in.), similar to that used in Solid 

Rocket Booster (SRB) closeout activity, which was found adhered to the +Y side of the 

ET LH2 aft dome (Y-Y axis approximately 1 ft aft of station XT2058) during the L-1 day 

walk down. There was no visible Thermal Protection System (TPS) damage noted in 

the vicinity of the tape. The tape was accepted to use as-is via the Material Review 

Board (MRB) process. The rationale was that the tape was limited in size and mass, 

presented no adverse effect to the TPS performance, and was outside of the critical 

debris zone since it was located on the very bottom part of the ET. 

The third item was IPR 107V-0106, Booster Bond Jumper Sleeve Not Removed. This 

IPR was written for a part marking identification sleeve found on the systems tunnel 

ground strap 5 feet below the aft web of the right booster ET attach ring near the 

booster factory joint Xb-1577. The small plastic sleeve was accepted via Material 

Review (MR) board to use as-is, because the sleeve and strap would not be affected 

by aero heating, and if the sleeve melted or tore away during ascent, its trajectory 

would be outside the orbiter debris zone. 
3.0 LAUNCH 

3.1 INTRODUCTION 

This section discusses the launch and ascent phases of STS-107 in four separate 

sections. The first section outlines some general launch conditions and an introduction 

to the ET bipod foam impact, including photographic and debris transport analyses, as 

well as RCC impact testing and analyses. The next section discusses several key 

MADS measurement signatures from the ET foam impact timeframe. This is followed 

by a summary of launch and ascent radar, and corresponding analyses. The final 

section is a detailed discussion of several orbiter Guidance, Navigation, and Control 

(GNC) system events of interest from the ascent timeframe. These include wind shear, 

ascent loads, ET propellant slosh, and SSME and SRB nozzle positions. The 

discussion centers around possible correlation of these events with other families of 

flights, including the family of flights where it is known that ET bipod foam loss 

occurred. 

3-1 
3.2 LAUNCH DEBRIS IMPACT OBSERVATION 

3.2.1 Launch/Ascent Conditions 

Launch occurred at the Kennedy Space Center (KSC), launch pad 39A, on 

January 16, 2003, at 10:39 EST (see Figure 3-1). The weather at pad 39A, 60-foot 

level was: temperature 65 degrees Fahrenheit, relative humidity 68 percent, dew point 

59 degrees Fahrenheit, with calm winds. Figure 3-2 shows the STS-107 reconstructed 

altitude data and Figure 3-3 shows the mach number and dynamic pressure during first 

stage, prior to Solid Rocket Booster (SRB) separation, as a function of Mission Elapsed 

Time (MET). 

Figure 3-1. Launch of STS-107 at pad 39A at Kennedy Space Center 

3-2 
Mission Elapsed Time, MET 

Altitude (104 ft) 

Figure 3-2. STS-107 reconstructed altitude during first stage (prior to SRB 

separation) 

400 

Mach 

Qbar_psf 

4.0 800 

3.5 700 

3.0 600 

2.5 500 

2.0 

1.5 300 

1.0 200 

0.5 100 

0 

0 20 40 60 80 100 120 140 

0.0 

MET (secs) 

Mach Number 

Dynamic Pressure (psf) 

Figure 3-3. STS-107 flight reconstruction data for mach number and dynamic 

pressure (Q-bar) prior to SRB separation. Note that Q-bar is highest during first 

stage (prior to SRB separation), and reduces to a very small number after SRB 

separation 

3-3 
3.2.2 Launch Debris Impact Area 

Postlaunch photographic analysis determined that one major piece of foam and at least 

two minor pieces departed the External Tank (ET) left bipod ramp area approximately 

82 seconds after launch. The primary foam piece impacted Columbia in the vicinity of 

the lower left wing Reinforced Carbon-Carbon (RCC) panels 5 through 9 at 81.86 

seconds after launch. There were no indications that any of the minor pieces impacted 

the left wing based on their post-separation trajectories. The orbiter was at an altitude 

of ~65,860 feet, traveling at Mach 2.46 at time of impact. 

Several approaches were taken to assess the area of left wing damage. The efforts 

included launch video and photograph analysis, review of launch MADS data, debris 

transport analysis, forensic analysis of debris found in Texas, wire bundle burn through 

analysis, and aero/thermal modeling of the entry. The data indicate that the area of the 

highest probability of damage to the left wing was between RCC panels 5 and 9, with 

the most likely damage occurring on the lower side of RCC panel 8 or an adjacent Tee 

seal. The damage was most likely equivalent in size to a 6 to 10 inch diameter hole or 

area broken from the RCC panel or an adjacent Tee seal. Figure 3-4 shows the area of 

highest probability of wing damage and Table 3-1 shows the methods used to 

determine the damage. 

RCC Panel: 9 

8 

7 

6 

5 

Figure 3-4. Area of most likely wing damage 

3-4 
Table 3-1. Wing damage analysis methods and results 

WING DAMAGE 

ANALYSIS 

METHOD 

PREDICTED 

DAMAGE AREA 

COMMENTS DISCUSSION 

FOUND IN 

SECTION 

LAUNCH VIDEO 

AND PHOTO 

ANALYSIS 

RCC 5 through 9 Most likely area of impact was RCC 

panels 6 through 8. 

3.2 

ASCENT MADS 

DATA 

RCC 6 through 8 Unusual temperature sensor data 

observed on spar behind RCC panel 9, 

and temperature rise matches a thermal 

math model of a 10 inch diameter hole in 

RCC panel 8. 

3.3 

DEBRIS 

TRANSPORT 

ANALYSIS 

RCC 5 through 8 Most likely area of impact was RCC 

panels 6 through 8. 

3.2 

HARDWARE 

FORENSICS DATA 

RCC 8 or 9 Fragments of RCC panels 8 and 9 

showed extreme temperature indications, 

knife edge heat erosion patterns, and 

heavy amounts of slag deposited on the 

insides of those panels. 

5.3 

ENTRY MADS 

DATA 

RCC 8 or 9 First unusual indication observed during 

entry was a strain gauge behind RCC 

panel 9 (could be due to a strain behind 

adjacent panel 8). 

5.4 

WIRE BUNDLE 

BURN THROUGH 

RCC 7 through 9 Burn through from locations forward of 

panel 7 or aft of panel 9 are very unlikely 

based on sensor data loss timing. 

5.4 

ENTRY 

AERO/THERMAL 

MODELING 

RCC 8 or 9 Based on wind tunnel test results and 

CFD analysis. 

5.5 

3-5 
3.2.3 Launch Photo and Transport Analysis 

Photographic analysis of the debris impact event included participation from the 

Johnson Space Center, the Marshall Space Flight Center, the Kennedy Space Center, 

Lockheed Martin Management and Data Systems, Boeing NASA Systems, the 

Eastman Kodak Company, and the National Imaging and Mapping Agency. 

Video and Computer-Aided Design (CAD) analysis determined that the most likely 

impact location was leading edge RCC panels 6 through 8 (Figure 3-5). Due to the 

foam size, RCC panels 5 and 9 must also be included in this impact zone. The best 

estimate of the foam size, based on imagery measurements, is 21 to 27 inches long 

and 12 to 18 inches wide. The precise foam shape and thickness cannot be 

determined from the available imagery; however, a reasonable estimate is that it was a 

plate-like shape and several inches thick. The foam tumbled at a minimum rate of 18 

times per second based on the imagery, although the actual rate may never be known 

more accurately. Figure 3-6 illustrates a portion of the photographic analysis 

techniques used to determine the size of the foam. 

The most useful video analysis was performed using two cameras that are part of the 

Eastern Launch Range imaging system. Camera E212 (film), located on the Cape 

Canaveral Air Force Station, was approximately 17 miles from the orbiter at the time of 

foam impact and Camera ET208 (video), located in Cocoa Beach, Florida, was 26 

miles from the orbiter. The overall camera geometry relative to the launch pad and 

ascent flight path is shown in Figure 3-7. Camera E212 had a better view of the topside 

of the launch vehicle, while Camera ET208 had a better bottom side view. Figure 3-8 

depicts the view from each of the camera systems. A third camera, E208 (film), also 

recorded the launch but was blurred and contained no useful data for the investigation. 

There are no Launch Commit Criteria (LCC) regarding cameras, or camera views for 

ascent, for either onboard or ground. 

Figure 3-5. Multiple analyses determine foam impacted 

lower RCC panels 6 through 8 area 

3-6 
 

Figure 3-6. Photographic analysis techniques determined foam size: 

debris appears almost circular in frame 4914 and elongated in frame 4919 

E212 film 

Top-side view of 

Shuttle 

ET208 video 

Bottom-side view 

of Shuttle 26 miles 

17 miles 

Figure 3-7. Camera geometry for ascent video analysis; note that video camera 

ET208 is at same location as film camera E208 

3-7 
E212 ET208 

Figure 3-8. Orbiter view from Cameras E212 and ET208 

There is significant visual and debris trajectory information to implicate the left bipod 

ramp area as the source of debris. In Figure 3-9, the red line depicts the estimated 

foam trajectory as it moved from the bipod ramp area toward the left wing. In addition 

to locating the impact in the RCC panels 6 through 8 region, the video analysis has also 

shown that the impact was below the apex of the RCC panels since no foam or post 

impact debris was observed to traverse over the top of the wing. This is indicative of an 

impact below the wing leading edge aerodynamic stagnation line (Figure 3-10). The 

stagnation line, or dividing streamline, is the line along the leading edge of the wing 

where the airflow comes to rest; above this line, airflow moves over the upper wing 

surface and below this line, the airflow moves over the lower wing surface. 

Enhancements of the ascent video indicated there was no discernable damage to the 

orbiter wing leading edge or lower tile surface. Figure 3-11 is a sample of these video 

enhancements. The figure compares 30 pre-impact integrated video fields with 21 

post-impact integrated video fields. Based on these enhancements, photo experts 

have been unable to determine or quantify any damage to any portion of the orbiter 

vehicle as a result of the impact. 

3-8 
Figure 3-9. Multiple analyses indicate foam is from ET left bipod area. 

Red line depicts the estimated foam trajectory as it moved 

from the bipod ramp area toward the left wing. 

Figure 3-10. Video analysis shows impact is below wing 

leading edge stagnation line. Trajectories of particles 

are depicted after the impact. 

3-9 
Figure 3-11. Pre-impact vs. post-impact shows no observable damage 

within the resolution limits 

3.2.4 Debris Velocity and Size Assessment 

In addition to size and location of the foam impact, there are several other parameters 

necessary to complete the postflight analysis of possible impact damage. These 

include an estimate of the foams mass, relative velocity at impact, rotational energy, 

and the angle of impact with respect to the Shuttle wing at the point of impact. These 

parameters combine to determine the amount of impulse imparted at impact and are 

therefore critical to determine whether there was possible damage to the RCC panel, 

associated attach fitting hardware, or other leading edge structure. 

Photographic analysis was used to establish a range of relative impact velocities, from 

625 to 840 feet per second (416 to 573 miles per hour). This large uncertainty is due to 

the small number of video and film frames between release of the foam and impact with 

the wing, since the estimated time between the foam release and foam impact is only 

0.2 seconds. The predominant direction of motion is toward the aft of the orbiter along 

the X-axis, although the foam is moving slightly outboard at the time of impact with little 

to no motion in the Z-axis (see Figure 2-2 for vehicle coordinate system orientation). 

The direction of motion is from the ET bipod area toward the left wing at an angle of 2 

to 10 degrees with respect to the orbiter X-axis in the orbiter X-Y plane. The motion is 

slightly toward the wing surface at a 0 to 3 degree angle measured in the orbiter 

X-Z plane. 

Three-dimensional trajectories from the launch films and videos were refined using a 

physics-based trajectory fit that included a realistic flow field model generated using 

computational fluid dynamics (CFD) techniques. These results indicated that the 

relative velocity at impact was in the range of 775 to 820 feet per second. The CFD 

analysis used numerical methods to model the flow field around the orbiter/ET/Solid 

3-10 
Rocket Booster (SRB) stack including the SRB and Space Shuttle Main Engine (SSME) 

plumes. An example of this analysis is shown in Figure 3-12. 

The transport analysis was also used to estimate a range of sizes and corresponding 

weights for the foam, which are summarized in Table 3-2. For an impact velocity of 

820 feet per second, the estimated foam volume is approximately 1025 cubic inches 

with a weight of 1.42 pounds assuming the density of the foam was 2.4 pounds per 

cubic foot. Similarly, for a velocity of 775 feet per second, the estimated volume is 

1240 cubic inches, and the resulting weight is 1.72 pounds. Additional results produced 

with a more complex CFD model included lift forces and the unsteady rotation of the 

debris. Table 3-2 also lists the ET Working Group estimate of the bipod foam size and 

weight. This estimate was for one particular ET bipod ramp configuration and did not 

account for manufacturing variability. Thus, it was not used as the volume for the RCC 

impact testing, and more details are included in Section 11. 

Numerous factors could affect mass of the foam debris, and the exact volume and 

mass may never be known. For example, the BX-250 foam could have had a higher 

than predicted density of 2.4 pounds per cubic foot, since the density can range from 

1.8 to 2.6 lb/ft3

. Alternatively, a lower drag coefficient on the debris could also account 

for a higher ballistic number (BN). 

Figure 3-12. Sample CFD flow field with debris modeling 

3-11 
 

Table 3-2. Transport analysis and ET Working Group estimates of ET bipod 

debris size, weight, and volume 

Transport 

Analysis 

Lower Bound 

Transport 

Analysis 

Upper Bound 

ET Working 

Group Estimate 

RCC Impact 

Testing 

Target 

Ballistic 

Number (BN) 

1.2 1.45 1.0 1.45 

Velocity 

(ft/sec) 

820 775 850 775 

Volume (in3) 

@ 2.4 lb/ft3 

1025 1240 855 1200 

Weight (lbs) 

@ 2.4 lb/ft3 

1.42 1.72 1.19 1.67 

3.2.5 Impact Damage Testing and Analysis 

Analysis and experimental results were used to assess the potential for debris impact to 

damage Columbias wing leading edge. The overall concept was to replicate, to the 

greatest extent feasible, the debris impact event that occurred on Columbias left wing 

during ascent, by impacting flight-ready composite panel assemblies with a 

representative foam projectile fired from a compressed gas gun. The target panel 

assemblies had a flight history similar to that of Columbia, and were mounted on a 

support structurally equivalent to Columbias left wing. The attaching hardware and 

fittings were either flight certified, or built to Columbias drawings. BX-250 foam, without 

entrained ablator material, was used for the impacting projectile material because it 

represented the ascent event and provided a lower bound damage assessment. After 

significant study and consideration of all inputs by the NAIT and CAIB members, the 

parameters for representative impacts were established as: foam volume 1200 cubic 

inches, velocity 775 feet per second, and foam mass 1.67 pounds. 

Impact testing has been completed on full size fiberglass panels, an RCC panel 6, and 

an RCC panel 8 to obtain insight and experimental data important to the understanding 

and modeling of the response of the wing leading edge components. The RCC panel 6 

assembly was from Discovery and had flown 30 missions, and the RCC panel 8 was 

from Atlantis and had flown 27 previous missions. 

The test of the RCC composite panel assembly 6 demonstrated that a foam impact 

representative of the debris strike at 82 seconds was capable of damaging RCC 

material. A 5.5 inch crack was created, extending from a visible 3/4 inch diameter 

damage area on the outside of the panel to the rib inside the wing. The panel 6/7 Tee 

seal was also damaged with a 2.5-inch crack, and the Tee seal as well as panel 6 were 

shifted in position. In addition, a carrier panel on the upper side of the wing was 

chipped. 

3-12 
Subsequent engineering testing has demonstrated that the localized impact loads 

imposed on the panel 6 assembly would have been substantially higher with changes in 

foam impact orientation and location. These changes were included in the RCC panel 

8 assembly test and included a 30 degree clocking angle (orientation of the foam 

projectile relative to the target), a 22 degree angle relative to the impact surface, and an 

impact location lower and farther outboard relative to the panel 6 test. Impact target 

location was six inches farther down the trajectory track from the earlier tests. The test 

generated a 16 inch by 16 inch hole in the lower surface of panel 8, which is the most 

substantial damage to date in any RCC impact test. 

The exact flight damage is unknown but is believed to be bracketed by these two tests. 

The testing is important in that it confirms that the ET bipod foam can catastrophically 

damage the RCC. 

3-13 
3.3 LAUNCH MADS DATA 

There are two other indications that the foam impact occurred in the panels 6 through 8 

area. Two Modular Auxiliary Data System (MADS) lower surface pressure 

measurements behaved anomalously immediately after the time of the impact. Figure 

3-13 shows the location of these measurements along with possible areas for postimpact debris re-contact in the vicinity of the sensors. The unusual behavior of one of 

the sensors is shown in Figure 3-14. 

Possible Debris Re-contact Areas 

Figure 3-13. CFD surface flow with lower 

 left wing pressure sensors 

3-14 
 

Figure 3-14. Unusual behavior of pressure sensor V07P8074A 

Additionally, there is another MADS measurement that had an off-nominal signature 

during the ascent timeframe. The temperature sensor on the leading edge spar behind 

RCC panel 9 showed a slightly higher temperature rise than seen on any previous 

Columbia flight. Figure 3-15 shows the location of the temperature sensor behind the 

wing leading edge spar inside the wing. The slight temperature rise can be seen in 

Figure 3-16. Note that most flights show a small rise in this temperature during ascent 

due to aerodynamic heating. 

STS-107 had a 7.5 degree Fahrenheit rise that started very early during ascent (five to 

six minutes after launch). Although the data do not prove that the RCC was breached 

during ascent, the data are consistent with a possible flow path into the RCC cavity via 

damage in the RCC panels 6 through 8 area. A simplified thermal math model was 

constructed and verified with flight data from STS-5. The model was then correlated to 

the flight data from STS-107. Assuming the equivalent heating from a 10 inch diameter 

hole in RCC panel 8, this model nearly predicts both the ascent and entry temperature 

profiles for the wing leading edge spar temperature sensor. Figure 3-17 compares the 

model with the flight data for both ascent and entry. For comparison, Figure 3-18 

shows the overall heating rate of the STS-107 ascent and entry environments on RCC 

panel 9. As shown, the heating on the wing leading edge is much greater during the 

entry profile than during the ascent profile. 

3-15 
 

 

CERACHROME 

Figure 3-15. Close-out photo shows RCC panel 9 wing leading edge 

temperature measurement 

3-16 
 Figure 3-16. Three-bit rise (7.5 degrees F) on MADS wing leading edge spar 

temperature measurement (V09T9895A) during ascent 

Spar Temperature 

Ascent Heating 

Spar Temperature 

Entry Heating 

Model includes 10% Sneak Flow 

300 

67.5 

65.0 

 

Spar Temperature

Entry Heating - 10" Hole with Sneak Flow

Spar Temperature

Ascent Heating - 10" Hole with Sneak Flow

100 

150 Temperature (F) 

60.0 

62.5 Temperature (F) 

Flight 

data 

10 hole model 

200 300 400 500 600 700 800 900 

250 

Flight 

data 

10 hole model 

0 100 200 300 400 500 600 700 

200 

57.5 

50 

55.0 

Time from Launch (sec) Time from Entry Interface(sec) 

Figure 3-17. Correlation between simplified thermal math model 

and STS-107 ascent and entry flight data 

3-17 
Ascent

8:44:09 EST

Entry Interface

8:52 EST

8:59:32 EST

8:54 EST

Panel 9

Entry

81 sec.

60

50

40

30

20

10

0

Launch 200 400 E. I. 200 400 600 800 1000 1200

TIME (seconds)

HEATING RATE (Btu/ft^2-sec)

60 

50 

40 

30 

20 

10 

0 

Launch 200 400 E. I. 200 400 600 800 1000 1200 

TIME (seconds) 

Panel 9 

8:54 EST 

8:59:32 EST 

8:52 EST 

8:44:09 EST 

Entry 

Entry In 

Ascent 

terface 

81 sec. 

Figure 3-18. STS-107 ascent and entry heating environments on RCC panel 9 HEATING RATE (Btu/ft^2-sec) 

3-18 
3.4 LAUNCH AREA RADAR ANALYSIS 

STS-107 was tracked during ascent by the Eastern Range (ER) land-based C-Band 

radars, and identified debris was analyzed for time of separation, radar cross section 

(RCS), and range separation rate. In summary, the radars were unable to detect debris 

prior to SRB separation. Following SRB separation, from Launch + 150 to L + 230 

seconds (2:30 to 3:50 Mission Elapsed Time, MET), 46 items were catalogued, of which 

27 items are considered to be debris; however, the radar return signal was not of 

sufficient strength to determine the approximate shape, size, or rigidity of the debris. 

The radar analysis results are consistent with the debris analyses from previous STS 

missions. Table 3-3 lists the STS-107 catalogued radar detected events. 

The launch radar is optimized for range safety and vehicle trajectory determination, and 

not for small debris assessment. A better radar for small debris, the Multiple-Object 

Tracking Radar (MOTR) was not available for use on STS-107. The ER radars used on 

STS-107 were not designed for signature analysis and were not able to lock onto and 

track multiple targets simultaneously. Additionally, debris could remain undetected if 

the debris was emitted at a time and angle where it was shielded from the radar by the 

vehicle body. 

Detailed postlaunch radar debris analysis was performed on a regular basis until 

STS-57. There are reports available from previous flights, and typical observations 

include low strength radar returns from SRB separation to T + 300 seconds. 

In general, the strength of the radar (C-band, AN/FPQ-14 unit) return depends on 

distance to the object, size of the object, and reflectivity of the object. For the STS-107 

analyses, the distance to the objects is known but the object size and reflectivity are 

unknowns for all objects detected. As such, it was necessary to perform an 

exclusionary exercise to try to identify the objects. Some basic rules could be applied, 

such as knowledge that objects with very high separation speed are known to be part of 

the exhaust plume or products (such as SRB slag). In Table 3-3, items 30 and 31 were 

determined to be SRB slag. Moderate separation speed indicates solid objects being 

left behind. Separation rates can also be used to infer the density. There are limits to 

the debris size and shape that can be detected by the radar (see Figure 3-19 and 

Figure 3-20). 

3-19 
 

Table 3-3. STS-107 ascent radar events 

Catalog 

number 

Radar 

source 

(site no.)a 

First / last 

appearance 

(T + sec) 

MAX 

RCSb 

(dBsm)c 

Separation 

rate 

(m/sec) 

Lower 

RSRd 

(m/sec) 

Upper 

RSRd 

(m/sec)

 1 e 0.14 80.4/87 8 14 44 541 

29 e

 19.14 81.6/86.1 -1 30 36 688 

34 19.14 117/121 -15 771 10 1268 

30 f

 0.14 117.5/118 -8 1240 0 1162 

31 f 28.14 117/118.5 -11 1500 3 616 

32 28.14 118/119 -8 350 0 622 

35 19.14 121/122 -16 771 4 1286 

36 19.14 121/125 -16 372 6 1289 

37 19.14 121/123 -15 426 4 1286 

38 19.14 123/126 -14 424 1 1294 

39 19.14 124/126 -14 480 3 1297 

40 19.14 126/127 -12 490 2 1303 

41 19.14 126.5/128 -13 490 2 1306 

42 19.14 127/128 -14 476 2 1307 

43 19.14 128/129 -13 570 0 1310 

33 28.14 128/130 1 520 1 710 

44 19.14 129.5/131.5 -14 670 2 1320 

45 19.14 130/132.5 -15 371 4 1324 

46 19.14 130.5/131.5 -13 370 2 1320 

23 28.14 152/158.5 -12 187 13 947 

2 0.14 152.5/156 -10 210 9 1405 

3 0.14 152.5/162.5 -8 326 26 1405 

4 0.14 153/160 -9 229 104 1505 

24 28.14 154.5/162 -14 400 15 975 

5 0.14 156/170 -16 217 38 1465 

6 0.14 158.5/171 -17 309 34 1477 

7 0.14 164/170 -17 312 17 1493 

8 0.14 166.5/173 -21 357 19 1513 

3-20 
 

Table 3-3. STS-107 ascent radar events (concluded) 

Catalog 

number 

Radar 

source 

(site no.)a 

First / last 

appearance (T 

+ sec) 

MAX 

RCSb 

(dBsm)c 

Separation 

rate 

(m/sec) 

Lower 

RSRd 

(m/sec) 

Upper 

RSRd 

(m/sec) 

25 28.14 167/176.5 -18 221 22 1106 

9 0.14 167/184.5 -15 260 53 1557 

10 0.14 170/184.5 -15 265 44 1568 

11 0.14 174.5/180 -14 290 17 1568 

12 0.14 173/180 -16 206 21 1562 

13 0.14 174/175.1 -16 244 2 1546 

14 0.14 175.5/180 -15 180 14 1572 

15 0.14 178/180 -14 296 8 1583 

26 28.14 179/187.5 -10 884 22 1221 

16 0.14 184/190 -14 236 19 1643 

17 0.14 187/192.7 -11 649 19 1665 

27 g

 28.14 201/207 Low signal Low signal 18 1438 

28 g

 28.14 205/208.5 Low signal Low signal 11 1468 

18 g

 0.14 204.5/210 Low signal Low signal 20 1812 

19 0.14 204.5/214 -18 326 36 1829 

20 0.14 204.5/212 -17 166 28 1820 

21 0.14 206/212 -18 225 22 1827 

22 0.14 211.5/228 -17 219 66 1926 

a - Radar source: 0.14 = Patrick Air Force Base (PAFB), 19.14 = Kennedy Space Center (KSC), 

28.14 = Jonathan Dickinson Missile Tracking Annex (JDMTA) 

b - Radar cross section (RCS) 

c - Decibels relative to one square meter (dBsm) 

d - range separation rate (RSR) 

e - Objects 1 and 29 are explained as plume artifacts evident by low separation rates from vehicle 

f - Objects 30 and 31 are probably SRB slag ejection evident by high separation rates from vehicle 

g - Objects 27, 28, and 18 had indeterminable RCS and RSR due to low level of signal returns 

3-21 
20 

12 

10 

8 

6 

4 

2 

-10 

-20 

-30 

-40 

Limits of Detection for Three Simple Shapes 

18 Sphere 

16 

Plate 

Cylinder (2.5 cm diameter) 

14 

0 

0 50 100 150 200 250 300 350 400 450 500 

Time (T+ Seconds) 

Figure 3-19. Limits of dimensional detectability for three simple shapes 

Min Detectable RCS 

10 0.14 

19.14 

28.14 

0.14 debris 

0 28.14 debris 

19.14 debris 

-50 

0 40 80 120 160 200 240 280 320 360 400 440 480 

T+Time (seconds) 

Figure 3-20. Limits of radar cross section (RCS) detectability and 

measured STS-107 debris for three radar source sites 

3-22 
From Table 3-3, debris item numbers 1 and 29 appear from 80.4 - 87 seconds and 

81.6 - 86.1 seconds, respectively. This time coincides with the ET left bipod foam 

debris generation at 81.7 seconds. However, the low separation rate and relatively 

large RCS of the two radar objects indicate that they are most likely traveling with the 

vehicle and are flame (plume) artifacts. There are also several radar objects around 

the SRB separation time frame, ~126 seconds; however, the data are inadequate to 

determine the size, shape, or composition of the objects beyond that their moderate 

separation speed indicates solid objects being left behind. Some known debris objects 

at the time of SRB separation are the aft Booster Separation Motor (BSM) throat 

covers. It should be noted that the number and strength of the radar returns are typical 

as compared to previous Shuttle missions where no significant debris damage 

occurred. 

In an effort to identify the STS-107 launch debris, data was reviewed from a postSTS-27 radar calibration that was performed on several materials. These objects 

included many applicable Space Shuttle system materials, including various Orbiter 

thermal protection system tiles, various ET insulation foam types, as well as numerous 

SRB/SRM materials and potential debris sources. Table 3-4 lists the material samples 

tested for Orbiter, ET, and SRB/SRM elements. Additionally, data was used from the 

2003 Wright Patterson Air Force Base testing, including Orbiter Felt Reusable Surface 

Insulation (FRSI), High-Temperature Reusable Surface Insulation (HRSI), and HRSI 

with Room-Temperature Vulcanized (RTV) sealant and Strain Isolation Pad (SIP). 

Table 3-4. Material samples from post-STS-27 radar calibration tests 

Orbiter SRB/SRM 

Black tile MSA-1/TPS with Hypalon 

White tile MSA-2/TPS with Hypalon 

Cork with Hypalon 

ET Aft booster separation motor (BSM) cover 

PDL (closeout foam) SRM slag 

Ice plate Cork 

CPR 488 (acreage foam) K5NA 

Super Light Ablator (SLA) 561M Instafoam 

MA25 Inhibitor 

BX250 EA934 adhesive 

Instafoam Viton thermal curtain 

Quartz cloth blanket 

As a result of the testing, the minimum detectable size for each radar return for selected 

materials was determined and catalogued. These data were carefully screened and 

scrutinized, using some reasonableness tests and assumptions, in an attempt to 

identify STS-107 radar objects as Orbiter, ET, or SRB/SRM debris. 

3-23 
The radar data are inconclusive with respect to determining identity, size, or shape of 

any of the debris objects detected. The signal returns were weak and too close to radar 

noise to allow estimation of object shape. The number and strength of the returns on 

STS-107 are typical of previous Space Shuttle launches, including those where no 

debris damage occurred. 

3-24 
3.5 LAUNCH GUIDANCE NAVIGATION AND CONTROL 

Postflight analysis of the STS-107 ascent data revealed several events that were within 

the design capability of the Shuttle, but considered to be new flight experience. These 

events were reviewed in detail, primarily because they occurred prior to SRB 

separation, when the foam loss and wing impact were observed. The items considered 

new flight experience were environmental (wind relative) side-slip angle during the 

period of maximum dynamic pressure (Hi-Q), SSME yaw nozzle positions during Hi-Q, 

and SRB thrust mismatch during SRB tail-off. Other events observed during the flight 

that were not new flight experience, but were considered worthy of note included the 

presence of a negative orbiter body yaw rate at ET separation and a period of ET slosh 

during powered ascent. Each event was separated into the following categories for 

detailed study and evaluation: wind shear, predicted versus actual vehicle loads, ET 

slosh, nozzle positions, and ET separation yaw rate. 

Those parameters along with several other STS-107 ascent Guidance, Navigation, and 

Control (GNC) related points of interest were studied to determine if they were 

significant relative to the scenario. The study included integrated vehicle loads 

analysis, comparison of the STS-107 data with historical flight experience envelopes, 

and comparison of STS-107 data with specific families of flights. This section of the 

report summarizes the integrated GNC flight data review. 

3.5.1 Wind Shear, Day of Launch Wind Effects 

STS-107 experienced a wind shear during the period of maximum dynamic pressure 

starting at 57 seconds MET (Mach 1.27). The wind shear was due to a rapid change in 

the out-of-plane wind velocity of -37.7 feet per second over a 1200 foot altitude range 

starting at approximately 32,000 ft (as shown in Figure 3-21). Immediately after the 

vehicle flew through this altitude range, its side-slip angle began to increase in the 

negative direction, reaching a value of approximately 1.75 degrees at 60 seconds. 

This value of side-slip angle is a new flight experience value for MET 60 seconds (as 

shown in Figure 3-22). Post-flight data review indicates that the new flight experience 

side slip event not the result of the wind shear itself. Instead, it was the direct result of 

a difference in the L - 4:35 minutes balloon measurement, upon which orbiter guidance 

commands were updated on launch day, and the actual winds flown through by the 

orbiter during launch and ascent. Figure 3-21 highlights the difference in these two 

winds in this altitude region (a 25 foot per second increase in out-of-plane magnitude 

pre-launch compared to a 12 foot per second reduction in magnitude as experienced by 

the vehicle) 

The L - 4:35 minutes weather balloon is launched to measure atmospheric conditions at 

the launch site, which are then used as part of a standard process to update the orbiter 

guidance software to keep it within design limits and minimize loads during ascent. 

After the Day of Launch I-Load Update (DOLILU) software update but prior to launch, 

additional balloons are used to verify that the L - 4:35 minutes balloon atmospheric 

conditions are still valid and meet required tolerance checks required to commit for 

launch. All STS-107 balloon measurements taken on launch day after L - 4:35 minutes 

3-25 
satisfied the required launch commit criteria, and were subsequently verified by balloon 

data taken 15 minutes after launch. 

95% Confidence Envelope 

L - 4:30 Design Wind 

L + 15 Flight Data 

Mean 

Wind Shear 

Region of Maximum 

Side-Slip Angle 

Figure 3-21. Out-of-plane wind velocity 

3-26 
 

Mach No. 

