Matthew E. Melis, Duane M. Revilock, and Michael J. Pereira

Glenn Research Center, Cleveland, Ohio

Karen H. Lyle

Langley Research Center, Hampton, Virginia

Impact Testing on Reinforced Carbon-Carbon

Flat Panels With BX–265 and PDL–1034

External Tank Foam for the Space Shuttle

Return to Flight Program

NASA/TM—2009-213642/REV1

December 2009
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Matthew E. Melis, Duane M. Revilock, and Michael J. Pereira

Glenn Research Center, Cleveland, Ohio

Karen H. Lyle

Langley Research Center, Hampton, Virginia

Impact Testing on Reinforced Carbon-Carbon

Flat Panels With BX–265 and PDL–1034

External Tank Foam for the Space Shuttle

Return to Flight Program

NASA/TM—2009-213642/REV1

December 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Available from

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7115 Standard Drive

Hanover, MD 21076–1320

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Trade names and trademarks are used in this report for identifi cation 

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either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2009-213642/REV1 1

Impact Testing on Reinforced Carbon-Carbon Flat Panels 

With BX–265 and PDL–1034 External Tank Foam for the 

Space Shuttle Return to Flight Program 

Matthew E. Melis, Duane M. Revilock, and Michael J. Pereira 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Karen H. Lyle 

National Aeronautics and Space Administration 

Langley Research Center 

Hampton, Virginia 23681 

Summary 

Following the tragedy of the Orbiter Columbia (STS–107) on February 1, 2003, a major effort 

commenced to develop a better understanding of debris impacts and their effect on the space shuttle 

subsystems. An initiative to develop and validate physics-based computer models to predict damage from 

such impacts was a fundamental component of this effort. To develop the models it was necessary to 

physically characterize reinforced carbon-carbon (RCC) along with ice and foam debris materials, which 

could shed on ascent and impact the orbiter RCC leading edges. The validated models enabled the launch 

system community to use the impact analysis software LS–DYNA (Livermore Software Technology Corp.) 

to predict damage by potential and actual impact events on the orbiter leading edge and nose cap thermal 

protection systems. 

Validation of the material models was done through a three-level approach: Level 1—fundamental tests 

to obtain independent static and dynamic constitutive model properties of materials of interest, Level 2—

subcomponent impact tests to provide highly controlled impact test data for the correlation and validation of 

the models, and Level 3—full-scale orbiter leading-edge impact tests to establish the final level of 

confidence for the analysis methodology. 

This report discusses the Level 2 test program conducted in the NASA Glenn Research Center (GRC) 

Ballistic Impact Laboratory with external tank foam impact tests on flat RCC panels, and presents the data 

observed. The Level 2 testing consisted of 54 impact tests in the NASA GRC Ballistic Impact Laboratory on 

6- by 6-in. and 6- by 12-in. flat plates of RCC and evaluated two types of debris projectiles: BX–265 and 

PDL–1034 external tank foam. These impact tests helped determine the level of damage generated in the 

RCC flat plates by each projectile and validated the use of the foam and RCC models for use in LS–DYNA. 

Introduction 

On February 1, 2003, the Orbiter Columbia broke apart during reentry resulting in the loss of seven 

crewmembers. For the next several months an extensive investigation of the accident ensued, involving a 

nationwide team of experts spanning dozens of technical disciplines from NASA, industry, and academia. 

The Columbia Accident Investigation Board (CAIB), a group of experts assembled to conduct an 

investigation independent of NASA, concluded in August 2003 that the cause of the loss of Columbia and 

its crew was a breach in the left-wing leading-edge reinforced carbon-carbon (RCC) thermal protection 

system initiated by the impact of thermal insulating foam that had separated from the orbiter’s external fuel 

tank 81 seconds into that mission’s launch. During reentry, this breach allowed superheated air to penetrate 

behind the leading edge and erode the aluminum structure of the left wing, which ultimately led to the 

breakup of the orbiter. 

The CAIB report (ref. 1) made over two dozen recommendations to increase the overall safety of the 

shuttle for future launches. Prior to the Columbia accident, there were no sophisticated analysis tools in 
NASA/TM—2009-213642/REV1 2

existence to reliably quantify the debris impact damage threat to the shuttle system. As a consequence, 

CAIB recommendation R3.8–2 directed NASA to “Develop, validate, and maintain physics-based 

computer models to evaluate thermal protection system damage from debris impacts. These tools should 

provide realistic and timely estimates of any impact damage from possible debris from any source that 

may ultimately impact the orbiter. Establish impact damage thresholds that trigger responsive corrective 

action, such as on-orbit inspection and repair, when indicated.” In response to R3.8–2, an Agency debris 

assessment team, often informally referred to as the DYNA team, consisting of members from Glenn 

Research Center (GRC), Langley Research Center (LaRC), Johnson Space Center (JSC), and Boeing, was 

assembled to develop such a tool using LS–DYNA (ref. 2). LS–DYNA is a commercial finite element 

code that utilizes an explicit formulation (as opposed to the more common implicit formulation) to predict 

a wide range of transient dynamic phenomena. 

As a critical path element of NASA’s Return to Flight Program for the STS–114 mission, the primary 

objectives set for the DYNA team were to develop analysis models for potential debris and RCC 

materials. RCC is used as the thermal protection system on the Orbiter leading edge and nose cap. To 

address these objectives, the team established a three-level approach: (1) fundamental tests to obtain 

independent static and dynamic constitutive model properties of materials of interest, (2) subcomponent 

impact tests to provide highly controlled impact test data for the correlation and validation of the models, 

and (3) full-scale impact tests to establish the final level of confidence for the analysis methodology. The 

debris materials under primary consideration were external tank thermal protection foams BX–265 and 

PDL, and ice, which also might shed off of the external tank. 

All of the Level 2 impact testing for this program was conducted at the NASA GRC Ballistic Impact 

Laboratory in Cleveland, Ohio. The objective of this report is to provide details of the Level 2 impact test 

program and the experimental facilities used to conduct the tests, and present the observed data in an 

organized form (Appendixes A through F) for the Space Shuttle Program. Although there are a number of 

noted observations made about some of the data throughout, it is beyond the scope of this report to draw 

any technical conclusions about what is presented. For the Return to Flight program, both external tank 

foam and ice were tested on RCC. This report covers the foam impact testing program and the ice testing 

program is covered in reference 3. Although this report is largely comprehensive in outlining the details 

and procedures of the actual testing, a complete description can be obtained in the Orbiter RCC Flat Panel 

Impact Testing Plan (ref. 4). 

Test Program 

Program Objectives 

Initially, the objective of the foam impact test program was to establish projectile velocities for each 

material of a fixed size and shape that initiates the threshold of visual damage on simply supported 6- by 

6-in. and 6- by 12-in. flat RCC panels at impact angles of 90° and 45°. However as shuttle program 

requirements evolved, the focus changed to include establishing thresholds for damage detectable by 

nondestructive evaluation (NDE) with ultrasound and pulse thermography. In addition, RCC panel 

deformations were to be measured with a three-dimensional image correlation (digital photogrammetry) 

deformation measurement system throughout the entire spectrum of impact velocities for each test series. 

Observations made during each test were used to corroborate the validity of LS–DYNA deformation and 

damage RCC/projectile constitutive models to reliably predict actual or potential threats to the Space 

Shuttle Orbiter. 

