3.6 De-Orbit/Re-Entry

¶As Columbia re-entered Earthʼs atmosphere, sensors in the Orbiter relayed streams of data both to entry controllers on the ground at Johnson Space Center and to the Modular Auxiliary Data System recorder, which survived the breakup of the Orbiter and was recovered by ground search teams. This data – temperatures, pressures, and stresses – came from sensors located throughout the Orbiter. Entry controllers were unaware of any problems with re-entry until telemetry data indicated errant readings. During the investigation data from these two sources was used to make aerodynamic, aerothermal, and mechanical reconstructions of re-entry that showed how these stresses affected the Orbiter.

¶The re-entry analysis and testing focused on eight areas:

Analysis of the Modular Auxiliary Data System re-

corder information and the pattern of wire runs and sensor failures throughout the Orbiter.

Physical and chemical analysis of the recovered de-

bris to determine where the breach in the RCC panels likely occurred.

Analysis of videos and photography provided by the general public.
Abnormal heating on the outside of the Orbiter body.

Sensors showed lower heating and then higher heating than is usually seen on the left Orbital Maneuvering System pod and the left side of the fuselage.

Early heating inside the wing leading edge. Initially,

heating occurred inside the left wing RCC panels before the wing leading edge spar was breached.

Later heating inside the left wing structure. This analy-

sis focused on the inside of the left wing after the wing leading edge spar had been breached.

Early changes in aerodynamic performance. The Or-

biter began reacting to increasing left yaw and left roll, consistent with developing drag and loss of lift on the left wing.

Later changes in aerodynamic performance. Almost

600 seconds after Entry Interface, the left-rolling tendency of the Orbiter changes to a right roll, indicating an increase in lift on the left wing. The left yaw also increased, showing increasing drag on the left wing.

¶For a complete compilation of all re-entry data, see the

¶CAIB/NAIT Working Scenario (Appendix D.7) and the Reentry Timeline (Appendix D.9). The extensive aerothermal calculations and wind tunnel tests performed to investigate the observed re-entry phenomenon are documented in NASA report NSTS-37398.

¶Re-Entry Environment

¶In the demanding environment of re-entry, the Orbiter must withstand the high temperatures generated by its movement through the increasingly dense atmosphere as it deceler- ates from orbital speeds to land safely. At these velocities, shock waves form at the nose and along the leading edges of the wing, intersecting near RCC panel 9. The interaction between these two shock waves generates extremely high temperatures, especially around RCC panel 9, which experiences the highest surface temperatures of all the RCC panels. The flow behind these shock waves is at such a high temperature that air molecules are torn apart, or "dissoci- ated." The air immediately around the leading edge surface can reach 10,000 degrees Fahrenheit; however, the boundary layer shields the Orbiter so that the actual temperature is only approximately 3,000 degrees Fahrenheit at the leading edge. The RCC panels and internal insulation protect the aluminum wing leading edge spar. A breach in one of the leading-edge RCC panels would expose the internal wing structure to temperatures well above 3,000 degrees Fahrenheit.

¶In contrast to the aerothermal environment, the aerodynamic environment during Columbiaʼs re-entry was relatively be- nign, especially early in re-entry. The re-entry dynamic pressure ranged from zero at Entry Interface to 80 pounds per square foot when the Orbiter went out of control, compared with a dynamic pressure during launch and ascent of nearly 700 pounds per square foot. However, the aerodynamic forces were increasing quickly during the final minutes of Columbiaʼs flight, and played an important role in the loss of control.

¶Orbiter Sensors

¶The Operational Flight Instrumentation monitors physical sensors and logic signals that report the status of various Orbiter functions. These sensor readings and signals are telemetered via a 128 kilobit-per-second data stream to the Mission Control Center, where engineers ascertain the real-time health of key Orbiter systems. An extensive review of this data has been key to understanding what happened to STS-107 during ascent, orbit, and re-entry.

¶The Modular Auxiliary Data System is a supplemental instrumentation system that gathers Orbiter data for processing after the mission is completed. Inputs are almost exclusively physical sensor readings of temperatures, pressures, mechanical strains, accelerations, and vibrations. The Modular Auxiliary Data System usually records only the missionʼs first and last two hours (see Figure 3.6-1).

¶The Orbiter Experiment instrumentation is an expanded suite of sensors for the Modular Auxiliary Data System that was installed on Columbia for engineering development purposes. Because Columbia was the first Orbiter launched, engineering teams needed a means to gather more detailed flight data to validate their calculations of conditions the vehicle would experience during critical flight phases. The instrumentation remained on Columbia as a legacy of the development process, and was still providing valuable flight data from ascent, de-orbit, and re-entry for ongoing flight analysis and vehicle engineering. Nearly all of Columbiaʼs sensors were specified to have only a 10-year shelf life, and in some cases an even shorter service life.

¶Panel 10 Panel 9

¶Figure 3.6-1. The Modular Auxiliary Data System recorder, found near Hemphill, Texas. While not designed to withstand impact damage, the recorder was in near-perfect condition when recovered on March 19, 2003.