STS-107 Reconstruction 

STS-107 Flight Data 

Max Flight Experience 

Beta Angle (deg) 

Min Flight Experience 

Figure 3-22. Side-slip angle 

3-27 
Several theories consider this wind shear event and the difference between the balloon 

data to be significant. A negative side-slip angle places the wind vector on the left side 

of the orbiter, pushing the orbiter to the right, changing the complex aerodynamic flow 

pattern characteristics in the left ET bipod area. To better understand the conditions on 

the ET left bipod, several studies were conducted. The studies (1) compared flight data 

for missions that had ET bipod foam liberation; (2) compared flight data for missions 

that flew a Light Weight Tank (LWT) in combination with ascent Performance 

Enhancements (PEs), a package of vehicle software and hardware changes designed 

to increase overall weight to orbit capability for the ISS; (3) analyzed external 

aerodynamic loads on the ET forward attach bipod ramp; and (4) studied integrated 

orbiter/ET vehicle loads. 

The flight data correlation studies indicate that a negative side-slip angle during the 

period of maximum dynamic pressure alone could not explain the liberation of the bipod 

foam. For both families of flights in the study (LWT and PE flights, and bipod foam 

liberation flights), a negative side-slip angle was seen on almost every flight. Of the 

bipod foam loss flights, STS-90 was of particular interest. STS-90 had a larger 

negative side-slip angle in Hi-Q of -2.0 degrees, when compared to STS-107, yet STS90 did not lose bipod foam. When flights that shed bipod foam were studied as one 

family of flights, STS-112 is another outlier that does not support the negative side-slip 

angle theory. During the STS-112 ascent, video coverage shows the bipod foam 

liberation occurring prior to Hi-Q, yet the negative side-slip angle on STS-112 did not 

occur on that flight until after Hi-Q. The details of the flight data correlation studies are 

summarized in Sections 3.5.6 and 3.5.7 of this report. 

To understand the aerodynamic loads on the ET forward attach bipod ramp, a CFD 

loads assessment was performed. The resulting CFD loads, discussed in more detail in 

Section 3.5.2, demonstrated that the external aerodynamic loads were below the design 

requirement. 

To measure the orbiter/ET interface loads, an integrated orbiter/ET loads assessment 

was performed. The assessment, summarized in Section 3.5.2 of this report, also 

showed all integrated vehicle loads were below design limits. 

The day-of-launch wind effects (including the noted wind shear event and associated 

negative side-slip angle) alone did not cause the ET left bipod foam loss. 

3-28 
3.5.2 Predicted/Actual Loads 

Postflight reconstruction analysis of the STS-107 ascent loads characterized the effects 

of (1) RSRM thrust mismatch, (2) ET slosh dynamics, and (3) wind shear in Hi-Q. The 

integrated effects of these events were calculated through a flexible body loads 

assessment. This loads assessment used the STS-107 reconstructed ascent 

trajectory, and included ET slosh dynamic forces. The assessment produced (1) a wing 

loads summary, (2) an ET/orbiter interface loads summary, and (3) a summary of 

external aerodynamic loads on the ET forward attach bipod ramp. 

The wing loads analysis used a flexible body structural loads assessment that was 

validated by the MADS data. The wing loads analysis used reconstructed trajectory 

parameters to generate the loads on the orbiter wings during ascent. The assessment 

demonstrated that all orbiter wing loads were 50 to 60% of their design limit, or less, 

throughout the ascent. This includes the wind shear event at 57 seconds MET, and 

subsequent side-slip angle at 60 seconds MET (as shown in Figure 3-23). 

-1200

-1000

-800

-600

-400

-200

0

200

1.00 1.10 1.20 1.30 1.40 1.50

-1200

-1000

-800

-600

-400

-200

0

200

1.00 1.10 1.20 1.30 1.40 1.50

-1200 

-1000 

-800 

-600 

-400 

-200 

0 

200 

1.00 1.10 1.20 1.30 1.40 1.50 

-1200 

-1000 

-800 

-600 

-400 

-200 

0 

200 

1.00 1.10 1.20 1.30 1.40 1.50 

Wind shear at 57 seconds 

Beta at 60 seconds (M1.38) 

Wind shear at 57 seconds 

Beta at 60 seconds 

Left Wing Loads (lb/in)

Right Wing Loads (lb/in) 

A08L (LBS/IN)

A08R (LBS/IN)

Wind shear at 57 seconds

Beta at 60 seconds (M1.38)

Wind shear at 57 seconds

Beta at 60 seconds

Left Wing Loads (lb/in)

Right Wing Loads (lb/in)

Measured Load Reconstructed Load Certification Limit A08L (LBS/IN)

A08R (LBS/IN)

Mach Mach 

Measured Load Reconstructed Load Certification Limit 

Mach 

Figure 3-23. Wing loads during wind shear and side-slip angle 

3-29 
FTO2 (lbs)

20000

15000

10000

5000

0

5000

10000

15000

20000

30 35 40 45 50 55 60 65 70 75 80 85 90 95 100 105 110 115 120 125

Slosh ~0.6 Hz oscillation

Time (seconds)

Interface

Loads, 

FTO2, 

(lbf) -

-

-

-

The ET/orbiter interface loads were generated using reconstructed trajectory 

parameters that included the effects of wind shear/crosswind, side-slip angle, and ET 

liquid oxygen (LOX) slosh. The loads analysis demonstrated that the ET forward attach 

loads were within certification requirements at all times. The wind shear event had only 

a small effect on the overall ET loads relative to the required limits (as shown in Figure 

3-24), as did the ET liquid propellant slosh (as shown in Figure 3-25). The resulting 

load from the wind shear event was of the same order magnitude as the roll maneuver 

and other first stage events prior to SRB separation. 

20000 1200 

-10000 

Interface Loads, FTO2, (lbf) 

15000 

10000 

5000 

0 

-5000 

FTO2Load 

Q-Bet a 

Response to 

Wind Shear Event 

1000 

800 

600 

400 

200 

0 

-200 

-400 

-600 

Q-beta (psf-deg) 

-15000 -800 

20 30 40 50 60 70 80 

Time (sec) 

Figure 3-24. ET interface loads at forward attachment during wind shear 

and side-slip angle. Q-beta is side-slip angle multiplied by the dynamic pressure 

and represents the side-slip angle contribution of the interface load 

20000 

15000 

10000 

Interface 5000

Loads, 

FTO2, 0 

(lbf) 

-5000 

-10000 

-15000 

-20000 

Slosh ~0.6 Hz oscillation 

30 35 40 45 50 55 60 65 70 75 80 85 90 95 100 105 110 115 120 125 

Time (seconds) 

Figure 3-25. Slosh effect on ET interface loads 

3-30 
The external aerodynamic loads on the ET forward attach bipod were analyzed using a 

CFD simulation. The simulation produced axial, side-force, and radial loads as shown 

in Figure 3-26, Figure 3-27, and Figure 3-28, respectively. The CFD assessment of the 

bipod area indicated that the external air loads were below the design limit during the 

Hi-Q region and at the time of the bipod foam liberation. 

Flexible body simulation results indicate that all vehicle elements and associated loads 

were within required limits. The reconstruction loads analyses indicate that the ascent 

environment-induced loads alone did not cause the ET bi-pod foam loss. 

Figure 3-26. ET bipod axial aerodynamic loads 

3-31 
Figure 3-27. ET bipod side-force aerodynamic loads 

Figure 3-28. ET bipod radial aerodynamic loads 

3-32 
 

3.5.3 ET Liquid Oxygen Slosh 

The STS-107 ascent data indicate a 0.6 Hz actuator oscillation frequency that peaks in 

amplitude at 55 seconds, and again at 77 seconds MET and continues through SRB 

separation. The peaks directly correlate to peaks in 0.6 Hz wind content. A 0.6 Hz 

oscillation in the Flight Control System output is of interest since it can couple with the 

ET Liquid Oxygen (LOX) slosh mode. Slosh refers to the repeated side-to-side 

movement of the center of gravity of the liquid oxidizer propellant in the external tank. 

The slosh mode frequency and amplitude cannot be measured directly through vehicle 

data. In order to determine if ET LOX slosh is present, a post-flight process of 

reviewing the vehicle SRB and SSME actuator frequency content must be conducted, 

as well as that of the launch wind. When this post-flight process was conducted for 

STS-107, it revealed that this flight experienced more than typical 0.6 Hz frequency 

content in the SRB tilt actuators with moderate content in the rock actuators. Figure 

3-29 illustrates this point with the results of the SRB left tilt actuator frequency response 

as compared to previous Columbia flight history. 

 Frequency band between 0.5 

and 0.7 Hz is of concern due 

to potential for coupling with 

ET slosh 

 Shock Response Spectra 

computed from 30 to 100 

seconds to evaluate frequency 

|•Rigid Body Modes• | |• ET Slosh Modes• | |•Higher Freq. Rigid Body 

Figure 3-29. STS-107 SRB tilt actuators experienced more than 

typical 0.6 Hz content 

3-33 
 

 

 

 

Figure 3-30 shows the relative time variation of amplitudes of the 0.6 Hz frequency 

content in wind and actuator data. The close correlation between the peaks in the 

0.6 Hz content of the right and left actuator responses and the wind dynamics indicates 

that the actuators were responding primarily to wind rather than ET LOX slosh at this 

frequency through most of first stage (prior to SRB separation). As the 0.6 Hz content 

of the wind dynamics reduces in magnitude late in first stage, the remaining 0.6 Hz 

content in the actuator response may be attributed to a combination of the remaining 

wind dynamics and low-amplitude ET LOX slosh. STS-90 shows a similar wind 

frequency content. 

Comparison Plot of 0.6 hz Content from Wind and SRB Gimbals 

 STS-107 wind frequency content was a 2 

1.8 

1.6 

1.4 

1.2 

1 

0.8 

0.6 

0.4 

0.2 

0 

STS-107 Wind 

STS-90 Wind 

STS-107 Right SRB Actuators*10 

STS-107 Left SRB Actuators*10 

max at 75 seconds and sustained past 

110 seconds 

 STS-90 was similar in peak, but decayed 

sooner (PE, OV-102, 39o Inc) 

 STS-107 SRB Gimbal response peaks

were correlated with wind peaks 

 STS-107 gimbal response at 0.6 hz 

content continued past 110 seconds as a 

result of additional wind features at 95 & 

105 seconds 

40 45 50 55 60 65 70 75 80 85 90 95 100 105 110 115 

time (sec) 

max amplitude 

Figure 3-30. STS-107 SRB gimbal responses at 0.6 Hz frequency 

correlated to wind 

In general, ET LOX slosh is due to (1) commanded vehicle attitude transients, 

(2) additional wind dynamics after the start of ET LOX slosh, and (3) the 0.2 Hz rigid 

body vehicle mode. Note that a 0.6 Hz mode is the 3rd harmonic of the 0.2 Hz 

frequency, and is therefore subject to cross-coupling, and that some wind conditions 

can naturally contain a 0.6 Hz content. 

The data from the ET LOX slosh study indicate that the flight control system operated 

as designed, and that more than adequate slosh phase stability margin existed. When 

the ET LOX slosh data is combined with the integrated vehicle loads analysis results 

(reference Figure 3-25), data indicate that the ET LOX slosh did not result in excessive 

vehicle loads at the orbiter/ET interface. 

3-34 
3.5.4 Nozzle Positions 

A review of the STS-107 ascent data identified two discrete points in time when the 

SRB and SSME nozzle positions exceeded the flight experience envelope for those 

respective times in the ascent profile. The first event occurred when the center and 

right SSME yaw deflections exceeded the previous flight experience envelope during 

the period of maximum dynamic pressure, as a result of the differences between 

predicted and actual flight wind conditions (as shown in Figure 3-31 and Figure 3-32). 

This nozzle yaw event was coincident with a wind-inducted positive lateral acceleration, 

as sensed via the body mounted accelerometer assemblies and a positive orbiter body 

yaw rate, as sensed by the orbiter rate gyro assemblies. The yaw event follows the 

period of greatest change in out-of-plane wind velocity (e.g., the wind shear previously 

shown in Figure 3-21). 

The large offset in the Center and Right SSME yaw positions at 62 seconds MET was 

the reaction of the flight control system to the wind shear event and day-of-launch wind 

differences as compared to the DOLILU design. The nozzle motion was within the 

capability of the Shuttle flight control system, and the system operated as designed. As 

discussed in Section 3.5.1, the reconstruction loads analyses indicate that the ascent 

environment-induced loads alone did not cause the ET bipod foam loss. 

The second nozzle motion event occurred when the SRB and SSME Thrust Vector 

Control (TVC) pitch and yaw deflections exceeded the previous flight experience 

envelope during SRB tail-off (as the SRB thrust diminished). The new flight experience 

envelope for the SSME and SRB nozzle positions was primarily due to (1) low Reusable 

Solid Rocket Motor (RSRM) performance that caused a time shift of the SRB tailoff 

events relative to previous flight experience, as indicated by a low burn rate shown in 

Figure 3-33, (2) a thrust mismatch between the left and right SRB caused by lower than 

normal thrust on the right SRB during tail-off, the final seconds of SRB burn (as shown 

in Figure 3-34), (3) a small bias in the left SRB pitch actuator that shifted the actuator 

positions farther toward the edge of the flight experience envelope, and (4) flight control 

trim characteristics unique to PE flights (as shown in Figure 3-35). 

3-35 
Actuator Command (deg) 

Seconds MET 

Upper Flight 

Lower Flight 

STS-107 Reconstruction 

STS-107 Flight 

Figure 3-31. Center SSME yaw position 

Upper Flight History 

Lower Flight History 

STS-107 Reconstruction 

STS-107 Flight Data 

Seconds MET 

Actuator Command (deg) 

Figure 3-32. Right SSME yaw position 

3-36 
 

SRB Separation Cue (Pc=50psia) Time Correlation with Burn Rate at Flight PMBT 

Trend Line: y = -481.57x + 298.37 

R2

 = 0.9246 

116.00 

117.00 

118.00 

119.00 

120.00 

121.00 

122.00 

123.00 

124.00 

125.00 

Separation Cue (MET) 

RSRM Flight Data 

Linear (RSRM Flight Data) 

STS-107 

0.3620 0.3640 0.3660 0.3680 0.3700 0.3720 0.3740 0.3760 0.3780 

Burn Rate at Flight PMBT 

Figure 3-33. RSRM burn rate at propellant mean bulk temperature (PMBT) 

3-37 
Figure 3-34. SRB thrust mismatch 

3-38 
Non-PE 

flights 

PE flights 

Figure 3-35. SRB nozzle position for PE flights versus non-PE flights 

The RSRM burn rate is a temperature dependent function, and are determined based 

on pre-flight tests of small samples of the actual motor propellant at a reference 

temperature. These values are then adjusted based on a Predicted Mean Bulk 

Temperature (PMBT) based on the actual weather conditions prior to launch day. For 

STS-107, the pre-flight predicted motor performance was very close to that determined 

by post flight reconstruction. A low RSRM burn rate does not affect the total impulse 

produced by the RSRM during first stage; it only affects the amount of time the RSRMs 

must burn to achieve the same level of impulse. 

The SRB thrust mismatch observed during tail-off was well within the design margin of 

the flight control system, and similar occurrences have happened numerous times 

during previous flights. 

Due to flight control gain settings unique to PE flights, PE flights have a nozzle position 

closer to zero inches deflection from 85 to 110 seconds MET. The flight data that 

coincides with the STS-107 data are all from PE flights, seen clearly in the 85 to 

110 seconds MET timeframe in Figure 3-33. The other grouping of flights in this same 

timeframe (85 to 110 seconds MET) are all non-PE flights and have larger pitch nozzle 

deflections. 

To examine if SRB thrust mismatch during tail-off contributed to the loss of the ET 

bipod foam, several studies were conducted. The studies included data correlation of 

3-39 
(1) flights that used LWT and PEs, and (2) flights that shed ET left bipod foam. The 

data correlation showed that for both families of flights, SRB thrust mismatches were 

observed on the majority of flights. The only flights to not have significant SRB thrust 

mismatches near SRB tail-off were STS-87 and STS-90. The study of the two families 

of flights are summarized in Sections 3.5.6 and 3.5.7 of this report. 

The data indicate that the SRB thrust mismatch on STS-107 was a direct result of SRB 

burn rate differences between the left and right SRB. The thrust mismatch observed on 

STS-107 and the new flight experience nozzle positioning occurred after the foam 

shedding event. The SRB thrust mismatch occurred on the majority of flights in both 

families of flights, including those that did not shed foam. 

3.5.5 ET Separation Yaw Rate 

A higher than typical negative yaw rate was observed at ET separation during STS-107. 

The yaw rate, shown in Figure 3-36, was approximately -0.12 deg/sec and near the 

edge of the flight experience envelope. The negative yaw rate is noteworthy because it 

does not correspond to the flight control systems thruster activity, known vent forces, or 

any other explained mission activity. Furthermore, the rate appears at the time of 

physical separation between the orbiter and external tank, which indicates that it is 

related to the structural release between the two objects. Although the negative yaw 

rate was unusual, it was well within the design and certification envelope for ET 

separation. This rate was also well within the flight limits (± 0.7 deg/sec) for ET 

separation to occur. 

Flight Data 

Post-Flight Reconstruction 

w/ moments 

Post-Flight Reconstruction 

Yaw Rate (deg/sec) 

Seconds MET 

Figure 3-36. ET separation yaw rate 

A fault tree analysis narrowed down the cause of the negative yaw rate to a release of 

strain energy at ET separation due to either (1) a misalignment of the orbiter and 

external tank at structural mating, or (2) a build up of strain energy in the ET structure 

3-40 
 

and associated orbiter attachment strut areas due to thermal differences. STS-107 was 

one of six flights to have a negative yaw rate at ET separation. The other flights with 

negative yaw rates at ET separation were STS-2, STS-70, STS-80, STS-92, and 

STS-98. None of these flights are known to have had bipod foam loss. The ET sep 

yaw rate on STS-2 was identical to that on STS-107, within the 0.02 deg/sec accuracy 

limits of the sensors and related data and signal noise. 

Data indicate that the yaw rate at ET separation did not result in re-contact between the 

orbiter and the ET after separation. This observation is based on high rate telemetry 

(25 Hertz) orbiter body rate data, and MADS accelerometer data analyzed post-flight. 

The yaw rate was within the expected range of vehicle body rates when taking into 

account all known error sources, including rate sensor noise. The data indicates that 

no correlation exists between observed yaw rate at ET separation and bipod ramp foam 

loss. 

3.5.6 Data Correlation of Flights that Used a LWT and PEs 

To determine if any of the items considered new flight experience were unique to the 

use of Light Weight (LWT) ET and Performance Enhancements (PEs), an evaluation 

was performed to compare the STS-107 flight data to other flights using LWT and PEs. 

The STS-107 data was compared to flights STS-87, STS-89, STS-90, and STS-99 (as 

shown in Table 3-5). 

Table 3-5. LWT and PE flights 

Flight Tank 

type ET # Orbiter Flight software Inclination Launch date 

STS-87 LWT 89 Columbia OI-26A 28.45 deg 11/19/97 

STS-89 LWT 90 Endeavour OI-26A 51.6 deg 1/22/98 

STS-90 LWT 91 Columbia OI-26B 39.0 deg 4/17/98 

STS-99 LWT 92 Endeavor OI-27 57.0 deg 2/11/00 

STS-107 LWT 93 Columbia OI-29 39.0 deg 1/16/03 

LWT and PEs were also used on STS-85 and STS-86. Neither flight was included in 

the LWT and PE flight data correlation study because the flights were the first to use 

the new PE flight software, and as such had very few of the PEs active. One of the 

most significant PEs not active for STS-85 and STS-86 was the first stage pitch parallel 

change. Not having the first stage pitch parallel PE in place resulted in STS-85 and 

STS-86 being outliers when compared to the other five flights, due to significantly 

different SRB and SSME nozzle positioning during first stage (as previously shown in 

Figure 3-35). 

The LWT and PE flight data correlation study compared the STS-87, STS-89, STS-90, 

STS-99, and STS-107 flight data for parameters that were considered new flight 

experience for STS-107. The LWT and PE flight data correlation included a 

3-41 
comparison of environmental side-slip angle during Hi-Q, SSME yaw position during 

Hi-Q, SRB thrust mismatch during Hi-Q, ET separation yaw rate during Hi-Q, and ET 

slosh. 

The LWT and PE flight correlation study indicated that negative side-slip angles of 

-0.75 degrees or more occurred on all flights, including STS-90, which had the second 

largest side-slip angle of any flight in Hi-Q of -2.0 degrees. Evaluation of SSME yaw 

positions during Hi-Q indicated that only STS-90 had a similar signature. The STS-90 

SSME yaw was primarily due to a large wind shear on that flight. Evaluation of SRB 

thrust mismatch shows a similar thrust mismatch and corresponding SRB and SSME 

TVC gimbal activity on STS-89 and STS-99 only. Within this family of flights, the 

negative yaw rate at ET separation and ET slosh characteristics were only observed on 

STS-107. 

Of all of the flights studied, STS-90 and STS-107 were the most similar. Both flights 

were flown on Columbia, on a 39.0-degree inclination trajectory, used LWT and PEs, 

were daytime launches, and had a SPACEHAB module as the primary payload. 

Furthermore, STS-90 and STS-107 flew through a large wind shear during the Hi-Q 

region. 

The data is inconclusive as to whether ascent GNC parameters/events correlated for 

flights using a combination of LWT and PEs. 

3.5.7 Data Correlation of Flights with ET Bipod Foam Liberation 

To examine if any of the items considered new flight experience for STS-107 

contributed to the ET bipod foam liberation, a flight data correlation study was 

performed for all flights known to have lost ET bipod foam during ascent. The flights 

compared to STS-107 below included STS-7, STS-32, STS-50, STS-52, STS-62, and 

STS-112 (as summarized in Table 3-6). These are the only flights to have definitive 

photographic information to show ET bipod foam loss between liftoff and ET separation. 

An estimate of the ET bipod foam volume obtained from this photographic evaluation 

can also be found in Table 3-6. It should be noted that STS-32 is under review as a 

flight that lost ET bipod foam. It is known that STS-32 lost ET foam, but it is not clear at 

this time if it was acreage foam or bipod foam. 

The ET bipod foam liberation flights were compared for parameters that were 

considered new flight experience for STS-107. The data correlation study included a 

comparison of environmental side-slip angle during Hi-Q, SSME yaw position during 

Hi-Q, SRB thrust mismatch during thrust tail-off, ET separation yaw rate, and ET slosh. 

The negative side-slip angle of -1.5 degrees or more occurred on all flights in this 

family, and STS-62 had the largest side-slip angle of any flight in first stage (prior to 

SRB separation) at -2.5 degrees. Evaluation of SSME yaw positions in first stage show 

similar signatures occurred on STS-50, STS-52, and STS-62 (all primarily due to large 

wind shears). Evaluation shows that similar thrust mismatch and corresponding SRB 

and SSME TVC gimbal activity occurred on all flights in this family. STS-107 is the only 

3-42 
flight in this family to have a negative yaw rate at ET separation. The ET slosh 

characteristic was present on STS-32, STS-50, STS-52, STS-62, and STS-107. 

In summary, the negative side-slip angle and SRB thrust mismatch were evident for all 

flights on which ET bipod foam loss was observed. For other parameters within this 

family of flights, no correlations are evident. It is noteworthy that five of the seven 

flights in the foam loss family were Columbia missions, all with the ET slosh 

characteristic. Finally, the data are inconclusive as to whether any of the new flight 

experience parameters (individually, or in some combination) by themselves caused 

bipod foam loss. 

Table 3-6. STS flights with ET left bipod foam liberation 

FLIGHT STS-7 STS-32 STS-50 STS-52 STS-62 STS-112 STS-107 

BIPOD FOAM 

LIBERATED 

ON ASCENT 

YES Under 

Review 

YES YES YES YES YES 

APPROX. 

DEBRIS 

VOLUME 

(cu. in.) 

404 295 707 15 1 202 1200 

ET # 06 25 45 55 62 115 93 

ET TYPE SWT LWT LWT LWT LWT SLWT LWT 

ORBITER Challenger Columbia Columbia Columbia Columbia Atlantis Columbia 

INCLINATION 28.45 deg 28.45 deg 28.45 deg 28.45 deg 39.0 deg 51.6 deg 39.0 deg 

LAUNCH 

DATE 

06/18/83 01/09/90 06/25/92 10/22/92 03/04/94 10/07/02 01/16/03 

LAUNCH 

TIME 

(LOCAL) 

07:33:00 

AM EDT 

07:35:00 

AM EST 

12:12:23 

PM EDT 

1:09:39 

PM EDT 

08:53:00 

AM EST 

3:46:00 

PM EDT 

10:39:00 

AM EDT 

SIDE-SLIP 

ANGLE 

DURING 

FIRST STAGE 

YES YES YES YES YES YES YES 

NOZZLE YAW 

DURING 

FIRST STAGE 

No No YES YES YES No YES 

SRB THRUST 

MISMATCH 

YES YES YES YES YES YES YES 

ET SLOSH No YES YES YES YES No YES 

NEGATIVE 

YAW RATE 

AT ET SEP 

No No No No No No YES 

3-43 
4.0 ORBIT 

4.1 INTRODUCTION 

While Columbia was on-orbit, there was no indication of damage from either the ascent 

foam impact or a micrometeoroid/orbital debris (MMOD) hypervelocity debris impact 

based on orbiter telemetry, crew downlinked video and still photography, or crew reports. 

Multiple comprehensive postflight reviews of the same data indicated that there was 

nothing unusual with any of Columbias systems or structure. This included a detailed 

review of orbiter Inertial Measurement Unit (IMU) accelerometer, body rates, and jet 

firing data to determine if there were indications of an orbital debris hypervelocity impact. 

The results of this analysis show that there were no indications of an orbital debris 

impact, although there are several unexplained events. Data from an additional 

accelerometer package, known as Space Acceleration Measurement System (SAMS), 

was used to determine if this more sensitive system was able to detect any unusual 

activity during these timeframes. Details of the orbital debris analysis can be found in 

Section 4.2 and the flight day 2 debris event will be discussed in Section 4.3. 

4.2 ORBITAL DEBRIS 

4.2.1 Orbital Debris Risk Assessment 

There were multiple payload constraints on this mission, which resulted in 239 attitude 

maneuvers, or orientation changes. For each Shuttle mission the complement of 

attitude maneuvers is analyzed for orbital debris risk of a critical penetration due to an 

on orbit hypervelocity impact. This same analysis, performed post-flight, determined 

that the probability of no critical penetration was 0.9972, which is well below the 

guideline for critical penetrations. The analysis also included specifics for critical 

penetrations of the left wing. The results show that the overall probability for no critical 

penetration is 0.9996 for the entire left wing and 0.9999 for the left wing leading edge 

RCC. 

4.2.2 Micrometeoroid or Orbital Debris Detection 

Postflight, a NASA JSC team consisting of members from Mission Operations, 

Engineering, and Space and Life Sciences with the support of Draper Labs, participated 

in an effort to use downlisted data to identify any external forces or torques that could 

be correlated with an MMOD impact. This task was divided into four different areas: 

1. Build an inclusive, detailed activity timeline that includes all known Shuttle and 

payload events (venting, waste control system activities, LiOH canister change out, 

payload bay door operations, and SPACEHAB systems operations) that would 

cause attitude and rate errors or momentum changes detectable by the orbiter 

systems. 

2. Review the orbiter IMU rate data for net changes in angular momentum, which 

would be indicative of an MMOD strike. 

4-1 
 

3. Screen the 20,000 plus orbiter Vernier Reaction Control System (VRCS) jet firings 

with an algorithm to determine whether or not each firing was due to the control 

system response to normal attitude changes or disturbances, or in response to an 

MMOD strike. 

4. Examine SAMS payload experiment data for potential signs of an MMOD strike. 

4.2.2.1 IMU Rate Data Review 

This study reviewed all orbiter data from various sensors and systems. The only data 

useful for this study were the orbiter body axis rate data, which are derived from IMU 

attitude data by the Guidance Navigation and Control (GNC) flight software. This 

analysis assumed rigid body dynamics; flexural response was covered in the SAMS 

data analysis (see below). 

The entire orbit portion of the mission, from the orbit transition (1 hour MET) to four 

hours prior to the deorbit, was examined. A total of 238 events of interest were 

identified which required further examination. All but 13 of these events were correlated 

to either a known forcing function, or the signature did not match the expected dynamic 

response of an externally applied impulse (MMOD strike). The remaining 13 

unexplained events were analyzed in significantly greater detail. 

Additional analysis included the evaluation of the rate transients and a time integration 

of the change in angular momentum across the event of interest. The guiding principle 

of this analysis is that unless there is an external force or torque applied to the vehicle, 

conservation of angular momentum will always apply. This study resulted in the 

elimination 10 of the 13 events that did not fit the expected response for an externally 

applied impulse. One event was inconclusive due to the low resolution of the data, and 

the remaining two events have the potential to be caused by an MMOD strike; however, 

other causes are also possible (unknown venting, etc.). The orbiter rate data cannot be 

used to explicitly determine mass, velocity, or point of impact of an MMOD object. 

Table 4-1 provides an overview of the original 13 events of interest. 

4-2 
 

 

 

 

 

 

Table 4-1. Summary of analysis of 13 rate events 

Event EST* 

(Day:hour:min:sec) 

MET* 

(Day:hour:min:sec) 

External torque, 

unknown 

venting, or 

potential 

MMOD** strike 

Angular 

momentum 

conserved 

(crew motion, 

other, or 

unknown) 

H2O 

dump 

Inconclusive 

1 18/11:45:00 2/01:06:00 X 

2 19/12:45:50 3/02:06:50 X 

3 19/20:02:20 3/09:23:20 X 

4 19/21:31:30 3/10:52:30 X 

5 24/16:45:10 8/06:06:10 X 

6 25/04:19:20 8/17:40:20 X (possible) 

7 25/05:08:00 8/18:29:00 X 

8 26/03:53:20 9/17:14:20 X 

9 29/00:02:00 12/13:23:00 X 

10 29/15:48:30 13/05:09:30 X 

11 29/17:40:10 13/07:01:10 X 

12 31/11:07:00 15/00:28:00 X 

13 32/02:02:30 15/15:23:30 X 

* Times are approximate 

** Micrometeorite or orbital debris 

4.2.2.2 Lower Bound of IMU MMOD Detection Threshold 

Two separate techniques were evaluated to attempt to bound the lowest MMOD mass 

and velocity that could be detected using the orbiter IMU data. The first used measured 

angular rate data, while the second used the accelerometers to measure a change in 

velocity. 

There were two assumptions for this angular rate analysis. First, the lowest value of 

angular rate change that can be detected by the Shuttle IMUs is 0.002 deg/sec, based 

on an evaluation of body rates and engineering judgment. Second, to bound the 

minimum mass of an MMOD object, the efficiency of transfer of linear momentum of the 

striking object was assumed to be 100% with optimal geometry. The resulting transfer 

of the linear momentum is a change of orbiter angular momentum. 

The bounding of the lower limit of the linear momentum and/or mass of a potential 

strike object is not a one-dimensional exercise. Several assumptions must be made to 

perform this analysis. Strike location on the orbiter is significant. For a fixed orbiter rate 

change from a strike, the radius from the orbiter center of gravity (CG) to the strike 

location is inversely proportional to the linear momentum of the striking object. Also, 

once the linear momentum of the striking object is defined, the mass of the object is 

inversely proportional to the velocity. The examples shown in Table 4-2 are three of 

many possible solutions; however, they have been selected to be representative of a 

strike location roughly associated with the main landing gear door and forward through 

the leading edge of the wing. 

4-3 
Table 4-2. Summary of analysis of the lower bound of MMOD 

(based on body rate data) 

Body 

Axis 

Body 

Rate 

(degrees 

per 

second) 

Angular 

Momentum 

(slug*ft2

/sec) 

Assumed Strike Location of 

MMOD 

Assumed 

Velocity 

of MMOD 

(nmi/sec) 

Lower Bound of the 

Mass of MMOD 

(Assumes optimal 

geometry & 100% 

momentum transfer) 

(gram) 

Roll 0.002 36 Outside edge of the main 

landing gear door, or about 

14 ft Y c.g. offset. 

5 1 

Pitch 0.002 273 The forward most portion of 

the wing structure, or about 

23 ft in front of the X c.g. 

5 6 

Yaw 0.002 285 The forward most portion of 

the wing structure, or about 

23 ft in front of the X c.g. 

5 6 

The lowest value of velocity change that can be detected is 0.0344 feet per second 

based on the minimum integrated acceleration (velocity) pulse size from the IMUs. In 

order to determine the minimum possible detectable MMOD mass for this orbiter 

velocity detection capability, the following assumptions were used: conservation of 

linear momentum, a 100% momentum transfer from the striking object, object impact at 

the orbiter center of mass, and a relative debris velocity of 5 nmi/sec. Based on these 

assumptions, the lowest detectable MMOD mass is 127 grams. From this momentum 

analysis, it is apparent that the orbiter being struck by an approximately one-quarter 

pound object (at 5 mi/sec) assuming 100% momentum transfer would most likely be 

noticeable by the crew. Therefore, IMU accelerometers are not considered of 

significant value in the search for an MMOD strike on-orbit. 

4.2.2.3 Vernier Thruster Firing Algorithm 

The review of orbiter data accounted for momentum changes due to VRCS jet firings. 