Test Series Description 

The external tank foam impact testing was divided into six series: BX–265 foam on 6- by 6-in. at 90° 

and 45° impact angles, BX–265 foam on 6- by 12-in. at 90° and 45° impact angles, and PDL–1034 foam 

on 6- by 6-in. panels at 90° and 45° impact angles. 
NASA/TM—2009-213642/REV1 3

Each RCC flat panel was visually inspected before and after each impact test, in order to provide an 

indication of coating loss (spalling) due to the impact event. Another objective of this visual examination 

was to provide guidance for static and NDE testing. Posttest NDE inspection and digital photography 

were then performed on each panel. 

Success Criteria

A NASA GRC Quality Assurance (QA) officer was present and provided quality assurance and 

configuration control for all tests. The QA officer verified all test plan requirement compliance and 

ensured that the resulting test data sheets were accurate and complete and that all success criteria were 

met for each test. 

Each individual test was deemed successful provided that the projectile remained intact upon exit of 

the gun barrel and prior to impact on the RCC target panel and that interpretable high-speed digital video 

was acquired from the Vision Research Phantom (Vision Research, Inc.) cameras for velocity and 

deformation measurements. In addition, it was highly desirable that each of the following parameters 

were met: 

 All projectile impact velocity parameters called for in this plan were kept within 20 ft/s 

(as measured from digital high-speed video) 

 Projectile impacted the expected target area within 0.25 in. 

 Projectile rotational motion at impact was 5°, as determined by engineering review of video imagery 

after each shot 

 Vacuum of 0.20.1 psi was maintained in test chamber for each test 

 Load cell response data was not compromised by any extraneous vibratory loads independent of the 

projectile impact event 

Test Facilities 

All of the Level 2 RCC impact tests were conducted in the large vacuum gun (fig. 1) at the NASA GRC 

Ballistic Impact Laboratory in Cleveland, Ohio. The large vacuum gun is a single-stage compressed helium 

type with a 0.33 ft3

 pressure vessel. Barrels from 12 to 25 ft long, and inner diameters of 1.25 to 1.5 in., 

were fit to the chamber depending on the velocity and projectile requirements for each test. 

Figure 1.—Large vacuum gun with 2-in. barrel. 
NASA/TM—2009-213642/REV1 4

The large vacuum chamber has an inside dimension of 5 by 4 by 4 ft. It currently has provisions for 

16 instrumentation feedthroughs (with ability to easily add additional ones). Viewing access ports on the 

front, side, top, and back allow for photo instrumentation with high-speed digital cameras. Four 120-V 

feedthroughs provide power for high-intensity lighting inside the chamber required for the high-speed 

digital imagery. The barrels for the large vacuum gun protrude into the vacuum chamber via a plate 

attachment that creates a circular opening with an o-ring seal to the chamber. Each barrel (each barrel has 

its own plate attachment) used with the vacuum chambers fit snugly against the o-ring to seal it with the 

chamber itself. This gun utilizes a Mylar (DuPont) burst disk system to release the helium propellant gas. 

The burst disk is in contact with an electronically heated nichrome wire, which melts the Mylar, releasing 

the propellant. Impact tests are conducted by simultaneously drawing a vacuum in the vacuum chamber 

and behind the projectile, to avoid pulling the projectile down the gun barrel prior to shooting. 

Experimental Fixtures 

Panel fixtures and mounts made from aluminum were fabricated for use in the large vacuum chamber 

to hold the RCC panels at 90° and 45° to the axis of the gun barrel. The construction was designed to be 

massive to minimize structural ringing due to the impacts, which could adversely affect load cell and 

deformation data. Figure 2 shows the panel fixture and mounts of the inside of the large vacuum chamber 

in its 45° configuration with the lighting system on and the high-speed cameras in place. The side of the 

panel fixture, shown in figures 3 and 4, is called out in the photograph. 

Alignment of the test article with respect to the gun barrel shot line was accomplished through the use 

of a center-bore laser alignment tool inserted in the barrel such that the laser beam represented the 

centerpoint axis of the gun. The beam projected onto the test targets and was used to establish the correct 

fixture alignments and aim points for any given test conditions. 

Figure 2.—RCC flat panel fixture assembly mounted in large vacuum 

chamber for 45° impact tests. 
NASA/TM—2009-213642/REV1 5

Figure 3.—Overview of reinforced carbon-carbon (RCC) flat panel holder fixture with load cells at corners. 

Figure 4.—Exploded view of reinforced carbon-carbon (RCC) flat panel holder fixture. 

Two panel fixtures were fabricated to accommodate the 6- by 6-in. and 6- by 12-in. RCC panels. 

Figures 3 and 4 depict the 6- by 6-in. fixture with a panel in assembled and exploded views, respectively. 

Load cells, discussed in the next section, are shown at the corners of the fixture. The 6- by 6-in. panels 

were simply supported on all four edges in their mounts by being clamped in the frame between halfround aluminum bar stock around the complete perimeter of the panels and secured by 9 bolts torqued to 

8 in.-lb using a calibrated torque wrench. This torque limit was established with a study to establish the 

safe clamping pressure without crushing the RCC material or coating. The 6- by 12-in. panels were only 

restrained (simply supported) on the 6-in. edges at the same torque limit. The frames were designed to 

contact the RCC panels just inside of each panel edge. Hence, the actual distance between the center lines 

of the half-round aluminum bar supports was 5 7/8 in.2

 for the 6- by 6-in. frame, and 11 7/8 in. for the 6- 

by 12-in. frame. 
NASA/TM—2009-213642/REV1 6

Instrumentation and Data Acquisition System 

Load-time histories in three directions (one axial or normal and two shear to the panel face) were 

acquired from four piezoelectric three-axis load cells for each test using load cells mounted in the fixture 

used to hold the flat panels for testing. The fixture had a load cell mounted at each corner of the target 

panels shown in figures 3 and 4. Measurements taken from the load cells were averaged and filtered to 

obtain force time histories of each impact event. 

The load cells were Kistler Model 9067 piezoelectric three-axis load washers (Kistler). They have a 

measurement range of 10,000 lb in the normal axis and 4500 lb in two shear directions. The load cell 

sensitivities are 17 picocoulomb/lb in the normal direction and 35 picocoulomb/lb in the shear direction. 

At the onset of the test program, the full usefulness of the load cell data had yet to be identified, and it 

was recorded for each test essentially because it was relatively straightforward to do so. However, 

because of the complex dynamic response of the overall large vacuum gun test structure, the load cell data 

did not accurately represent the impact loads on the RCC and ultimately used for validating the 

LS–DYNA models. Consequently, load cell data is not presented or discussed further in this report. 

Accelerometers were attached at several locations on the test frame in the large vacuum gun in order 

to identify and resolve potential response anomalies from the impact tests as well as quantify the degree 

of movement that the test frame experienced. Fortunately, there were no anomalies that arose during the 

entire foam test series, and the test frame displacements were within an expected range. 

The application of strain gages to the RCC panels was considered initially as the plan for this test 

program was being developed; however, it was determined that strain gage data on RCC could be 

misleading and of low quality. This was established through laboratory level tests, which indicated that 

strain gages produce erratic data largely due to the craze cracking of the silicon carbide (SiC) coating on 

the panels. The full-field three-dimensional deformation measurements taken with the ARAMIS system 

(discussed below) ultimately provided full-field strain data that was of high value to the validation effort. 