¶At 22 years old, the majority of the Orbiter Experiment instrumentation had been in service twice as long as its specified service life, and in fact, many sensors were already failing. Engineers planned to stop collecting and analyzing data once most of the sensors had failed, so failed sensors and wiring were not repaired. For instance, of the 181 sensors in Columbiaʼs wings, 55 had already failed or were producing questionable readings before STS-107 was launched.

¶Re-Entry Timeline

¶Times in the following section are noted in seconds elapsed from the time Columbia crossed Entry Interface (EI) over the Pacific Ocean at 8:44:09 a.m. EST. Columbiaʼs destruction occurred in the period from Entry Interface at 400,000 feet (EI+000) to about 200,000 feet (EI+970) over Texas. The Modular Auxiliary Data System recorded the first indications of problems at EI plus 270 seconds (EI+270). Because data from this system is retained onboard, Mission

¶Control did not notice any troubling indications from telemetry data until 8:54:24 a.m. (EI+613), some 10 minutes after

¶Entry Interface.

¶Left Wing Leading Edge Spar Breach (EI+270 through EI+515)

¶At EI+270, the Modular Auxiliary Data System recorded the first unusual condition while the Orbiter was still over the Pacific Ocean. Four sensors, which were all either inside

Sensor 1 WLE WLEStrain Strain- -V12G9921A V12G9921

Sensor 4 WLE WLESpar SparTemp Temp- V09T9895A - V09T9895

Sensor 2 WLE WLEClevis Clevis- -V09T9910A V09T9910

Sensor 3 Looking Aft AftPanel Panel99Lower LowerSurface SurfaceTemp Temp- -V09T9666A V09T9666 Forward

¶Figure 3.6-2. Location of sensors on the back of the left wing leading edge spar (vertical aluminum structure in picture). Also shown are the round truss tubes and ribs that provided the structural support for the mid-wing in this area.

¶or outside the wing leading edge spar near Reinforced Carbon-Carbon (RCC) panel 9-left, helped tell the story of what happened on the left wing of the Orbiter early in the re-entry. These four sensors were: strain gauge V12G9921A (Sensor 1), resistance temperature detector V09T9910A on the RCC clevis between panel 9 and 10 (Sensor 2), thermocouple V07T9666A, within a Thermal Protection System tile (Sensor 3), and resistance temperature detector V09T9895A (Sensor 4), located on the back side of the wing leading edge spar behind RCC panels 8 and 9 (see Figure 3.6-2).

¶V12G9921A – Left Wing Leading Edge Spar Strain Gauge

¶48:39

¶STS - 107

STS - 073 STS - 090

¶Strain (micro-in./in.)

¶STS - 109

¶-250

-500 First off nominal indication -750

¶-1000

0 100 200 300 400 500 600 700 800 900 — 1000

44:09 59:09 Time (seconds from EI)

¶Figure 3.6-3. The strain gauge (Sensor 1) on the back of the left wing leading edge spar was the first sensor to show an anomalous reading. In this chart, and the others that follow, the red line indicates data from STS-107. Data from other Columbia re-entries, similar to the STS-107 re-entry profile, are shown in the other colors.

¶V07T9910A – Left Wing Leading Edge Spar Temperature

¶600 48:59 EI+487

¶STS - 107

STS - 073 STS - 090 STS - 109

¶Temperature (0F)

¶-100

-200 First off nominal indication -300

0 100 200 300 400 500 600 700 800 — 900

44:09 Time (seconds from EI) 59:09

¶Figure 3.6-4. This temperature thermocouple (Sensor 2) was mounted on the outside of the wing leading edge spar behind the insulation that protects the spar from radiated heat from the RCC panels. It clearly showed an off-nominal trend early in the re-entry sequence and began to show an increase in temperature much earlier than the temperature sensor behind the spar.

¶Sensor 1 provided the first anomalous reading (see Figure 3.6-3). From EI+270 to EI+360, the strain is higher than that on previous Columbia flights. At EI+450, the strain reverses, and then peaks again in a negative direction at EI+475. The strain then drops slightly, and remains constant and negative until EI+495, when the sensor pattern becomes unreliable, probably due to a propagating soft short, or "burn-through" of the insulation between cable conductors caused by heating or combustion. This strain likely indicates significant damage to the aluminum honeycomb spar. In particular, strain rever- sals, which are unusual, likely mean there was significant high-temperature damage to the spar during this time.

¶At EI+290, 20 seconds after Sensor 1 gave its first anomalous reading, Sensor 2, the only sensor in the front of the

70 Clevis Temperatures 10" Hole with Sneak Flow

TEMPERATURE (F)

0 50 100 150 200 250 300 350 400 — 450

¶Time (seconds from EI)

¶Figure 3.6-5. The analysis of the effect of a 10-inch hole in RCC panel 8 on Sensor 2 from EI to EI+500 seconds. The jagged line shows the actual flight data readings and the smooth line the calculated result for a 10-inch hole with some sneak flow of superheated air behind the spar insulation.