However, the possibility existed that a debris strike with enough energy or striking the 

orbiter at the right time could have caused the On-Orbit Digital Auto-Pilot (DAP) to 

command a jet firing due to a rate deadband exceedance. 

The On-Orbit DAP will command jets to fire to maintain attitude errors within attitude 

deadbands and rate errors within rate deadbands. During periods of attitude hold, the 

majority of jet firings are due to the attitude deadband. Rate deadband firings typically 

occur at the beginning and end of attitude maneuvers, and during maneuvers due to 

changes in the desired rate. Figure 4-1 depicts changes in the vehicle rates due to jet 

firings and normal gravity gradient forces. 

4-4 
 

Jet Firings

Slow Rate

From

Gravity Gr

Response to Rate Change from Venting

Response to Rate Change from Jet Firing

Rate Change from Jet Firing Rate Change from Jet Firing 

Response to Rate Change from Jet Firing 

Response to Rate Change from Venting 

Change

adient

Rate Change from Venting

Jet Firings Gravity Gr

Slow Rate Change

From 

adient 

Rate Change from Venting 

deg/sec 

Figure 4-1. Jet firing example for vehicle rates 

Analysis was undertaken to examine every jet firing and determine the cause of the 

firing. An algorithm was built to screen all nominal attitude deadband related firings. 

The remaining firings were examined to determine cause. 

The algorithm assumed a VRCS jet was firing any time the downlist (telemetry) 

indicated a command to fire any one or more of the six VRCS jets. Also, instances of 

VRCS firings when the DAP attitude error (downlisted at 1 Hz) was less than 95% of the 

estimated attitude deadband were flagged for further investigation. 

A total of 747 jet firings out of 28,779 were identified by the screening process for 

further investigation. Of these, 19 were due to faulty driver indications (data hits). 

These were verified via no change in slope of attitude rates, DAP attitude errors, and 

the six vernier jet fuel and oxidizer injector temperatures. 

The remainder were examined and determined to be caused by (1) rate limit firings at 

the start and stop of attitude maneuvers, (2) rate limit firings that occurred during 

maneuvers due to changes in the desired rate, and (3) attitude deadband firings not 

screened. The final result was that there were no unexplainable jet firings in the STS107 on-orbit data. 

4.2.2.4 SAMS Data Analysis 

After a review of the available payload sensors, it was determined that the SAMS data 

package would be the only suitable sensor that could provide additional data to aid in 

the detection of an MMOD strike. SAMS provides tri-axial accelerometers to measure 

the vibratory and transient portion of the microgravity environment. Those vibratory and 

transient accelerations are composed of disturbances that originate in STS equipment, 

scientific experiment, and crew operations. The vibratory/transient accelerations are on 

the order of milli-gs and are sampled at 100 Hz. While the Shuttle IMUs are designed 

to measure the rigid body accelerations and attitude, SAMS measures the 

vibratory/transient portion of the micro-gravity environment. The vibratory portion is the 

4-5 
dominant part of the SAMS data. Three SAMS sensor sets were aboard STS-107; 

however, only one had data that was downlinked during the flight. This sensor was 

located in the SPACEHAB Module near the Combustion Module 2 experiment. 

SAMS data was used to support the aforementioned IMU rate data review. Anomalies 

in rates from manual review of orbiter body rates were compared to SAMS 

measurements to help identify sources. Figure 4-2 provides a sample plot of SAMS 

data and the response signature to an IMU alignment and the Enhanced Orbiter 

Refrigerator/Freezer (EORF) operation, as well as downlisted telemetry in the 

Operational Data Retrieval Complex (ODRC) system. SAMS data was also scanned for 

large transients to identify potential strikes (the assumption is that a hypervelocity 

impact would ring the structure). Various frequencies from nonstructural items were 

identified, so that they could be filtered out of the data. Figure 4-3 shows the frequency 

response of several items such as the EORF refrigerator and the Ku-band antenna. A 

detailed structural model that identifies frequencies of primary structure was developed. 

This model was used to screen for vibrational transients associated with orbiter wing 

strikes. 

Figure 4-2. Sample data from SAMS and ODRC 

4-6 
Figure 4-3. SAMS data frequency content 

The review of SAMS data has not uncovered any events that could be correlated to a 

hypervelocity debris strike from micrometeoroids or orbital debris. 

4.3 FLIGHT DAY 2 EVENT 

4.3.1 Radar Tracking of Flight Day 2 Object 

Air Force Space Command post-flight evaluation of radar tracking data indicated an 

object in the vicinity of the orbiter on flight day 2. The object remained on-orbit for 

approximately two and a half days, and reentered the atmosphere. Multiple 

government agencies participated in complex post-mission analysis of this object. 

These agencies include the Department of Defense Columbia Investigation Support 

Team, United States Strategic Command, Air Force Research Labs (AFRL) at WrightPatterson Air Force Base, Air Force Space Command (AFSPC), Lincoln Laboratory at 

the Massachusetts Institute of Technology, and NASAs Johnson and Kennedy Space 

Centers. 

The AFSPC Space Analysis Center estimated the departure time for the object was 

January 17, between 10:00 and 11:15 EST. Because there was no direct radar 

observation at the exact time of departure from the orbiter, analysis indicated that the 

most likely window of departure was between 10:30 EST and 11:00 EST. The analysis 

was complicated by the high drag profile, making it difficult to determine the precise 

time when the object left the vicinity of the orbiter. 

The calculated departure velocity was relatively low and was estimated to be 0.7 to 

3.4 miles per hour with the lower velocity being more likely. An exact departure 

4-7 
direction relative to the orbiter could not be determined. Multiple ground sensors 

including Eglin Air Force Base (AFB), Beale AFB, Cape Cod Air Force Station (AFS), 

and the Navy Space Surveillance fence radar tracked the object. The object reentered 

the atmosphere on January 19 between 20:45 EST and 23:45 EST. Figure 4-4 depicts 

the tracking of the object including various sensor passes. 

Shuttle Orbit 

2003-003B 

(FD2 Object 

aka 90626) 

-1400 

-1200 

-1000 

-800 

-600 

-400 

-200 

0 

200 

DELTA TIME (SECONDS) . 

BEALE 

CAPE COD 

EGLIN 

NAV 

SHUTTLE 

2003-003B Separation 

Track grouping shows CONUS overflights 

17-00:00:00 18-00:00:00 19-00:00:00 20-00:00:00 21-00:00:00 

GMT (DAY-HH:MM:SS) 

Figure 4-4. Tracking of flight day 2 object through 

various sensor passes 

Based on the observed radar cross sections, the object appeared initially to have a 

minimal to no tumble/rotation rate, but it gradually developed a rate over the next two 

days. During a Cape Cod AFS sensor pass on January 18 at 15:29 EST, the 

tumble/rotation rate had a period of seven seconds. Later, on January 19 at 10:39 EST 

during another Cape Cod AFS pass, the apparent tumble/rotation rate had increased 

and the period was approximately three seconds. Figure 4-5 depicts the 

tumble/rotation rates during the timeframe that the object remained in orbit. The exact 

physical size and mass of the object are unknown, although it appeared to be a 

lightweight piece based on the observed ballistic coefficient. 

4-8 
 

 

Cape Cod track on day 17, Cape Cod track on day 18, 0.70 0.5000 

0.4500 0.60 

Radar Cross Section (m2 ) 

RADAR CROSS SECTION (m2 ) 

Slow or no tumble 

0 10 20 30 40 50 60 70 

Time (Sec.) 

Tumble Period: ~7 sec. 

0.4000 

0.50 0.3500 

0.3000 0.40 

0.30 

0.20 

0.2500 

0.2000 

0.1500 

0.1000 

0.10 0.0500 

0.0000 0.00 

0 5 10 15 20 25 

TIME (sec.) ) 

0.8 

 Piece started to rapidly 0.7 

tumble/rotate between day 17 and 0.6 

day 18, rate increases on day 19 

 RCS varies from ~0.1 to ~0.7 m2 

RCS (m2) 0.5 

0.4 

0.3 

0.2 

0.1 

Tumble Period: ~3 sec. 

Cape Cod track on day 19, 0

0.0 5.0 10.0 15.0 20.0 25.0 30.0 

Time (Sec.) 

Figure 4-5. On-orbit RCS shows increased tumble/rotation rate over time 

4.3.2 Analysis of Mechanisms for Object Release 

The timeframe of estimated departure has been reviewed in detail. There were no 

unusual crew events, telemetry data, or accelerations in orbiter or payload 

accelerometer data that can account for the ejection of an object matching this 

description. SAMS, IMU, and jet firing data indicate that there was no orbital debris 

impact during the timeframe. Additional reviews indicate that no external mechanical 

systems such as the radiators or FREESTAR experiment canister doors were active 

during the time of interest. The port radiator was deployed on January 16 at 13:47 EST 

and was stowed on January 19 at 17:39 EST. 

Crew commentary in the air-to-ground voice transmission during this window was 

routine and there was no mention of an object being observed. There was no video 

downlink at the time of interest, but subsequent surveys of downlinked video and still 

imagery did not reveal any items missing from the payload bay or visible exterior of the 

vehicle. 

The orbiter did not perform any translational maneuvers during this timeframe. Two 

attitude maneuvers or orientation changes were accomplished using the small, 24 lb 

vernier attitude control thrusters. The first maneuver was a 48-degree yaw maneuver to 

a biased tail forward bay-to-earth attitude that occurred from 09:42 to 09:46 EST. Near 

the window of estimated departure, there was a maneuver back to the bay-to-earth tail 

forward attitude from 10:17 to 10:21 EST. 

4-9 
A manual fuel cell purge was performed later at 11:25 EST, outside the window of 

probable object departure. The first orbiter water dump occurred approximately two 

days after this event. Table 4-3 lists the chronology of relevant events. 

Table 4-3. Chronology of events related to flight day 2 object 

Date/Time (EST hh:mm) EVENT 

January 17 9:42 to 9:46 Attitude maneuver to a biased tail forward bay to 

earth attitude (biased -ZLV, -XVV ) 

January 17 10:17 to 10:21 Maneuver back to the bay to earth tail forward 

attitude (ZLV, -XVV) 

January 17 10:30 to 11:00 Best estimate of object departure window 

January 17 11:25 Manual fuel cell purge 

January 19 16:39 First water dump 

January 19 20:45 to 23:45 Object re-enters atmosphere 

Data indicate that in the timeframe of the object departure there were no unusual forces 

or mechanisms for liberating the debris that were not also present prior to this 

timeframe. The orbiter had encountered a more severe loading environment during the 

ascent and post-insertion timeframe than on-orbit as depicted in Table 4-4. The orbiter 

was using the large 870 lb primary reaction control system thrusters for attitude 

maneuvers until the small 24 lb vernier thrusters were activated about two hours after 

launch. One theory is that 16 orbits of thermal cycling (day/night transitions) caused 

stored energy from an object in the payload bay or on the orbiter structure to be 

released. Another theory is that attitude maneuvers in this timeframe could have 

assisted the object in obtaining the opening rate from the orbiter. The data is 

inconclusive in determining the cause of the object departing on flight day 2. 

4-10 
 

 

 

Table 4-4. Summary of nominal launch day events 

MET 

(hh:mm:ss) 

EST 

(hh:mm:ss) 

EVENT 

0:00:00 10:39:00 Columbia launch 

0:01:21.7 10:40:21.7 Foam departs ET left bipod ramp 

0:01:21.9 10:40:21.9 Foam impacts orbiter left wing RCC panels 6 through 8 

0:02:06.6 10:41:06.6 SRB separation 

0:07:23.6 10:46:23.6 3-G throttling of Shuttle Main Engines 

0:08:22.5 10:47:22.5 Main Engine cutoff command 

0:08:33 10:47:33 Zero thrust 

0:08:43.7 10:47:43.7 ET separation translation 

0:08:57 10:47:57 Crew +X translation for ET photography 

0:10:24-

0:12:24 

10:49:24-

10:51:24 

Main Propulsion System dump 

0:13:44-

0:14:33 

10:52:44-

10:53:33 

Manual pitch maneuver for ET photography 

0:29:52-

0:34:24 

11:08:52-

11:13:24 

Attitude maneuver to Orbital Maneuvering System 

(OMS)-2 burn attitude 

0:41:24-

0:43:24 

11:20:24-

11:22:24 

OMS-2 burn using left and right OMS engines 

~01:15:00 11:54:00 Attitude maneuver to payload bay door opening 

~0:02:00 12:39 Configure for vernier attitude control (six small, 

24 lb thrusters) 

4.3.3 Radar Cross Section and Ballistics Testing 

In addition to the careful inspection of downlinked orbiter payload bay video and still 

photography, radar testing and ballistics analysis of various thermal protection system 

items and thermal blankets have been conducted in an attempt to identify the flight day 

2 object. The AFRL Advanced Compact Range Facility at Wright-Patterson AFB in 

Ohio tested a total of 32 items for radar cross section (RCS) at the Ultra-High 

Frequency (UHF) frequency of 433 MHz. These items comprise nearly the entire 

external surface of the orbiter as well as the exposed surfaces in the cargo bay, RCC 

panels, and carrier panels. The items tested also included four pieces of recovered 

RCC debris from Columbia to better understand the radar characteristics of partial Tee 

seals and RCC panels. 

4-11 
The results of this radar testing and ballistics analysis have excluded all external Shuttle 

materials with the exception of 1) a whole Tee seal, 2) a Tee seal fragment that 

includes an attachment flange and/or apex segment, or 3) RCC panel acreage no less 

than 90 square inches and roughly square in shape (+/- 20%), although curvature is 

possible, with a thickness on the order of 0.33 inches. An RCC panel segment 

matches the RCS and ballistic performance characteristics observed during the STS107 mission. 

A Tee seal fragment with an apex segment matched the RCS characteristics extremely 

well in any spin orientation; however, the ballistic match required a very specific spin 

orientation that was shown to be feasible in one analytical simulation. Therefore, it is 

possible that the flight day 2 object was either a partial Tee seal or RCC panel acreage 

piece. The Incoflex spanner beam ear muff insulation was also a good match for both 

ballistics and RCS. Because the ear muff is situated behind the RCC panel, it is 

excluded from being considered a very likely candidate because of the lack of a 

mechanism for exposing it to the space environment. If the damage to the wing were 

actually a 10-inch diameter, uniformly round hole, then an ear muff would be a more 

plausible candidate. However, it is considered unlikely that the wing damage was a 10-

inch diameter round hole. The damage is considered to be the equivalent of that which 

would provide the same thermal response during entry as a 10-inch diameter hole did in 

the analyses and simulation. It is not likely that the actual wing damage was 

geometrically uniform. The damage was more likely a combination of cracks and holes, 

or a slot, such as a Tee seal or partial Tee seal missing or displaced. Therefore, the 

ear muff is not considered to be a good candidate for the flight day 2 object. Figure 4-6 

shows the three leading edge components that match both RCS and ballistics analysis. 

It should be noted that a full Tee seal and RCC panel are shown in these photos while 

there are specific partial Tee seal and RCC panel configurations that match the test 

results. 

Reinforced Carbon-Carbon (RCC) 

Leading Edge Panel (Flight Hardware) 

Incoflex Ear Muff Spanner Beam 

Insulator 

Reinforced Carbon-Carbon 

(RCC) T-Seal 

Figure 4-6. Leading edge structural subsystem components 

matching RCS and ballistics 

4-12 
4.3.4 KSC Lost and Found Items 

A review was conducted of the lost-and-found items from the Columbia processing 

flows of STS-107, STS-109, and the last Columbia OMM-J3. The largest tools that 

were lost and not found are listed in Table 4-5; other smaller items (e.g., washers and 

nutplates) are not listed. The item, size, location, and Problem Report (PR) number are 

noted. The largest item documented on a Lost and Found (LAF) PR is a piece of a 

blanket 6" x 3" lost in the payload bay (LAF-2-27-0611) during STS-109 processing. 

These items were screened using the ballistic coefficient and RCS criteria. All of the 

items failed the RCS screening and their RCS is too low to be a candidate for the flight 

day 2 object. 

Table 4-5. Lost tools in Columbia processing for STS-107, STS-109, 

and OMM J3 

Processing 

Flow 

Tool Description Location PR # 

OMM J3 Allen Socket, 2"x 1/2" Mid-body (LAF-2-J3-0550) 

OMM J3 Plastic Extraction Tool, 

22 gage 

Flight Deck (LAF-2-J3-0567) 

OMM J3 Pliers, 7 3/4" Hypergolic 

Maintenance 

Facility (HMF) 

(LAF-RPO5-15-0004) 

OMM J3 Screwdriver, 11" HMF (LAF-RPO5-15-0005) 

OMM J3 Screwdriver, 7" HMF (LAF-RPO5-15-0006) 

OMM J3 Screwdriver, 8" HMF (LAF-RPO5-15-0007) 

STS-109 Mini Flashlight, 6" Forward Reaction 

Control System 

(FRCS) 2 

(LAF-FRC2-27-0005) 

STS-107 Socket, 7/16"x 5/16" Aft Compartment (LAF-2-28-0632) 

4.4 ORBIT SUMMARY 

Extensive data review provided no conclusive indication of damage from either the 

ascent foam impact or an MMOD hypervelocity impact based on orbiter telemetry, crew 

downlinked video and still photography, or crew reports. 

Orbiter IMU and jet firing data have been reviewed, and this review confirmed that the 

IMUs were not designed for MMOD detection and data available to detect an MMOD 

strike is coarse. This data review found 13 events that required additional analysis. 

After this additional analysis, only two events remained that could not be ruled out as 

MMOD strikes. An examination of all VRCS jet firings was conducted and showed no 

unexplainable jet firings during STS-107. 

4-13 
SAMS data were also used in the analysis of the 13 events detected using IMU rate 

data. SAMS sensor data were also screened for large transients indicative of an 

MMOD strike; however, none were found. A model was developed that identifies the 

modal frequencies of the Shuttle structure (including wing modes) to further screen of 

the SAMS data for MMOD strikes. 

A review of the flight day 2 event has been performed including RCS testing and 

ballistics analysis of 41 items, including TPS. The analysis performed to date indicates 

that a full Tee seal, a partial Tee seal, and RCC panel are the only tested items that 

have not been excluded. 

It is possible that another untested object could match the RCS and ballistics and have 

departed the orbiter on flight day 2. Objects have departed the payload bay on 

previous Shuttle missions. The data are inconclusive as to whether the ET ascent foam 

debris event and the flight day 2 event are related. 

4-14 
5.0 DEORBIT/ENTRY 

5.1 INTRODUCTION 

The deorbit and entry section is divided into four distinct areas. The first section 

discusses the upper atmosphere weather including high altitude winds, and the 

Kennedy Space Center (KSC) Shuttle Landing Facility (SLF) weather for the STS-107 

landing. The next section includes a detailed discussion of forensics data obtained 

from testing and analysis of the key items in the recovered debris. The third section is 

a narrative of the entry events from February 1, 2003, and the fourth section is a brief 

discussion of key elements of the aerodynamic reconstruction. 

5.2 WEATHER 

5.2.1 Upper Atmosphere Weather 

As the Shuttle entered the atmosphere, it descended from about 400,000 feet when 

located over the central Pacific Ocean to roughly 200,000 feet over Texas. The 

Goddard Space Flight Center Data Assimilation Office (DAO) provided the GEOS-4 

model analysis for the investigation in order to provide a best estimate of the density, 

temperature, and wind along the entry trajectory. The GEOS-4 model assimilates a 

wide variety of data sources to produce an integrated 3-dimensional analysis of the 

atmosphere from the Earth's surface to about 250,000 feet. The Global Reference 

Atmosphere Model (GRAM) was used to provide information about the atmosphere 

from Entry Interface to the top of the GEOS-4 analysis. In general, the entry 

environment was characterized by a lower than average density and higher than 

average winds prior to the vehicle breakup. Comparison of the GEOS-4 analysis to 

GRAM indicates that the estimated density and winds were within the expected 

climatology for the upper atmosphere. Figure 5-1 shows the wind profile that was 

developed by the DAO as part of the STS-107 investigation. 

5.2.2 Landing Weather 

On the morning of February 1, 2003, there was a concern for ground fog formation at 

KSC for the first STS-107 landing opportunity. This concern is not uncommon for a 

morning landing at KSC during the winter months. The landing time for the first KSC 

opportunity was 9:16 EST. The forecast called for the fog to burn off as the sun rose, 

producing mixing in the lower levels of the atmosphere. The Shuttle Training Aircraft 

(STA), which is used for weather reconnaissance, flew approaches to both the KSC-15 

and KSC-33 runways to determine the best runway for landing. The STA is used to 

evaluate touchdown conditions, visibility, turbulence, crosswind, and overall pilot 

workload. At the time of deorbit decision, runway visibility was reported as 4 miles in 

light fog with winds 5 knots from the west. Visibility on final approach was slightly better 

for Runway 33. The final landing runway decision was not made at that time. 

5-1 
600.0 350 

300 

500.0 

250 

400.0 

200 

300.0 

150 

200.0 

100 

100.0 

50 

0.0 0 

STS-107 Wind Data 

47:00.0 48:00.0 49:00.0 50:00.0 51:00.0 52:00.0 53:00.0 54:00.0 55:00.0 56:00.0 57:00.0 58:00.0 59:00.0 00:00.0 

Time (min:sec) 

Wind Speed 

Wind Azimuth 

Vehicle Yaw 

Figure 5-1. Wind profile developed by DAO as part of the STS-107 

investigation (time referenced to 8:min:sec EST) 

Leading up to the deorbit decision time, the fog had been the main point of discussion 

until some clouds developed to the northwest of the landing area. Satellite imagery 

indicated an area of broken clouds (5/8 to 7/8 sky coverage) with bases at 

approximately 4,000 feet above ground level between 20 and 25 nautical miles 

northwest of the runway. The forecast was for those clouds to erode as they 

approached the SLF producing scattered clouds (3/8 or 4/8 sky coverage) at landing 

time. The Spaceflight Meteorology Group (SMG) stated that if erosion did not occur, 

the clouds reaching the SLF would be covering the runway for the first landing 

opportunity. No low clouds were being reported at the SLF at deorbit burn decision 

time and no surrounding observing sites were reporting low ceilings. The final forecast 

update was for a few clouds at one thousand feet and scattered clouds at three 

thousand feet, and the forecast remained Go per the flight rules. 

At the actual deorbit decision time and at the actual deorbit time, the landing weather 

satisfied all criteria per the documented Flight Rules, resulting in a Go observation and 

a Go forecast. At 9:10 EST, approximately five minutes prior to the expected landing 

time, the weather observation at the SLF reported a broken ceiling at 3,500 feet with 

6/8 sky coverage and visibility 7 miles. The ceiling remained 3,500 broken until 9:25 

EST at which time the SLF observer reported scattered clouds with 3/8 sky coverage. 

5-2 
 

The cloud deck at landing time was below the Flight Rule ceiling minimum requirement 

of 8,000 feet. Therefore, the commander would have relied in part on computer 

instrumentation and visible geographic references of the airfield, flying a Microwave 

Scanning Beam Landing System (MSBLS) approach until breaking out of the clouds at 

3,500 feet, a procedure regularly practiced in several landing simulators. The 

incorporation of MSBLS data provides very accurate onboard navigation, allowing for 

more accurate instrument information and facilitating instrument approach capability. 

The opinion among several experienced astronaut commanders, including the Chief of 

the Astronaut Office, is that the landing would likely not have been affected by this 

ceiling, when considering all other conditions of the day. 

5-3 
 

Very Heavy 5

4

3

2

Very Light 1

0

1 3 5 7 9 11 13 15 17 19 21

5.3 HARDWARE FORENSICS 

As discussed earlier in Section 3, Columbia entered the upper atmosphere with 

unknown damage to a Reinforced Carbon-Carbon (RCC) panel or Tee seal in the left 

wing RCC panels 6 through 9 area. The panel 8/9 area is the most likely area of 

damage as determined by hardware forensics testing, and analysis of MADS entry 

temperature and strain measurements on the left wing leading edge structure. This 

damage area is also consistent with the location of the ascent foam impact, and 

includes the Tee seals adjacent to panel 8, Tee seals 7 and 8. 

The forensic data indicate that the panel 8/9 area was subjected to extreme entry 

heating over a long period of time leading to RCC rib erosion, severely slumped carrier 

panel tiles, and a substantial slag deposition on the upper portion of RCC panels 8 and 

9. Figure 5-2 shows the slag deposition (both metallic and oxide) in the RCC panel 8/9 

area relative to the other parts of the wing leading edge, and Figure 5-3 shows samples 

of the severe slag deposition on the panel 8 rib. A review of all recovered debris 

indicates that this is the most probable area of a breach into the wing since there are no 

other debris pieces that exhibit the unique characteristics observed in this area. 

Very Heavy 5 

4 

3 

2 

Very Light 1 

0 

1 3 5 7 9 11 13 15 17 19 21 

Figure 5-2. Slag deposition in the RCC panel 8/9 area relative to the other parts of 

the left wing leading edge 

5-4 
Outboard Of Proposed Breach Location

Inboard Of Proposed Breach Location

RCC 8 Rib

Lower Item # 61143

Spar Spar

FWD FWD

Lower

Lower 

RCC 8 Rib 

Item # 61143 

Outboard Of Proposed Breach Location 

Spar Spar 

FWD FWD 

Inboard Of Proposed Breach Location 

Lower 

Figure 5-3. Samples of severe slag deposition on the panel 8 rib 

Based on the slag deposition on the upper RCC panel 8 and the rib erosion at the 

panel 8/9 interface, the most likely area of damage was the bottom portion of RCC 

panel 8. The outboard apex on the panel 8 upper inboard rib shows knife edge erosion, 

and the rib tapers from a design thickness of 0.365 inches to 0.05 inches. The surface 

of the panel 8 outboard rib and matching heel piece show a similar sign of erosion, as 

does the panel 9 upper inboard rib. The erosion on both the panel 8 and 9 rib is on the 

inboard side, indicating that flow is coming from the panel 8 location. Additionally, 

several lower carrier panel tiles in the RCC panel 9 area also show significant slumping 

and erosion that is consistent with a hole or breach in the lower part of RCC panel 8. 

Figure 5-4 and Figure 5-5 show an example of the rib erosion and the flow on the lower 

carrier panel 9. 

Figure 5-6 is a CAD drawing of the recovered debris showing overall slag deposition 

and erosion patterns. The drawing is a view from behind the RCC panels since this 

provides the best view of the erosion and slag deposition. Three full Tee seals can be 

seen in this drawing; the leftmost, Tee seal 9, divides panel 9 and 10; Tee seal 8 in the 

center divides panel 8 and 9; the rightmost, Tee seal 7, is the division between panel 7 

and 8. The drawing shows the heavy slag deposition of the upper portion of panel 8 

indicating that the probable breach area was the bottom of panel 8. The severely 

eroded RCC ribs are also visible near the RCC panel 8/9 Tee seal. The heavy slag on 

5-5 
Item # 49619 Tee Seal Between RCC 8, 9

Remainder Of RCC 8 Close-ups of knife edge, note fibers not visible 

on internal surface of panel due to deposits. Rib 

tapers from design thickness of .365 to .05.

Panel 8 RCC Outboard Rib

these inner surfaces indicate flow from the panel 8 direction toward panel 9. This is 

also consistent with the knife-edge erosion shown in Figure 5-4 below. The proposed 

flow direction leading to the erosion and slag deposition on the lower carrier panel 9 

tiles can be seen in this view as well. The detailed flow, erosion, and deposition are 

best viewed in Figure 5-5. The last significant feature in Figure 5-6 is the heavy 

slumping that is observed on the upper carrier panel 8 tile (50336T) in the upper right 

portion of the drawing. 

Item # 49619 Tee Seal Between RCC 8, 9 

Lower

Upper 

Remainder Of RCC 8 Close-ups of knife edge, note fibers not visible 

on internal surface of panel due to deposits. Rib 

tapers from design thickness of .365 to .05. 

Panel 8 RCC Outboard Rib 

Figure 5-4. Example of rib erosion 

5-6 
 

22571

50338

Tile V070-191026

Directly Inboard of C/P 9

61143

49619

Knife Edges From

Thermal Erosion From

Inboard Source

View looking up 

15523

Spar Side

16692 57754

Panel 10 RCC Side Panel 8 

Slag deposits

Depressed/eroded regions

50336T

Heavy Slump

29741

17957

29233

26014

Heavy Slag On

Inner Surfaces

35201 (Location

not positively 24724

identified) 58291

Heavy Slag 

On Inner 

Surfaces Slumping & Erosion 

Patterns Suggest

Plasma Flow Exiting At

C/Ps Of RCC 9 w/ Flow

Outboard Across Tiles

Upper 

Tile V070-191026

Directly Inboard of C/P 9 View looking up 

15523 

Spar Side 

16692 57754 

Panel 10 RCC Side Panel 8 

Slag deposits 

Depressed/eroded regions 

22571 

50338 

` 

Figure 5-5. Flow on the lower carrier panel 9 tiles 

Tee Seal 8 Tee Seal 7 

Tee Seal 9 RCC 8 

RCC 9 Looking Fwd & 

Outboard 50336T 

Heavy Slump 

29741 

61143 

49619 

17957 

29233 

26014 

Heavy Slag On

Inner Surfaces 

35201 (Location

not positively 24724 

identified) 58291 

Heavy Slag

On Inner 

Surfaces 

Knife Edges From 

Thermal Erosion From 

Inboard Source 

Slumping & Erosion 

Patterns Suggest

Plasma Flow Exiting At

C/Ps Of RCC 9 w/ Flow 

Outboard Across Tiles Evidence Of Hot Gas Exiting Design Slot

Indicates Significant Breach Was Into Panel 8

Evidence Of Hot Gas Exiting Design Slot 

Indicates Significant Breach Was Into Panel 8 

Figure 5-6. CAD drawing of the recovered debris showing overall slag 

deposition and erosion patterns 

5-7 
The data shown in Figure 5-6 is important when combined with the analysis of the slag 

deposition. The slag deposition on the upper RCC panel 8 was analyzed using 

sophisticated cross sectional optical and scanning electron microscopy, microprobe 

analysis, and x-ray diffraction to determine the content and layering of the slag 

deposition. This analysis indicated that the materials in this area were exposed to 

extremely high temperatures, since Cerachrome insulation was deposited first and its 

melting temperature is greater than 3200 degrees Fahrenheit. The analysis also 

showed no presence of Aluminum 286 in the slag indicating that the RCC attach fittings 

were not in the direct line of the breach and that the Inconel 718 spanner beam was 

one of the first internal items to be subjected to heating. Inconel slag was prevalent in 

much of the analyzed slag indicating melting of the spanner beam, foil, and associated 

insulation. Aluminum was found in the last deposited layer indicating the wing 

honeycomb spar was the last area to be subjected to hot gas flow. 

Analysis of the slag deposition on the lower carrier panel 9 tiles was also performed. 

Materials on these tiles are consistent with wing leading edge materials (Aluminum, 

Inconel, Nickel Alloy, and Carbon) indicating an outflow from the panel 8 area across 

the tiles. Tile slumping in this area is indicative of temperatures in excess of 3000 o

F. 

The upper carrier panel 8 tile was also analyzed and the results were similar to lower 

carrier panel 9 except that this tile appeared to have more Cerachrome and Nextel fiber 

deposits. These materials are consistent with the insulator that protects the wing 

leading edge spar and with flow moving toward the upper wing surface through the vent 

between the upper carrier panel and RCC. 

This forensics analysis further corroborates the breach location to be the lower portion 

of RCC panel 8 below the apex, approximately midway between the apex and where 

the RCC panel meets the carrier panel. Based on the flow patterns, the breach was in 

an area that caused the flow to impact the spanner beam associated with Tee seal 8 

and create the knife edge erosion shown in this area in Figure 5-6. 

The flow appears to have entered through this breach and into the lower aft corner, 

exiting through a slot toward carrier panel 9. The flow burned through the horse collar 

and eroded and slumped the carrier panel tiles. The flow continually grew the hole in 

panel 8 as time progressed and it eroded the remaining aft flange part of RCC panel 8 

and the forward flange on RCC panel 9. Although the lower carrier panel 9 tiles are 

slumped and eroded, there must have been an RCC rib protecting the adjacent carrier 

panel 8 tiles since there is no erosion or slumping of these tiles. Compared to the 

severely eroded carrier panel 9 tiles, the three recovered carrier panel 8 tiles are in 

relatively pristine condition, and likely separated due to backside heating with no 

indications of mechanical damage occurring prior to vehicle break-up. 

As time progressed, the Cerachrome and Inconel wing spar insulators were eroded, 

and eventually hot gas flow impinged on the wing leading edge, burning through the 

honeycomb spar. Figure 5-7 depicts the possible flow direction and deposition of 

various metals as determined by this analysis. 