Two data acquisition systems were utilized to record signals from the load cells and accelerometers 

during testing: The first was a Spectral Dynamics model VX2805D 8-channel, 16-bit, 5- Msample/s/ 

channel system with signal conditioning capabilities. The second was an IOtech WaveBook 516E, 16-bit, 

1-Msample/s system with eight analog input channels, eight strain gage conditioning channels, and 8 ICP 

sensor channels. Dual-mode Kistler model 5010B charge amps were used to power the load cells. 

High-Speed Cameras 

High-speed digital Phantom cameras were used to document each impact test. For the majority of the 

program, both Phantom 5 and Phantom 7 cameras from Vision Research (ref. 5) were used and for a 

select number of tests, Photron (ref. 6) FASTCAMS were used. The cameras measured projectile velocity 

and captured the damage and deformation dynamics resulting in each test in addition to being utilized 

with the three-dimensional displacement measurement system. Typically, five high-speed cameras would 

document one of these impact tests. 

Various resolutions and frame rates were used depending on the requirements for each test as there is 

a direct tradeoff between frame rates and image resolution with the Phantom cameras. Higher frame rates 

result in lower resolution images. Typically frame rates of around 29000 frames per second or more were 

used during these tests at resolutions of 256 by 256 or 256 by 128 pixels. It was a test goal to have 

exposure times of 2-20 (driven by light and viewing angle) μs to try to limit motion blur to 2 pixels. 

The Phantom cameras record a continuous 1- to 2-s loop on an internal memory chip until stopped 

and do not require a trigger system to start recording. They are triggered manually to stop data acquisition 

at the sound of the gun blast, thus capturing the impact event in its entirety. Once the cameras were 

triggered, all of the recorded information was downloaded to laptop computers. All high-speed video 

captured in this test program has been archived by the Shuttle Program at Johnson Space Center and is 

easily accessible for viewing to the interested reader (contacts: Justin Kerr or Jim Hyde). 
NASA/TM—2009-213642/REV1 7

Optical Full-Field Three-Dimensional Displacement Image Correlation Measurement System 

Deformation and computed strains of the back side (opposite of the impact side) of the RCC panels 

were obtained using the ARAMIS system built by GOM mbH (ref. 7) acquired from Trilion Optical Test 

Systems (refs. 8 to 10). ARAMIS is a three-dimensional image correlation photogrammetry system that 

captures full field displacement measurements under static, quasistatic, and ballistic impact loading. 

The system works off image pairs taken simultaneously with two Phantom 7 cameras set up on a 

fixed beam, which viewed the back of the RCC panels undergoing impact. The image pairs were typically 

taken at 37.17 increments (~27 000 frames/s) with the exception of those taken in appendix D, which 

were 69.44 increments (~14 400 frames/s). The camera pair was set up outside the large vacuum chamber 

for the 90° impact tests and inside the chamber for the 45° tests. Inside mounting was necessary for the 

45° tests to optimize the view of the panel backs. The outside mount configuration is shown in figure 5. 

In order for the ARAMIS system to make its measurements, a painted speckle or spot pattern must be 

applied to the field area of interest. The backside of each panel was painted before the impact testing was 

performed. Testing was performed on small samples of RCC that were painted for use with the ARAMIS 

system. A scanning electron microscope was used to confirm that the paint did not excessively wick into 

the RCC panels and potentially change the RCC material performance and NDE techniques or flash 

thermography, and ultrasound tests revealed no adverse interference or artifacts due to the speckled paint 

pattern on the backside of the panels. The speckled pattern can be seen on the post-impact photographs in 

the appendices of this report. As the test program proceeded, an optimum paint pattern was adapted going 

from a “spatter” pattern to a pattern spray painted through a uniform stencil. This accounts for the two 

types of patterns seen in the photographs. The paint patterns were created by applying paint to the back of 

each test panel in two layers. The first layer is white paint that has a high reflectivity rating followed by 

black spray paint applied in a random speckle pattern or through the stencil. The inset in the lower righthand corner of figure 5 shows the uniform paint pattern on the back of an RCC panel. 

Figure 5.—ARAMIS Three-Dimensional Deformation/Strain Measurement System 

shown with static camera assembly setup at large vacuum gun viewing ports. 

Inset in lower right corner depicts paint pattern on back of 

the reinforced carbon-carbon panel. 
NASA/TM—2009-213642/REV1 8

Figure 6.—Example output plots from the ARAMIS displacement measurement system. 

Using photogrammetric principals, the three-dimensional coordinates of the surface of the specimen 

can be calculated precisely from observing the paint pattern from two known points of view. On the basis 

of the three-dimensional coordinates, the three-dimensional displacements, the strains, and shape of the 

specimen were calculated with a high degree of accuracy and resolution. The results can be rapidly postprocessed after each test and visualized in similar fashion to finite element results. Three types of 

ARAMIS output are presented in the appendices of this report: Displacement cross sections at various 

times on a panel, center point displacements as a function of time, and full-field color displacement fringe 

plot at maximum displacement of each panel. An example of these plots is shown in figure 6. 

Velocity Measurement 

Posttest evaluation of the high-speed Phantom camera video established impact velocities for each 

test as well as provided verification of projectile integrity and orientation in flight. 

To set up for obtaining projectile velocities, a camera was mounted orthogonal to the path of the 

projectile and a calibration bar was seated in the end gun barrel and projected into the test chamber along 

the centerline axis of the barrel. The calibration bar had precisely machined markings 0.5 in. apart on a 

section the same diameter as the foam projectiles to be shot. Given this device with known markings in 

the same plane of the projectile’s motion, the relationship between physical distance in the plane of the 

projectile’s trajectory and camera screen pixels was determined. The impact velocity of a projectile in a 

given test was then calculated from the following equation: 

NF

FR Vp  SF PIX

where Vp is the velocity of the projectile, SF is the scaling factor in units of length per pixel, PIX is the 

number of pixels a mark on the projectile travels in a given number of frames (NF), and FR is the frame 

rate. 

Pretest and Posttest RCC Panel Observations 

Material Pedigrees 

RCC Panel Pedigree and Fabrication 

The RCC material used in this test program was obtained from Lockheed Martin Missiles & Fire 

Control in Dallas, Texas, through the Shuttle Program Office. Typically the material was provided in 12-

by 12-in. panels and cut to 6- by 6-in. or 6- by 12-in. sizes at Southern Research Institute (SRI) in 

Birmingham, Alabama. The cut plan for the panels is discussed in appendix C of reference 4. In addition, 

it should be noted that due to material loss during the cutting process, the panels were delivered to GRC 

nominally at 5.9 by 5.9 in. and 5.9 by 11.9 in. but are referred to in whole numbers for easier referencing. 

SRI supported a significant number of engineering efforts relating to the RCC material studies. Much of 

this work was complementary and of great use to the LS–DYNA RCC constitutive model development. 
NASA/TM—2009-213642/REV1 9

Figure 7.—Test numbering scheme for reinforced carbon-carbon flat panel impact tests. 

Final reports from two of these studies may be of noteworthy interest to those reading this report: 

“Silicon Carbide Thickness Measurements of Various RCC Panels and Plates” (ref. 11), which provides 

further detail on the panels tested in this report, and “Correlation of RCC Substrate Properties” (ref. 12), 

which sheds some additional light on the weighted contributions of substrate RCC properties to the 

overall properties of coated RCC panels. In addition, it should be noted that significant work also was 

performed to obtain static and dynamic mechanical properties of the RCC material and is presented in 

references 13 and 14. 