¶left wing leading edge spar, recorded the beginning of a gradual and abnormal rise in temperature from an expected

¶30 degrees Fahrenheit to 65 degrees at EI+493, when it then dropped to "off-scale low," a reading that drops off the scale at the low end of the sensorʼs range (see Figure 3.6-4). Sensor 2, one of the first to fail, did so abruptly. It had indicated only a mild warming of the RCC attachment clevis before the signal was lost.

¶A series of thermal analyses were performed for different sized holes in RCC panel 8 to compute the time required to heat Sensor 2 to the temperature recorded by the Modular

¶Auxiliary Data System. To heat the clevis, various insula1000 tors would have to be bypassed with a small amount of leakage, or "sneak flow." Figure 3.6-5 shows the results of these calculations for, as an example, a 10-inch hole, and demonstrates that with sneak flow around the insulation, the temperature profile of the clevis sensor was closely matched by the engineering calculations. This is consistent with the same sneak flow required to match a similar but abnormal ascent temperature rise of the same sensor, which further supports the premise that the breach in the leading edge of the wing occurred during ascent. While the exact size of the breach will never be known, and may have been smaller or larger than 10 inches, these analyses do provide a plausible explanation for the observed rises in temperature sensor data during re-entry.

¶Investigators initially theorized that the foam might have broken a T-seal and allowed superheated air to enter the wing between the RCC panels. However, the amount of T-seal debris from this area and subsequent aerothermal analysis showing this type of breach did not match the observed damage to the wing, led investigators to eliminate a missing T-seal as the source of the breach.

¶Although abnormal, the re-entry temperature rise was slow and small compared to what would be expected if Sensor 2 were exposed to a blast of superheated air from an assumed breach in the RCC panels. The slow temperature rise is at-

¶V07T9666A – Left Wing Lower Surface Temperature

¶OSH

¶Degrees F

¶1000 STS - 107

STS - 073 STS - 090

EI+496 STS - 109

¶500 0 OSL

0 100 200 300 400 500 600 700 800 900 — 1000

¶Time (seconds from EI)

¶Figure 3.6-6. As early as EI+370, Sensor 3 began reading significantly higher than on previous flights. Since this sensor was located in a thermal tile on the lower surface of the left wing, its temperatures are much higher than those for the other sensors.

¶tributed to the presence of a relatively modest breach in the RCC, the thick insulation that surrounds the sensor, and the distance from the site of the breach in RCC panel 8 to the clevis sensor.

¶The readings of Sensor 3, which was in a thermal tile, began rising abnormally high and somewhat erratically as early as EI+370, with several brief spikes to 2,500 degrees

¶Fahrenheit, significantly higher than the 2,000-degree peak temperature on a normal re-entry (Figure 3.6-6). At EI+496, this reading became unreliable, indicating a failure of the wire or the sensor. Because this thermocouple was on the wing lower surface, directly behind the junction of RCC panel 9 and 10, the high temperatures it initially recorded were almost certainly a result of air jetting through the damaged area of RCC panel 8, or of the normal airflow being disturbed by the damage. Note that Sensor 3 provided an external temperature measurement, while Sensors 2 and 4 provided internal temperature measurements.

¶Sensor 4 also recorded a rise in temperature that ended in an abrupt fall to off-scale low. Figure 3.6-7 shows that an abnormal temperature rise began at EI+425 and abruptly fell at EI+525. Unlike Sensor 2, this temperature rise was extreme, from an expected 20 degrees Fahrenheit at EI+425 to 40 degrees at EI+485, and then rising much faster to 120 degrees at EI+515, then to an off-scale high (a reading that climbs off the scale at the high end of the range) of 450 degrees at EI+522. The failure pattern of this sensor likely indicates destruction by extreme heat.

¶The timing of the failures of these four sensors and the path of their cable routing enables a determination of both the timing and location of the breach of the leading edge spar, and indirectly, the breach of the RCC panels. All the cables from these sensors, and many others, were routed into wiring harnesses that ran forward along the back side of the leading edge spar up to a cross spar (see Figure 3.6-8), where they passed through the service opening in the cross spar and then ran in front of the left wheel well before reaching interconnect panel 65P, where they entered the fuselage. All sensors with wiring in this set of harnesses failed between EI+487 to EI+497, except Sensor 4, which survived until

¶EI+522. The diversity of sensor types (temperature, pressure, and strains) and their locations in the left wing indicates that they failed because their wiring was destroyed at spar burn-through, as opposed to destruction of each individual sensor by direct heating.