5-8 
Proposed Breach Location & Plasma 

Flow Based On Slag Results Cerachrome 

Cerachrome 

Globules 

Flow Exiting Through RCC 8

To Carrier Panel 9 Hot Gas Flow 

{

>3200 oF 

{ 

{

Tee Seal 8 Tears 

Inconel 

Spheroids 

RCC 8 

 

 

 

 

>3200 oF

{

Tears 

Inconel 

Spheroids 

e

Proposed Breach Location & Plasma

Flow Based On Slag Results Cerachrome

Cerachrom

Globules 

Flow Exiting Through RCC 8 

To Carrier Panel 9 Hot Gas Flow

Figure 5-7. Analysis results show possible flow direction 

and deposition of metals 

In addition to the slag deposition and flow analysis, there are two other significant 

pieces of data that point to a breach in the RCC 8/9 area as the initial damage. The 

first item is the location of the leading edge RCC in the debris footprint. Figure 5-8 

shows the recovered RCC for both the left and right wing and its location in the debris 

footprint. The eroded RCC pieces from panels 8 and 9 are found in the westernmost 

part of this debris footprint near Waxahachie, Texas, along with other pieces of RCC 

panel 8. Left wing RCC panel 9 and other aft panels appear to have been lost relatively 

early in the break-up sequence since their footprint spans the western to center part of 

the footprint. This is indicative of a left wing breach in the panel 8/9 area. The forward 

portion of the RCC panels on both the left and right wings (panels 1 through 7) are 

found from the center to eastern part of the debris footprint possibly indicating that 

these were lost in a secondary aerodynamic break-up. 

5-9 
Corsicana 

North

 Left Wing RCC 

 Left Wing Eroded RCC 

 Right Wing RCC 

Figure 5-8. RCC panel debris location 

5-10 
The second additional piece of data is an upper left wing tile recovered near Littlefield, 

Texas. Littlefield is a small town near the Texas/New Mexico border along Columbias 

ground track. This tile is the westernmost piece of debris that has been found to date in 

the debris recovery efforts. Due to the unique features of the tile (thickness, shape, 

paint, etc.), the tile must be from the upper wing area in the RCC panel 9 area. Figure 

5-9 shows the only three possible locations for this tile. 

The tile departed the orbiter more than one minute prior to final break-up due to 

prolonged internal heating of the upper wing skin in the area shown in Figure 5-9. The 

tile shows indications of backside heating and an RTV debond. It was not a failure in 

the densification layer, which would have been caused by mechanical loading. This 

piece of recovered debris is not very significant on its own merit; however, it is 

consistent with the previously discussed forensics data (rib erosion, carrier panel 9 tile 

slumping, etc.) and other events that will be discussed later in Section 5.4. 

V070-195003-097 

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V070-195003-150 

V070-195003-101 

Upper L/H LESS # 9 

Figure 5-9. Three possible orbiter locations of the Littlefield tile on left wing 

5-11 
5.0 

5.4 ENTRY EVENTS TIMELINE 

5.4.1 Early Entry Heating Events 

Columbia successfully completed the deorbit burn at 8:18:08 EST over the Indian 

Ocean. The deorbit burn and entry targeting were accomplished using well-established 

Mission Control Center procedures, and there were no problems identified with this 

process. Both the left and right Orbital Maneuvering System (OMS) engines performed 

nominally and the post burn residuals were less than 0.2 feet per second indicating a 

precise burn. The maneuver to the Entry Interface (EI) attitude, the Forward Reaction 

Control System Dump, and remaining Auxiliary Power Unit (APU) start (APU 1 and 

APU 3) were accomplished nominally. 

At 8:44:09 EST, Columbia reached EI, the transition between orbital and atmospheric 

flight. The altitude was 400,000 feet and the orbiter was traveling Mach 24.6 in wings 

level (zero degree bank) attitude with a nominal 40-degree angle of attack. The orbiter 

guidance had been moded to OPS 304 nominally at five minutes prior to entry interface. 

OPS 304 is the name given to the entry flight software that contains the aerojet digital 

auto-pilot control mode. It is used from five minutes prior to EI through Mach 2.5. 

Figure 5-10 is a plot of dynamic pressure and stagnation heating from EI to vehicle 

break-up. The plot shows that both heating and dynamic pressure were very low during 

the two to three minutes (120-180 seconds) after EI. The heating rate shown is 

stagnation heat flux that is the allowable heat flux that could be achieved by the gas if 

all its thermal and kinetic energy were available. For this plot and others that follow in 

Section 5, EI occurred at 8:44:09 EST, which corresponds to zero seconds on the plots. 

This is a convenient reference point for many of the entry events that will be discussed. 

At approximately 8:48:39 EST (EI + 270 sec.), a left wing leading edge spar strain 

gauge began a small off-nominal increase. Figure 5-11 shows the STS-107 response 

of this strain measurement along with three other previous Columbia missions. Figure 

5-12 shows the location of this sensor (WLE Strain V12G9921) and others on the wing 

leading edge. The damage in lower RCC panel 8 is believed to be the cause of this 

strain increase. The breach allowed hot gas intrusion onto the panels 8 through 9 wing 

leading edge spar area leading to extreme heating and thermally induced strain. The 

strain increase grew over time and reached a maximum at approximately 8:50:09 EST 

(EI + 360 sec.). Thermal and structural analyses indicate that a breach would need to 

be within approximately 15 inches of the strain gauge to create the observed strain 

increase. 

5-12 
 

First off nominal indication

STS-107

STS-073

STS-090

STS-109

48:39 44:09

0.0 

10.0 

20.0 

30.0 

40.0 

50.0 

60.0 

70.0 

80.0 

90.0 

100.0 

Stagnation Heat Flux (BTU/ft^2*sec)

Dynamic Pressure (lb/ft^2) 

Stagnation Heat Flux - BTU/ft^2*sec 

Dynamic Pressure - lbf/ft^2 

LOS - EI + 923s 

0 100 200 300 400 500 600 700 800 900 1000 

Time from Entry Interface (EI) - secs 

Figure 5-10. STS-107 stagnation heat flux and dynamic pressure. 

Note that EI was at 8:44:09 EST. 

59:09

STS-107

STS-073

STS-090

STS-109

48:39

Strain (micron-in./in.) 

First off nominal indication 

44:09 59:09 

TIME (seconds from EI ) EST 

Figure 5-11. Left wing RCC panel 9 strain gauge is first measurement 

to indicate an off-nominal event. Note that EI was at 8:44:09 EST. 

5-13 
 

 

 

 

 

 

 

 

WLE Spar Temp - V09T9895

WLE Strain - V12G9921

WLE Clevis - V09T9910

Aft Panel 9 Lower Surface Temp - V09T9666

Panel 10 Panel 10 Panel 9 Panel 9 

WLE Spar Temp - V09T9895 

WLE Strain - V12G9921 

WLE Clevis - V09T9910 

Aft Panel 9 Lower Surface Temp - V09T9666 

Figure 5-12. MADS sensors inside left wing 

Twenty seconds later at 8:48:59 EST (EI + 290 sec.), the left wing lower attach clevis 

temperature sensor (between RCC panel 9 and 10) began an early off nominal 

temperature trend. Figure 5-13 shows the abnormal temperature response when 

compared to other Columbia missions. This temperature rise is consistent with an early 

entry of hot gas into the RCC cavity. This Modular Auxiliary Data System (MADS) 

measurement (V09T9910) is the only temperature measurement located in the RCC 

cavity along the left wing leading edge. It is positioned on the lower attach fitting 

between panel 9 and 10 and is well protected thermally by Inconel foil insulation. The 

sensor is also thermally isolated since it sits on the attach fitting away from other 

structure as shown in Figure 3-15. In order to get an early temperature rise for this 

sensor, unlike that observed on any other flight, there must be a path in the RCC cavity 

to allow hot gas to reach the sensor. 

5-14 
First off nominal indication

STS-107

STS-073

STS-090

STS-109

48:59 44:09

A thermal analysis was performed with heating rates from various hole sizes in the 

bottom of RCC panel 8 in an attempt to match this temperature rise. The analysis used 

a thermal math model of the wing leading edge (Inconel Cerachrome insulation, 

Inconel 718 and A-286 steel attach fittings, and aluminum honeycomb spar). The 

results indicated that the heating equivalent of a 6 to 10 inch diameter hole with a 

10 percent sneak flow around the insulation would be required to match the thermal 

response of the clevis temperature. In the same timeframe several MADS lower 

surface temperatures on the left wing showed a slight off nominal early temperature rise 

when compared to previous flights of Columbia of the same inclination. 

59:09

STS-107

STS-073

STS-090

STS-109

48:59

Temperature (o F) 

First off nominal indication 

44:09 59:09 

Time (seconds from EI) 

EST 

Figure 5-13. Left wing RCC panel 9/10 clevis temp sensor is second 

measurement to indicate an off-nominal event 

5-15 
 

5.0 

Reduced, off-nominal heating

First off nominal indication

Left OMS Pod Surface Mounted Tile Temperature on Forward Looking Face

STS-107

STS-073

STS-090

STS-109

49:49

44:09 59:09

-

5.4.2 First Roll Maneuver Through Wing Spar Breach 

Columbia executed a nominal roll to the right at 8:49:32 EST (EI + 323 sec.) as the 

entry guidance software began to actively control energy (i.e., closed loop guidance) to 

land at KSC. This initial roll command is also timed to ensure atmospheric capture by 

reducing the lift on the vehicle. Within 17 seconds of this maneuver, at 8:49:49 EST 

(EI + 340 sec.), four left OMS pod surface temperature measurements showed an offnominal trend with lower temperature rises when compared to similar Columbia 

missions. A sample of these measurements compared to other Columbia missions is 

shown in Figure 5-14, and the location of these measurements on the left OMS pod 

forward face can be found in Figure 5-15. 

Left OMS Pod Surface Mounted Tile Temperature on Forward Looking Face 

STS-107 

STS-073 

STS-090 

STS-109 

First off-nominal indication 

Reduced, off-nominal heating 

44:09 59:09

49:49 

EST 

Figure 5-14. Typical off-nominal OMS pod thermocouple (V07T9220A) 

5-16 
V07T9913A V07T9925A

V07T9972A; V07T9220A;

V07T9976A; V07T9978A

Left OMS Pod Temps 

Off Nominal temp (cooler)

V07T9953A

1

0 1

0

V07T9913A V07T9925A

V07T9972A; V07T9220A; 

V07T9976A; V07T9978A 

Left OMS Pod Temps 

Off Nominal temp (cooler) 

V07T9953A

1

0 1

0

Figure 5-15. Location of OMS pod thermocouples off-nominal low 

5-17
Wing leading edge damage perturbs leeward flow separation and re attachment locations

as well as leeward embedded shocks (not shown)

leeside

vortex field

Flow

Flow

separation Pe ed

eparation

Canopy/leeside

vortex Perturbed

re-attachment Recirculating flow Perturbed flow

Baseline

-

The reduced heating is not completely understood since the weak aerodynamic flow 

field on the upper surface of the orbiter is difficult to model and is extremely sensitive to 

disturbances. The best explanation for this reduced heating is that flow into the RCC 

cavity was venting through to the upper surface of the wing through an existing 0.1-inch 

vent between the RCC and upper surface carrier panels. This vent exists all along the 

leading edge from RCC panel 1 through panel 22 and has an approximate area of 66 

square inches. This upper surface RCC venting and the flow disturbance created by 

the panel 8 and upper carrier panel 8 damage caused the vortices from the canopy or 

area where the wing meets the orbiter fuselage to move from their normal positions, 

thus reducing the heating on the OMS pod. Figure 5-16 depicts the change in the 

upper wing surface vortices and the weak upper surface flow. 

Damaged leading edge

rturb

flow s

leeside 

vortex field 

Flow 

Flow 

separation 

Damaged leading edge 

Perturbed 

flow separation 

Canopy/leeside 

vortex Perturbed 

re-attachment Recirculating flow Perturbed flow 

Baseline 

Wing leading edge damage perturbs leeward flow separation and re-attachment locations 

as well as leeward embedded shocks (not shown) 

Figure 5-16. Postulated orbiter leeside flow field associated 

with wing leading edge damage 

In order to verify this theory of a weak upper surface flow being disturbed from venting 

on the upper surface, several wind tunnel tests were performed in the NASA Langley 

Research Center Mach 6 Tetraflouromethane (CF4) Wind Tunnel. The use of CF4 as 

the gas for the flow analysis is required to best replicate the Mach number environment 

during this timeframe. These wind tunnel tests used a ceramic model and a 0.01-inch 

leading edge vent to mimic the postulated venting. Nitrogen gas was allowed to flow 

through this upper surface vent via a gas supply line. A picture of this model is shown 

in Figure 5-17. Results of this testing show that it is feasible to obtain reduced heating 

on both the left OMS pod and the left fuselage as a result of flow through the RCC 

upper surface vent. Figure 5-18 shows the change in heating as the vent velocity is 

altered. 

5-18 
 

Ceramic

coated

resin model

Baseline

mr ≈ 0 0.60 0.90 1.09

ρvent (Vvent)2

CF4 γ eff = 1.13 α = 40 deg β = 0 deg mr =

ρ∞ (V∞)2

Ceramic 

coated 

resin model 

Shuttle 

model in 

tunnel 

Gas supply 

line 

Figure 5-17. Orbiter wind tunnel model with vent gap along wing leading edge 

Baseline 

mr ≈ 0 0.60 0.90 1.09 

ρvent (Vvent)2 

CF4 γ eff = 1.13 α = 40 deg β = 0 deg mr = 

ρ∞ (V∞)2 

Figure 5-18. Wind tunnel model results for sensitivity of orbiter side fuselage and 

OMS pod heating patterns to mass addition along WLE leeside vent gap 

5-19 
V07T9913A

V07T9925A

Left PLBD Surface Temp

Off Nominal temp (cooler)

V07T9953A

1

0 1

0

Over the next 43 seconds, there were five communications dropouts beginning at 

8:50:00 EST and ending at 8:50:43 EST (EI + 351 through 394 sec.). It is possible that 

hot gas in the RCC cavity had begun to erode the Inconel and Cerachrome insulation 

along the wing leading edge spar. Molten materials could have been ejected into the 

environment around the orbiter creating multi-path signal scattering with the link 

between the orbiter and the western Tracking and Data Relay Satellite (TDRS). The 

best parallel for this explanation is chaff used by some military aircraft to confuse 

opposing radar systems. 

Forensic analysis of the recovered left OMS pod debris indicates that molten 

Inconel 718 and A-286 cress were sprayed onto the left OMS pod during entry. This 

OMS pod debris and the left side of a recovered vertical tail debris piece were 

significantly pitted by this metallic spray supporting the concept that there was 

vaporized metal in the environment around Columbia. These materials must have 

originated from the RCC panel 8 wing spar damage area since Inconel 718 is used as 

the wing leading edge insulator and A-286 is used for the RCC attach fittings. 

In the same timeframe, at 8:50:09 EST (EI + 360 sec.), a left payload bay fuselage 

MADS surface temperature measurement (Figure 5-19) showed an off-nominal 

temperature trend. This trend is a reduced rise rate when compared to other previous 

Columbia missions, as shown in Figure 5-20. The previously discussed theory of 

venting and or disturbed flow due to panel 8 damage, causing a shift in the vortices on 

the upper surface of the wing, is also believed to be the cause of this off-nominal 

behavior. The flow field and venting on the upper surface rate are constantly increasing 

since the mass flow rate into the RCC breach is increasing as the orbiter descends 

lower into the atmosphere. 

V07T9913A

V07T9925A 

Left PLBD Surface Temp 

Off Nominal temp (cooler) 

V07T9953A

1

0 1

0

Figure 5-19. Location of left sidewall temperature sensor 

5-20
Reduction in heating

First off nominal indication

Fuselage Sidewall Surface mounted Tile Temperature

STS-107

STS-073

STS-090

STS-109

50:09

44:09 59:09

Fuselage Sidewall Surface mounted Tile Temperature 

STS-107 

STS-073 

STS-090 

STS-109 

First off nominal indication 

Reduction in heating 

44:09 59:09

50:09 

EST 

Figure 5-20. Off-nominal temperature indication on the left sidewall 

5-21 
 

First off nominal

indication

50:19

59:09

At 8:50:19 EST (EI + 370 sec.), a lower surface thermocouple showed signs of 

off nominal, increased heating. The best explanation for the increased heating in this 

area is severely disturbed, turbulent flow caused by the leading edge damage on the 

bottom of RCC panel 8 and flow from the lower corner of this panel as discussed in 

Section 5.3. Langley Research Center wind tunnel testing has confirmed that wing 

leading edge damage (notch or protuberance) near panel 8 will cause increased 

heating to the lower wing surface. As previously discussed, the eroded lower carrier 

panel tiles on panel 9 indicate this strong flow from panel 8. This is consistent with flow 

patterns observed on many recovered lower surface wing acreage tiles along the flow 

lines aft of the RCC panel 8 area. Figure 5-21 shows this temperature response as 

compared to other Columbia missions. The location of the sensor can be seen in 

Figure 5-22. 

44:09

50:19 

Temperature (deg. F) 

First off nominal 

indication 

EST 

TIME (seconds after EI) 59:09

Figure 5-21. Temperature rise on tile surfaces aft of RCC panel 9 

5-22 
1040 Spar Strain

V12G9049A, V12G9048A

V08D9Elevon 

MSID, Start Time Of Signal Loss (From EI) 

All LW MADS 

Sensors Shown 

Nominal 

Off Nominal 

Off Line 

Indeterminate 

V07P8024A, 490 V07P8073A, 492 V07P8023A, 489 V07P8072A, 490 

V07P8022A, 492 V07P8071A, 491 V07P8010A, 497 V07P8058A, 495 

1040 Spar Strain 

V12G9049A, V12G9048A 

V09T9910A, 493 - Panel 9/10 Clevis TV12G9921A, 495  Panel 9 WLE Spar V07T9666A, 496  Lower Surface TheV12G9169A, Snap Shot Data 

8

9

V08D9Elevon V07P8

V07P8

V07P

V07P8026

V07P8025A, 

V07P8074A, 492 

7 V07T9895A, 522 - Panel 9 WLE Spar 

6

Figure 5-22. Left wing MADS sensors, including Measurement Stimulation Identification 

and start time of loss of signal (EI + sec.) 

5-23 
By 8:51:14 EST (EI + 425 sec.), the wing leading edge spar temperature began an 

early, off-nominal rise, shown in Figure 5-23. At the same time, the clevis temperature 

in the RCC panel 9/10 region continued to increase. The initial spar temperature rise 

was relatively slow and was caused by conduction since this measurement is on the 

backside of the spar in the wing cavity. Eventually, the rise rate increased dramatically 

as first the Inconel and Cerachrome insulation and later the aluminum honeycomb were 

destroyed. The exact time of the spar breach is unknown; however, it is estimated to 

have occurred between 8:51:14 to 8:51:59 EST (EI + 425 to 470 sec.) based on the 

observed wing leading edge linear decrease in strain during this timeframe. A more 

detailed discussion of the method used to bracket the time of the wing spar breach is 

contained in the following section. 

500 

V12G9921A STS-107 

Panel 9 WLE Spar Strain 

V09T9910A STS-107 

Panel 9/10 Clevis Temp 

V09T9895A STS-107 

Panel 9 WLE Spar Temp 500 

400 

450 

400 

Strain (uin/in) 

300 350 

300 

200 

250 

100 

200 

150 0 

Burn through 100 

-100 

50 

-200 0 

100 150 200 250 300 350 400 450 500 550 600 

TIME (seconds after EI) Time (EI Seconds)

51:14 

Temperature (degF) 

52:14  Increase heating at V09T9895A 

51:59  Strain Assessment Burn Through Time EST 

Figure 5-23. RCC panel 9 MADS strain and temperature measurements, STS-107 

5-24 
At 8:51:49 EST (EI + 460 sec.), one of the left OMS pod measurements (V07T9972A) 

began to show an increased temperature rise, indicating that the upper surface flow has 

changed again. This is shown in Figure 5-24 along with this measurement on other 

Columbia missions. It is evident that the measurement rises to a higher temperature 

and at a faster rate than has been observed on previous missions within the next few 

minutes. Figure 5-25 shows other left OMS pod and fuselage temperature 

measurements that exhibit an off-nominal rise later in time. The sensor locations on the 

left side of the orbiter are also shown in Figure 5-25. 

STS-073 

STS-090 

STS-109

STS-107 

8:51:49 EST 

Figure 5-24. Off-nominal low OMS pod thermocouple (V07T9972A) 

5-25 
STS-107 MADS1 Entry Data

1800 V07T9219A - OMS-L POD HRSI SURF T1-AFT

V07T9220A - OMS-L POD LRSI SURF TEMP-FW D

V07T9222A - OMS-L POD HRSI SURF T2-AFT 1600 V07T9223A - OMS-L POD HRSI SURF T3-AFT

V07T9972A - OMS-L POD THERMOCOUPLE BP0749T

V07T9976A - OMS-L POD THERMOCOUPLE BP0731T

1400 V07T9978A - OMS-L POD THERMOCOUPLE BP0732T

V07T9253A - FUSLG SIDE SURF TEMP BP3605T

V07T9925A - FUSLG SIDE SURF TC BP3703T

1200

Start of off-nominal 1000

response

800

600

400

200

0

Temperature (DEG F)

GMT (032_HH:MM:SS)

9 9 9 

09 

9 

09 

9 9 9 

09 

9 

9 

9 9 

0 0 0 0 0 0

:0:0

0 0 0

4:

: :

8:

:

1:4:5:

:

7:

: :

45

6 7 0 2 3 6 :4 : :4 :4 :4 :49

:5 :5 :5 :5 :5 :5 :5 :5

1313

3 3

13

3 3

13

3 3

1313

3 1 1 1 1 1 1 1

13

STS-107 MADS1 Entry Data 

1800 V07T9219A - OMS-L POD HRSI SURF T1-AFT 

V07T9220A - OMS-L POD LRSI SURF TEMP-FW D 

V07T9222A - OMS-L POD HRSI SURF T2-AFT 1600 V07T9223A - OMS-L POD HRSI SURF T3-AFT 

V07T9972A - OMS-L POD THERMOCOUPLE BP0749T 

V07T9976A - OMS-L POD THERMOCOUPLE BP0731T 

1400 V07T9978A - OMS-L POD THERMOCOUPLE BP0732T 

V07T9253A - FUSLG SIDE SURF TEMP BP3605T 

V07T9925A - FUSLG SIDE SURF TC BP3703T 

1200

Start of off-nominal 1000 

response 

800 

600 

400 

200 

0 

Temperature (DEG F) 

0 60 120 180 240 300 360 420 480 540 600 660 720 

TIME (seconds after EI)

GMT (032_HH:MM:SS) 

Figure 5-25. Left side fuselage/OMS pod off-nominal responses 

indicate increased heating 

These off-nominal temperature rises were caused by a change in the flow field along 

the upper portion of the left wing as the wing leading edge (WLE) and wing spar 

damage progressed. The mechanism was a shift in the vortices due to venting on the 

upper portion of the wing and is similar to the previous discussions on lower than 

expected OMS pod temperatures. Wind tunnel testing and CFD analysis confirmed 

that WLE damage on RCC panels 8 and 9 could cause increased heating to the OMS 

pod due to disturbed upper surface flow or flow impingement caused by re-directed flow 

from the lower surface. Additionally, the molten Inconel 718 and A-286 spray onto this 

area are indications that this flow is the cause of these temperature rises. Figure 5-26 

(from wind tunnel test results) is a representative example of increased heating on the 

left fuselage and OMS pod due to RCC panel 9 damage from wind tunnel test results. 

Figure 5-27 shows a similar picture of the left side as determined by CFD analysis. 

Note that neither the wind tunnel testing nor the CFD analysis accounts for increased 

heating that is likely due to molten metal contacting the OMS pod. 

5-26
RCC 9

Mach 6 Air 

RCC 9

Mach 6 CF4

RCC 9

Mach 24.9

CFD Case #1

Smooth WLE

Lower 1⁄2 Panel Out Full Panel Out

Solid Cavity Side Walls Solid Cavity Side Walls

RCC 9 

Mach 6 Air 

RCC 9 

Mach 6 CF4 

Figure 5-26. Wing tunnel test results for RCC panel 9 missing 

and resulting in increased heating to OMS pod 

RCC 9 

Mach 24.9 

CFD Case #1 

Smooth WLE 

Lower 1⁄2 Panel Out Full Panel Out 

Solid Cavity Side Walls Solid Cavity Side Walls 

Figure 5-27. CFD results for no damage, partial damage, and full damage to RCC 

panel 9 show increased heating on side fuselage and OMS pod 

The increased heating implies that damage/erosion to the RCC panel 8 area had 

increased or that at least one carrier panel had been lost. The increased heating is 

also consistent with an increased mass flow rate due to the nature of the entry 

environment. The dynamic pressure and stagnation heat flux had more than doubled 

by this timeframe when compared to a few minutes after EI. There are three additional 

communication dropouts that follow at approximately 8:52:09 EST to 8:52:31 EST 

(EI + 480 to 502 sec.). Again, these dropouts are consistent with molten aluminum and 

other metals being released into the environment surrounding Columbia. 

5-27 
5.0 

First Significant Aileron Rate Change - Left Roll 

Left Lower/Upper Wing S kin Temp O SL 

Start of 2 Yaw Jets Firing (R2R,R3R) 

Start of Slow Aileron Trim Change 

Start o f S harp Aileron Trim Increase 

LMG Brake Line Temp Increase (3) 

Max Observed A irleron Trim (-2.3o

) 

Start of Roll Trim in Elevons 

1215 

M-FUS LT BL Temp at

Start of Peak Heating 

Entry Interface 

5.4.3 Wing Breach and Wire Failures 

The next significant event was a breach through the left wing leading edge spar leading 

to many wire measurement failures and eventual deformation of the left wing. As 

discussed in Section 5.4.2, by approximately 8:52:00 EST (EI + 471 seconds), strain 

and temperature measurements indicated that hot gas had begun to weaken the wing 

spar. Figure 5-28 shows that the orbiter has entered the peak-heating region during 

this timeframe and would remain there for many of the ensuing events that will be 

discussed later. 

1.0 Reference Heating Temperature ( RHOo

F) 

0.9 3000 

0.8 2700 

2400 

0.7 2100 

1800 

0.6 1500 

0.5 1200 

900 

0.4 600 

0.3 300 

0 

0.2 

0.1 

0.0 

13:44

 0 

13:50

 360 

13:55

660 

14:00

960 

14:05

1260 

14:10 

1560 

Heating Rate (Normalized) 

TIME (seconds from EI) Time (GMT) 

Figure 5-28. STS-107 entry heating rate profile 

The precise timing of the wing spar breach is difficult to determine, and three different 

techniques were used in an attempt to bound the breach time. These techniques 

included a structural assessment using the wing leading edge spar measurement, use 

of a thermal model to predict the time required to burn through the spar insulators and 

then the honeycomb structure, and a wire failure assessment. 

The first technique used the strain response, shown earlier in Figure 5-23. On this plot, 

the initial strain rise that began at about 8:49:09 EST (EI + 300 sec.) is due to thermal 

elastic strain. It appears that the spar structural softening occurs at approximately 

8:50:09 EST (EI + 360 sec.), followed by loss of the structural integrity, or breach, at 

5-28 
approximately 8:51:14 EST (EI + 425 sec.). This appears to be completed by 

approximately 8:52:00 EST (EI + 471 sec.). A detailed structural model that attempted 

to reproduce the thermal strain response observed during this timeframe confirmed this 

timing. This analysis determined that the location of the spar breach must be within 

about 15 inches of the spar strain measurement. This would locate the spar breach 

near the intersection of panel 8/9, as shown in Figure 5-29. The flow through the RCC 

breach maintained some directionality although it was influenced by the shape of the 

hole, remaining RCC structure, attach hardware, and leading edge insulators. Overall, 

this strain response is consistent with an RCC breach in the lower part of panel 8 as 

previously discussed in Section 5.3. Although it is difficult to pinpoint a precise location 

of the RCC breach, this analysis supports the argument that the breach was closer to 

the panel 8/9 intersection and Tee seal 8. 

V07P8010A 

V07P8058A 

V09T9895A 

Panel 9 Panel 8 

EI + 522 

EI + 497 

EI + 497 

? 

Remainder 

EI + 487 to 496 

Figure 5-29. Cable routing on wing leading edge and wheel well wall 

5-29 
3500 WLE PANEL9 LOWER TEMPERATURE PROFILE

STS-107 DAMAGE ANALYSIS FOR 6 INCH HOLE

3000 EARMUFFINCON

EARMUFFNEXTEL

2500 EARMUFFCER

SPARFIXINSUL

Temperature (°F)

2000 Inconel melt@ 350 Sec

Nextel Melt @ 460Sec

1500 Cerachrome Melt @ 470 Sec

Spar Melt @ 490 Sec

1000

500

0

Inconel melt

Nextel melt

Cerachrome melt

Spar melt

0 100 200 300 400 500

Time (EI + Seconds)

The second technique used a detailed thermal model to determine the time required to 

burn through the various insulations immediately in front of the wing leading edge spar 

(Inconel, Nextel, Cerachrome) and then the honeycomb spar. This model assumed the 

equivalent heating of a six-inch diameter hole in the bottom of RCC panel 8. Using the 

expected aero heating rate and the various material properties, the spar burn through 

occurred at 8:52:19 EST (EI + 490 sec.). This time would be accelerated slightly for a 

larger diameter hole. The results of this thermal model are shown in Figure 5-30. 

3500 WLE PANEL9 LOWER TEMPERATURE PROFILE 

STS-107 DAMAGE ANALYSIS FOR 6 INCH HOLE 

3000 EARMUFFINCON 

EARMUFFNEXTEL 

2500 EARMUFFCER 

SPARFIXINSUL 

Temperature (°F) 

2000 Inconel melt@ 350 Sec 

Nextel Melt @ 460Sec 

1500 Cerachrome Melt @ 470 Sec 

Spar Melt @ 490 Sec 

1000 

500 

0 

Inconel melt 

Nextel melt 

Cerachrome melt 

Spar melt 

0 100 200 300 400 500 

Time (EI + Seconds) 

Figure 5-30. Thermal model prediction of wing spar burn through 

The final technique used was an examination of the wire failures on the wing leading 

edge spar. These wire runs are shown in Figure 5-29. The first measurement loss was 

a MADS upper left wing pressure measurement, which failed at approximately 

8:52:16 EST (EI + 487 sec.). This measurement is contained in the upper wire bundle 

in the left photo in Figure 5-29. The combination of these three separate and distinct 

analyses results in a range of wing spar breach times as early as 8:51:14 EST and as 

late at 8:52:16 EST (EI + 425 to 487 sec.). 

5-30 
Immediately after the breach, hot gas inside the wing began to heat the remaining wire 

bundles that contained real-time telemetry and the recorded MADS measurements. 

Figure 5-29 shows an inside-the-wing view of the approximate breach location based 

on this wire failure analysis. The view on the left is of the panel 8/9 spar location. This 

area is not visible in the right photo, which contains the transition spar and a view of the 

three wire bundles along the outside of the wheel well wall. Figure 5-31 shows the 

overall spar breach location with respect to the rest of the wheel well. The distance 

from the spar to the wheel well wall in this area is approximately 56 inches, as shown in 

Figure 5-32. Figure 5-33 shows a sketch of the venting of the left wing into the payload 

bay, driving the direction of the internal flow depicted in Figure 5-31. Note the 

142 square inch vent into the midbody fuselage located forward of the wheel well. 

Figure 5-31. Hot gas begins to fill left wing 

5-31 
9 

8 

7 

6 

Wing Leading

 Edge Spar 105 in. 

12 in. 

56.5 in. 

Figure 5-32. Columbia LH wing and wheel well geometry 

5-32 
 

 

Fuzz leakage at Hyd lines

Wheel well to Mid fuselage Gap behind retract actuator 2.4 IN2

Fuzz leakage at Hyd lines 

Wheel well to Mid fuselage 2 

MAIN GEAR WELL INTERNAL VENT

19.60 IN2 THROUGH SCREEN (5 dia)

Xo 1040, Yo 129, Zo 306

180 IN2

thru Xo 1040 spar

(Based on CAD model)

WING/MID FUSELAGE

INTERNAL VENT 142 IN2

HINGE BOXES to WING

4.5 IN2 fwd

2.5 IN2 mid

2.0 IN2 aft

Leakage behind stiffeners 0.20 IN2

0.02 in.2 each (0.16 in.2 total)

12 -.0.25  DIA DRAIN HOLES

ON Xo 1040 SPAR

MAIN GEAR WELL INTERNAL VENT 

19.60 IN2 THROUGH SCREEN (5 dia) 

Xo 1040, Yo 129, Zo 306 

180 IN2 

thru Xo 1040 spar 

(Based on CAD model) 

WING/MID FUSELAGE 

INTERNAL VENT 142 IN2 

HINGE BOXES to WING 

4.5 IN2 fwd 

2.5 IN2 mid 

2.0 IN2 aft 

Leakage behind stiffeners 0.20 IN2 

0.02 in.2 each (0.16 in.2 total) 

12 -.0.25  DIA DRAIN HOLES 

ON Xo 1040 SPAR 

Figure 5-33. Columbia LH wing and wheel well vent model (wheel well leak paths 

based on Atlantis test comparison) 

One hundred sixty-three other measurements failed very quickly over the next two 

minutes. Figure 5-34 is a plot of the percentage of measurements lost as a function of 

time from EI. The first measurements to fail are all located on the wing leading edge 

spar as shown by the purple line on the plot. Bundles 1, 3, and 4 along the wheel well 

(shown in Figure 5-35) began to fail about 14 seconds after the first spar 

measurements. This is indicative of a plume impinging on the wire bundles on the 

wheel well wall. 