In the early stages of this test program, little RCC was available for testing and a number of panels 

were “found” to be in storage at Lockheed, which were made available for these impact tests. These 

panels lacked traceability documents typically associated with any RCC fabricated by Lockheed but were 

considered acceptable for this program. Test results on some of the found panels were demonstrated to be 

suspect and are noted appropriately in the appendices. Found RCC material was only used in the test 

series in appendices A and B (Lots 2 and 3, set 6). For the remaining panels in the test program, Lockheed 

newly manufactured additional 12- by 12-in. panels, which had full traceability documentation. 

RCC material pedigrees to be used in this test program were identified through shipping documents 

that accompanied the deliveries to GRC. Reference to relevant shipping documents was recorded by the 

quality officer for each test to ensure traceability. 

For the purpose of identifying test panels, each flat panel was labeled with the unique identifier 

relating to the 12- by 12-in. parent panel and then assigned an additional number identifying the order in 

which it was cut from the parent panel. The application of this numbering scheme can be seen in the 

tables at the beginning of each appendix under the first column labeled “Test No.” Expanding upon this, a 

test number identifier was assigned to each test which incorporated the assigned panel numbers. It 

included the test number (counted sequentially), the angle of impact, and the projectile material. An 

example of the flat panel/test identification is given in figure 7. This example identifies the first impact 

test in the RCC impact testing was performed on subpanel 1 cut from parent T8015 and shot with BX–

265 at a 90° impact angle. It should be noted that the NASA GRC Ballistic Impact Laboratory maintained 

its own test numbering scheme along with the aforementioned one. Test numbers were assigned 

sequentially after the prefix GRCC. It is presented in the appendices for clarity and completeness as the 

numbering used for the data archives follows this format. 

Foam Pedigree and Fabrication 

For foam impact testing, two materials were used: BX–265 and PDL–1034. All projectiles were 

fabricated foam sprayed and fabricated at NASA Michoud Assembly Facility (MAF) per STP–1540 

(Lockheed standard for the spraying of BX–265) and STP–1532 (Lockheed standard for the spraying of 

PDL–1034). Projectiles were rough cut from the 2-ft by 2-ft by 4-in.-thick base panels in the “Parallel to 

Rise” direction per figure 8. Rough-cut blocks were then mounted into a lathe and turned down to a final 
NASA/TM—2009-213642/REV1 10

Figure 8.—External tank foam cut direction nomenclature. 

diameter of 1.25 in. (0, –0.015 in.) for 90° tests and 1.50 in. (0, –0.015 in.) for 45° tests and cut to a 

final length of 3.000.01 in. Further machining of the length as necessary was performed to produce 

projectiles with consistent weight (0.001 lb). 

Digital Photography 

Digital photographs were taken of the front, back, and all four edges of each RCC panels both prior to 

and after impact testing to document any changes in the panel condition after testing. All of the posttest 

images are presented in this report for each panel. Both pretest and posttest images have been archived in 

digital form by the Shuttle Program and are easily accessible to the interested reader (contacts: Justin Kerr 

or Jim Hyde). In addition, these photographs have been all assigned NASA GRC C-numbers (seen in the 

lower right-hand corner of the appendices images) and will remain in the GRC archives, accessible 

through the NASA GRC Imaging Technology Center. 

Nondestructive Evaluation 

Nondestructive evaluation (NDE) was performed on each RCC panel before and after testing to 

baseline the specimens and to evaluate the damage due to impact. The NDE not only characterized any 

damage due to the impact testing, but the findings from each panel’s evaluation fed back into defining the 

parameters of the panel testing to follow. All NDE outlined in this report was performed by the NDE 

group at NASA GRC. Two NDE methods will be utilized for this program: pulse, or flash, thermography 

and through-transmission ultrasound. 

Pulsed, or flash, thermography involves the heating of a specimen with a short duration pulse of 

energy and monitoring the transient thermal response of the surface of the specimen with an infrared 

camera. The thermal energy on the surface conducts into the cooler interior of the sample. In turn, there is 

a reduction of the surface temperature over time. This surface cooling will occur in a uniform manner as 

long as the material properties are consistent throughout the specimen. Subsurface defects that possess 

different material properties (e.g., thermal conductivity, density, or heat capacity) will affect the flow of 

heat in that particular region. This resistance in the conductive path causes a different cooling rate at the 

surface directly above the defect, when compared to the surrounding, defect-free material. The change in 

the subsurface conduction is seen as a nonuniform surface temperature profile as a function of time. Since 
NASA/TM—2009-213642/REV1 11

Figure 9.—Nondestructive evaluation facilities at NASA Glenn: through transmission ultrasonic immersion tank 

with reinforced carbon-carbon panel undergoing scanning (left) and thermal imaging setup to perform pulse 

thermography (right). Insets in the left corners of each photo depict sample output from each technique. 

the method depends on the interaction of the defect with the advancing thermal front, defects that are 

located at greater depths will show up later. Due to lateral diffusion, deeper defects will tend to have less 

contrast than near-surface flaws. Therefore, the critical flaw size capability of a thermographic inspection 

system is a function of the defect size, depth, and the material properties of the component being tested. 

Analysis of thermographic data involves examination of images based on the temperature-time data or 

derivatives calculated from the original data sets. Anomalous areas can then be identified based on 

deviations in the cooling behavior. Figure 9 (right) shows the thermal imaging setup used to conduct 

pulse thermography with an example output image. 

Through-transmission ultrasonic inspection utilizes two transducers, placed on opposite sides of a 

material for interrogation. One transducer sends an ultrasonic pulse through the material where it is 

received by the second. In scanning mode, the transducer pair is moved across the area of interest, and an 

image based on the amplitude of the received waveform is generated. Defects and other significant 

variations will result in the additional attenuation and scattering of the ultrasonic signal as it passes 

through the material, thus reducing the signal amplitude. Flaws are located in the image based on this 

decrease in signal amplitude. Minimum flaw resolution is a function of the wavelength of the ultrasonic 

signal flaw orientation. Resolution, in general, increases with increasing frequency. Figure 9 (left) shows 

the immersion ultrasonic tank and relevant hardware to perform the through-transmission ultrasonic 

inspection with a sample image. 

Through-transmission ultrasound evaluation will require that the RCC panels be immersed in water 

during the scan process. Each RCC panel was weighed before each scan and then vacuum dried until the 

after-scan weight is that of the initial weight to ensure that all water absorbed from the immersion was 

removed from the panel. The vacuum drying process was accomplished referencing Southern Research 

standard specification for drying RCC panels. This specification used 180 °F heat while pulling a vacuum 

of approximately 10–3 torr. The weight of the flat panel was monitored until stabilization occurs. 

Testing Summaries 

Below are brief summaries of the external tank foam impact test program as organized into impact 

angle and foam type. Comprehensive data sets reside in each of the six appendices in the back of this 

report. 
NASA/TM—2009-213642/REV1 12

BX–265 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 90° Impact Angle 

(Appendix A) 

Thirteen shots were conducted in this test series with impact velocities ranging from 1388 to 

2109 ft/s. The RCC material shot in this test series was “found” material. 