¶Examination of wiring installation closeout photographs (pictures that document the state of the area that are normally taken just before access is closed) and engineering drawings show five main wiring harness bundles running forward along the spar, labeled top to bottom as A through E (see Figure 3.6-8). The top four, A through D, are spaced 3 inches apart, while the fifth, E, is 6 inches beneath them. The separation between bundle E and the other four is consistent with the later failure time of Sensor 4 by 25 to 29 seconds, and indicates that the breach was in the upper two-thirds of the spar, causing all but one of the cables in this area to fail between EI+487 to EI+497. The breach then expanded vertically, toward the underside of the wing, causing Sensor 4 to fail 25 seconds

¶V09T9895A – Left Wing Front Spar Panel 9 Temperature

¶OSH

¶400 EI+522

¶STS - 107

STS - 073 STS - 090

¶STS - 109

¶Degrees F

¶-100

-200 OSL

¶-300

0 100 200 300 400 500 600 700 800 900 — 1000

¶Time (seconds from EI)

¶Figure 3.6-7. Sensor 4 also began reading significantly higher than previous flights before it fell off-scale low. The relatively late reaction of this sensor compared to Sensor 2, clearly indicated that superheated air started on the outside of the wing leading edge spar and then moved into the mid-wing after the spar was burned through. Note that immediately before the sensor (or the wire) fails, the temperature is at 450 degrees Fahrenheit and climbing rapidly. It was the only temperature sensor that showed this pattern.

¶later. Because the distance between bundle A and bundle E is 9 inches, the failure of all these wires indicates that the breach in the wing leading edge spar was at least 9 inches from top to bottom by EI+522 seconds.

Sensor 5 V07P8010A V07P8010A

¶Panel

Panel — 98

¶9 Panel Panel

Panel — 88

¶9 Panel

A B C D

E Sensor 4 V09T9895A V09T9895A

Sensor 6 V07P8058A V07P8058A

¶Figure 3.6-8. The left photo above shows the wiring runs on the backside of the wing leading edge behind RCC panel 8 – the circle marks the most likely area where the burn through of the wing leading edge spar initially occurred at EI+487 seconds. The right photo shows the wire bundles as they continue forward behind RCC panels 7 and 6. The major cable bundles in the upper right of the right photo carried the majority of the sensor data inside the wing. As these bundles were burned, controllers on the ground began seeing off-nominal sensor indications.

¶Also directly behind RCC panel 8 were pressure sensors V07P8010A (Sensor 5), on the upper interior surface of the wing, and V07P8058A (Sensor 6), on the lower interior surface of the wing. Sensor 5 failed abruptly at EI+497. Sensor 6, which was slightly more protected, began falling at EI+495, and failed completely at EI+505. Closeout photographs show that the wiring from Sensor 5 travels down from the top of the wing to join the uppermost harness, A, which then travels along the leading edge spar. Similarly, wiring from Sensor 6 travels up from the bottom of the wing, joins harness A, and continues along the spar. It appears that Sensor 5ʼs wiring, on the upper wing surface, was damaged at EI+497, right after Sensor 1 failed. Noting the times of the sensor failures, and the locations of Sensors 5 and 6 forward of Sensors 1 through 4, spar burn-through must have occurred near where these wires came together.

¶Two of the 45 left wing strain gauges also recorded an anomaly around EI+500 to EI+580, but their readings were not erratic or off-scale until late in the re-entry, at EI+930. Strain gauge V12G9048A was far forward on a cross spar in the front of the wheel well on the lower spar cap, and strain gauge V12G9049A was on the upper spar cap. Their responses appear to be the actual strain at that location until their failure at EI+935. The exposed wiring for most of the left wing sensors runs along the front of the spar that crosses in front of the left wheel well. The very late failure times of these two sensors indicate that the damage did not spread into the wing cavity forward of the wheel well until at least EI+935, which implies that the breach was aft of the cross spar. Because the cross spar attaches to the transition spar behind RCC panel 6, the breach must have been aft (outboard) of panel 6. The superheated air likely burned through the outboard wall of the wheel well, rather than snaking forward and then back through the vent at the front of the wheel well. Had the gases flowed through the access opening in the cross spar and then through the vent into the wheel well, it is unlikely that the lower strain gauge wiring would have survived.

Left OMS Pod Surface Mounted Tile Temperature on Forward Looking Face

¶49:49

¶STS - 090

¶V07T9913

V07T9913 V07T9913

V07T9913 V07T9913 V07T9913

¶Degrees F

0 100 200 — 300

¶44:09 Time (seconds from EI)

¶Figure 3.6-9. Orbital Maneuvering System (OMS) pod heating was initially significantly lower than that seen on previous Columbia missions. As wing leading edge damage later increased, the OMS pod heating increased dramatically. Debris recovered from this area of the OMS pod showed substantial pre-breakup heat damage and imbedded drops of once-molten metal from the wing leading edge in the OMS pod thermal tiles.

¶Finally, the rapid rise in Sensor 4 at EI+425, before the other sensors began to fail, indicates that high temperatures were responsible. Comparisons of sensors on the outside of the wing leading edge spar, those inside of the spar, and those in the wing and left wheel well indicate that abnormal heating first began on the outside of the spar behind the RCC panels and worked through the spar. Since the aluminum spar must have burned through before any cable harnesses attached to it failed, the breach through the wing leading edge spar must have occurred at or before EI+487.