Arc-jet testing was performed in a facility at JSC to demonstrate how quickly a hole in a 

0.1 inch thick aluminum plate would grow in an attempt to determine the feasibility of a 

rapid spar burn through followed by rapid wire measurement failures. The test 

configuration had an initial 1-inch diameter hole, and a stagnation heat rate of 

12.13 BTU per square foot per second (equivalent to the flight environment at about 

8:50 EST, EI ~ 351 sec). During the test, the aluminum plate hole grew from 1 inch in 

diameter to 5 inches in diameter in approximately 13 seconds. This was consistent with 

thermal math models developed to analytically predict hole growth. A wire bundle 

placed 15 inches from the heat source showed very rapid erosion after the hole grew to 

5 inches in diameter. The measurement losses in this bundle were very consistent with 

those observed for wire bundle number 3 (see Figure 5-34). The hole would have 

grown to a larger diameter; however, the test set-up provided a heat sink that 

eliminated further hole growth. Overall, the test indicated that a hole in the aluminum 

honeycomb would grow rapidly, allowing a substantial plume to destroy the wire 

bundles on the wheel well wall. 

5-33 
 Quantity of Sensor Signals Starting Loss - % 

100 

90 80 70 60 50 40 30 20 10 

0 

450 500 550 600 650 700 7

Bundle 1 (9 of 9) 

Bundle 4 (25 of 25) 

Bundle 3 (115 of 117) 

Leading Edge (18 of 18) 

V09T9895A 

Wheel Well Bit 

Flip 

L Elevon 

Accel Response 

1st OI Starts Failure 

1st Wheel Well Temp Rise 

(1700,1702 Bit change) 

L Elevon Accel Fail 

5th OI Starts Failure 

1st Orbiter Debris Event 

Reversal of Roll Moment and Start of Slow Aileron Trim Change 

6th OI Starts Failure 7th OI 

V07P9197A 

V07P80

Fl at Por l t i n o 3 B nd es Po ti on fo r 3 B und le s

Start LMG Struct Actuator Temp Rise 

Time from Entry Interface Sensors with Cables Along Leading Edge 

Start Loss 14 sec Earlier Than the 3 Bundles 

Figure 5-34. MADS data failure due to wire burning 

5-34 
Wing spar 

Hinge 

Bundle 

1 

4 

3 

Note: This is the 

bundle order along 

the sidewall only 

 

Figure 5-35. View of cables running along outside of wheel wall cavity bulkhead 

Four additional measurements failed after 8:54:16 EST (EI + 605 sec.) with the last 

starting to fail at 8:56:24 EST (EI + 735 sec.). This last measurement failure took over 

a minute to fail completely. Figure 5-34 and Figure 5-35 show the measurement loss 

timing and the wire bundle runs along the wheel well. Although almost all 

measurements internal to the wing failed, two strain measurements on the 1040 spar 

(forward of the wheel well) were unaffected and produced data until loss of all MADS 

data at 9:00:13 EST (EI + 964 sec.). Figure 5-36 shows the location of the sensors on 

the forward wheel well spar. The fact that these measurements are available until loss 

of data are important indicators that the RCC breach could not be forward on the panel 

6 area since these measurements would have been lost as well (see Figure 5-32). 

These measurements also record key changes in strain that help indicate some of the 

changes that the left wing underwent as the remainder of the entry events unfolded. 

5-35 
V12G9049A 

V12G9048A 

Figure 5-36. Strain measurements on 1040 spar 

A comprehensive evaluation of all MADS measurement failures in the left wing was 

performed to attempt to determine the spar breach progression. Each measurements 

failure signature was evaluated to determine both the start time of the failure and when 

the wire failure was complete. As the wire burns, the short between the twisted pairs of 

wiring increases thereby producing the time delay effect between failure initiation and 

complete failure. Eighteen measurements routed on the wing leading edge and over 

100 measurements in three wire bundles routed on the outboard side and external to 

the wheel well were used for this assessment. Seventeen of the 18 wing leading edge 

measurements failed in 10 seconds starting at 8:52:16 EST (EI + 487 sec.). No other 

MADS measurements failed during this time. The one measurement that did not fail in 

this time span was located in the bottom of five harnesses on the WLE. The remaining 

17 measurements are in the upper four harnesses with the majority being in the top 

harness. This would indicate that the spar breach started near the top wire bundle and 

worked toward the bottom of the spar. Additionally, two of the 18 measurements join 

the harnesses at the panel 7 to 8 interface, making the wing spar breach outboard of 

panel 8 very unlikely. 

5-36 
5.0 

5.4.4 Aerodynamic Events 

Coincident with the spar breach at approximately 8:52:09 EST (EI + 480 sec.), the nose 

cap RCC attach clevis temperature had an off-nominal change in rise rate when 

compared to previous missions. This rise rate was abnormal for approximately 

30 seconds and ended at 8:52:39 EST (EI + 510 sec.). The exact cause of this 

abnormal temperature rise is not known, but the timing is coincident with the breach of 

the wing leading edge spar. An explanation is that the abnormal rise is an 

instrumentation anomaly caused by multiple wire failures in this timeframe. This 

measurement response is shown in Figure 5-37 along with the same measurement for 

several other Columbia missions. This figure also shows the location of this internal 

nose cap measurement on the very forward portion of the left fuselage. 

8:52:09 8:52:39 EST 

V09T9889A 

Figure 5-37. Off-nominal temperature rise rate in nose cap RCC attach clevis 

5-37 
 

Nominal Sensor

Off Nominal Sensor

Left Main Gear Offline Sensor

Brake Line Temp D Expected Discrete Value

V58T1703A Unexpected Discrete Value

V58T0841A

Left Main Gear V58T0405A V58T0842A Brake Line Temp C

V58T1702A V58T0125A

V51X0125E

V51T0575A, 

V51P0571A, V51X0100X

V51P0573A

Inboard Tire Left Main Gear V58T1700A, Measurements Brake Line Temp A V58T1701A

V51X0130X

V51T0574A, Outboard Tire V51P0570A, Measurements V51P0572A

V51X0116X

Until approximately 8:52:17 EST (EI + 488 sec.), there was no indication of any offnominal situation that could be observed by the MCC or the crew in real-time. The flight 

control response and all orbiter telemetry measurements were nominal. The first 

indication of any anomalous Operational Instrumentation (OI) was a small increase in 

the left main gear brake line temperature D measurement at this time (see Figure 5-38). 

Within one second of this time, there is a lower 1040 spar (forward wall of the wheel 

well) strain measurement that shows an off-nominal increase in strain (see Figure 

5-39). These are both indications that the breach in the wing leading edge spar had 

allowed hot gas to reach the wheel well area, most likely through vent paths around the 

hinge covers, which allowed hot gas into the wheel well cavity. 

Nominal Sensor 

Off Nominal Sensor 

Left Main Gear Offline Sensor 

Brake Line Temp D Expected Discrete Value 

V58T1703A Unexpected Discrete Value 

V58T0841A 

Left Main Gear V58T0405A V58T0842A Brake Line Temp C

V58T1702A V58T0125A 

V51X0125E 

V51T0575A, 

V51P0571A, V51X0100X 

V51P0573A 

Inboard Tire Left Main Gear V58T1700A, Measurements Brake Line Temp A V58T1701A 

V51X0130X 

V51T0574A, Outboard Tire V51P0570A, Measurements V51P0572A 

V51X0116X 

Figure 5-38. Location of sensors in the LH wing wheel well 

5-38 
First off nominal indication at 52:17 

Strain (

μin. / in.) 

TIME (seconds after EI) 

EST 

Figure 5-39. Strain rise in lower 1040 spar cap 

At approximately 8:52:25 EST (EI + 496 sec.), the left outboard elevon wideband 

accelerometer showed an unusual 2 g peak-to-peak acceleration (see Figure 5-40). 

This is consistent with the timing of the many wire failures within the left wing and the 

timeframe when the spar breach occurs. Additionally, there are two unexplained 

communication dropouts in this same timeframe. Another 3 g peak-to-peak 

acceleration anomaly was observed at 8:52:31 EST (EI + 502 sec). Additional 

temperature measurements (left main gear brake line temp A and C) in the wheel well 

area (see Figure 5-38) began an off-nominal rise at 8:52:41 EST (EI + 512 sec.). 

5-39 
Time 

2.928

 0.00

 2.000

Acceleration (gs) 

EI + 496 sec. 

V08D9729 

 

 

 

Time (seconds) 1 : 9729 

Time (seconds) 2 : 1302 

LEOUTB ELE

RIG OUTB ELE

TRANSIENT (EI+496sec.- TRANSIENT (EI+502sec.-

Left 

Outboard 

Elevon 

-2.185495.00 496.00 498.00 500.00 502.00 503.00 

TIME (EI + sec.) Transient Transient 

- EI + 502 sec.-

(8:52:25 EST) (8:52:31 EST) Time 

2.928

 2.000 

Right 

Outboard 0.00 Elevon 

-2.185 

TIME (EI + sec.) 

Figure 5-40. Outboard elevon accelerometer responses at 8:52:25 and 

8:52:31 EST (EI + 496 and 502 sec.) 

The elevon response event was followed by a change in the rise rate for two supply 

nozzle temperatures and the vacuum vent nozzle temperatures at 8:52:32 EST (EI + 

503 sec.). The off-nominal rise rate occurred for approximately 15 seconds for the 

supply nozzle temperatures and 23 seconds for the vacuum vent temperatures. The 

location of these nozzles on the left side of the fuselage is shown in Figure 5-41, and 

the arrow indicates that they could have been located along a line of disturbed flow. 

The short, abnormal rise rate is unexplained, has not been observed on any previous 

Shuttle missions, and may not be associated with the upper wing disturbed flow caused 

by the leading edge damage. Unlike the RCC nose cap clevis temperature, the 

instrumentation appears to have been valid for both nozzles. Plots of this off-nominal 

temperature rise for one of the supply nozzle temperatures and the vacuum vent nozzle 

are shown in Figure 5-42. 

Immediately following these events, at 8:52:44 EST (EI + 515 sec.), the aerodynamic 

roll and yaw coefficients that have been extracted from the flight data showed a slight 

negative trend (see Figure 5-43). This is indicative of more drag and decreased lift on 

the left wing. Acceleration (gs) 

V08D9737 

495.00 496.00 498.00 500.00 502.00 503.00 

5-40 
Vacuum Vent 

Supply Dump 

Waste Dump 

No temp rise observed 

Vacuum Vent 

Supply Dump 

Waste Dump No temp rise observed 

Figure 5-41. Location of supply dump and vacuum vent nozzles 

Figure 5-42. Off-nominal temperature for supply nozzle and vacuum vent nozzle 

5-41 
o

e

AlActi

HydEl

Temp

Off-Nominal Ro

Delta Roll/Yaw Moment Coefficient 

STS 107 Delta Rolling/Yawing M ment Coefficients 

0.0025 0.0020 0.0015 0.0010 0.0005 0.0000 

-0.0005 -0.0010 -0.0015 -0.0020 -0.0025 

Delta Cll (Roll Moment) Delta Cln (Yaw Moment) Delta Cll Aero Model Delta Cln Aero Model 

Left OMS Pod TC BP0731T 

- Start Off Nominal Trend - Reduced Rise Rate - 13:49:49 

Left OMS Pod LRSI Surface Temp Left OMS Pod TC BP0732T Left OMS Pod TC BP0749T 

- Start Off Nominal Trend - Reduced Rise Rate - 13:49:59 

Left PLBD Surface TC BP3703T 

- Start Off Nominal Trend - Reduced Rise Rate - 13:50:09 

Initial Roll 

- 13:49:32 

Wing Frnt Spar Panel 9 TempWing LE 55 LWR Att. Clevis RCC 10 

- Start Off Nominal Trend

 - 13:51:14 

Left Wing Lower Surface TC 

- Start Off Nominal

 Temp Increase 

- 13:50:19 

LMG Brake Line Temp (D) 

- Start Off Nominal Trend - 13:52:17 

Left Wing Spar Cap 

- Off Nominal Strain Increase - 13:52:18 

LMG Brake Line Temps Start Off Nominal Trend

 - 13:52:41 

2 Temp Sensors Begin Off Nominal Response

 - V09T9895A - Wing Front Spar Panel 9 - V09T9849A - OB Elevon, Lower Surfac - 13:52:49.5 / 51.4 

Left INBD Elevon Lower Skin Temp 

- Start of Off Nominal Trend

 - 13:52:56 

4 Left OMS Pod Surf Temp 

- Change in Existing

 Off Nominal Trend

 - 13:52:39/ 53:09 

pha Mod 

ve 

- 13:53:31 

Debris #1 

- 13:53:44 / 48 

Debris #2

 - 13:53:46 / 50 

Debris #3 

- 13:53:54 / 58 

Debris #4

 - 13:54:00 / 04 

Debris #5

 - 13:54:07 / 11 

LMG Brake Line Temp B 

- Off Nominal Trend - 13:54:10 

Flash #1 

- 13:54:33.3 

Debris #6 

- 13:54:35 / 37 

Start Slow Aileron Trim Change 

Debris 

 13:54:20 

raulic System Left OUTBD / INBD 

evon Return Line Temps - OSL

 - 13:53:10 / 36 

erature Rise Rate Change 

- Hyd Sys 1 LMG UpLK Unlk Ln Temp

 - Sys 3 LMG Brake Ret Line Temp 

- LMG Brake Line Temp B,C

 - 13:56:16 / 22 

Left Upper Wing Skin Temp

 - Begin Off Nominal Trend - 13:56:24 

Debris #8 

- 13:55:21 / 27 - 13:57:19 / 29 

Debris #9, #10 

- 13:55:25 / 30 

Debris #12

 - 13:55:45 / 

Debris #15 

- 13:56:09 / 13 

Debris #13, #14

 - 13:55:55:59 

Debris #7

 - 13:55:04 / 10

 #11 

- 13:55:36 / 42 

Mid Fuselage Bondline Temp -&LH Aft Fus Sidewall Temp 

- Off Nominal - 13:54:22 

Debris #17 

Left Lower Wing Skin Temp

 - OSL - 13:57:28 

Left Upper Wing Skin Temp

 - OSL - 13:57:43 

Fuse Side Surf Temp Fuse Low Surf BF Temp

 - Start Off Nom Trend - 13:57:09 

Initiation1st Roll Reversal

 13:56:30 

Complete - 13:46:55 

MLG LH OB Tire Pressure #1, #2 

- Start Off Nom

 - 13:57:19 / 24 

Left Main Gear Strut Actuator Temp - Temp Rise Rate Chg. 

- 13:56:53 

Flare #2 

- 13:57:59.5 

/ 58:01.5 

Start Sharp Aileron Trim Increase 

- 13:58:03 

Flare #1, #2

 - 13:57:53.7 

- 13:57:59.5 

Fuse. Left PL

- Tem - 13:5

LeDo

- 

 - 

BFS FaulTire Pres

 - 13:58

MLG LH INBD Tire  - Pressu - 13:58:3

49:00.0 50:00.0 51:00.0 52:00.0 53:00.0 54:00.0 55:00.0 56:00.0 57:00.0 58:00.0 59:0

GM Time 13hrs:(min:sec) 

Figure 5-43. First noted off-nominal aero event (Greenwich Mean Time, GMT, is EST + 

5-42 
Hypersonic wind tunnel tests indicate that both the slight negative roll and yaw deltas 

can be explained by leading edge damage in the lower half of RCC panel 8. This will 

be discussed in more detail in Section 5.4.5. 

The flight control system compensated for the initial aerodynamic disturbance, and the 

aileron trim continued to match pre-entry predictions. To account for the increased 

drag the orbiter yawed slightly to the right to balance the yaw moments. The inertial 

sideslip exceeds flight history at 8:53:38 EST (EI + 569 sec.); however, this small 

departure was well within the vehicles capability to control. There was also another 

communication dropout in this timeframe (8:52:49 to 8:52:55 EST, EI + 520 to 

526 sec.). 

By 8:53:28 EST (EI + 559 sec.), Columbia had crossed the California coast. After this 

coastal crossing, there are indications of damage progression on the left wing since the 

temperature response on the upper surface changes and measurement losses continue 

in the left wing. At 8:53:29 EST (EI + 560 sec.), several left fuselage temperature 

measurements showed an unusual 400 degree temperature increase over the next 

minute. These measurement increases were accompanied by another short 

communications dropout. 

At 8:53:39 (EI + 570 sec.), four left OMS temperature measurements also exhibited an 

unusual temperature rise. This temperature rise is attributed to shifting of the left wing 

leading edge upper surface vortices due to interaction with disturbed flow caused by 

damage progression in the RCC panels 8 through 9 area. During this same timeframe, 

the upper cap strain gauge on the 1040 spar began an off nominal increase indicating 

continued heating internal to the left wing. 

By this point in the entry almost all measurements in the left wing had been lost and 

there appeared to be a temporary decrease in the measurement failure rate. It is 

possible that the lull in measurement failures was caused by a release of upper wing 

skin and FRSI. The breach in the upper wing surface and resulting pressure 

differentials internal to the wing would shift the internal plume impingement location 

relative to the wiring. In this same timeframe, there were multiple debris events 

captured on video by public ground observers beginning at 8:53:46 EST (EI + 577 sec., 

20 seconds after California coastal crossing) and ending at 8:54:11 EST (EI + 

602 sec.). The source of the debris may be upper wing skin and other thermal 

protection system (TPS) elements. 

Since the wing had been ingesting hot gas for over two minutes, it is quite probable that 

significant internal damage to the wing occurred over this timeframe. The aluminum 

structure in the wing was not designed for high heating and many of the components 

are unlikely to survive this heating environment. For example, aluminums melting point 

is ~1200 o

F, but the ingested gas into the wing may have been up to 8000 oF near the 

breach. There were other communication dropouts in this timeframe as well (8:53:32 to 

8:53:34 EST, EI +563 to 565 sec.). 

5-43 
 

In the 8:54:10 to 8:54:35 EST (EI + 601 to 626 sec.) timeframe, several key events 

occurred. The first event at 8:54:10 EST (EI + 601 sec.) was an indication that the 

temperatures in the wheel well had a greater rise rate, indicating the sneak flow or 

breach into the wheel area had increased. The roll moment trend changed sign at 

8:54:11 EST (EI + 602 sec.) almost simultaneously with the change in wheel well 

temperature rise rate as shown in Figure 5-44, followed by a debris flash event at 

8:54:33 EST (EI + 624 sec.). A large debris item, labeled debris event six, was seen 

leaving the orbiter two seconds later. This debris event and the roll moment trend 

change are believed to be created by growing damage to Columbias left lower wing. 

There are several theories that attempt to explain this response including a change in 

the wing camber or shape due to deformation and a lower wing recession, caused by 

the loss of much of the support structure internal to the wing. Wing recession here is in 

reference to structural deformation of the wing surface. In this case, one or more areas 

on the lower wing form a more concave shaped depression of wing skin and tiles as a 

result of the structural support in those areas being weakened or lost. A structural 

analysis of both wing deformation due to the loss of three ribs internal to the wing, and 

the wing recession concept was performed. 

Structural and aerodynamic analysis of wing deformation was performed without any 

type of recession in the wing lower surface. An assumption of 70 pounds per square 

foot was used for wing loading. Heat transfer coefficients were updated from previous 

coupled venting and thermal models for the left wing. A structural model was used 

along with the heating equivalent of a 10 inch diameter breach in the wing spar 

beginning at 8:52:16 EST (EI + 487 sec.) assuming 100 percent of the energy from this 

hole was transferred to the wing interior. Two different cases were analyzed: one with 

no breach in the upper wing skin and another with a 5 inch diameter breach in the 

upper skin at 8:54:37 EST (EI + 628 sec.) when a significant visual flash event was 

observed. Both cases showed that the temperatures of the wing skins, wing spars, and 

the wheel well wall were high enough by 8:58:19 (EI + 850 sec) that significant damage 

to the wing structure would occur. Figure 5-45 shows the potential area of damage and 

that significant deformation of the intermediate wing area and/or a recession in the 

lower surface is possible. 

It is difficult to postulate the exact wing deformation that occurred. One case is 

localized leading edge damage, resulting in global wing deformation. Local deformation 

was relatively small, less than 1.0 inch in the damage area with a 0.25-inch global 

increase in wing tip deflection. Delta rolling and yawing moments were calculated for 

this case, and they were very small, approximately +0.0001 for roll and -0.0001 for yaw. 

Another case looked at deformation resulting from the loss of three internal ribs. Again, 

the local deformation was small (approximately 5 inches) and the resulting aerodynamic 

moments were small. These delta moments are in the correct direction for change in 

the aerodynamic moments, but are not nearly large enough in magnitude when 

compared to the flight derived moments for the time after the roll moment trend 

reversal. To achieve the aerodynamic response observed in the flight data, more 

significant damage to the wing would have been required. 

5-44 
 

ffi

ilero

UTps

 Rat

G

rak

e 22

ng S

Off

:24

 #9,

5:25

Delta Roll/Yaw Moment Coefficient 

STS 107 Delta Rolling/Yawing Moment Coe cients 

-0.0025 -0.0020 -0.0015 -0.0010 -0.0005 

0.0000 0.0005 0.0010 0.0015 0.0020 0.0025 

Delta Cll (Roll Moment) Delta Cln (Yaw Moment) Delta Cll Aero Model Delta Cln Aero Model 

Left OMS Pod TC BP0731T

 - Start Off Nominal Trend 

- Reduced Rise Rate - 13:49:49 

Left OMS Pod LRSI Surface Temp Left OMS Pod TC BP0732T Left OMS Pod TC BP0749T

 - Start Off Nominal Trend 

- Reduced Rise Rate - 13:49:59 

Left PLBD Surface TC BP3703T

 - Start Off Nominal Trend 

- Reduced Rise Rate - 13:50:09 

Initial Roll

 - 13:49:32 

Wing Frnt Spar Panel 9 Temp Wing LE 55 LWR Att. Clevis RCC 10 

- Start Off Nominal Trend

 - 13:51:14 

Left Wing Lower Surface TC

 - Start Off Nominal 

 Temp Increase 

- 13:50:19 

LMG Brake Line Temp (D) 

- Start Off Nominal Trend - 13:52:17 

Left Wing Spar Cap 

- Off Nominal Strain Increase - 13:52:18 

LMG Brake Line Temps Start Off Nominal Trend

 - 13:52:41 

2 Temp Sensors Begin Off Nominal Response 

- V09T9895A - Wing Front Spar Panel 9

 - V09T9849A - OB Elevon, Lower Surface - 13:52:49.5 / 51.4 

Left INBD Elevon Lower Skin Temp 

- Start of Off Nominal Trend

 - 13:52:56 

4 Left OMS Pod Surf Temp

 - Change in Existing 

Off Nominal Trend

 - 13:52:39/ 53:09 

Alpha Mod Active 

- 13:53:31 

Debris #1

 - 13:53:44 / 48 

Debris #2

 - 13:53:46 / 50 

Debris #3

 - 13:53:54 / 58 

Debris #4 

- 13:54:00 / 04 

Debris #5

 - 13:54:07 / 11 

LMG Brake Line Temp B

 - Off Nominal Trend - 13:54:10 

Flash #1

 - 13:54:33.3 

Debris #6

 - 13:54:35 / 37 

Start Slow A n Trim Change 

 13:54:20 

Hydraulic System Left O BD / INBD Elevon Return Line Tem - OSL

 - 13:53:10 / 36 

Temperature Rise e Change 

- Hyd Sys 1 LM UpLK Unlk Ln Temp - Sys 3 LMG B e Ret Line Temp - LMG Brak Line Temp B,C

 - 13:56:16 / 

Left Upper Wi kin Temp 

- Begin Nominal Trend - 13:56 

Debris #8 

- 13:55:21 / 27 

Debris #10 

- 13:5 / 30 

Debris #12 

- 13:55:45 / 

Debris #15

 - 13:56:09 / 13 

Debris #13, #14 

- 13:55:55:59 

Debris #7 

- 13:55:04 / 10 

Debris #11 

- 13:55:36 / 42 

Mid Fuselage Bondline Temp &LH Aft Fus Sidewall Temp 

- Off Nominal - 13:54:22 

Debris #17 

- 13:57:19 / 29 

Left Lower Wing Skin Temp

 - OSL - 13:57:28 

Left Upper Wing Skin Temp

 - OSL - 13:57:43 

Fuse Side Surf Temp Fuse Low Surf BF Temp

 - Start Off Nom Trend - 13:57:09 

1st Roll Reversal Initiation - 13:56:30 

Complete - 13:46:55 

MLG LH OB Tire Pressure #1, #2

 - Start Off Nom - 13:57:19 / 24 

Left Main Gear Strut Actuator Temp - Temp Rise Rate Chg. 

- 13:56:53 

Flare #2 

- 13:57:59.5 

/ 58:01.5 

Start Sharp Aileron Trim Increase

 - 13:58:03 

Flare #1, #2

 - 13:57:53.7 

- 13:57:59.5 

Fuse. Side Surf TC & Left PLBD Surface TC 

- Temp Increase to O - 13:59:29 

Left Main Gear Downlock Indicati - Transferred ON - 13:59:06 

BFS Fault Message (4) Tire Pressures 

- 13:58:40 / 56 

MLG LH OutBD & INBD Tire Pressure #1 - Pressure Trend to OSL - 13:58:32 

Sys 2 LH BVlv Return  - Start of s

 Dwnrd. 

 - 13:59:22 

Sharp Change in Rol

49:00.0 50:00.0 51:00.0 52:00.0 53:00.0 54:00.0 55:00.0 56:00.0 57:00.0 58:00.0 59:00.0 00

GM Time 13hrs:(min:sec) 

Figure 5-44. Sharp change in rolling moment (GMT is EST + 5 hours) 

5-45 
 

Modulus Reduced to 1% Modulus Reduced by 50% Modulus Reduced by 30% 

-Intermediate Rib Struts -Upper & Lower Intermediate -Xo1191 Spar 

-Upper H/C Panels Xo1040 to Xo1140 Wing Skin Panels -Yo167 Rib 

-WLE Web Aft of RCC Panels 8 & 9 -Intermediate Wing Rib Caps 

Figure 5-45. Modeling results show potential area of damage 

 and that significant deformation of the intermediate wing area 

 and/or a recession in the lower surface are possible 

Based on the structural analysis, it appears feasible that a wing recession occurred and 

resulted in the large positive slope in the delta rolling moment. The recession was 

caused by severe prolonged heating internal to the left wing that melted many of the 

support struts. Once the struts were lost, the wing skin lost structural support, a 

concave cavity developed, and some lower surface tiles may have been lost. Wind 

tunnel testing has shown that this type of cavity can cause the change in delta aero 

moments derived from the aerodynamics reconstruction in this timeframe. The aero 

moment change is a negative yaw moment due to increased drag and a positive roll 

moment due to increased lift on the left wing. This testing will be discussed in more 

detail in the next section. 

The flash event in this timeframe could indicate a loss of either upper surface wing skin 

or thermal blankets or a release of molten material into the environment around the 

orbiter. More than 10 debris events followed and were observed by various public 

videos in the 8:55:04 to 8:56:00 EST (EI + 655 to 720 sec.) timeframe. Several of 

these events were large, consisting of a shower of particles, and lead to a brightening of 

the plasma trail. Another communication dropout followed at 8:56:00 EST (EI + 

720 sec.). 

5-46 
 

 

 

 

 

 

5.0 

5.4.5 Wheel Well Gas Penetration and Final Aerodynamic Events 

By 8:56:16 EST (EI + 727 sec.), hot gas had penetrated the wheel well wall as indicated 

by off-nominal rates rise rates in several hydraulic line temperatures (see Figure 5-46). 

Preliminary analysis indicates that it is feasible to have had some gas intrusion into the 

wheel well area as early as 8:52:39 EST (EI + 510 sec.) since a honeycomb access 

panel could melt as quickly as 22 seconds after the wing spar is breached. Additionally, 

there are various vent paths into the wheel well around the landing gear door hinge 

covers. 

50 V58T1700A - Brakeline A 

V58T0125A Upr Act Unlock Line 

V58T0405A Strut Actuator 

Normalized to 0 Port-side data from 13:50 to LOS 

V51T0574A Outboard Wheel Rim 

45 V51T0575A Inbd Wheel Rim 

V58T1701A Brakeline B 

V58T1702A Brakeline C Start of Significant Trend Upward 

40 

V58T1703A Brakeline D (Under Cover) 

V58T0841A Sys2 Sw itch Vlv Rtn Line (Under Cover) 

V58T0842A Sys3 Sw itch Vlv Rtn Line 

Supply Dump Nozzle Temp Rise 13:52:32 

V09T1006 Elevon Skin OSL 13:52:59 

Large breach into Wheel Well 

Debris 1 13:53:36 

Debris 2 13:53:47 

Debris 3 13:53:56 

35 Debris 4 13:54:03 

Debris 5 13:54:07 

Start of Slow Elevon Trim 13:54:20 

V34T1106 LH Struct x1215 Rapid Rise 13:54:22 

Flash 1 13:54:32 

Normalized Value 

Debris 6 13:54:35 

30 Start of Sharp Elevon Trim 13:58:03 

25 

Sharp Elevon Trim 

20 

15 

10 

Supply Nozzle 

V09T1006 Elevon Skin OSL 

Slow Elevon Trim

Flash 1 ise V34T1106 LH Temp R 

Debris 1-5 

5 

0 

13:52:19GMT 13:53:46 13:55:12 13:56:38 13:58:05 13:59:31 14:00:58 

54:10 56:16

Figure 5-46. Temperature data in left wheel well trends up (GMT is EST + 5 hours) 

The centerline of the plume contained enough energy to begin melting the exterior of 

the wheel well wall by approximately 8:54:00 EST (EI + 594 sec.). The modeling 

described in the previous sub-section assumed the heating equivalent of a 5 inch 

diameter hole in the leading edge spar, but does not include the complex thermal 

interaction with struts and other wing spar structure internal to the wing. The intent of 

the analysis is to show that it is feasible to obtain the temperature response shown in 

Figure 5-46, including the early response that was seen in the 8:54:10 to 8:55:10 EST 

(EI + 601 to 666 sec.) timeframe, when several left main gear brake line temperatures 

and strut actuator temperatures began an off-nominal rise. 

Immediately after the wheel well wall was breached, a hot gas plume began to flow on 

to the left main gear strut (depicted in Figure 5-47), leading to excessive strut erosion. 

A wheel well wall breach in this area is consistent with the erosion pattern observed on 

5-47 
 

 

the recovered left main gear strut. Even after damage had significantly progressed into 

the wheel well, the orbiter initiated the first roll reversal at 8:56:30 EST (EI+741 sec.). 

The maneuver was completed by 8:56:55 EST (EI+766 sec.), and the vehicle was in a 

normal left bank. The guidance and flight control systems were performing nominally, 

although the aileron trim continued to slowly change to counteract the additional drag 

and lift from the left wing. 

Figure 5-47. Hot gas breaches the wheel well 

5-48 
 

Imaged at 8:57:14 EST (EI + 785 sec.), the Kirtland photo could indicate a flow 

disturbance on the leading edge of the left wing and/or flow leaving the leading edge 

of the left wing (see Figure 5-48). It also appears to show a disturbed flow leaving 

the trailing edge of the left wing. Other images, not shown here, also show 

disturbed flow on the upper side of the left wing, indicating that the damage and 

venting through the upper RCC vent was deflecting the flow upward. The Kirtland 

photo is a digital still image taken by off-duty employees of the Starfire Optical 

Range at Kirtland Air Force Base, New Mexico, during the STS-107 entry using a 

3.5 inch telescope through a computer controlled 1 meter rotating mirror. 

Figure 5-48. Kirtland photo 

5-49 
At 8:58:32 EST (EI + 863 sec.), there was a change in stress measured on the left wing 

1040 spar (main landing gear forward wall spar), as indicated by strain gauge 

measurements that began trending up at 8:52:18 EST (EI + 489 sec.). At 8:58:32 EST 

(EI + 863 sec.) the lower cap strain (near the bottom of the 1040 spar) essentially 

returned to a normal measurement when compared to previous flights (see previous 

Figure 5-39). The upper strain measurement continued to increase during this 

timeframe indicating that the heating was different on the upper and lower portions of 

the wheel well and 1040 spar. Although the response of the 1040 spar strain is not 

completely understood, structural analysis indicates that the strain response can be 

completely explained by thermal stresses caused by severe heating of the wheel well 

wall and internal wing components. 

As shown in Figure 5-49, this strain response appears to be consistent with another 

sharp change in the slope of the derived delta rolling moment coefficient that occurred 

slightly earlier at 8:58:03 EST (EI + 834 sec.), along with several additional debris 

events. The vehicle responded to this event with a sharp change in the aileron trim. 