Panel A–146 is considered to be an anomalous panel in this test series. For reasons not fully 

quantified, this panel was prone to severe damage and lower impact velocities than other panels and was 

considered defective. Test results from Panel A–146 were consequently not considered valid but are 

presented in this report for completeness. 

BX–265 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact Angle 

(Appendix B) 

Thirteen shots were conducted in this test series with impact velocities ranging from 1244 to 

2440 ft/s. The RCC material shot in this test series was “found” material. 

Panels from parent R1–47 is considered suspect panels in this test series. For reasons not fully 

quantified, tests with these panels yielded inconsistent results as compared to other panels in the test 

series. Test results from the R1–47 panels were consequently not considered as valid but are presented in 

this report for completeness. 

BX–265 on 6- by 12-in. Reinforced Carbon-Carbon (RCC) Panels at 90° Impact Angle 

(Appendix C) 

Seven shots were conducted in this test series with impact velocities ranging from 1054 to 1849 ft/s. 

The RCC material for this test series was tested in the as-manufactured condition. 

Tests GRCC 169 and GRCC 170 were shot on the same panel, 54–1. GRCC 169, the foam projectile, 

came out damaged, crooked, and slow. An impromptu experiment was then conducted to determine if 

damage does accumulate. The second shot, GRCC 170 produced NDE detectable damage, despite its low 

velocity. Time constraints in the test lab did not allow for ultrasound or thermography to be performed 

between tests. Although not definitive in nature, this experiment did draw interest to the concept that 

damage might accumulate in the RCC material as a consequence of sequential hits that singularly would 

not damage the RCC. 

All of the 6- by 12-in. panel tests were only simply supported on two ends. As a consequence, the 

wave propagation in these panels were dramatically different and less controlled and damped than in the 

6- by 6-in. panels resulting in additional modes of vibration in the panels being significant which can be 

identified in the ARAMIS data presented in the appendix. 

BX–265 on 6- by 12-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact Angle 

(Appendix D) 

Five shots were conducted in this test series with impact velocities ranging from 1381 to 2000 ft/s. 

The RCC material for this test series was tested in the as-manufactured condition. 

All of the 6- by 12-in. panel tests were only simply supported on two ends. As a consequence, the 

wave propagation in these panels were dramatically different and less controlled and damped than in the 

6- by 6-in. panels resulting in additional modes of vibration in the panels being significant which can be 

identified in the ARAMIS data presented in the appendix. 
NASA/TM—2009-213642/REV1 13

PDL–1034 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 90° Impact Angle 

(Appendix E) 

Eight shots were conducted in this test series with impact velocities ranging from 960 to 1825 ft/s. 

The RCC material for this test series was tested in the as-manufactured condition. 

PDL–1034 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact Angle 

(Appendix F) 

Eight shots were conducted in this test series with impact velocities ranging from 1122 to 2105 ft/s. 

The RCC material for this test series was tested in the as-manufactured condition. 

Concluding Remarks 

The Level 2 reinforced carbon-carbon (RCC) flat panel impact test program at the NASA Glenn 

Research Center Ballistic Impact Laboratory was successfully completed on time supporting NASA’s 

Return to Flight with the STS–114 mission. The data in the appendices of this report present the results 

from 54 external tank foam impact tests (38 with BX–265 and 16 with PDL–1034), which were used to 

demonstrate the validity of BX–265, PDL, and RCC models developed and implemented in the LS–

DYNA impact analysis program. As a point of interest, the final validation step of the LS–DYNA models 

was through correlation with observations from follow-on full-scale orbiter wing leading edge and nose 

cap impact tests performed at Southwest Research Institute (SwRI). 

In preparation for the STS–114 launch, virtually hundreds of analyses with LS–DYNA were 

performed to establish certified impact damage thresholds for RCC thermal protection systems on the 

orbiter helping to recertify the shuttle system for flight. For the interested reader, references 15 through 25 

provide additional details on much of the analysis development process with LS–DYNA, and references 

26 through 29 highlight the efforts preformed to evaluate ice impacts on RCC. Reference 3 is the 

companion report to this one which provides all of the post-impact test data on the RCC panels from that 

program. 
NASA/TM—2009-213642/REV1 14

Appendices 

Data from the external tank foam impact tests are presented in the following six appendices 

(A through F) as follows: 

Appendix A—BX–265 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Flat Panels at 90° 

Impact Angle 

Appendix B—BX–265 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact 

Angle 

Appendix C—BX–265 on 6- by 12-in. Reinforced Carbon-Carbon (RCC) Panels at 90° Impact 

Angle 

Appendix D—BX–265 on 6- by 12-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact 

Angle 

Appendix E—PDL–1034 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 90° Impact 

Angle 

Appendix F—PDL–1034 on 6- by 6-in. Reinforced Carbon-Carbon (RCC) Panels at 45° Impact 

Angle 

In each appendix, the data are organized in the following fashion: 

1. A table summarizing the appendix test series. 

2. Baseline and posttest imagery from ultrasound and thermography nondestructive evaluations 

arranged in order of ascending impact velocities. (It should be noted that in the thermography 

imagery, black spots appear in the upper left or right corners of the panels. This is an artifact of 

the fixturing to hold the panels during evaluation.) 

3. Displacement contours output from ARAMIS showing the displacements along a cross-sectional 

line on the back of the panels at uniform increments of time as the panels were impacted. Note 

the sectional line for the 6- by 12-in. panels were taken along the long edge of the panels. 

4. Displacement trace plots output from ARAMIS of the panel centerpoints as a function of time. 

5. ARAMIS full-field color fringe plots taken from the point of maximum displacement of the panel 

resulting from the impact. (Note that in some of these plots, coating loss on the backside of the 

panels obscured ARAMIS from measuring accurate displacements. In these cases, the plots 

shown were selected just before or after failure of the panels. Knowing maximum displacement 

for a panel subject to failure is less meaningful for analysis correlation since the structure of the 

panel is compromised.) 

6. High-resolution digital photography of each panel show front, back, and edge composite views as 

well as isometric views of both front and back of each panel. 
NASA/TM—2009-213642/REV1 15

Appendix A.—Test Data 

BX–265 External Tank Foam Impact Testing at 90° Angle on 6- by 6-in. Reinforced 

Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix A Test Series 

1. The RCC material used in this test series was “found” material. 

2. For the first several impact tests in this test series, no digital photographs were taken of the panels 

from angles other than 90°. Only front, back, and side images were taken. Shortly into the program, it 

was determined that photos taken from an angle other than 90° would help to further visually 

distinguish damage on the panels. As a consequence, several of the tests in Appendix A will not have 

these images included. 

3. Panel A–146 is considered to be an anomalous panel in this test series. For reasons not fully 

quantified, this panel was prone to severe damage and lower impact velocities than other panels and 

was considered defective. Test results from panel A–146 were consequently not considered valid but 

are presented in this report for completeness. 

4. Lot 2 in the following test series table refers to remnant RCC material from the original qualification 

of ENKA (American Enka Corp.) fabric. 