¶Other abnormalities also occurred during re-entry. Early in re-entry, the heating normally seen on the left Orbital Maneuvering System pod was much lower than usual for this point in the flight (see Figure 3.6-9). Wind tunnel testing demonstrated that airflow into a breach in an RCC panel would then escape through the wing leading edge vents behind the upper part of the panel and interrupt the weak aerodynamic flow field on top of the wing. During re-entry, air normally flows into these vents to equalize air pressure across the RCC panels. The interruption in the flow field behind the wing caused a displacement of the vortices that normally hit the leading edge of the left pod, and resulted in a slowing of pod heating. Heating of the side fuselage slowed, which wind tunnel testing also predicted.

¶To match this scenario, investigators had to postulate damage to the tiles on the upper carrier panel 9, in order to allow sufficient mass flow through the vent to cause the observed decrease in sidewall heating. No upper carrier panels were found from panels 9, 10, and 11, which supports this hypothesis. Although this can account for the abnormal temperatures on the body of the Orbiter and at the Orbital Maneuvering System pod, flight data and wind tunnel tests confirmed that this venting was not strong enough to alter the aerodynamic force on the Orbiter, and the aerodynamic analysis of mission data showed no change in Orbiter flight control parameters during this time.

¶During re-entry, a change was noted in the rate of the temperature rise around the RCC chin panel clevis temperature sensor and two water supply nozzles on the left side of the STS - 107 STS - 073 fuselage, just aft of the main bulkhead that divides the crew cabin from the payload bay. Because these sensors were well forward of the damage in the left wing leading edge, it is still

¶STS - 109 unclear how their indications fit into the failure scenario.

¶Sensor Loss and the Onset of Unusual Aerodynamic

¶Effects (EI+500 through EI+611)

¶Reduced, off-nominal heating Fourteen seconds after the loss of the first sensor wire on the First off nominal indication wing leading edge spar at EI+487, a sensor wire in a bundle

400 500 600 700 800 900 — 1000

¶59:09 of some 150 wires that ran along the upper outside corner of the left wheel well showed a burn-through. In the next 50 seconds, more than 70 percent of the sensor wires in three cables in this area also burned through (see Figure 3.6-10). Investigators plotted the wiring run for every left-wing sensor, looking for a relationship between their location and time of failure.

¶Only two sensor wires of 169 remained intact when the Modular Auxiliary Data System recorder stopped, indicat-Percent Loss of Sensor Signals Versus Time In Left Wing and Wing Leading Edge Wire Bundles

¶V09T9895A

Leading Edge (18 of 18)

¶Wheel Well

¶Quantity of Sensor Signals Lost - %

¶70 Bit Flip

Left Elevon 60 Accel Response

1st Wheel Well Temp Rise 50 (1700,1702 Bit change)

¶1st OI Starts Failure

450 — 500

Sensors with Cables Along Loading Edge Start Loss 14 sec Earlier Than the 3 Bundles

¶Figure 3.6-10. This chart shows how rapidly the wire bundles in the left wing were destroyed. Over 70 percent of the sensor wires in the wiring bundles burned through in under a minute. The black diamonds show the times of significant timeline sensor events.

¶ing that the burn-throughs had to occur in an area that nearly every wire ran through. To sustain this type of damage, the wires had to be close enough to the breach for the gas plume to hit them. Arc jet testing (in a wind tunnel with an electrical arc that provides up to a 2,800-degree Fahrenheit airflow) on a simulated wing leading edge spar and simulated wire bundles showed how the leading edge spar would burn through in a few seconds. It also showed that wire bundles would burn through in a timeframe consistent with those seen in the Modular Auxiliary Data System information and the telemetered data.

¶Later computational fluid dynamics analysis of the mid-wing area behind the spar showed that superheated air flowing into a breached RCC panel 8 and then interacting with the internal structure behind the RCC cavity (RCC ribs and spar insulation) would have continued through the wing leading edge spar as a jet, and would have easily allowed superheated air to traverse the 56.5 inches from the spar to the outside of the wheel well and destroy the cables (Figure 3.6-11). Controllers on the ground saw these first anomalies in the telemetry data at EI+613, when four hydraulic sensor cables that ran from the aft part of the left wing through the wiring bundles outside the wheel well failed.

¶Aerodynamic roll and yaw forces began to differ from those on previous flights at about EI+500 (see Figure 3.6-12). Investigators used flight data to reconstruct the aerodynamic forces acting on the Orbiter. This reconstructed data was then compared to forces seen on other similar flights of Columbia

Bundle 3 (115 of 117) V07P8049A

¶Bundle 1 Bundle 4

¶(25 of 25) (9 of 9)

¶V07P9197A lure 5th OI Starts Failure

¶L Elevon Accel Fail 1st Orbiter Debris Event

¶Reversal of Roll Moment and start of Slow Aileron Trim Change

¶Start LMG Struct Actuator Temp Rise

¶6th OI Starts Failure 7th OI Starts Failure

¶Flat Portion for 3 Bundles

550 600 650 700 750 — 800

¶Time (seconds from EI) mph

¶Flow

Contours of Velocity Magnitude (fps) Jun 10, 2003 FLUENT 6.1 (2d, coupled imp, ske)

Figure 3.6-11. The computational fluid dynamics analysis of the speed of the superheated air as it entered the breach in RCC panel 8 and then traveled through the wing leading edge spar. The dark- est red color indicates speeds of over 4,000 miles per hour. Temperatures in this area likely exceeded 5,000 degrees Fahrenheit. The area of detail is looking down at the top of the left wing.