These events indicate that there was another significant change to the left wing 

configuration at this time. Wing deformation and an increase in the lower surface 

recession along with a loss of additional bottom tiles are possible explanations for this 

behavior. 

At the same time as the stress was relieved on the lower 1040 spar, two left main gear 

outboard tire pressures began trending toward an off-scale low reading. This was 

preceded by a slight upward trend at 8:57:19 EST (EI + 790 sec.) for both pressure 

measurements. This is an indication of extreme heating of both the left outboard tire 

and the surrounding instrumentation. The tire has significant thermal mass and 

substantial heating would be required to produce the slight temperature rise. By 

8:58:56 EST (EI + 887 sec.), all left main gear inboard and outboard tire pressure and 

wheel temperature measurements were lost indicating a rapid progression of damage 

or wire burning inside of the wheel well. Figure 5-38 shows the location of these 

pressure sensors. 

At 8:59:06 EST (EI + 897 sec.), the left main gear downlocked position indicator 

changed state. There are indications that the gear did not come down until after Loss 

of Signal (LOS) because the left main gear uplock position indicator still showed the 

gear in the stowed position, and the left main landing gear door latch position indicator 

showed that the door was still closed. Additionally, there are several measurements on 

the strut that produce valid data until final loss of telemetry in the MCC. 

5-50 
/Yaw

0.0025

Delta Cll (RolDelta Cln (YaDelta Cll Aero

Model

 Spar Panel 9 Temp

 LWR A tt. Clevis

 Nominal

se

9

 13:52:56

d Surf Temp

 in Existing

inal Trend

/ 53:09

BD

13:5

Left Up

n Off No

56:24

Debris #9, #10

13:55:25 /

 - 13:57:09

Press ur

 Start 13:5

13:5

Significant Changes in Aerodynamics Final Changes in Aero

STS 107 Delta Rolling ing Moment Coefficients 

Delta Roll/Yaw Moment Coefficient 

0.0020 0.0015 0.0010 0.0005 0.0000 

-0.0005 -0.0010 -0.0015 -0.0020 -0.0025 

l Moment) 

w Moment) 

 Model 

Delta Cln Aero 

Le ft OMS Po d TC BP07 31T

 - Sta rt Off N omin al Tren d 

- Red uce d Ri se Ra te

 - 13:49:49 

Le ft OMS Po d LR SI Su rfa ce Temp Le ft OMS Po d TC BP07 32T Le ft OMS Po d TC BP07 49T

 - Sta rt Off N omin al Tren d 

- Red uce d Ri se Ra te

 - 13:49:59 

Left PLBD S urface TC BP3 703T

 - Sta rt Off N omin al Tren d 

- Red uce d Ri se Ra te

 - 13:50:09 

Initial R oll 

-

 13:49:32 

Wi ng FrntWing LE 55 RCC 10

 - Start Off Nominal Trend 

- 1 3:51:14 

L eft Wi ng Lo wer Su rfa ce TC 

- Start Off Nomi nal 

Temp Increase 

- 13 :5 0:19 

LMG B rake Line Temp (D)

 -

 Sta rt Off No min al Tren d

 -

 13:52 :1 7 

Left Wing S par C ap

 -

 Off Nomi nal Strain Increase

 -

 13:52 :1 8 

LMG Brak e Line Temps Start Off Tren d 

-

 13 :5 2:41 

2 Temp Senso rs Begin Off No minal Respon

- V09T9895A - Win g Fro nt S par Panel - V09T9849A - OB El evon, Lower S urface

 - 13 :5 2:49 .5 / 51.4 

Le ft IN BD El evon Low er Skin Te mp 

- Start o f Off Nomin al Tren d

 -

4 Le ft OMS Po

 - Change

 Off Nom

 - 13:52:39 

Alpha Mod Active 

- 13 :53:31 

Debris #1

 - 13 :5 3:44 / 48 

De bris #2 

- 13:53:46 / 50 

Deb ris #3 

- 13 :53:54 / 58 

Deb ris #4 

- 13:54:00 / 0 4 

De bris #5

 - 13:54 :0 7 / 1 1 

LMG Brake Lin e Temp B 

- Off No min al Tren d - 13:54:10 

Fla sh #1 

- 13 :5 4:33 .3 

Debris #6 

Te mpe rature R ise R ate Ch ang e

- 13 :5 4:35 / 37 

Start Slo w Aileron Trim Change

Deb ris 

  13 :5 4:20 

H ydrauli c S yste m Left OUTBD / IN E levo n Re tu rn Li ne Temp s - OSL

 -

13:53:10 / 36

 - H yd Sys 1 LMG U pL K U nl k Ln Temp 

- S ys 3 LMG Brake Ret Li ne Tem p - LMG Brake Line Tem p B,C

 - 6:16 / 22 

per Wing Skin Temp 

- B egi m ina l Trend - 13: 

De bri s # 8 De bris #1 7

 - 13 :5 5:21 / 27

 - 3 0 

De bris # 12

 - 13:5 5:45 / 

De bris # 15 

- 13 :5 6:09 / 13 

Debris #13, #14 

- 1 3:55:55 :59 

Deb ris #7 

- 13:55 :04 / 10

 #1 1 

&LH Aft Fus Sidewall Tem p 

- Off Nomin al - 13 :5 4:22 

-

-

13:55:36 / 42

Mid Fu sela ge Bo nd lin e Temp 

13 :5 7:19 / 29 

Left Lo wer Wing Ski n Temp 

- OSL - 13:57 :28 

Left Upper Wing Ski n Temp 

- OSL - 13:57 :43 

Fuse Side Surf Temp Fu se Lo w Surf B F Temp

 - Sta rt Off N om Trend 

1st Roll Reversal Initiation - 13:56:30 Complete - 13:46:55 

MLG LH OB Tire 

e #1, #2 

- Off Nom - 7:19 / 24 

Left Main Gear Strut Actu ator Temp - Temp Rise Rate Chg. 

- 6:53 

Fl are #2 

- 13 :5 7:59.5 

/ 58:01.5 

Start Sharp Ai le ron Trim Increase 

- 13 :58:03 

Fl are # 1, # 2

 - 13:5 7:53.7 -

 13:5 7:59.5 

Fus e. Side SuLeft PLBD Su - Temp Incre

 - 13:59:29 

Left Mai n D own lock  - Tra nsfe - 13:59:0

BFS Faul t Mes sTire P ressures 

- 1 3:58:40 / 5

MLG LH OutBD IN BD Tire Pre ss

- Pre ssure Tren 

 - 13 :58:32 

SV

- 

 - 

49:00.0 50:00.0 51:00.0 52:00.0 53:00.0 54:00.0 55:00.0 56:00.0 57:00.0 58:00.0 59:00.0 

GM Time 13hrs:(min:sec) 

Figure 5-49. Increased wing deformation and wing recession leads to signific

aerodynamic changes (GMT is EST + 5 hours) 

5-51 
As shown in Figure 5-49, there was another abrupt change in the vehicle aerodynamics 

caused by the continued progression of left wing damage at 8:59:26 EST (EI + 917 

sec.). This change was a significant increase in the positive delta roll moment and 

negative delta yawing moment, indicating increased drag and lift from the left wing. 

Columbia attempted to compensate for this by firing all four right yaw jets. By this point 

the MCC had lost all telemetry data at 8:59:23 EST (EI + 914 sec.). Even with all four 

right yaw jets and a maximum rate of change of the aileron trim, Columbia was unable 

to control the side-slip angle that was slightly negative (wind on the left side of the 

fuselage) during much of the entry. The side-slip angle changed sign at 8:59:36 EST 

(EI + 927 sec.) indicating that vehicle loss of control was imminent (side-slip angle is an 

aerodynamics terms for the angle between the relative wind velocity and the vehicle 

direction of motion, or velocity vector). 

A large piece of debris was observed leaving the orbiter at approximately 8:59:46 EST 

(EI + 937 sec.). Five additional debris events and two flash events were observed over 

the next thirty seconds. MADS recorder data was lost at approximately 9:00:14 EST 

(EI + 965 sec.) and main vehicle aerodynamic break-up occurred at 9:00:18 EST 

(EI + 969 sec.), based on video imagery. 

In the Mission Control Center, the Entry Flight Control Team waited for tracking data 

from the Eastern Range and communication link handover to the Merritt Island Launch 

Area (MILA) ground station. There was no radio frequency (RF) communication 

received from the vehicle at MILA and no valid tracking data was ever produced at the 

Eastern Range since the vehicle never crossed the KSC area horizon. Shortly 

thereafter, Mission Controls Entry Flight Director implemented contingency action 

procedures. 

5-52 
 

 

 

 

5.0 

5.5 AERODYNAMIC RECONSTRUCTION 

As previously discussed, the flight-derived aerodynamic moments use the high altitude 

winds and atmosphere developed by the DAO and represent the most accurate 

reconstruction that is possible based on available data. Many different scenarios were 

proposed to define the damage necessary to match this reconstruction using wind 

tunnel testing and CFD analyses at facilities across the United States. These scenarios 

include individual and multiple full or partial missing RCC panels, a missing landing 

gear door, a deployed left main landing gear, missing lower surface tiles, holes through 

the wing, lower surface deformation, and several others. The details for all of the 

options that did not match the flight-derived data will not be discussed here and are 

beyond the scope of this document. 

Figure 5-50 through Figure 5-52 show the flight derived delta roll, yaw, and pitch 

aerodynamic moments, respectively, along with the tested damaged configuration 

results that are consistent with the flight data. The intent of the remainder of this 

section is to briefly discuss the CFD analysis and wind tunnel testing used to replicate 

the derived delta aero moments. Note the time scale on these plots is in seconds from 

EI. delta roll 

0.003 

0.002 

0.001 

0.000 

-0.001 

-0.002 

200 400 600 800 1000 

 flight - sim f 

lower rcc 8 

lower rcc 8 + slot 

full rcc 9 

full rcc 9 + long dim ple 

full rcc 9 + 2x long dimple 

full rcc 9 + 3x long dimple 

roll trend reverses as lower surface depression is included 

800 s - subjective 

300 s 500 s 875 s - subjective 

615 s  increased heating on side fuselage 

(RTD 1106, 1724) 

900 s - subjective 

time from EI, sec 

Figure 5-50. Wind tunnel testing configurations that match delta roll 

moment data 

5-53 
 

 

 

 

 

delta pitch 

 flight - sim f 

-0.003 

-0.002 

-0.001 

0.000 

0.001 

0.002 

delta yaw

lower rcc 8 

lower rcc 8 + slot 

full rcc 9 

full rcc 9 + long dimple 

full rcc 9 + 2x long dimple 

full rcc 9 + 3x long dimple 

yaw trend continues as lower 

surface depression is included 

800 s - subjective 500 s 

300 s 

875 s - subjective 615 s  increased heating on side fuselage 

(RTD 1106, 1724) 

900 s - subjective 

200 300 400 500 600 700 800 900 1000 

time from EI, sec 

Figure 5-51. Wind tunnel configurations that match delta yaw moment data 

0.003 

0.000 

-0.003 

-0.006 

 flight - sim f, modified 

lower rcc 8 

lower rcc 8 + slot 

full rcc 9 

full rcc 9 + long dim ple 

full rcc 9 + 2x long dimple 

full rcc 9 + 3x long dimple 

pitch trend continues as lower 

surface depression is included 

800 s - subjective 

300 s 500 s 

875 s - subjective 

615 s  increased heating on side fuselage 

(RTD 1106, 1724) 

flight residual has bias 

removed (+0.0015) 

900 s - subjective 

200 400 600 800 1000 

time from EI, sec 

Figure 5-52. Wind tunnel configurations that match delta pitch 

moment data 

5-54 
As discussed previously, the reconstructed aerodynamic moments showed little to no 

change due to damage through 8:52:29 EST (EI + 500 sec.). Based on the forensics 

data discussed in Section 5.3, the most likely region of initial damage was in the lower 

part of RCC panel 8. Wind tunnel testing in the Langley Research Center (LaRC) CF4 

tunnel indicated that a missing bottom part of RCC panel 8 (from the apex to the lower 

carrier panel) matches the initial aerodynamic increments, which show a minimal effect 

on the overall vehicle aerodynamics. In fact, even a full missing panel 8 produces only 

a small change to the roll and yaw aero moments. Figure 5-53 and Figure 5-54 show 

the results of this LaRC evaluation. 

0.002 

Re/ft = 0.46 

Re/ft = 0.33 

full panels out 

0.001 

0.000 

-0.001 

-0.002 

3 4 5 6 7 8 9 10 

panel number 

Figure 5-53. Delta roll for lower half and full panel RCC panel missing 

delta roll 

5-55 
delta yaw 

0.002 

Re/ft = 0.46 

Re/ft = 0.33 

full panels out 

0.001 

0.000 

-0.001 

-0.002 

3 4 5 6 7 8 9 10 

panel number 

Figure 5-54. Delta yaw for lower half and full panel RCC panel missing 

The left wing spar was breached in the 8:51:14 EST (EI + 425 sec.) to 8:52:15 EST 

(EI + 487 sec.) timeframe. Initially, the spar breach had little to no effect on the derived 

aero moments. Over time the leading edge damage progressed and a slot or upward 

deflection of the flow through the upper carrier panel 8 developed. The combination of 

flow through the wing leading edge and flow through a slot onto the upper carrier panel 

is consistent with the first observed aerodynamic response, which occurs at 8:52:29 

EST (EI + 500 sec.). This can be observed in Figure 5-50 and Figure 5-51 as a slow 

negative trend in delta roll and yaw. Figure 5-55 shows the wind tunnel test results for 

three different cases: missing lower RCC panel 8, missing panel 8 combined with a 4 

inch diameter hole in the upper carrier panel 8, and missing RCC panel 8 with a slot 

through the upper carrier panel. The slot was shown to produce the increased delta 

yaw observed during flight as well as the upper surface flow disturbances on the OMS 

pod and left side fuselage, which were also observed during flight by abnormal 

temperature rise rates. 

5-56 
Lower RCC 8 only 

 Flow through slot in 

upper carrier panel (panel 8) affects both leeside side fuselage and OMS pod. 

 w/ Hole only Small 

effect on OMS pod 

 Slot increases Delta 

Yawing moment 

Hole diameter 4.0 in Hole area 12.6 in2 

Lower RCC 8 + hole through carrier panel 

delta yaw 

Lower RCC 8 + slot throug

0.002 0.001 0.000 

-0.001 -0.002 

1 - lo2 - lo3 - lo

0 1 2 

damage scenario 

Figure 5-55. Wind tunnel testing results for missing lower carrier panel 8 and a slot and hole 

5-57 
 

It is possible that a hole through the upper wing developed in the 8:54:11 EST (EI + 

600 sec.) to 8:54:31 EST (EI + 630 sec.) timeframe. This is consistent with a lull in the 

measurement failure rate in the wire bundle along the wheel well that was discussed 

earlier in Section 5.4.3. Wind tunnel testing and CFD analysis have shown that sizable 

holes through the wing have little to no effect on the aerodynamics and heating on the 

left side of the fuselage and OMS pod. A representative sample of this work is shown 

in Figure 5-55 as damage scenario number 2. 

The next significant aerodynamic event occurred at 8:54:11 EST (EI + 602 sec.) when a 

dramatic reversal of the aero rolling moment trend occurs. By this time hot gas and an 

internal plume environment had severely degraded the structural integrity of some of 

the intermediate wing support structure leading to wing deformation. Three different 

configurations were tested to validate the theory of wing deformation. The first involved 

global wing deflection of up to 0.79 inches due to damage and is depicted in Figure 

5-56. CFD analysis of this configuration showed extremely small aero moment 

response for yaw and roll that does not match the flight derived data. More substantial 

local wing skin deformation with a maximum deflection of 5.1 inches due to three ribs 

missing along with other internal wing damage was examined using CFD tools. Again, 

these results produced only small aerodynamic moment changes, which did not match 

the flight-derived data. 

 Case 4 Aerodynamic Increments   Hand Calculation Increments  

Global wing deflection due to damage Local wing deformation due to damage 

Max RMS Delta ∆Wing Deflection Max ∆Wing Deflection, Zo = 5.1 

(Nominal  Damaged structure) is 0.79 • ∆Cl = 0.000, ∆Cn = -0.0001  Newtonian 

• ∆Cl = 0.0001, ∆Cn = -0.0001 • ∆Cl = 0.00005, ∆Cn = -0.00014  Cart3D 

• ∆Cl = 0.00012, ∆Cn = -0.0001  FELISA 

Figure 5-56. CFD analysis of wing deformation 

The third and final configuration that was tested was a depression in the lower surface 

of the wing caused by the significant structural damage caused by the hot gas plume 

environment internal to the wing. A previous section (5.5.5) discussed the internal 

5-58 
 

MLG door

RCC 9

structural damage that was most probable in this timeframe. LaRC wind tunnel testing 

and CFD analysis were performed for several different configurations of lower surface 

recessions shown in Figure 5-57. The data shows that it is feasible for a recession to 

cause the change in the rolling moment sign when combined with some portion of RCC 

panel 9 missing at 8:54:11 EST (EI + 602 sec.). 

Windward Surface Dimples RCC 9 

Depth 

5.33 full scale (0.040 model) 

Three Lengths 

L/h = 10 (X = 1093 to 1141) 

L/h = 19 (X = 1093 to 1191) 

L/h = 28 (X = 1093 to 1243) 

Three Widths starting with center box 

Single Width Y = -205 to -228 

Double Width Y = -182 to -228 

Triple Width Y = -182 to -249 

Note: the bottom of the dimple was milled 

parallel t the XY plane. Due to wing taper 

& dihedral, the inboard dimple was slightly 

deeper and outboard more shallow 

MLG door 

* Test with and without WLE RCC Panel 9 

Figure 5-57. LaRC wind tunnel testing of lower surface depressions 

Initially, the recessed area would have been relatively small; however, it would gradually 

grow over time to cause the delta roll moment to increase. By 8:57:29 EST (EI + 

800 sec.) wind tunnel testing showed that the depression has to be on the order of 

20 feet long, two feet wide, and 5.3 inches deep along with panel 9 missing in order to 

duplicate the delta roll coefficient shown in Figure 5-50. This configuration provides a 

delta yaw moment that is slightly larger than was observed, but is consistent with a 

decreasing negative delta yaw moment observed in this timeframe (Figure 5-51). 

A little more than a minute later, at 8:58:44 EST (EI + 875 sec.), the width of this 

recession would need to have increased by another two feet to match the aerodynamic 

delta roll and yaw moments. At this point, the rate of change of the aerodynamic 

moments and damage progression is so great that it likely grew by about an additional 

two feet in width over the next 25 seconds at 8:59:09 EST (EI + 900 sec.). Additionally, 

the delta pitch moment was now observed to deviate from previous mission 

reconstructions in this timeframe. Figure 5-52 shows that the recession concept is 

5-59 
consistent with the delta pitch moment reconstruction. Previous structural analysis 

indicates that by 8:58:19 EST (EI + 850 sec.) there is large-scale wing deformation and 

thus the possibility of a large recession is plausible in this timeframe. 

In summary, the latest aerodynamic wind tunnel testing and CFD analysis performed to 

date indicate that the initial damage was probably relatively small, like a hole and/or 

missing part of the bottom of RCC panel 8. A slot then developed so that there is 

upward flow through the RCC vent and across the upper 8 carrier panel. Later, more of 

RCC panel 8 and/or panel 9 is lost along with some substantial wing deformation 

probably involving a locally depressed area on the lower wing surface. The wing 

deformation and lower surface recession gradually increased over time, and eventually 

the yaw and roll moments were too great for the flight control system to manage, 

leading to a loss of vehicle control and aerodynamic break-up. 

Although this aerodynamic reconstruction represents a reasonable sequence of vehicle 

configurations that led to loss of control during entry, it is not meant to be interpreted as 

an exact literal sequence of events. The wind tunnel testing and analysis was 

performed using representative geometries; however, the actual specific vehicle 

damage is unknown and may never be known completely. The sequence of events 

discussed here is consistent with the reconstructed aero moments, MADS data, and 

forensics data and provides the best aerodynamic, thermal, and structural 

understanding possible for the eventual loss of Columbia. 

5-60 
6.0 RE-USABLE SOLID ROCKET MOTOR 

All Reusable Solid Rocket Motor (RSRM) investigation fault tree legs have been closed 

for the STS-107 RSRM set, RSRM-88. All Contract End Item (CEI) performance 

specifications were met including all flight individual and paired motor requirements. 

Postflight inspections revealed a tear in the right-hand nozzle flex boot that is 

considered an IFA (STS-107-M-01), but the tear is attributed to thrust tail-off or 

splashdown events. A slightly low out-of-family thrust level was observed for the righthand motor in the 113.5 to 114.5 second interval during thrust tail-off, but the resultant 

thrust imbalance was still within family experience and CEI limits. The new experience 

has been reviewed and accepted as being within the statistical expectations for the 

RSRM motor population and is attributed to the increased population sample size (see 

Section 3.4 for more details). 

6-1 
 

7.0 SOLID ROCKET BOOSTER 

The Solid Rocket Booster (SRB) fault tree for the STS-107 SRB set, SRB BI116, 

remains open due to possible debris sources at the forward SRB/External Tank (ET) 

separation bolt catcher assembly and the forward Booster Separation Motors (BSM). 

The STS-107 SRBs performed nominally and there were no reported SRB IFAs. 

Four blocks on the STS-107 SRB fault tree remain open pending completion of forward 

bolt catcher testing. The bolt catcher, shown in Figure 7-1, was not qualified as an 

assembly, and structural qualification testing was not representative of the current flight 

configuration. The exact magnitude of loads transmitted to the bolt catcher housing 

cannot be determined based on available data. The SLA-561 thermal protection 

system (TPS) material on the bolt catcher was qualified by test and analysis for general 

ET application, but no pyrotechnic shock testing was performed. There is no test data 

available on the bolt catcher honeycomb dynamic crush strength versus separation bolt 

velocity, and random pressure loading from the NASA Standard Initiator (NSI) ejection 

was not included in the original qualification tests. Lastly, the running torque/breakaway torque was not measured during STS-107 bolt catcher fastener and ET range 

safety system (RSS) fairing installation, which is used to verify the insert locking feature 

is in place. A review has determined that the bolt catchers and RSS fairings were 

installed and secured for flight with the correct bolts and final torque. Testing is in work 

to close the four remaining fault tree blocks, but initial static tests results show failure 

below the required safety factor of 1.4. 

Two other blocks on the STS-107 SRB fault tree remain open that pertain to potential 

debris from the forward BSMs. Inspection of the forward BSMs found no indication of 

unburned propellant or any indication that the BSMs contained any Foreign Object 

Debris (FOD). The two debris related fault tree blocks will remain open pending 

transport and impact analysis. 

7-1 
Figure 7-1. Details of SRB/ET forward separation bolt 

catcher assembly 

7-2 
8.0 SPACE SHUTTLE MAIN ENGINE 

All Space Shuttle Main Engine (SSME) investigation fault tree legs have been closed. 

The STS-107 Block II SSMEs (center #2055, left #2053, and right #2049) performed 

nominally and there were no reported SSME In-Flight Anomalies (IFAs). 

8-1 
 

 

 

 

 

ET Age at Launch (days) 2000 

1500 

1000 

500 

0 

9.0 ENVIRONMENTAL FACTORS 

9.1 INTRODUCTION 

A survey was conducted of the relevant environmental factors during STS-107/ET-93 

processing to determine if a correlation could be drawn between those factors and ET 

bipod foam loss observed in flight. The data are inconclusive as to whether any 

correlation can be shown between environmental factors and ET bipod foam loss. The 

review considered ET age and exposure time, as well as weather factors such as 

rainfall, temperature, and humidity. 

9.2 AGE AND EXPOSURE 

The ET age was compared for various flights, presented in Figure 9-1. As shown in 

Figure 9-2, the ET age for STS-107/ET-93, 806 days, falls above the 95% confidence 

interval upper limit for the average age of all tanks, mean value 689 days, as well as the 

average age for all tanks with known bipod foam loss. STS-107/ET-93 also falls within 

the 95% confidence limit for missions with known bipod foam loss. Although the upper 

bound of the 95% confidence interval of the age of missions with bipod foam loss 

appears to be greater than the other groups in Figure 9-2, the 95% confidence interval 

limits of the different groups overlap each other. Therefore, data are inconclusive as to 

whether a correlation can be drawn about ET age and bipod foam loss. 

2500 

Average for All Missions: 689 days 

STS-107 (ET-93): 806 days 

LEGEND

 Confirmed Foam Loss

 Confirmed Bipod Foam Loss

 Confirmed No Foam Loss

 Unknown (no/poor imagery) 

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99

101106

929798

102100104105108109110111112113107 

STS (Mission Number) 

Figure 9-1. ET age for all STS missions 

9-1 
0 

200 

400 

600 

800 

1000 

1200 

1400 

Age at Launch(days) 

Average of All ETs 

Average of ETs with Foam Loss 

Average of ETs w/Bipod Foam Loss 

STS-107 with ET-93 

Average of ETs w/o Foam Loss 

Average of ETs w/unknown Foam Loss 

- Bars show +/- 95% Confidence Intervals 

- Overlapping Confidence Intervals Indicate

 No Significant Difference in Average Values 

Figure 9-2. ET age for STS-107 compared to ET age for missions with and 

without bipod foam loss 

9-2 
180 

Exposure Time (days) 160 

140 

120 

100 

80 

60 

40 

20 

0 

A similar comparison was made relating ET exposure time and bipod foam loss across 

the flight history, shown in Figure 9-3. Note that the STS-107/ET-93 exposure time, 39 

days, is the same as the mean value for all STS flights. As shown in Figure 9-4, the 

STS-107/ET-93 exposure time falls within the 95% confidence limit of all missions ET 

exposure time, as well as the time confidence limits for flights with or without known 

bipod foam loss. The STS-107/ET-93 exposure time is larger than the 95% confidence 

upper bound for missions with known bipod foam loss. However, as stated above when 

discussing ET age, the 95% confidence limits of the different groups in Figure 9-4 

overlap each other, and data are inconclusive as to whether ET exposure time and 

bipod foam loss can be correlated. 

Average for All Missions: 39 days 

STS-107 (ET-93): 39 days 

LEGEND 

Confirmed Foam Loss 

Confirmed Bipod Foam Loss 

Confirmed No Foam Loss

 Unknown (no/poor imagery) 

 

 

 

 

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99

101106

929798

102100104105108109110111112113107 

STS (Mission Number) 

Figure 9-3. ET exposure time (to weather) prelaunch for all STS missions 

9-3 
0 

10 

20 

30 

40 

50 

60 

70 

ET Exposure (days) 

Average of All ETs 

Average of ETs with Foam Loss 

Average of ETs w/Bipod Foam Loss 

STS-107 with ET-93 

Average of ETs w/o Foam Loss 

Average of ETs w/unknown Foam Loss 

- Bars show +/- 95% Confidence Intervals 

- Overlapping Confidence Intervals Indicate

 No Significant Difference in Average Values 

Figure 9-4. ET exposure time (to weather) for STS-107 compared to ET 

exposure time for missions with and without bipod foam loss 

9-4 
 

 

 

 

Total Pre-Launch Rainfall (inches) 20.00 

15.00 

10.00 

5.00 

0.00 

9.3 WEATHER FACTORS 

An extensive review of the relevant weather at Kennedy Space Center (KSC) was 

conducted in order to determine if a correlation could be derived for the weather 

conditions impact on ET bipod foam loss. The precipitation review examined total 

rainfall, maximum one-day rainfall, average daily rainfall, launch day rainfall, and L-5 

days through liftoff total rainfall. Figure 9-5 shows the total prelaunch rainfall for all STS 

missions. As shown in Figure 9-6, although the STS-107 value for total prelaunch 

rainfall, 12.78 inches, is greater than the mean value for all mission, 5.45 inches, the 

data are inconclusive as to whether a correlation can be made for ET bipod foam loss 

as a function of total rainfall prelaunch. The 95% confidence limit of the missions with 

ET bipod foam loss overlaps the confidence interval for all missions, as well as 

missions with no foam loss. 

25.00 

Average for All Missions: 5.45 inches 

STS-107 (ET-93): 12.78 inches 

LEGEND

 Confirmed Foam Loss

 Confirmed Bipod Foam Loss

 Confirmed No Foam Loss

 Unknown (no/poor imagery) 

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99

101106

929798

102100104105108109110111112113107

STS (Mission Number) 

Figure 9-5. Total prelaunch rainfall for all STS missions 

9-5 
 

 

0 

2 

4 

6 

8 

10 

12 

14 

Total Rain (inches) 

Average of All ETs 

Average of ETs with Foam Loss 

Average of ETs w/Bipod Foam Loss 

STS-107 with ET-93 

Average of ETs w/o Foam Loss 

Average of ETs w/unknown Foam Loss 

- Bars show +/- 95% Confidence Intervals 

- Overlapping Confidence Intervals Indicate 

No Significant Difference in Average Values 

Figure 9-6. Total rainfall for STS-107 compared to total rainfall for missions with 

and without bipod foam loss 

Similarly, the other rainfall parameters studied (e.g., average daily, day-of-launch) 

reveal no correlations for ET bipod foam loss. Figure 9-7 and Figure 9-8 show the data 

correlation for average daily prelaunch rainfall. The STS-107 value, 0.33 inches, and 

the mean value for missions with bipod foam loss are greater than the average mission 

value, 0.14 inches. However, the confidence intervals overlap each other, and the data 

are inconclusive as to whether average daily rainfall and ET bipod foam loss can be 

correlated. 

9-6 
Average Daily Rainfall (inches) 0.45 

0.40 

0.35 

0.30 

0.25 

0.20 

0.15 

0.10 

0.05 

0.00 

Average for All Missions: 

0.14 inches 

STS-107 (ET-93): 0.33 inches 

LEGEND

 Confirmed Foam Loss

 Confirmed Bipod Foam Loss

 Confirmed No Foam Loss

 Unknown (no/poor imagery) 

1 

Rainfall, Average Daily Prelaunch (inches) 2

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99

101106

929798

102100104105108109110111112113107 

STS (Mission Number) 

Figure 9-7. Average daily rainfall prelaunch for all STS missions 

0.40 

0.35 

0.30 

0.25 

0.20 

0.15 

0.10 

0.05 

0.00 

Average of All ETs 

Average of ETs with Foam Loss 

Average of ETs w/Bipod Foam Loss 

STS-107 with ET-93 

Average of ETs w/o Foam Loss 

Average of ETs w/unknown Foam Loss 

- Bars show +/- 95% Confidence Intervals 

- Overlapping Confidence Intervals Indicate

 No Significant Difference in Average Values 

Figure 9-8. Average daily rainfall prelaunch for STS-107 compared to average 

daily rainfall for missions with and without bipod foam loss 

9-7 
Average Temperature (°F) 100 

80 

60 

40 

20 

0 

In addition to rainfall, the study also reviewed average, minimum, and maximum 

temperature, dew point, and humidity for both prelaunch and day of launch. Figure 9-9 

shows the day-of-launch average temperature. The STS-107 day-of-launch average 

temperature, 58 o

F, was less than the mean value for all missions, 71 oF, but no 

correlation can be made between day-of-launch average temperature and ET bipod 

foam loss. Similar comparisons made for other temperature samplings, dew point 

(Figure 9-10), and humidity (Figure 9-11), yielded no correlations either. 

Average for All Missions: 71°F STS-107 (ET-93): 58°F 

LEGEND 

Confirmed Foam Loss 

Confirmed Bipod Foam Loss 

Confirmed No Foam Loss 

Unknown (no/poor imagery) 

 

 

 

 

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99

101106

929798

102100104105108109110111112113107 

STS (Mission Number) 

Figure 9-9. Day-of-launch average temperature for all STS missions 

9-8 
Average Humidity (%rh) Average Dewpoint Temperature (°F) 100 

90 

80 

70 

60 

50 

40 

30 

20 

10 

0 

1 1 2 2

3 3 4 4 5 5 6 6 

100 

90 

80 

70 

60 

50 

40 

30 

20 

10 

0 

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103

99

101106 

Average for All Missions: 76% STS-107 (ET-93): 75% 

LEGEND

 Confirmed Foam Loss

 Confirmed Bipod Foam Loss

 Confirmed No Foam Loss

 Unknown (no/poor imagery) 

 

 

 

 

 

 

 

 

STS (Mission Number) 

Figure 9-10. Prelaunch average dewpoint for all STS missions 

789

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99

101 

Average for All Missions:64°F 

STS-107 (ET-93): 50°F 

LEGEND

 Confirmed Foam Loss

 Confirmed Bipod Foam Loss

 Confirmed No Foam Loss

 Unknown (no/poor imagery) 

STS (Mission Number) 

Figure 9-11. Prelaunch average humidity for all STS missions 

9-9 

9297 

106

92 

98 97 

102 98 100 102 104 100 105 104 108 105 109 108 110 109 111 110 112 111 113 112 107 113

107 
10.0 LEFT WING PROCESSING AND RCC DESIGN 

10.1 INTRODUCTION 

This section summarizes the processing effort performed on the left wing of the Space 

Shuttle Columbia from the last Orbiter Major Maintenance (OMM) period through the 

launch of STS-107, and provides some background on the design of the RCC panels 

used on the orbiter. The processing includes all work done on Columbia from the major 

maintenance period (Columbia J3-OMM) through the flight of the STS-109 mission and 

all the normal preflight work done in preparation for the STS-107 mission. 