5. Lot 3, set 6 in the following test series table was used by Lockheed to indicate any piece of RCC that 

they had that could support the Return to Flight Program. They were at the RCC–3 Bimatrix 

condition, and resumed processing for silicon carbide (SiC) conversion coating, TEOS (tetra-ethylortho-silicate), and type A. 
NASA/TM—2009-213642/REV1 16

 

Appendix A Test Series
NASA/TM—2009-213642/REV1 17

Figure A1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) 

at a 90° angle.
NASA/TM—2009-213642/REV1 18

Figure A1–2.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) 

at a 90° angle. 
NASA/TM—2009-213642/REV1 19

Figure A2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels measured at 37-μs increments undergoing impact with BX–265 foam cylinders (nominally 1.25 in. 

in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 20

Figure A2–2.—ARAMIS out-of-plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels plotted at 37-s increments undergoing impact with BX–265 foam cylinders (nominally 1.25 in. in diameter 

by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 21

Figure A3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 90° angle.
NASA/TM—2009-213642/REV1 22

Figure A3–2.—ARAMIS centerpoint out-of-plane deformation versus time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 23

Figure A4–1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.25 in. in 

diameter by 3 in.) at a 90° angle.
NASA/TM—2009-213642/REV1 24

Figure A4–2.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.25 in. in 

diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 25

Figure A5–1.—Digital photography of edges and faces of panel T8015–2 at 1388 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 17. 
NASA/TM—2009-213642/REV1 26

Figure A6–1.—Digital photography of edges and faces of panel T8015–3 at 1717 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 19. 
NASA/TM—2009-213642/REV1 27

Figure A7–1.—Digital photography of edges and faces of panel A146–4 at 1741 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 25. 
NASA/TM—2009-213642/REV1 28

Figure A8–1.—Digital photography of edges and faces of panel A146–3 at 1845 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 23. 
NASA/TM—2009-213642/REV1 29

Figure A9–1.—Digital photography of edges and faces of panel T8015–4 at 1907 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 20. 
NASA/TM—2009-213642/REV1 30

Figure A10–1.—Digital photography of edges and faces of panel A150–18 at 1952 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 60. 
NASA/TM—2009-213642/REV1 31

Figure A10–2.—Digital photography front (impact side) face of panel A150–18 at 1952 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 60. 
NASA/TM—2009-213642/REV1 32

Figure A10–3.—Digital photography of back face of panel A150–18 at 1952 ft/s with a BX–265 foam 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 60. 
NASA/TM—2009-213642/REV1 33

Figure A11–1.—Digital photography of edges and faces of panel A150–19 at 1978 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 61. 
NASA/TM—2009-213642/REV1 34

Figure A11–2.—Digital photography front (impact side) face of panel A150–19 at 1978 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 61. 
NASA/TM—2009-213642/REV1 35

Figure A11–3.—Digital photography of back face of panel A150–19 at 1978 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 61. 
NASA/TM—2009-213642/REV1 36

Figure A12–1.—Digital photography of edges and faces of panel A150–17 at 2000 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 59. 
NASA/TM—2009-213642/REV1 37

Figure A12–2.—Digital photography front (impact side) face of panel A150–17 at 2000 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 59. 
NASA/TM—2009-213642/REV1 38

Figure A12–3.—Digital photography of back face of panel A150–17 at 2000 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 59. 
NASA/TM—2009-213642/REV1 39

Figure A13–1.—Digital photography of edges and faces of panel A146–2 at 2002 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 22. 
NASA/TM—2009-213642/REV1 40

Figure A14–1.—Digital photography of edges and faces of panel A146–1 at 2015 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 21. 
NASA/TM—2009-213642/REV1 41

Figure A15–1.—Digital photography of edges and faces of panel T8015–1 at 2054 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 15. 
NASA/TM—2009-213642/REV1 42

Figure A16–1.—Digital photography of edges and faces of panel P20L–23 at 2077 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 57. 
NASA/TM—2009-213642/REV1 43

Figure A16–2.—Digital photography front (impact side) face of panel P20L–23 at 2077 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 57. 
NASA/TM—2009-213642/REV1 44

Figure A16–3.—Digital photography of back face of panel P20L–23 at 2077 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 57. 
NASA/TM—2009-213642/REV1 45

Figure A17–1.—Digital photography of edges and faces of panel P20L–24 at 2109 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 58. 
NASA/TM—2009-213642/REV1 46

Figure A17–2.—Digital photography front (impact side) face of panel P20L–24 at 2109 ft/s with a long BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 58. 
NASA/TM—2009-213642/REV1 47

Figure A17–3.—Digital photography of back face of panel P20L–24 at 2109 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 58. 

NASA/TM—2009-213642/REV1 49

Appendix B.—Test Data 

BX–265 External Tank Foam Impact Testing at a 45° Angle on 6- by 6-in. 

Reinforced Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix B Test Series 

1. The reinforced carbon-carbon material used in this test series was “found” material. 

2. Panels from parent R1–47 are considered suspect panels in this test series. For reasons not fully 

quantified, tests with these panels yielded inconsistent results as compared to other panels in the test 

series. Test results from the R1–47 panels were consequently not considered as valid but are 

presented in this report for completeness. 

3. Lot 2 in the following test series table refers to remnant reinforced carbon-carbon material from the 

original qualification of ENKA (American Enka Corp.) fabric. 

4. Lot 3, set 6 in the following test series table was used by Lockheed to indicate any piece of reinforced 

carbon-carbon that they had that could support the Return to Flight Program. They were at the 

reinforced carbon-carbon–3 Bimatrix condition, and resumed processing for silicon carbide (SiC) 

conversion coating, TEOS (tetra-ethyl-ortho-silicate, and type A. 
NASA/TM—2009-213642/REV1 50

 

Appendix B Test Series
NASA/TM—2009-213642/REV1 51

Figure B1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 45° 

angle. 
NASA/TM—2009-213642/REV1 52

Figure B1–2.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 45° 

angle. 
NASA/TM—2009-213642/REV1 53

Figure B2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels measured at 37-μs increments undergoing impact with BX–265 foam cylinders (nominally 1.25 in. in 

diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 54

Figure B2–2.—ARAMIS out-of -plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels plotted at 37- μs increments undergoing impact with BX–265 foam cylinders (nominally 1.25 in. in diameter 

by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 55

Figure B3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 56

Figure B3–2.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.25 in. in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 57

Figure B4–1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.25 in. in 

diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 58

Figure B4–2.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.25 in. in 

diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 59

Figure B5–1.—Digital photography of edges and faces of panel R285–9 at 1244 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 41. 
NASA/TM—2009-213642/REV1 60

Figure B5–2.—Digital photography front (impact side) face of panel R285–9 at 1244 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 41. 
NASA/TM—2009-213642/REV1 61

Figure B5–3.—Digital photography of back face of panel R285–9 at 1244 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 41. 
NASA/TM—2009-213642/REV1 62

Figure B6–1.—Digital photography of edges and faces of panel R147–14 at 1471 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 30. 
NASA/TM—2009-213642/REV1 63

Figure B6–2.—Digital photography front (impact side) face of panel R147–14 at 1471 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 30. 
NASA/TM—2009-213642/REV1 64

Figure B6–3.—Digital photography of back face of panel R147–14 at 1471 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 30. 
NASA/TM—2009-213642/REV1 65

Figure B7–1.—Digital photography of edges and faces of panel R285–10 at 1837 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 43. 
NASA/TM—2009-213642/REV1 66

Figure B7–2.—Digital photography front (impact side) face of panel R285–10 at 1837 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 43. 
NASA/TM—2009-213642/REV1 67

Figure B7–3.—Digital photography of back face of panel R285–10 at 1837 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 43. 
NASA/TM—2009-213642/REV1 68