¶STS 107 Delta Rolling/Yawing Moment Coefficients

Off-Nominal Roll & Yaw 0.0025 LMG Brake Line Temps Start Off Nominal Trend Delta Cll (Roll Moment) - 13:52:41 Delta Cln (Yaw Moment) 2 Temp Sensors Begin Off Nominal Response 0.0020 Delta Cll Aero Model - V09T9895A - Wing Front Spar Panel 9 Delta Cln Aero Model - V09T9849A - OB Elevon, Lower Surface - 13:52:49.5/51.4 Initial Roll

¶Delta Roll/Yaw Moment Coefficient

¶Wing Frnt Spar Panel 9 Temp

0.0015 - 13:49:32 Wing LE 55 LWR Att. Clevis Left INBD Elevon Lower Skin Temp - Start of Off Nominal Trend

RCC 10 - Start Off Nominal Trend - 13:52:56 - 13:51:14

4 Left OMS Pod Surf Temp 0.0010 - Change in Existing

¶Off Nominal Trend

13:52:39/ 53:09

0.0005 Alpha Mod

¶Active

13:53:31

¶0.000

-0.0005 Left Wing Lower Surface TC - Start Off Nominal Temp Increase - 13:50:19 Debris #1

13:53:44/48

-0.0010 Left OMS Pod TC BP 0731T - Start Off Nominal Trend Debris #2

Reduced Rise Rate - 13:53:46/50

13:49:49 Debris #3 - 13:53:54/58 -0.0015 Left OMS Pod LRSI Surface Temp Debris #4 Left OMS Pod TC BP0732T Left OMS Pod TC BP0749T - 13:54:00/04

Start Off Nominal Trend LMG Brake Line Temp (D) Debris #5 - Reduced Rise Rate - Start Off Nominal Trend - 13:54:07/11 Debris #6 -0.0020 - 13:49:59 - 13:52:17

Left Wing Spar Cap Left PLBD Surface TC BP3703T - Off Nominal Strain Increase LMG Brake Line Temp B

Start Off Nominal Trend - 13:52:18 - Off Nominal Trend

Reduced Rise Rate -0.0025 - 13:50:09

¶49:00.0 50:00.0 51:00.0 52:00.0 53:00.0 54:00.0

¶Figure 3.6-12. At approximately EI+500 seconds, the aerodynamic roll and yaw forces began to diverge from those observed on previous flights. The blue line shows the Orbiterʼs tendency to yaw while the red line shows its tendency to roll. Nominal values would parallel the solid black line. Above the black line, the direction of the force is to the right, while below the black line, the force is to the left.

¶and to the forces predicted for STS-107. In the early phase of fight, these abnormal aerodynamic forces indicated that Columbiaʼs flight control system was reacting to a change in the external shape of the wing, which was caused by pro- gressive RCC damage that caused a continuing decrease in lift and a continuing increase in drag on the left wing.

¶Between EI+530 and EI+562, four sensors on the left inboard elevon failed. These sensor readings were part of the data telemetered to the ground. Noting the system failures, the Maintenance, Mechanical, and Crew Systems officer notified the Flight Director of the failures. (See sidebar in Chapter 2 for a complete version of the Mission Control Center conversation about this data.)

¶At EI+555, Columbia crossed the California coast. People on the ground now saw the damage developing on the Orbiter in the form of debris being shed, and documented this with video cameras. In the next 15 seconds, temperatures on the fuselage sidewall and the left Orbital Maneuvering System pod began to rise. Hypersonic wind tunnel tests indicated that the increased heating on the Orbital Maneuvering System pod and the roll and yaw changes were caused by

Temperature Rise Rate Change - Hyd Sys 1 LMG UpLK UnIK Ln Temp

Sys 3 LMG Brake Ret Line Temp - LMG Brake Line Temp B, C Left Lower Wing Skin Temp - 13:56:16/22 - OSL - 13:57:28 Left Upper Wing Skin Temp Left Upper Wing Skin Temp - OSL - 13:57:43 - Begin Off Niminal Trend

13:56:24

Left Main Gear Strut Start Sharp Aileron Trim

¶Hydraulic System Left OUTBD / INBD Actuator Temp Increase Elevon Return Line Temps - OSL - Temp Rise Rate Chg - 13:58:03 - 13:53:10 / 36 - 13:56:53 Sys 2 LH Brake

¶Viv Return Temp Start Slow Alllegron Debris #11 MLG LH OB Tire - Start of sharp Trim Change Dwnrd. Temp