10.2 LEFT WING PROCESSING (PALMDALE, J3-OMM) 

Columbia was in Palmdale, California, for its most recent OMM from September 1999 

through March 2001. The work performed on the left wing included work on the 

electrical power and distribution system, instrumentation, mechanisms, structures, and 

the Thermal Protection System (TPS). There were 29 Problem Reports (PRs) on the 

electrical system, mostly wire lead discrepancies and wire stow issues. Two 

pyrotechnic connectors were found out of configuration and repaired. Instrumentation 

sensors and wire splices accounted for 20 PRs on the left wing and all were 

appropriately resolved. In the mechanisms area, a main landing gear door rotational 

pin inspection was partially performed at Palmdale and subsequently completed at the 

KSC. Slight damage to the chromium plating of the forward inboard gear door hook 

was repaired. The gear downlock bungee was sent to the vendor for refurbishment. 

Palmdale logged 62 PRs to the left wing structure that addressed elevon cove 

corrosion, elevon flipper door modification (material change from Inconel to Aluminum), 

and minor work on the main landing gear door. 

All Reinforced Carbon-Carbon (RCC) upper and lower wing Leading Edge Structural 

Subsystem (LESS) access panels, spar insulators, ear muff insulators, wing leading 

edge RCC panels, and spar fittings (see Figure 10-1 and Figure 10-2) were removed 

and inspected for discoloration and damage. Visual pinhole inspections were 

performed on each RCC panel and the wing leading edge spar was inspected for 

damage. Oversized pinholes were originally reported in RCC panels 8 and 19, but after 

further evaluation with an optical comparator, it was determined that the pinholes were 

acceptable. No other significant damage was noted. Leading edge RCC panels 6 and 

13 through 17 were sent to the vendor (Vought) for refurbishment. New shims were 

installed to accommodate the reinstallation of the spar insulators. 

The panels and spar fittings were reinstalled and all step and gap measurements were 

taken. At that time, gaps were found to be unacceptable in numerous locations. Wing 

leading edge RCC panels 11, 12, 17, and 18 were removed and additional anomalies 

were noted, which included insufficient step and gap, spar fitting shims not per design 

(too small), and the lower access panel nutplates debonded and/or with low running 

torque. The low torque was due to a combination of the shim problem and a procedural 

error on the torque sequence. All 22 RCC panels were removed a second time. The 

10-1 
nutplate issues were resolved by removing and replacing the nutplates that were 

accessible and securing with safety wire those that were not accessible. All anomalies 

identified were repaired, reworked, or accepted by Material Review (MR). 

Ear Muff 

Insulation 

Spar 

Insulator 

Spar Attach 

RCC Panel Fittings (4) 

Wing Leading 

Edge Spar 

Figure 10-1. RCC components 

10-2 
 

Figure 10-2. RCC panel assembly 

Tiles are attached to a strain isolation pad and then to the orbiter structure by a Room 

Temperature Vulcanized (RTV) adhesive. The outer tile surfaces must be flush with 

one another to preclude steps that would lead to excessive heat damage of surrounding 

tiles due to aerodynamic heating (Figure 10-3). Gaps present between adjacent tiles 

must be adequately sealed. There were 200 tile PRs worked for step and gap, gap 

fillers, and repair on the elevon cove area tile. One hundred thirty one (131) upper and 

lower wing surface tiles were replaced for various reasons, including baseline removal 

and replacements, damaged tile, instrumentation problems, and structural inspections. 

Wear and tear accounted for 27 maintenance items. Tile gap filler replacements 

numbered 58 with no issues noted. There were 100 discrepancy reports for minor tile 

putty repairs. Six chits (change items) were worked on the left side, mostly in the 

landing gear area. The main landing gear rotational pins, wheel well wire, and landing 

gear structural components were all inspected. The left inboard brake interference was 

slightly out of tolerance, but was corrected. One chit addressed the application of 

corrosion protection coating to the forward wing spar. 

10-3 
0.045" Tile to 

Reaction-cured Tile Gap 

Filler bar 

Koropon-primed 

structure 

Strain isolation 

Tile densi

glass (RCG) coating 

fied layer 

RTV 

adhesive 

pad (SIP) 

Figure 10-3. Typical tile installation 

Twenty Master Change Records were incorporated during the period. They included 

the replacement of the aluminized Mylar tape that lines the wheel well walls, the 

deletion of some non-functional acoustic sensors, and the removal of inactive Modular 

Auxiliary Data System (MADS) instrumentation. An additional part of that effort was the 

modification of the elevon columbium seal springs, some wing leading edge protective 

shielding, and enhancements to various gap fillers. 

The 22 left wing Line Replaceable Units (LRUs) that were replaced included the RCC 

panels previously mentioned, a hydraulic retract valve, the landing gear extend isolation 

valve, left main landing gear bungee, Tee seals, and an Inconel (Incoflex) insulator. 

All items not completed at Palmdale were dispositioned and transferred to KSC for 

completion. 

10.3 LEFT WING PROCESSING (STS-109) 

Once Columbia was delivered to KSC in March 2001, the outstanding main landing 

gear work was completed. This work included the left inboard axle rework to improve 

brake clearance and the completion of the rotational pin inspections. The elevon flipper 

doors had a few PRs for Wear and tear issues that were resolved. The lower elevon 

cove columbium seals/springs were out of configuration as compared to drawing 

requirements, and minor adjustments were successfully made. When it was discovered 

that there was excessive corrosion protection coating applied to the elevon cove area, 

10-4 
work was done to clean that area. There was minor corrosion in the elevon cove area 

that was removed as well. 

There were also numerous tile inspections and verifications performed during the 

processing flow for STS-109. No work was done on the wing leading edge RCC panels 

or Tee seals after Columbia returned from OMM and prior to STS-109. During that flow 

work was done on the LESS lower access panels 3, 6, 15, 17, 21, and 22 for step and 

gap issues and frayed horse collar gap concerns. Upper access panel 14 was replaced 

due to out-of-tolerance gap and out-of-tolerance Strain Isolation Pad (SIP). No lower 

tile acreage was replaced during the STS-109 flow, but the upper wing area had a few 

minor repair areas. All the leading edge and trailing edge panels for the left inboard 

elevon cove were replaced. Discrepancies were noted at Palmdale and corrected at 

KSC for the primary and secondary sealing circuits in the elevon cove seal assembly. 

The seals were operating within acceptable limits, but work was performed to repair 

leak paths and improve flow rate. There were 1,481 tiles that were suspect and had a 

manual deflection test performed on them in support of the corrective action required 

after one wing lower surface tile was found missing/debonded after the STS-103 

(orbiter Discovery) mission. Thirteen thermal barriers were replaced in the main landing 

gear door area. There were 14 total MR items for STS-109. 

10.4 LEFT WING PROCESSING (STS-107) 

During the STS-107 flow, damage was noted to the left main landing gear axle sleeve 

and axle nut. A review of the entire shuttle fleet revealed similar conditions on other 

vehicles. The tire separation harness for the tire temperature and pressure 

measurements was found caught in the brake mechanism and had to be removed. 

The tires were deflated and removed in order to inspect the wheel half-tie bolts. Due to 

the discovery of corrosion in the tie bolt holes in wheels throughout the fleet, wheels 

with sleeved tie bolt holes were installed. 

The angle seal at RCC panel 1 on the left wing leading edge (see Figure 10-4) was 

removed to support the evaluation of the horse collar gap filler between the adjacent 

tiles. During the removal attempt, the upper bushing remained bound with the shipside 

clevis. During subsequent attempts, the angle seal was manually manipulated to try 

and remove the preload. During the KSC paper review, structures engineers realized 

that the load applied to the angle seal was specified to be kept below 20 pounds, but 

was never recorded in the paper. The RCC specification requires that RCC panel loads 

be kept below 30 pounds. Subsequent tests at KSC verified that the angle seal load 

was below the 30-pound requirement. The LESS prevention and resolution team is 

addressing the issue of how to measure the load and how to support the seal in future 

operations. Work continued on the elevon flipper doors. Flipper door 1 blade seal was 

not making contact with the rub channel, potentially leading to excessive venting from 

the elevon cove area. This issue surfaced twice during this flow and was Material 

Review (MR) accepted to fly as is. Modification was made to the Inconel trailing edge 

seal and bulb seal on the elevon. Additional work was performed on the elevon cove 

corrosion protection again to reduce the excessiveness of the application. 

10-5 
Columbia Discovery, Atlantis, Endeavour 

Figure 10-4. Wing leading edge RCC 

An uncharacteristic number of access panels were removed during the STS-107 

processing flow. Most of the upper (14 of 22) and lower (13 of 22) access panels were 

removed due to misinterpreted requirements to check for excessive movement in the 

panels. Wing leading edge upper and lower access panels at RCC 15 had to be 

removed to retrieve a burned ball of tape that had been inadvertently left from previous 

work performed during OMM. Upper access panel 18 was removed to investigate the 

possibility of water intrusion from a water deluge system mishap in the orbiter 

processing facility, but no damage was noted. Tee seal 10 was removed and shipped 

to the vendor for repair. No other wing leading edge RCC panels were removed in 

preparation for the STS-107 mission. Only one access panel was replaced on the left 

outboard elevon cove area, but there were three other minor tile repairs performed on 

the left elevons. There were four tiles replaced on the under side of the left wing in 

front of the left outboard elevon because a gap filler had protruded 0.8 inches. This 

caused charred filler bar, SIP damage, and instrumentation wiring damage. None of 

the tiles in these areas is believed to affect the failure scenario. There were 36 total 

MRs for STS-107. 

10-6 
10.5 RCC DESIGN 

The RCC material is the basic structure of the wing leading edge panels (Figure 10-4), 

the nose cone, the chin panel between the nose cone and the nose landing gear door, 

and the forward external tank attach fitting cover plate on the orbiter. Its purpose is to 

protect the orbiter from local temperatures in excess of 2300 °F. Most RCC panels are 

designed with a 100-mission fatigue life, but RCC panels 8 through 12 have reduced 

lives due to higher temperature exposures. Panel 9 has the shortest mission life of 61 

missions because it has the highest heating load during entry. RCC panel 17 has the 

highest aerodynamic load. 

The panels were originally arc jet tested. Test data indicated that the multi-use 

temperature limit of 2960 deg F could be sustained for approximately 600 seconds. 

They were structurally tested up to a 1.2 factor of safety and eventually certified by 

analysis up to a 1.4 factor of safety. With these parameters, the panels are certified to 

140% of their expected load up to the ultimate strength of the panel. Other significant 

testing of the RCC panels was not performed due to lack of sufficient time to 

accomplish the testing prior to the first flight of Columbia. RCC panels show no obvious 

aging effects due to calendar life, but the panels normally lose mission life due to the 

combined effects of oxygen, high temperature, and high pressure during the entry of 

each mission. 

On Columbia, the structure supporting the RCC panels consisted of four attach fittings 

to mount each RCC panel to the aluminum honeycomb wing leading edge spar. In an 

effort to reduce the orbiter weight, wing components affecting the RCC installation were 

redesigned on subsequent vehicles. Beginning with the orbiter Discovery, the RCC 

attachment was accomplished using a single titanium attach fitting. The wing leading 

edge spar became a corrugated aluminum structure. Additional insulation was installed 

behind each RCC panel on all orbiters to shield the underlying structure from radiative 

heat damage from the high temperatures that the RCC reaches during entry. 

The RCC is composed of a carbon-based substrate (see Figure 10-5) that provides 

essentially all of the RCC strength. It is composed of graphitized rayon fabric 

impregnated with a phenolic resin called Tetraethyl Orthosilicate (TEOS) to provide 

internal protection against porosity within the laminate. The substrate is covered with a 

silicon carbide coating also enhanced with TEOS and sealed with a sealant to protect it 

from oxidation within the substrate. The silicon carbide coating provides no thermal 

protection for the RCC. 

10-7 
SILICON CARBIDE COATING 

CARBON SUBSTRATE 

WITH TEOS IMPREGNATION 

CRAZE CRACK WITH TYPE A SEALANT 

Approx. 

1/4 to 1/2 

Thickness 

Figure 10-5. RCC cross section 

During the manufacturing process, the silicon carbide surface acquires surface craze 

cracks due to differential contraction during the cooling process. The silicon carbide 

coating cools faster and contracts more than the carbon substrate during the cooling 

process. The craze cracks sometimes extend completely through the silicon carbide 

coating to the carbon substrate. A sodium silicate solution called Type A Sealant is 

applied to the silicon carbide coating to decrease porosity in the surface and fill the 

crazing cracks. Any erosion of the type A sealant and/or the silicon carbide coating 

could lead to direct exposure of the carbon fibers in the substrate. This provides a path 

for oxidation and can potentially lead to subsequent burn through of the RCC panel 

during entry. Development tests for the RCC never identified a susceptibility to 

oxidation; therefore, Columbia was not treated with the type A sealant until after the first 

five flights. Beginning in 1992, a double type A (DTA) sealant program was instituted 

on all vehicles to further enhance the corrosion protection on the wing RCC. 

Each time a vehicle returns from space, the entire RCC and Thermal Protection System 

(TPS) are visually inspected to determine the extent of any damage. Inspections look 

for RCC impact damage and any indications of flow in the interface between the TPS 

(tiles) and adjacent RCC. There also exists a test method whereby the RCC panel is 

pressed with a gloved hand in the vicinity of RCC cracks to determine the integrity of 

the panel and the existence of potential unacceptable subsurface oxidation. This same 

test is always performed on RCC panels 6 through 17 near each of the adjoining Tee 

seals. During each OMM, all RCC components are visually inspected including all the 

attachment hardware and underlying attachment structure. 

During the inspections, a determination is made to either repair, refurbish, or replace 

the panel as necessary. Repairs are required when there is noticeable damage to the 

surface of the panel. Field repairs can be made at KSC or Palmdale unless the carbon 

substrate is exposed. In that case the panels must be sent to the vendor for repair. 

Refurbishment is required at regular intervals to recoat the panels to increase their 

resistance to oxidation and mass loss. Occasionally, complete replacement of RCC 

panels is necessary due to unrepairable damage. Each wing leading edge RCC panel 

is paired with an associated Tee seal and both of these components are generally 

replaced/refurbished as a unit. 

10-8 
 

Columbia has only had three panels/Tee seals replaced over its history. Panels 12R 

and 10L were removed for destructive testing and pinhole evaluations. Panel 11L had 

fit problems and was sent to spares. Also, over Columbias lifetime, seven RCC panels 

and six seals on the left wing were repaired, and 11 panels and 12 seals were 

refurbished. All of the Columbia RCC panels were within their predicted mission life 

limits, and most were original panels. 

10.6 RCC IMPACT RESISTANCE 

The RCC was not considered part of the TPS for the purposes of impact resistance. 

The TPS was designed to accommodate particle impacts, such as from hail, rain, 

runway debris, etc., whose impact energy did not exceed 0.006 foot-pounds to the 

surface. The wing leading edge RCC impact resistance allowed no damage to the RCC 

with the application of up to16 inch-pounds of energy. Figure 10-6 shows RCC impact 

resistance ranging from 4 to 26 inch-pounds depending on the increasing thickness of 

the RCC element. Different tests including low velocity and hypervelocity tests have 

been conducted to determine the actual impact resistance of the RCC. Test projectile 

materials have included nylon, glass, aluminum, steel, lead, and ice and have taken 

shapes of spheres, bullets, and cylinders. The test results vary widely and appear to be 

significantly dependent on impact velocity, projectile type, and angle of incidence of 

impact. Because of the variability of the test results, no actual impact resistance could 

be defined. 

Min Wing 

Leading Edge 

RCC Thickness 

Figure 10-6. RCC impact resistance 

10-9 
In support of the STS-107 investigation, RCC impact testing was performed at 

Southwest Research Institute in San Antonio, Texas, by propelling a large piece of 

foam at high velocity at a previously flown RCC panel. These tests were described 

earlier in Section 3 and show that RCC material can be damaged by ET foam at impact 

velocities matching STS-107 debris impact conditions. 

10.7 RCC CORROSION 

The RCC panels are subject to mass loss due to loss of sealant that can be caused by 

normal entry heating, impact damage, or even undetected chemical attack. Mass loss 

results in a decrease in strength, burn resistance, and RCC mission life. Under the high 

temperatures of entry, the sealant may become molten in the vicinity of pinholes or 

debris impact areas and migrate, allowing an active oxidation process to begin at the 

surface. Some mass loss occurs normally during each mission. Mass loss is 

cumulative over mission life and is determined by analysis. Previously damaged RCC 

panels have been measured for mass loss using computer tomography, and that data 

is used in the analysis for all other RCC panels mass loss determinations. When 

analysis shows that the 1.4 factor of safety can no longer be maintained, the RCC 

panel is removed from service. The silicon carbide sealant does not prevent mass loss, 

but it does help increase corrosion resistance. The sealant must be refurbished 

periodically, but is usually performed during the most convenient OMM that does not 

violate the limits listed in Table 10-1. 

Table 10-1. RCC refurbishment limits 

Panels Refurbishment Interval 

1-5, 20-22 As required based on visual inspection 

6-17 16-18 missions, no calendar limit 

18-19 32-36 missions, no calendar limit 

Nose Cap 29 missions, no calendar limit 

Subsurface oxidation has been discovered beneath the silicon carbide surface cracks in 

the sealant and coating which allow the oxidation process to thrive. This process is 

considered to be an impact to RCC mission design life, but is not generally considered 

to be a safety of flight issue. This oxidation process (Figure 10-7) starts with the 

breakdown of the coating due to entry heating. Surface craze cracks allow oxygen to 

migrate to the subsurface carbon fibers and react with them. This increase in oxidation 

develops into larger crazed areas, which eventually allow pieces to become dislodged 

due to vibration, aerodynamic, or thermal loads. Once the pieces dislodge, they leave a 

large path for the oxidation process to continue. 

Dry ultrasonic and real-time radiographic inspections have been performed on the 

panels in the past to look at coating damage. More recently, special non-destructive 

10-10 
examinations are being evaluated which include infrared thermography to determine the 

extent of coating loss. 

New RCC Increasing Interfacial Oxidation 

With Flight Exposure 

SiC Debond 

SiC 

Carbon 

Craze cracks Small amounts of Increasing oxidation Silicon carbide 

provide an oxygen penetrate areas decrease the craze squares may 

oxidation path to the coating and adherence of the become dislodged 

the silicon carbide react with the silicon carbide craze by external forces 

to carbon carbon matrix and squares to the (airloads, impact, 

interface fibers, resulting in a substrate acoustics, flexure) 

decrease in strength 

Figure 10-7. RCC corrosion process 

Each wing leading edge RCC panel shares a Tee seal that is used to close the gap 

between adjacent RCC panels. Following STS-43 (Atlantis) in August 1991, routine 

inspections identified cracks in the web of a Tee seal. The cracks were in the silicon 

carbide coating and occasionally in the substrate, and were due to normal shrinkage. 

They were typically less than 1/2 inch long, were not visible to the naked eye, and 

usually occurred in the web of the seal, on the backside of the seal (Figure 10-8) near 

the apex rather than on the leading edge. Further examination of the remainder of the 

shuttle fleet identified 20 (of 132) cracked Tee seals. Columbia had 11 Tee seals 

identified with possible cracks. Detailed inspections determined that all the cracks were 

typical of the surface craze cracks in the coating. The Tee seal cracks were determined 

to be caused by warping of the substrate fabric during lay-up during the original build. 

The Tee seal cracking (Figure 10-9) leads to a reduction in mission life and loss of 

oxidation protection. All the seals were refurbished with new coating and sealant and 

were reinstalled. Failure analysis showed that cracks would form after excessive 

wishbone loading (bending) caused the brittle coating to crack. Crack testing was 

performed in 1991 on Tee seal 10 (attached to RCC panel 9) from the left wing of 

Columbia to try to determine the crack mechanism. The Tee seal was cycled 400 times 

in bending up to 70% of its ultimate load and no discernable damage was noted. After 

an engineering evaluation was performed on the health and strength of the Tee seal, it 

was subsequently reinstalled on Columbia. 

10-11 
Crack Area 

COLUMBIA 

DISCOVERY, ATLANTIS, ENDEAVOUR 

Figure 10-8. Tee seal crack location 

10-12 
Figure 10-9. Tee seal cracking 

Another phenomenon, discovered first on Columbia after STS-50 landed, was the 

existence of pinholes (Figure 10-10) in the RCC panels. The pinholes were found 

primarily in the wing leading edge RCC and were subsequently identified on all orbiters. 

Testing has shown that the pinholes are most likely the result of the accelerated 

oxidation process involving zinc oxide and the silicon carbide coating. The reaction of 

the zinc oxide and the silicon carbide produces a silica (glass) exudate that flows out of 

the pinhole area. The presence of zinc oxide is theorized to originate from the paint 

primer used to recondition the launch pad after each mission and is considered an 

accelerator to the oxidation process. The zinc-based contamination accumulates on 

the wing leading edge RCC as rainwater drips off of the launch pad. This 

contamination rests on the RCC without reacting to the surface material while at 

ambient conditions at the pad. All of the damaging oxidation occurs once the RCC is 

exposed to the high temperatures, pressures, and excess oxygen of re-entry. Only a 

few pinholes have been observed on the nose cap RCC, most likely because the nose 

10-13 
cap remains under a protective cover while at the launch pad. It is also believed that 

sodium chloride contributes to the oxidation process, but to a much lesser extent than 

the zinc oxide. 

In 1997, pinhole acceptance criteria were established. Pinholes with surface 

dimensions less than 0.040 inches discovered during routine processing flows are 

acceptable to fly as is for up to 16 missions unless the carbon substrate is exposed, in 

which case the panel must be repaired. Pinholes discovered at OMM greater than 

0.040 inches are unacceptable. Although the pinholes themselves constitute only a 

small mass loss, they are not considered to be a safety-of-flight issue by themselves. 

Analysis has identified that the sustainable thru-hole size in-flight due to orbital debris is 

0.25 inches in the lower surfaces of RCC panels 5-13. A hole under 1 inch in diameter 

anywhere else in the RCC is considered survivable for a single mission. 

Silicon 

Carbide 

Conversion 

Coating 

Carbon 

Substrate 

Oxidation Cavity in Carbon Substrate 

Pinhole Through 

Silicon Carbide 

Coating 

Type A Sealant Glass Exudate 

Glass Exudate 

0.010 to 0.040 Inch Diameter Pinholes 

Through Silicon Carbide Coating 

RCC Pinhole Cross-Section RCC Pinhole Surface View 

Figure 10-10. RCC pinholes 

There have been damaged RCC panels that were discovered after the vehicle returned 

from space on various missions. Some of the impact damage was only to the surface, 

but some even caused damage to the coating on the backside of the panel. In 1992, 

after STS-45, significant impact damage (overall length of ~1.75 inches) was noted on 

RCC panel 10R on Atlantis. The damage (Figure 10-11) was theorized to come from 

Orbital debris or micrometeorite impacts during the mission. This type of RCC damage 

is of particular concern in that a significant impact could cause a hole in the RCC large 

enough to lead to wing spar burn through and subsequent loss of crew and vehicle. At 

that time, the maximum acceptable hole size (0.040) criterion was established for 

10-14 
processing flows and advanced wing leading edge internal insulation was modified to 

reduce the risk should hot gas penetrate the RCC. 

Impact Side 

Coating Damage to 

Backside of Panel 

Figure 10-11. RCC impact damage 

There has been a history of loose bolts on access panels on all orbiters. Following 

STS-87, the right-hand lower access panel 4 had a loose bolt. All other installations 

were inspected and several additional bolts were found with low torque. All bolts were 

subsequently torqued to their proper values. During STS-95, the OMS pod Y-web door 

area had some damaged insulation. It was determined post-flight that there were bolts 

in the area that had low torque. A review of other orbiters identified low torque bolts on 

Discovery and Endeavour. Low torque bolts were also found during Columbias last 

OMM. The low torque was attributed to the performance of an improper torque 

sequence. All attach fittings were removed and reinstalled using the correct torque 

sequence. 

10-15 
11.0 EXTERNAL TANK 

11.1 INTRODUCTION 

The External Tank (ET) used for STS-107 was Light Weight Tank (LWT) number 

ET-93. This tank was the first LWT to be used with a cluster of three Block-II Space 

Shuttle Main Engines (SSMEs). As discussed in Section 3, there is significant visual 

and debris trajectory data to implicate the left bipod ramp area as the source of debris. 

Contributors to forward bipod thermal protection system (TPS) foam loss were: (1) the 

design, verification, and process validation did not encompass all material and 

processing variability or adequately address all failure modes, and (2) the acceptance 

testing and inspection techniques and procedures were not designed to be capable of 

rejecting ramps with adverse as-built features which would threaten the TPS integrity. 

11.2 TPS REQUIREMENTS 

During prelaunch, the ET TPS minimizes ice formation and maintains the quality of 

cryogenic propellant. During ascent, the ET TPS maintains the structure within design 

temperature limits. Program requirements (NSTS 07700, Vol. X, Book I, Paragraph 

3.2.1.2.14) indicate that the ET shall be designed to preclude the shedding of ice 

and/or other debris that would jeopardize the flight crew, vehicle, mission success, or 

would adversely impact turnaround operations. During ET entry, the TPS assures a 

predictable, low altitude ET break-up that meets the ET entry impact footprint boundary 

limits. 

The ET TPS itself is designed to have low density to maximize Shuttle payload 

capacity, high adhesion to cryogenic surfaces (-423 °F), resistance to thermal abrasion 

and degradation from aerodynamic shear, consistency (material qualified is the material 

flying), and environmental resistance to ultraviolet radiation, rain, etc. The application 

of ET TPS materials includes computer controlled automatic spray cells and manual 

application in normal working environments. 

11-1 
 

11.3 HISTORY OF FOAM CHANGES AND DEBRIS EVENTS 

The ET TPS history is marked by multiple material and configuration changes resulting 

from ET TPS and ice loss events, design enhancements, environmental regulations 

(especially blowing agent changes), and supplier changes. The history of foam changes 

is outlined in Figure 11-1, and Table 11-1 lists the ET flight history, as well as age and 

exposure data. Thousands of tests have been conducted to develop and qualify the ET 

TPS. There were no first time ET TPS changes on STS-107/ET-93 except for rework of 

the TPS on the upper aft ET/Solid Rocket Booster (SRB) fitting fairing (following SRB 

demate) using BX-265. Basic bipod TPS materials had not changed from the beginning 

of the program until after ET-93. The bipod TPS configuration has been stable since 

1983, when with ET-14 the ramp angle was changed. At ET-76 in 1995, there was one 

minor change to the ramp intersection with the ET intertank area. At ET-116 in 2002, 

the bipod material was changed to BX-265, but ET-93 had been constructed with 

BX-250. No indication has been found that any specific ET TPS foam change or any 

combination of historical ET TPS foam changes alone caused the bipod foam loss on 

STS-107/ET-93. 

ET-93 

Aft Fairing 

Rework 

Figure 11-1. History of foam changes. Blowing agent shown in 

parentheses, no changes to SLA 

11-2 
 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

Table 11-1. STS-Orbiter-ET configuration, age, and exposure 

STS 

1 

STS aka OV 

Columbia 

ET 

1 

ET wt. 

ET

Date 

04/12/81 

ET Age @ 

Launch (days) 

653 

ET Exposure @ 

Launch (days) 

105 

2 Columbia 2 ET 11/12/81 258 74 

3 Columbia 3 ET 03/22/82 175 35 

4 Columbia 4 ET 06/27/82 161 33 

5 Columbia 5 ET 11/11/82 169 52 

6 Challenger 8 LWT 04/04/83 208 126 

7 Challenger 6 ET 06/18/83 327 24 

8 Challenger 9 LWT 08/30/83 230 29 

9 Columbia 11 LWT 11/28/83 206 43 

11 41B Challenger 10 LWT 02/03/84 339 23 

13 41C Challenger 12 LWT 04/06/84 259 19 

14 41D Discovery 13 LWT 08/30/84 352 79 

17 41G Challenger 15 LWT 10/05/84 295 23 

19 51A Discovery 16 LWT 11/08/84 286 17 

20 51C Discovery 14 LWT 01/24/85 448 20 

23 51D Discovery 18 LWT 04/12/85 353 16 

24 51B Challenger 17 LWT 04/29/85 409 33 

25 51G Discovery 20 LWT 06/17/85 347 14 

26 51F Challenger 19 LWT 07/29/85 431 31 

27 51I Discovery 21 LWT 08/27/85 398 22 

28 51J Atlantis 25 LWT 10/03/85 287 35 

30 61A Challenger 24 LWT 10/30/85 348 15 

31 61B Atlantis 22 LWT 11/26/85 459 15 

32 61C Columbia 30 LWT 01/12/86 208 42 

33 51L Challenger 26 LWT 01/28/86 319 38 

26R Discovery 28 LWT 09/28/88 1261 87 

27R Atlantis 23 LWT 01/02/89 1561 62 

29R Discovery 36 LWT 03/13/89 1189 39 

30R Atlantis 29 LWT 05/04/89 1450 44 

28R Columbia 31 LWT 08/08/89 1484 25 

34 Atlantis 27 LWT 10/18/89 1723 51 

33R Discovery 38 LWT 11/22/89 1317 27 

32R Columbia 32 LWT 01/09/90 1609 43 

36 Atlantis 33 LWT 02/28/90 1597 35 

31R Discovery 34 LWT 04/24/90 1674 40 

41 Discovery 39 LWT 10/06/90 1635 32 

38 Atlantis 40 LWT 11/15/90 1609 88 

35 Columbia 35 LWT 12/02/90 1850 164 

37 Atlantis 37 LWT 04/05/91 1906 22 

39 Discovery 46 LWT 04/28/91 1327 49 

40 Columbia 41 LWT 06/05/91 1776 35 

43 Atlantis 47 LWT 08/02/91 1323 39 

11-3 
 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

Table 11-1. STS-Orbiter-ET configuration, age, and exposure 

(continued) 

STS 

48 

STS aka OV 

Discovery 

ET 

42 

ET wt. 

LWT 

Date 

09/12/91 

ET Age @ 

Launch (days) 

1829 

ET Exposure @ 

Launch (days) 

32 

44 Atlantis 53 LWT 11/24/91 846 33 

42 Discovery 52 LWT 01/22/92 994 35 

45 Atlantis 44 LWT 03/24/92 1840 34 

49 Endeavour 43 LWT 05/07/92 2005 56 

50 Columbia 50 LWT 06/25/92 1333 23 

46 Atlantis 48 LWT 07/31/92 1561 51 

47 Endeavour 45 LWT 09/12/92 1923 19 

52 Columbia 55 LWT 10/22/92 994 27 

53 Discovery 49 LWT 12/02/92 1577 25 

54 Endeavour 51 LWT 01/13/93 1440 42 

56 Discovery 54 LWT 04/08/93 1256 25 

55 Columbia 56 LWT 04/26/93 1082 79 

57 Endeavour 58 LWT 06/21/93 979 55 

51 Discovery 59 LWT 09/12/93 900 80 

58 Columbia 57 LWT 10/18/93 1180 33 

61 Endeavour 60 LWT 12/02/93 889 36 

60 Discovery 61 LWT 02/03/94 842 25 

62 Columbia 62 LWT 03/04/94 773 23 

59 Endeavour 63 LWT 04/09/94 737 22 

65 Columbia 64 LWT 07/08/94 718 24 

64 Discovery 66 LWT 09/09/94 591 23 

68 Endeavour 65 LWT 09/30/94 697 47 

66 Atlantis 67 LWT 11/03/94 535 25 

63 Discovery 68 LWT 02/03/95 546 25 

67 Endeavour 69 LWT 03/02/95 484 23 

71 Atlantis 70 LWT 06/27/95 495 63 

70 Discovery 71 LWT 07/13/95 435 58 

69 Endeavour 72 LWT 09/07/95 433 59 

73 Discovery 73 LWT 10/20/95 381 54 

74 Atlantis 74 LWT 11/12/95 360 33 

72 Endeavour 75 LWT 01/11/96 342 37 

75 Columbia 76 LWT 02/22/96 330 25 

76 Atlantis 77 LWT 03/22/96 303 24 

77 Endeavour 78 LWT 05/19/96 307 34 

78 Columbia 79 LWT 06/20/96 281 23 

79 Atlantis 82 LWT 09/16/96 188 38 

80 Columbia 80 LWT 11/19/96 368 35 

81 Atlantis 83 LWT 01/12/97 262 34 

82 Discovery 81 LWT 02/11/97 390 26 

83 Columbia 84 LWT 04/04/97 291 25 

84 Atlantis 85 LWT 05/15/97 281 22 

94 Columbia 86 LWT 07/01/97 266 21 

85 Discovery 87 LWT 08/07/97 246 25 

11-4 
 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

Table 11-1. STS-Orbiter-ET configuration, age, and exposure 

(concluded) 

STS 

86 

STS aka OV 

Atlantis 

ET 

88 

ET wt. 