Figure B8–1.—Digital photography of edges and faces of panel R284–21 at 1910 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 52. 
NASA/TM—2009-213642/REV1 69

Figure B8–2.—Digital photography front (impact side) face of panel R284–21 at 1910 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 52. 
NASA/TM—2009-213642/REV1 70

Figure B8–3.—Digital photography of back face of panel R284–21 at 1910 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 52. 
NASA/TM—2009-213642/REV1 71

Figure B9–1.—Digital photography of edges and faces of panel A112–26 at 1935 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 55. 
NASA/TM—2009-213642/REV1 72

Figure B9–2.—Digital photography front (impact side) face of panel A112–26 at 1935 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 55. 
NASA/TM—2009-213642/REV1 73

Figure B9–3.—Digital photography of back face of panel A112–26 at 1935 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 55. 
NASA/TM—2009-213642/REV1 74

Figure B10–1.—Digital photography of edges and faces of panel A112–25 at 1940 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 54. 
NASA/TM—2009-213642/REV1 75

Figure B10–2.—Digital photography front (impact side) face of panel A112–25 at 1940 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 54. 
NASA/TM—2009-213642/REV1 76

Figure B10–3.—Digital photography of back face of panel A112–25 at 1940 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 54. 
NASA/TM—2009-213642/REV1 77

Figure B11–1.—Digital photography of edges and faces of panel R147–15 at 1947 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 49. 
NASA/TM—2009-213642/REV1 78

Figure B11–2.—Digital photography front (impact side) face of panel R147–15 at 1947 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 49. 
NASA/TM—2009-213642/REV1 79

Figure B11–3.—Digital photography of back face of panel R147–15 at 1947 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 49. 
NASA/TM—2009-213642/REV1 80

Figure B12–1.—Digital photography of edges and faces of panel R147–16 at 1987 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 48. 
NASA/TM—2009-213642/REV1 81

Figure B12–2.—Digital photography front (impact side) face of panel R147–16 at 1987 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 48. 
NASA/TM—2009-213642/REV1 82

Figure B12–3.—Digital photography of back face of panel R147–16 at 1987 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 48. 
NASA/TM—2009-213642/REV1 83

Figure B13–1.—Digital photography of edges and faces of panel R284–22 at 2035 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 53. 
NASA/TM—2009-213642/REV1 84

Figure B13–2.—Digital photography front (impact side) face of panel R284–22 at 2035 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 53. 
NASA/TM—2009-213642/REV1 85

Figure B13–3.—Digital photography of back face of panel R284–22 at 2035 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 53. 
NASA/TM—2009-213642/REV1 86

Figure B14–1.—Digital photography of edges and faces of panel R284–20 at 2230 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact. Test GRCC 51. 
NASA/TM—2009-213642/REV1 87

Figure B14–2.—Digital photography front (impact side) face of panel R284–20 at 2230 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 51. 
NASA/TM—2009-213642/REV1 88

Figure B14–3.—Digital photography of back face of panel R284–20 at 2230 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 51. 
NASA/TM—2009-213642/REV1 89

Figure B15–1.—Digital photography of edges and faces of panel R285–12 at 2241 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 46. 
NASA/TM—2009-213642/REV1 90

Figure B15–2.—Digital photography front (impact side) face of panel R285–12 at 2241 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 46. 
NASA/TM—2009-213642/REV1 91

Figure B15–3.—Digital photography of back face of panel R285–12 at 2241 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 46. 
NASA/TM—2009-213642/REV1 92

Figure B16–1.—Digital photography of edges and faces of panel R147–13 at 2371 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 29. 
NASA/TM—2009-213642/REV1 93

Figure B16–2.—Digital photography front (impact side) face of panel R147–13 at 2371 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 29. 
NASA/TM—2009-213642/REV1 94

Figure B16–3.—Digital photography of back face of panel R147–13 at 2371 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 29. 
NASA/TM—2009-213642/REV1 95

Figure B17–1.—Digital photography of edges and faces of panel R285–11 at 2440 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 45. 
NASA/TM—2009-213642/REV1 96

Figure B17–2.—Digital photography front (impact side) face of panel R285–11 at 2440 ft/s with a BX–265 foam 

cylinder (nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 45. 
NASA/TM—2009-213642/REV1 97

Figure B17–3.—Digital photography of back face of panel R285–11 at 2440 ft/s with a BX–265 foam cylinder 

(nominally 1.25 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 45. 

NASA/TM—2009-213642/REV1 99

Appendix C.—Test Data 

BX–265 External Tank Foam Impact Testing at 90° Angle on 6- by 12-in. 

Reinforced Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix C Test Series 

1. The RCC material used in this test series was tested in the as-manufactured condition. 

2. Tests GRCC 169 and 170 were shot on the same panel, 54–1. GRCC 169 the foam projectile 

came out damaged, crooked, and slow. An impromptu experiment was then conducted to 

determine if damage does accumulate. The second shot, GRCC 170, produced NDE detectable 

damage, despite its low velocity. Time constraints in the test lab did not allow for ultrasound or 

thermography to be performed between tests. 

3. The 6- by 12-in. panel tests were only simply supported on two ends. As a consequence, the wave 

propagation in these panels was dramatically different and less controlled and damped than in the 

6- by 6-in. panels resulting in additional modes of vibration in the panels being significant. This 

behavior can be seen in both the C2–1 and C3–1 figures. In C2–1, this is noted by the irregular 

deformation contours plot as a function of time. In C3–1 the secondary modes of vibration are 

seen in the first loading cycle and dampen out in the second cycle. 

4. Lot 3, (37–57) in the following test series table refers to Lockheed Set 3A, processing Batch 37–

57, made specifically for the Return to Flight Program. 
NASA/TM—2009-213642/REV1 100

 

Appendix C Test Series
NASA/TM—2009-213642/REV1 101

Figure C1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 12-in. flat panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) 

at a 90° angle. 
NASA/TM—2009-213642/REV1 102

Figure C1–2.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 12-in. flat panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) 

at a 90° angle. 
NASA/TM—2009-213642/REV1 103

Figure C2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 12-in. reinforced carbon-carbon 

flat panels measured at 37-s increments undergoing impact with BX–265 foam cylinders (nominally 1.5 in. 

in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 104

Figure C3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 12-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 105

Figure C4–1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 12-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.5 in. in 

diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 106

Figure C5–1.—Digital photography of edges and faces of panel 54–1 at 1194 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 170. 
NASA/TM—2009-213642/REV1 107

Figure C5–2.—Digital photography front (impact side) face of panel 54–1 at 1194 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 170. 
NASA/TM—2009-213642/REV1 108

Figure C5–3.—Digital photography of back face of panel 54–1 at 1194 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 170. 
NASA/TM—2009-213642/REV1 109

Figure C6–1.—Digital photography of edges and faces of panel 41–1 at 1262 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 161. 
NASA/TM—2009-213642/REV1 110

Figure C6–2.—Digital photography front (impact side) face of panel 41–1 at 1262 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 161. 
NASA/TM—2009-213642/REV1 111

Figure C6–3.—Digital photography of back face of panel 41–1 at 1262 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 161. 
NASA/TM—2009-213642/REV1 112

Figure C7–1.—Digital photography of edges and faces of panel 50–2 at 1264 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 168. 
NASA/TM—2009-213642/REV1 113