13:55:36/42 Pressure #1, #2 - 13:54:20 - 13:59:22

Start Off Nom Debris #15 - 13:57:19/24

¶Debris #9, #10 - 13:56:09/13 MLG LH OutBD &

13:55:25/30 INBD Tire Pressure #1

Pressure Trend to OSL - 13:58:32 Fuse Side Surf Temp Fuse Low Surf BF Temp Fuse Side Surf TC & - Start Off Nom Trend Left PLBD Surface TC - 13:57:09 - Temp Increase to OSH - 13:59:29

BFS Fault Message (4) Tire Pressures

¶Debris #7 - 13:58:40/56

13:55:04/10

1st Roll Reversal Initiation - 13:56:30

¶Mid Fuselage Bondline Temp Left Main Gear

¶Complete - 13:46:55 & LH Aft Fus Sidewall Temp Downlock Indication

Off Nominal - Transferred ON
13:54:22 - 13:59:06

¶Debris #13, #14 Flash #1

13:54:33.3 - 13:55:55/59 Flash #1, #2 is #6 - 13:57:53.7
13:54:35/37

Debris #12 - 13:57:59.5 - 13:55:45/ Flash #2

¶p B Debris #8 - 13:57:59.5

¶Debris #7 Trend - 13:55:21/27 58:01.5

13:57:19/29 - 13:54:10

¶55:00.0 56:00.0 57:00.0 58:00.0 59:00.0 00:00.0

¶Time (min:sec) substantial leading edge damage around RCC panel 9. Data on Orbiter temperature distribution as well as aerodynamic forces for various damage scenarios were obtained from wind tunnel testing.

¶Figure 3.6-13 shows the comparison of surface temperature distribution with an undamaged Orbiter and one with an entire panel 9 removed. With panel 9 removed, a strong vortex flow structure is positioned to increase the temperature on the leading edge of the Orbital Maneuvering System pod. The aim is not to demonstrate that all of panel 9 was missing at this point, but rather to indicate that major damage to panels near panel 9 can shift the strong vortex flow pattern and change the Orbiterʼs temperature distribution to match the Modular Auxiliary Data System information. Wind tunnel tests also demonstrated that increasing damage to leading edge RCC panels would result in increasing drag and decreasing lift on the left wing.

¶Recovered debris showed that Inconel 718, which is only found in wing leading edge spanner beams and attachment fittings, was deposited on the left Orbital Maneuvering System pod, verifying that airflow through the breach and out of the upper slot carried molten wing leading edge material back to the pod. Temperatures far exceeded those seen on previous re-entries and further confirmed that the wing leading-edge damage was increasing.

¶By this time, superheated air had been entering the wing since EI+487, and significant internal damage had probably occurred. The major internal support structure in the mid-wing consists of aluminum trusses with a melting point of 1,200 degrees Fahrenheit. Because the ingested air may have been as hot as 8,000 degrees near the breach, it is likely that the internal support structure that maintains the shape of the wing was severely compromised.

¶As the Orbiter flew east, people on the ground continued to record the major shedding of debris. Investigators later scrutinized these videos to compare Columbiaʼs re-entry with recordings of other re-entries and to identify the debris. The video analysis was also used to determine additional search areas on the ground and to estimate the size of various pieces of debris as they fell from the Orbiter.

¶Temperatures in the wheel well began to rise rapidly at EI+601, which indicated that the superheated air coming through the wing leading edge spar had breached the wheel well wall. At the same time, observers on the ground noted additional significant shedding of debris. Analysis of one of these "debris events" showed that the photographed object could have weighed nearly 190 pounds, which would have significantly altered Columbiaʼs physical condition.

¶At EI+602, the tendency of the Orbiter to roll to the left in response to a loss of lift on the left wing transitioned to a right-rolling tendency, now in response to increased lift on the left wing. Observers on the ground noted additional significant shedding of debris in the next 30 seconds. Left yaw continued to increase, consistent with increasing drag on the left wing. Further damage to the RCC panels explains the increased drag on the left wing, but it does not explain the sudden increase in lift, which can be explained only by some other type of wing damage.

¶Investigators ran multiple analyses and wind tunnel tests to understand this significant aerodynamic event. Analysis showed that by EI+850, the temperatures inside the wing

¶Run 12

¶Baseline smooth

¶RCC #9

TC 1106 removed TC 1724

¶Effect of Missing RCC Panel on Orbiter RELATIVE HEATING RATE Mid-Fuselage Thermal Mapping

¶0 0.1 0.2 0.3 0.4 0.5

¶Figure 3.6-13. The effects of removing RCC panel 9 are shown in this figure. Note the brighter colors on the front of the OMS pod show increased heating, a phenomenon supported by both the OMS pod temperature sensors and the debris analysis.

THE KIRTLAND IMAGE

¶As Columbia passed over Albuquerque, New Mexico, during re-entry (around EI+795), scientists at the Air Force Starfire

¶Optical Range at Kirtland Air Force Base acquired images of the Orbiter. This imaging had not been officially assigned, and the photograph was taken using commercial equipment located at the site, not with the advanced Starfire adaptiveoptics telescope.