LWT 

Date 

09/25/97 

ET Age @ 

Launch (days) 

251 

ET Exposure @ 

Launch (days) 

39 

87 Columbia 89 LWT 11/19/97 146 22 

89 Endeavour 90 LWT 01/22/98 167 35 

90 Discovery 91 LWT 04/17/98 154 26 

91 Discovery 96 SLWT 06/02/98 141 32 

95 Discovery 98 SLWT 10/29/98 147 39 

88 Endeavour 97 SLWT 12/04/98 249 47 

96 Discovery 100 SLWT 05/27/99 183 24 

93 Columbia 99 SLWT 07/23/99 360 47 

103 Discovery 101 SLWT 12/19/99 24 37 

99 Endeavour 92 LWT 02/11/00 298 61 

101 Atlantis 102 SLWT 05/19/00 473 56 

106 Atlantis 103 SLWT 09/08/00 444 26 

92 Discovery 104 SLWT 10/11/00 498 31 

97 Endeavour 105 SLWT 11/30/00 503 31 

98 Atlantis 106 SLWT 02/07/01 418 30 

102 Discovery 107 SLWT 03/08/01 455 25 

100 Endeavour 108 SLWT 04/19/01 434 29 

104 Atlantis 109 SLWT 07/12/01 435 22 

105 Discovery 110 SLWT 08/10/01 380 40 

108 Endeavour 111 SLWT 12/05/01 258 36 

109 Atlantis 112 SLWT 03/01/02 358 38 

110 Atlantis 114 SLWT 04/08/02 294 28 

111 Endeavour 113 SLWT 06/05/02 401 38 

112 Atlantis 115 SLWT 10/07/02 376 28 

113 Endeavour 116 SLWT 11/23/02 360 43 

107 Columbia 93 LWT 01/16/03 805 39 

ET debris has been observed throughout program history, including both ET TPS and 

ice debris. Since STS-1, imagery was available on about 80 missions, and debris has 

been confirmed on at least 62 missions. At least six missions lost portions of the left 

bipod ramp (see Section 3.5). TPS loss on the right bipod ramp has never been 

observed. A portion of the left bipod ramp was lost during STS-112 ascent and 

impacted the left SRB Integrated Electronics Assembly. No changes were made to 

STS-113 or STS-107 bipod ramp configurations after this event. 

The majority of ET debris events have been limited to small mass (< 0.2 lbs). A 

definitive correlation to orbiter damage is difficult except for major debris events such as 

STS-27R, which was identified as SRB ablator debris, and STS-87, which was 

attributed to ET intertank foam loss. Based on available historical data, the bipod ramp 

represents the source of the largest pieces of ET debris (estimated > 1.0 lbs), and LO2 

feedline bellows ice is second (estimated < 0.3 lbs). 

11-5 
11.4 STS-107/ET-93 CHRONOLOGY 

Ascent film indicates that the origin of STS-107 ET TPS loss was from the forward 

bipod area (see Section 3). Image-based size estimates support this to be the bipod 

ramp rather than flange or acreage foam. The history of bipod TPS loss provides 

additional supporting evidence. Available data supports the bipod ramp as the most 

probable point of origin of STS-107 debris. 

11.4.1 Bipod Ramp TPS Configuration 

The forward bipod TPS configuration includes a complex combination of foams, Super 

Light Ablator (SLA), and underlying bipod structural substrate elements. The bipod 

ramp configuration has been essentially stable since early in the program. There have 

been no changes in material until after ET-93 and only minimal changes to 

configuration, processing, and personnel certification and training. The BX-250 ramp 

angle has been constant since 1983, when with ET-14, the ramp angle was changed to 

30° maximum with a 5.0 ±1.0 inch radius at the forward edge (changed from 45° ± 5.0° 

with no radius at the forward edge). This was changed as a result of suspected foam 

debris on STS-7/ET-6. For ET-76 in 1995, there was one minor change to the forward 

ramp intersection with the ET intertank area; the 5.0 ±1 inch radius was changed to a 

straight termination line with a 0.25-inch step allowed. At ET-116 in 2002, the bipod 

material was changed to BX-265, but ET-93 was BX-250. There has been no indication 

that the bipod ramp configuration changes affected the observed STS-107/ET-93 bipod 

foam loss. 

11.4.1.1 Left and Right Bipod Ramp Differences 

TPS loss on the right bipod ramp has never been observed. Launch/ascent imagery 

from ground assets is less favorable for seeing right bipod foam loss as compared to 

the left bipod, and post-ET separation crew imagery is random between imaging the left 

or right bipod ramps. 

There is no flight or test data to explain why the -Y (left) bipod looses foam and the +Y 

(right) does not. The Shuttle Program only provides the -Y ramp air loads as a worst 

case for ET project analysis. There are several bipod configuration differences that 

may contribute to foam not coming off the +Y ramp. First, the foam ramp is configured 

slightly differently to accommodate the inboard strut for the LO2 feed line support 

structure (see Figure 11-2). Second, the proximity of the right bipod to the LO2 feedline 

could potentially influence local surface pressure causing a lower internal to external 

pressure differential (see Figure 11-3). Finally, the outboard and aft facing surface of 

the -Y bipod may experience lower surface pressure due to flow separation and other 

local effects relative to the +Y side. If the internal ramp pressure was high due to 

adverse as-built features in the ramp, this could lead to a higher differential pressure 

on the -Y versus the +Y ramp. However, the aerodynamic loads analysis reviewed in 

Section 3.5 shows that the loads on both ramps are below their design requirements. 

11-6 
+Y Bipod with ramp complete showing feed line and 

location of foam cut out to accommodate strut 

Figure 11-2. Right (+Y) bipod ramp 

JSC CFD Results: Mach No. = 2.46, Alpha =2.08o 

Figure 11-3. Left and right bipod ramp flow differences, CFD results 

11-7 
 

11.4.2 Bipod Ramp Certification 

The BX-250 foam was supplied to the ET project as a flight verified material from the 

Saturn Program. A review of material properties certification indicated forward bipod 

materials met material performance requirements including thermal recession 

properties at design ascent heating rates, thermal conductivity to preclude ice formation 

and to maintain cryogenic propellant quality during prelaunch, and mechanical 

properties. 

Process validation was performed by similarity of flight-like design substrate 

configurations. There had been no specific bipod dissection prior to the STS-107/ET93 investigation. Dissection of the bipod ramps from the production flow provided 

insight that the bipod ramp could contain unique adverse as-built features. The 

features identified during these recent dissections could potentially reduce the strength 

of the foam and result in foam failure and subsequent debris. 

For the bipod ramp, there was no robust evaluation of the manual spray process. The 

complexity of the manual spray process of the forward bipod TPS closeout leads to 

unique defects in this area including voids, rollovers, and TPS discontinuities. The 

configuration of the forward bipod BX-250 foam was verified based on similarity to the 

Protuberance Air Load (PAL) ramp, which did not address all aspects and failure 

mechanisms in combination with critical environments (adverse as-built features). 

The interaction of the underlying SLA configuration interfacial boundary and the 

potential effects of cryopumping were not considered. The design, verification, and 

process validation did not encompass all material and processing variability or 

adequately address all failure modes. 

11.4.3 Bipod Ramp Build Process 

Experienced certified practitioners performed the ET-93 bipod ramp BX-250 sprays, 

each with over 20 years experience. No indications of sprayer error were found. 

Procedures were followed and documented, and processes were within control limits 

(e.g., material specifications, temperature, and humidity) except that the process plan 

review found no Quality Control (QC) verification of overlap timing. There is no 

requirement to verify the overlap timing, and the impact of the overlap timing verification 

is not known. 

Dissection results of five ET TPS configurations demonstrated the forward bipod as the 

configuration with the most significant defects. Defects are driven by the variable 

manual spray process and complex contour substrate. This creates the potential for a 

combination of large voids or defects at critical locations needed to produce a 

significant foam loss. The designed-in process plan controls related to QC buy-off of 

critical parameters did not preclude introduction of adverse as-built features resulting 

from the complex and variable forward bipod manual spray operation. There is also 

variability in the response of the foam based on inherent randomness of the foam cell 

structure. It may not be possible to control a manual process well enough to preclude 

defects in the bipod ramp. 

11-8 
11.4.4 Bipod Ramp Foam Acceptance/Non-Destructive Evaluation 

ET BX-250 ramp foam build acceptance processes include localized plug pull and core 

tests of the ramp material prior to final trim configuration. The plug-pulls are taken from 

trimmed-off over-spray lead-in/lead-out areas (witness, or sample panels) on either side 

of the ramp to provide density, final visual inspections, and dimensional features. Postbuild inspection techniques are limited to visual inspections only. There were no 

anomalies found with the STS-107/ET-93 forward bipod ramp using inspection and 

acceptance techniques available at Michoud Assembly Facility (MAF). 

Previous efforts to implement robust foam Non-Destructive Evaluation (NDE) 

techniques were unsuccessful (a variety of techniques were attempted). Some 

progress was made in certain areas, but it never reached a fully qualified approach and 

the MAF effort was discontinued in 1993. Development efforts found many false 

positives and many missed defects. NDE methods in use at MAF were not able to 

identify adverse as-built features in the forward bipod BX-250 ramp, which could 

combine with nominal environments and create debris. Acceptance testing and 

inspection techniques and procedures were not designed to be capable of rejecting 

ramps with adverse as-built features that would threaten the TPS integrity. 

11.4.5 ET Shipping and Handling 

Post-build activities include storage at MAF, shipment to Kennedy Space Center (KSC), 

storage at KSC, and mating to the SRBs and orbiter. Extensive documentation governs 

the steps taken to care for the ET. Documentation review found no issues in ET-93 

processing paperwork. Storage took place in locked, limited-access facilities. The 

tanks were shipped pressurized with nitrogen to 6.0 ± 0.5 psi per requirements. At 

KSC, the LH2 tank pressurant is changed to helium and the pressure on each tank is 

checked at least twice per week per requirements. Visual inspections were performed 

every 90 days while in storage. ET-93 was inspected seven times between arrival at 

KSC and launch, not counting additional daily inspections when mated to the SRBs. 

Processes were in place and followed to ensure that shipping and handling were 

performed in a manner that minimizes damage to the ET. 

11.4.6 KSC Processing Activities 

The shuttle flight manifest was delayed due to cracks found during inspections of Main 

Propulsion System feedline flow liners on Atlantis in June 2002. The final manifest 

moved STS-112 and STS-113 ahead of STS-107. ET-93 was de-mated from SRBs BI114/ RSRM-86 and later mated to SRBs BI-116/RSRM-88, and SRBs BI-114/RSRM-86 

were used for STS-113. All mate/de-mate operations were carried out in accordance 

with standard procedures, and are outlined in Figure 11-4. There are no indications 

that KSC ET processing (ET shipping, handling, and processing) contributed to the 

bipod foam loss on STS-107/ET-93. 

11-9 
ET/Bipod Mate 06/24/02 

ET/Bipod Demate 08/01/02 

Figure 11-4. ET-93 processing timeline at KSC 

11.4.6.1 ET-93 Mate/De-Mate/Re-Mate 

The ET-93 bipod struts were installed, and then later removed during de-mate and reinstalled. This process has been performed at least six times during the Shuttle 

program. ET-11 was used on STS-9, but there is no imagery to confirm ET foam loss. 

ET-13 was used on STS-14 (41D), but again there is no imagery to confirm ET foam 

loss. ET-23 was used on STS-27R with handheld video imagery available that does not 

show the bipod ramps, but no foam loss was observed elsewhere. ET-23 was mated 

and de-mated during checkout of the Vandenberg Air Force Base facilities. STS-27R 

had a great amount of tile damage thought to be due to the loss of SRB ablator during 

launch. ET-37 was used on STS-38 but there is no imagery to confirm ET foam loss. 

ET-80 was used on STS-80, and there were two lost divots on the flange under the 

bipod and one 10-inch diameter divot on the intertank forward of the bipods. ET-86 

was used on STS-94 and the left bipod strut was installed upside down then re-installed 

correctly. The bipod ramps were visible and no bipod foam loss was noted. 

11.4.6.2 ET-93 Crushed Foam 

On ET-93, crushed foam (1.5" x 1.25" x 0.187") was seen after the -Y strut removal at 

the clevis. The thickness of foam in this area is 2.187 inches. Exposed crushed foam 

is not permissible outside of specific acceptance criteria, so a Problem Report (PR 

VG-389216) was written to evaluate the condition. The crushed foam was essentially 

covered up after mating to a new set of bipod struts. No data is available to determine 

11-10 
 

if this section of foam could have been the source of, or contributed to, a void or leak 

path for liquid or gas. 

Inspection of the region after installation of the bipod struts showed that the crushed 

foam did not extend farther than 0.75 beyond the bipod fitting-clevis joint, which is 

within acceptable limits. Dye penetration testing with recreated conditions indicated 

that the damage extended 0.25 inches from the visible mark and 0.5 inches into the 

surface of the foam, where the damage stopped. 

The Material Review Board (MRB) decided to use as is, and STS-107 launched with 

crushed foam contained behind the -Y bipod strut clevis. Crushed foam in this area is a 

nominal configuration, and the PR was only written for documentation for bipod strut 

removal and future inspections. Available data indicates that every flight may have 

crushed foam beneath the bipod strut. Review of the ET-93 PR, MAF testing, and the 

ET-117 strut removal provided evidence that crushed foam had no impact on 

performance, both thermal and structural. Data are inconclusive as to whether the 

crushed foam and bipod foam loss are associated. 

11.4.7 ET Pre-launch Operations 

The electrical system performance was nominal based on evaluation of pre-launch data 

and post flight inspection of ground electrical interfaces and SRB hardware. No 

anomalous conditions were identified during STS-107 visual inspections during launch 

operations: preflight, tanking ice team, video surveillance, and postflight walk down. 

There are no indications that ET pre-launch operations at KSC contributed to bipod 

foam loss on STS-107/ET-93. 

11.4.8 Launch/Ascent 

The ET-93 propulsion system performance was within design limits based on preflight 

predictions and postflight reconstruction. Comparison to historical performance showed 

performance within flight history experience for LWT, Super Light Weight Tank (SLWT), 

and Block II SSME. The STS-107 trajectory was within design limits throughout ascent. 

There were no anomalous angles of attack or dynamic pressure indications (see 

Section 3.5). STS-107 reconstructed air loads were within design limits, and no unique 

observations were associated with STS-107/ET-93. It is unlikely that any significant 

bipod structural loads were associated with the 62 second wind shear event followed by a 

0.6 Hz RSRM gimbal reaction associated with LO2 slosh (see Section 3.5 and sections 

below). No anomalous structural loads have been identified. Best-estimated trajectory 

loads and flex body loads assessment reconstructions show the ET interfaces to be well 

within design limits. Bipod interface vehicle loads are not considered driving 

environments for the bipod foam ramp. Adjacent structural stiffness precludes significant 

induced bipod ramp deflections from the interface strut loads. The majority of flexural 

loading on the bipod ramp results from cryogenic shrinkage of the LH2 tank prior to liftoff. STS-107 ascent thermal environments were within design limits based on analysis of 

flight data and ET system performance. A higher LH2 tank ullage preflight pressurization 

pressure (pre-press) is required for flights with Block-II SSME clusters. This helps 

11-11 
reduce spikes in the high-pressure fuel turbopump turbine discharge temperatures 

during start. The LWT was certified for higher pre-press and approved for ET-92 and 

subsequent flights by Interface Revision Notice (IRN) IC-1432 on 8-28-98. 

Data are inconclusive as to whether the STS-107 ascent environments contributed to 

the bipod foam loss on ET-93. 

11.4.8.1 ET LO2 Slosh Baffle Changes 

Eight ET LO2 slosh baffles were used until ET-14 in 1983. Vehicle stability analysis 

based on development flight instrumentation confirmed minimum LO2 sloshing 

disturbances and Space Shuttle Program LO2 damping requirements were 

subsequently reduced. Analysis and sub-scale test showed the baffle count could be 

reduced from eight to two and still maintain margin, but a reduction to four was selected 

as a trade off between cost benefit and weight reduction (see Figure 11-5). 

One weight saving feature of the SLWT is the removal of one more slosh baffles 

section, as shown in Figure 11-6. This gives a predicted performance gain of 92 

pounds. This feature was incorporated into the LWT at ET-87 in 1996 to reduce weight 

and diminish the number of first time configuration changes for the subsequent first 

flight of the SLWT. Dynamics analysis showed available damping remained within 

requirements and propulsion and stress analysis also remained within requirements. 

Data are inconclusive as to whether the ET LO2 slosh baffle configuration alone caused 

bipod foam loss on STS-107/ET-93. 

11.4.9 Possible Contributors to Strain Energy at ET Separation 

The Space Shuttle Program Loads Panel is continuing to work actions to identify 

potential contributors to strain energy that could have led to the off-nominal yaw rate at 

ET separation described in Section 3.5. For induced loads during ground operations, 

KSC is reviewing handling and stacking (orbiter and Ground Ops). For loads that 

occurred during flight, Boeing GNC is looking into the left side thermal event at 

300 seconds MET and if the mechanical load overcomes the joint preload during 

ascent. The ET project is looking into loads induced through cryogenic and 

pressurization cycles and the effects on the ET, such as shrinkage of the diagonal strut 

and overall shrinkage of the ET affecting the forward and aft attachments. The data are 

inconclusive as to whether potential strain energy at ET separation can be associated 

with events that caused bipod foam loss on STS-107/ET-93. 

11-12 
Figure 11-5. ET LO2 slosh baffle changes  ET-14 

Figure 11-6. ET LO2 slosh baffle changes  ET-87 

11-13 
11.5 STS-107/ET-93 TPS BIPOD DEBRIS 

11.5.1 Bipod Foam Failure Modes and Contributors 

Four basic possible bipod failure modes have been identified (shown in Figure 11-7) 

and each may occur alone or act in combination with each other. However, due to lack 

of bipod instrumentation, it is impossible to know exactly why part of the left bipod foam 

came off ET-93 during STS-107 ascent. Cracking is a break in the foam, which does 

not exhibit material loss and is typically perpendicular to the substrate. Debond or 

delamination is a separation of the material running along the substrate or layer lines. 

A divot is a piece of material dislodged from the surface resulting in a cavity, which may 

or may not expose the substrate. Shear is the removal or separation of material within 

the cell structure and is not confined to the layer lines of the material, but is parallel to 

the substrate. 

Failure Mode Primary Contributors Examples 

Cracking Substrate Strain 

Substrate bending 

Differential Thermal Contraction 

 

Cryopumping** 

Debond / Differential Thermal Contraction 

Delamination Substrate bending 

Differential Pressure 

Divot Void or cavity 

Cryopumping** 

Shear Airloads 

**Cryopumping may contribute by adding to loads that induce the failure mode 

(but it is not a failure mode itself) 

Figure 11-7. Bipod foam failure modes 

Cryopumping could contribute to bipod foam loss, shown schematically in Figure 11-8. 

The mechanism that drives cryopumping is the transformation of a gas to a liquid at 

cryogenic temperatures. Gases may condense within a void or porous material at low 

temperatures. Air in cavities or porous material liquefies when in contact with structure 

below -297°F for oxygen or -320°F for nitrogen. Pressure is reduced locally due to the 

condensation. If a leak path exists, more air will be pumped, providing more gas to 

condense. When the structure warms, the consequence of cryopumping is that the 

liquefied air returns to the gaseous state with a local pressure increase. If the leak path 

is large, gas escapes with no detrimental effect. However, if the leak path is small, 

11-14 
cracks may form in the TPS to relieve pressure, or a rapid increase in pressure may 

result in a divot. In order for this to occur, the inlet source must be blocked off to avoid 

venting out the inlet. It should be noted that testing has been unable to demonstrate 

cryopumping in this application. 

Figure 11-8. Schematic of bipod ramp - potential cryopumping 

Ablator 

(SLA) 

Gaseous Nitrogen (GN2) 

in Intertank and Stringers 

Internal Voids 

N2 Leak 

Path 

Air Leak 

Path 

BX-250 

Foam 

11-15 
 

 

 

 

11.5.2 Test Results for Debris Assessment 

Bipod TPS static and dynamic coupon tests were performed (test ET-TR-003). The 

objectives of these tests were to evaluate the BX-250/SLA hand-pack (HP) bond line 

laminate mechanical properties and investigate whether liberated BX-250 material 

could pull or tear SLA HP material from the bipod region. Analysis shows the critical 

bipod Spray-On Foam Insulation (SOFI) ramp failure mode due to direct air load is 

shear failure between SOFI ramp and bipod fitting substrate. 

Testing shows that the potential loss of BX-250 does not liberate hand packed SLA due 

to impulse loading for cryogenic applications; the BX-250 fails before the SLA. For 

shear, testing shows BX-250 fails before SLA at all test temperatures. For tension, 

testing shows BX-250 fails before SLA when SLA temperatures are less than or equal 

to -100 oF (see Figure 11-9). 

Dynamic Lap Shear Test Results 

Maximum Failure Load (lb) 

4000 

3500 

3000 

2500 

2000 

1500 

1000 

500 

0 

- 4 in. thick SLA under 1 in. thick BX 250 

- All failures occurred in BX 250 

- 0.5 in. thick SLA under 1 in. thick BX 250 

- All failures occurred in BX 250 

-200 -150 -100 -50 0 50 100 

Temperature (deg F) 

Figure 11-9. Critical test results in debris assessment 

11-16 
11.5.3 Max Bipod SLA Temperatures (80 seconds MET) 

The maximum SLA temperatures possible at 80 seconds were analyzed to determine 

maximum worst-case multi-event material loss. No cryopumping or cryo ingestion was 

assumed in order to calculate temperatures as high as possible. STS-107 ambient 

environments were used. Results showed the maximum SLA temperature possible at 

80 seconds MET is less than or equal to -100 oF, as shown in Figure 11-10. Tension 

testing shows BX-250 fails before SLA at temperatures less than or equal to -100 o

F. 

-100 

-100 

-125 

-153 

-193 

-113 

Figure 11-10. Max bipod SLA temperatures (80 seconds MET), oF. 

11-17 
 

 

 

11.5.4 Bipod Ramp As-Built Hardware Assessment 

The dissection of six bipod ramps indicated similar patterns for geometry-induced 

defects in all ramps. Roll-overs were observed at complex substrate elements, and the 

majority of observations were associated with spraying over complex details at the 

substrate. Sporadic voids were also observed. One internal delamination and one 

weak plane at the knit line were observed. Critical locations, or areas at-risk for 

producing debris, were identified near the edge of the machined foam surface for both 

voids and roll-overs (see Figure 11-12). Vacuum pressure is the primary driver for divot 

formation; however, wind shear also contributes to flight loads. A combination of 

multiple large voids, geometry-induced defects, and critical locations is needed to 

produce significant foam loss. For example, a large interconnected void at close 

proximity to the surface plus a weakened plane (see Figure 11-13) may produce foam 

loss. 

ET 94/120/124 +Y

Composite View

ET 120

ET 94

ET 124

Composite View 

ET 120 

ET 94 

ET 124 

Large void critical location 

Only roll-over critical location (near surface) 

Figure 11-11. Defects found at critical locations 

11-18 
ET 94/120/124 -Y

Composite View

ET 120

ET 94

ET 124

ET 94/120/124 -Y 

Composite View 

ET 120 

ET 94 

Weakened plane 

Figure 11-12. Weakened plane defect found 

11.5.5 Multi-Failure Mode TPS Bipod Debris 

The ET Working Group conducted an analysis coupled with test data to estimate a 

multi-failure mode TPS bipod debris size and weight. The ET Working Group scenario 

includes seven simultaneous and interactive adverse events: A large rollover occurs at 

the inboard stringer interface immediately below the machined foam surface, side-toside thermal crack/weak knit line, a large void near the topmost surface one inch below 

machined foam surface, warm SLA environment, and foam machined to minimum 

tolerances (not a failure). The specific results are shown in Figure 11-13. 

The determination of the STS-107/ET-93 bipod TPS debris is based on evaluation of 

fault tree findings, possible TPS failure modes and contributors, and results to date 

from TPS debris test programs including dissection, foam loss secondary effects 

assessment (SLA/BX-250), and bipod TPS debris size analysis. The TPS bipod debris 

size was determined by the ET Working Group to be approximately 870 cubic inches 

and 1.3 pounds. 

The transport analysis presented in Section 3 suggests that the debris object may be 

heavier than average foam, but the ET Working Group analysis indicates there could 

not be ice or significant SLA in the debris and that the density of the foam is consistent. 

Also, the imagery analysis showed that not all the debris struck the wing, but it broke up 

prior to impact, with all debris passing beneath the wing, some without impacting the 

wing. Recall that the transport analysis presented in Section 3 states the bipod TPS 

debris would be 1026 cubic inches and 1.4 pounds for a 820 ft/sec velocity or 

1239 cubic inches and 1.7 pounds for a 775 ft/sec velocity. The RCC foam impact test 

conducted at Southwest Research Institute was performed at 775 ft/sec with a 1.67-

pound foam article. 

11-19 
 

 

 

 

 

TOT

Weight ~1.3 lbs

SLA

Vol. ~1.9 cu in

Weight ~0.02 lbs

Foam

Vol. ~867.2 cu in

Weight ~1.23 lbs

TOT 

Weight ~1.3 lbs 

SLA 

Vol. ~1.9 cu in 

Weight ~0.02 lbs 

Foam 

Vol. ~867.2 cu in 

Weight ~1.23 lbs 

Figure 11-13. Multi-failure mode bipod TPS debris estimated by the ET Working 

Group. Note that this size and weight were not used in the RCC impact testing as 

part of the STS-107 investigation. 

11-20 
12.0 SUMMARY 

During the first stage of ascent, before SRB separation, the left wing of Columbia was 

struck by debris from the ET -Y bipod foam ramp. Analysis of the bipod foam ramp 

design, material, and processes suggests that the probable contributing mechanisms 

for foam liberation were cracks, delamination or debonding, divots, shear loads, or 

some combination of these. Analytical and test estimates of foam debris size, 

trajectory, and impact location indicate that the foam struck the left Wing Leading Edge 

(WLE) in the area between Reinforced Carbon-Carbon (RCC) panels 5 and 9. The 

impact energy tests conducted at Southwest Research Institute support the theory that 

the left wing RCC (lower panel 8 area, and/or an adjacent Tee seal) was damaged by 

the debris impact. 

During ascent several new flight experience events occurred. These were all very near 

existing flight envelopes and well within the certified flight envelope for which the 

Shuttle was designed. The data indicate that all new flight experiences could be 

attributed to the winds aloft and SRB performance. The new flight experiences may 

have individually or collectively contributed to liberation of the bipod foam ramp, but 

data are inconclusive in this regard. 

Launch radar analyses are inconclusive in determining size, shape, or identity of the 

debris measured after SRB separation. The radar data and analyses are inconclusive 

as to whether any of the debris impacted the orbiter. 

There is data indicating that an object departed the orbiter on flight day 2 with a small 

relative separation velocity. Ballistics and Radar Cross Section (RCS) testing and 

analyses have excluded all tested objects except for a partial WLE Tee seal, a whole 

WLE Tee seal, or a partial WLE RCC panel. Data are inconclusive in determining the 

identity of the flight day 2 object, or whether the object was associated with the bipod 

foam debris impact. 

Analysis of the RCC damage location and size is consistent with data from ascent. 

Analyses from orbiter telemetry, Modular Auxiliary Data System (MADS), aerodynamic 

and aero-thermal reconstruction and simulation, and debris forensics suggest that the 

RCC was damaged prior to Entry Interface (EI). The best estimate of the damage 

location is in the panel 8 lower area. Indications from modeling are that the damage 

size could have produced heating equivalent to a 6 to 10 inch hole diameter in the lower 

panel 8 area, or in one of the Tee seals adjacent to RCC panel 8. 

The damage in the left wing RCC provided a pathway for hot gas to enter the left wing 

leading edge and support structure during entry. This resulted in significant damage to 

the left wing and the subsequent loss of vehicle control, leading to aerodynamic 

breakup. 

12-1 
 

APPENDIX A 

ACRONYMS AND ABBREVIATIONS 

AA Accelerometer Assembly 

AC Alternating Current 

AFB Air Force Base 

AFRL Air Force Research Labs 

AFS Air Force Station 

AFSPC Air Force Space Command 

APU Auxiliary Power Unit 

BN Ballistic Number 

BSM Booster Separation Motor 

CAD Computer-Aided Design 

CAIB Columbia Accident Investigation Board 

CEI Contract End Item 

CF4 Tetraflouromethane 

CFD Computational Fluid Dynamics 

CG Center of Gravity 

CSA Canadian Space Agency 

DAO Data Assimilation Office 

DAP Digital Auto Pilot 

dBsm Decibels Relative to One Square Meter 

DLR German Aerospace Research Establishment 

DOLILU Day of Launch I-Load Update 

DTA Double Type A 

EDO Extended Duration Orbiter 

EI Entry Interface 

EORF Enhanced Orbiter Refrigerator/Freezer 

ER Eastern Range 

ESA European Space Agency 

EST Eastern Standard Time 

ET External Tank 

FEP Front End Processor 

FOD Foreign Object Debris 

FRCS Forward Reaction Control System 

FREESTAR Fast Reaction Enabling Science Technology and Research 

FRSI Felt Reusable Surface Insulation 

GMT Greenwich Mean Time 

GNC Guidance Navigation and Control 

A-1 
GRAM Global Reference Atmosphere Model 

GSE Ground Support Equipment 

Hi-Q Maximum dynamic pressure 

HMF Hypergolic Maintenance Facility 

HP Hand Pack 

HRSI High-Temperature Reusable Surface Insulation 

IEA Integrated Electronics Assembly 

IFA In-Flight Anomaly 

IMU Inertial Measurement Unit 

IPR Interim Problem Report 

IRN Interface Revision Notice 

ISS International Space Station 

JDMTA Jonathan Dickinson Missile Tracking Annex 

KSC Kennedy Space Center 

LAF Lost and Found 

LaRC Langley Research Center 

LCC Launch Commit Criteria 

LCD Launch Countdown 

LESS Leading Edge Structural Subsystem 

LH2 Liquid Hydrogen 

LO2 Liquid Oxygen 

LOS Loss of Signal 

LOX Liquid Oxygen 

LPS Launch Processing System 

LRU Line Replaceable Unit 

LWT Light Weight Tank 

MADS Modular Auxiliary Data System 

MAF Michoud Assembly Facility 

MCC Mission Control Center 

MEIDEX Mediterranean Israeli Dust Experiment 

MET Mission Elapsed Time 

MILA Merritt Island Launch Area 

MLP Mobile Launch Platform 

MMOD Micrometeoroid or Orbital Debris 

MOTR Multiple-Object Tracking Radar 

MPS Main Propulsion System 

MR Management Review 

MRB Material Review Board 

MSBLS Microwave Scanning Beam Landing System 

A-2 
MSID Measurement Stimulation Identification 

NAIT NASA Accident Investigation Team 

NASA National Aeronautical and Space Administration 

NASDA Japanese National Space Development 

NDE Non-Destructive Evaluation 

NSI NASA Standard Initiator 

ODRC Operational Data Retrieval Complex 

OI Operational Instrumentation 

OMDP Orbiter Maintenance Depot Processing 

OMI Operations and Maintenance Instruction 

OMM Orbiter Major Maintenance 

OMS Orbital Maneuvering System 

PAFB Patrick Air Force Base 

PAL Protuberance Air Load 

PAPI Precision Approach Position Indicator 

PE Performance Enhancement 

PLB Payload Bay 

PMBT Propellant Mean Bulk Temperature 

PR Problem Report 

PRSD Power Reactants Storage Device 

QC Quality Control 

RCC Reinforced Carbon-Carbon 

RCS Radar Cross Section 

RDM Research Double Module 

RF Radio Frequency 

RGA Rate Gyro Assembly 

RSR Range Separation Rate 

RSRM Re-usable Solid Rocket Motor 

RSS Range Safety System 

RTV Room Temperature Vulcanized 

SAMS Space Acceleration Measurement System 

SIP Strain Isolation Pad 

SLA Super Light Ablator 

SLF Shuttle Landing Facility 

SLWT Super Light Weight Tank 

SMG Space Meteorology Group 

SOFI Spray-On Foam Insulation 

SRB Solid Rocket Booster 

SSME Space Shuttle Main Engine 

A-3 
STA Shuttle Training Aircraft 

STS Space Transportation System 

TDRS Tracking and Data Relay Satellite 

TEOS Tetraethyl Orthosilicate 

TPS Thermal Protection System 

TVC Thrust Vector Control 

UHF Ultra-High Frequency 

VRCS Vernier Reaction Control System 

WLE Wing Leading Edge 

A-4