¶The image shows an unusual condition on the left wing, a leading-edge disturbance that might indicate damage. Several analysts concluded that the distortion evident in the image likely came from the modification and interaction of shock waves due to the damaged leading edge. The overall appearance of the leading-edge damage at this point on the trajectory is consistent with the scenario.

¶were high enough to substantially damage the wing skins, wing leading edge spar, and the wheel well wall, and melt the wingʼs support struts. Once structural support was lost, the wing likely deformed, effectively changing shape and resulting in increased lift and a corresponding increase in drag on the left wing. The increased drag on the left wing further increased the Orbiterʼs tendency to yaw left.

¶Loss of Vehicle Control (EI+612 through EI+970)

¶A rise in hydraulic line temperatures inside the left wheel well indicated that superheated air had penetrated the wheel well wall by EI+727. This temperature rise, telemetered to Run 18 Mission Control, was noted by the Maintenance, Mechanical, and Crew Systems officer. The Orbiter initiated and completed its roll reversal by EI+766 and was positioned left-wing-down for this portion of re-entry. The Guidance and Flight Control Systems performed normally, although the aero-control surfaces (aileron trim) continued to counter- act the additional drag and lift from the left wing.

¶At EI+790, two left main gear outboard tire pressure sensors began trending slightly upward, followed very shortly by going off-scale low, which indicated extreme heating of both the left inboard and outboard tires. The tires, with their large mass, would require substantial heating to produce the sensorsʼ slight temperature rise. Another sharp change in the rolling tendency of the Orbiter occurred at EI+834, along with additional shedding of debris. In an attempt to maintain attitude control, the Orbiter responded with a sharp change in aileron trim, which indicated there was another significant change to the left wing configuration, likely due to wing deformation. By EI+887, all left main gear inboard and outboard tire pressure and wheel temperature measurements were lost, indicating burning wires and a rapid progression of damage in the wheel well.

¶Lower Left wing debris Lower Right wing debris Figure 3.7-1. Comparison of amount of debris recovered from the left and right wings of Columbia. Note the amount of debris recovered from areas in front of the wheel well (the red boxes on each wing) were similar, but there were dramatic differences in the amount of debris recovered aft of each wheel well.

¶At EI+897, the left main landing gear downlock position indicator reported that the gear was now down and locked. At the same time, a sensor indicated the landing gear door was still closed, while another sensor indicated that the main landing gear was still locked in the up position. Wire burn-through testing showed that a burn-induced short in the downlock sensor wiring could produce these same contra- dictions in gear status indication. Several measurements on the strut produced valid data until the final loss of telemetry data. This suggests that the gear-down-and locked indica- Figure 3.7-2. Each RCC panel has a U-shaped slot (see arrow) in tion was the result of a wire burn-through, not a result of the back of the panel. Once superheated air entered the breach the landing gear actually deploying. All four corresponding in RCC panel 8, some of that superheated air went through this proximity switch sensors for the right main landing gear re- slot and caused substantial damage to the Thermal Protection mained normal throughout re-entry until telemetry was lost. System tiles behind this area.

¶Post-accident analysis of flight data that was generated after telemetry information was lost showed another abrupt change in the Orbiterʼs aerodynamics caused by a continued progression of left wing damage at EI+917. The data showed a significant increase in positive roll and negative yaw, again indicating another increase in drag on and lift from the damaged left wing. Columbiaʼs flight control system attempted to compensate for this increased left yaw by firing all four right yaw jets. Even with all thrusters firing, combined with a maximum rate of change of aileron trim, the flight control system was unable to control the left yaw, and control of the Orbiter was lost at EI+970 seconds. Mission Control lost all telemetry data from the Orbiter at EI+923 (8:59:32 a.m.). Civilian and military video cameras on the ground documented the final breakup. The Modular Auxiliary Data System stopped recording at EI+970 seconds.

¶Findings:

¶F3.6−1 The de-orbit burn and re-entry flight path were normal until just before Loss of Signal.

¶F3.6−2 Columbia re-entered the atmosphere with a pre-existing breach in the left wing.

¶F3.6−3 Data from the Modular Auxiliary Data System recorder indicates the location of the breach was in the RCC panels on the left wing leading edge.

¶F3.6−4 Abnormal heating events preceded abnormal aerodynamic events by several minutes.

¶F3.6−5 By the time data indicating problems was telemetered to Mission Control Center, the Orbiter had already suffered damage from which it could not recover.

¶Recommendations:

¶R3.6-1 The Modular Auxiliary Data System instrumentation and sensor suite on each Orbiter should be maintained and updated to include current sensor and data acquisition technologies.

¶R3.6-2 The Modular Auxiliary Data System should be redesigned to include engineering performance and vehicle health information, and have the ability to be reconfigured during flight in order to allow certain data to be recorded, telemetered, or both, as needs change.