Figure C7–2.—Digital photography front (impact side) face of panel 50–2 at 1264 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 168. 
NASA/TM—2009-213642/REV1 114

Figure C7–3.—Digital photography of back face of panel 50–2 at 1264 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 168. 
NASA/TM—2009-213642/REV1 115

Figure C8–1.—Digital photography of edges and faces of panel 48–2 at 1461 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 165. 
NASA/TM—2009-213642/REV1 116

Figure C8–2.—Digital photography front (impact side) face of panel 48–2 at 1461 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 165. 
NASA/TM—2009-213642/REV1 117

Figure C8–3.—Digital photography of back face of panel 48–2 at 1461 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 165. 
NASA/TM—2009-213642/REV1 118

Figure C9–1.—Digital photography of edges and faces of panel 41–2 at 1538 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 163. 
NASA/TM—2009-213642/REV1 119

Figure C9–2.—Digital photography front (impact side) face of panel 41–2 at 1538 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 163. 
NASA/TM—2009-213642/REV1 120

Figure C9–3.—Digital photography of back face of panel 41–2 at 1538 ft/s with a BX–265 foam cylinder (nominally 

1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 163. 
NASA/TM—2009-213642/REV1 121

Figure C10–1.—Digital photography of edges and faces of panel 50–1 at 1849 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 167. 
NASA/TM—2009-213642/REV1 122

Figure C10–2.—Digital photography front (impact side) face of panel 50–1 at 1849 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in diam. by 3 in.) at a 90° impact angle. Test GRCC 167. 
NASA/TM—2009-213642/REV1 123

Figure C10–3.—Digital photography of back face of panel 50–1 at 1849 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 167. 

NASA/TM—2009-213642/REV1 125

Appendix D.—Test Data 

BX–265 External Tank Foam Impact Testing at 45° Angle on 6- by 12-in. 

Reinforced Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix D Test Series 

1. The RCC material used in this test series was tested in the as-manufactured condition. 

2. Note that in figures D2–1 and D2–2, the sampling rate is 69.44 s than the other test series taken 

at 37.17 s. This is due to the change in the viewing area for this series being photographed at 

higher resolution to capture the larger rectangular plates (this was not done for the test series 

detailed in appendix C). Sample rate is directly correlated to capture resolution on the Phantom 

high-speed digital cameras. 

3. The 6- by 12-in. panel tests were only simply supported on two ends. As a consequence, the wave 

propagation in these panels were dramatically different and less controlled than in the 6- by 6-in. 

panels resulting in additional modes of vibration in the panels being significant. This behavior 

can be seen in both the D2–1 and D3–1 figures. In D2–1, this is noted by the irregular 

deformation contours plot as a function of time. In D3–1 the secondary modes of vibration are 

seen in the first loading cycle and dampen out in the second cycle. 

4. In test GRCC 178, the foam projectile broke into two pieces which were tracked at 1970 and 

2000 ft/s respectively at impact on RCC panel 57–1. No visible or NDE damage was detected due 

to this test. 

5. Lot 3, (37–57) in the following test series table refers to Lockheed Set 3A, processing Batch 37–

57, made specifically for the Return to Flight Program. 
NASA/TM—2009-213642/REV1 126 Appendix D Test Series
NASA/TM—2009-213642/REV1 127

Figure D1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 12-in. flat panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) 

at a 45° angle. 
NASA/TM—2009-213642/REV1 128

Figure D1–2.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 12-in. flat panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) 

at a 45° angle. 
NASA/TM—2009-213642/REV1 129

Figure D2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 12-in. reinforced carbon-carbon 

flat panels measured at 69-s increments undergoing impact with BX–265 foam cylinders (nominally 1.5 in. 

in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 130

Figure D3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 12-in. reinforced carbon-carbon flat 

panels impacted with BX–265 foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 131

Figure D4–1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 12-in. 

reinforced carbon-carbon flat panels as they undergo impact with BX–265 foam cylinders (nominally 1.5 in. 

in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 132

Figure D5–1.—Digital photography of edges and faces of panel 54–2 at 1381 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 172. 
NASA/TM—2009-213642/REV1 133

Figure D5–2.—Digital photography front (impact side) face of panel 54–2 at 1381 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 172. 
NASA/TM—2009-213642/REV1 134

Figure D5–3.—Digital photography of back face of panel 54–2 at 1381 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 172. 
NASA/TM—2009-213642/REV1 135

Figure D6–1.—Digital photography of edges and faces of panel 55–1 at 1795 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 174. 
NASA/TM—2009-213642/REV1 136

Figure D6–2.—Digital photography front (impact side) face of panel 55–1 at 1795 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 174. 
NASA/TM—2009-213642/REV1 137

Figure D6–3.—Digital photography of back face of panel 55–1 at 1795 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 174. 
NASA/TM—2009-213642/REV1 138

Figure D7–1.—Digital photography of edges and faces of panel 55–2 at 1853 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 176. 
NASA/TM—2009-213642/REV1 139

Figure D7–2.—Digital photography front (impact side) face of panel 55–2 at 1853 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 176. 
NASA/TM—2009-213642/REV1 140

Figure D7–3.—Digital photography of back face of panel 55–2 at 1853 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 176. 
NASA/TM—2009-213642/REV1 141

Figure D8–1.—Digital photography of edges and faces of panel 57–2 at 2230 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 180. 
NASA/TM—2009-213642/REV1 142

Figure D8–2.—Digital photography front (impact side) face of panel 57–2 at 2230 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 180. 
NASA/TM—2009-213642/REV1 143

Figure D8–3.—Digital photography of back face of panel 57–2 at 2230 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 180. 
NASA/TM—2009-213642/REV1 144

Figure D9–1.—Digital photography of edges and faces of panel 57–1 at 1970 to 2000 ft/s with a BX–265 foam 

cylinder (nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 178. 
NASA/TM—2009-213642/REV1 145

Figure D9–2.—Digital photography front (impact side) face of panel 57–1 at 1970 to 2000 ft/s with a BX–265 foam 

cylinder (nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 178. 
NASA/TM—2009-213642/REV1 146

Figure D9–3.—Digital photography of back face of panel 57–1 at 1970 to 2000 ft/s with a BX–265 foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 178. 
NASA/TM—2009-213642/REV1 147

Appendix E.—Test Data 

PDL External Tank Foam Impact Testing at 90° Angle on 6- by 6-in. 

Reinforced Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix E Test Series 

1. The RCC material used in this test series was tested in the as-manufactured condition. 

2. Lot 3,(22–36) in the following test series table refers to Lockheed Set 3A, processing Batch 22–

36, made specifically for the Return to Flight Program. 

 
NASA/TM—2009-213642/REV1 148

 

Appendix E Test Series
NASA/TM—2009-213642/REV1 149

Figure E1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with PDL foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 150

Figure E2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels measured at 37.1-μs increments undergoing impact with PDL foam cylinders (nominally 1.5 in. in diameter 

by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 151

Figure E3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with PDL foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 152

Figure E4-1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in.

reinforced carbon-carbon flat panels as they undergo impact with PDL foam cylinders (nominally 1.5 in. in diameter

by 3 in.) at a 90° angle. 
NASA/TM—2009-213642/REV1 153

Figure E5–1.—Digital photography of edges and faces of panel 23–1 at 960 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 192. 
NASA/TM—2009-213642/REV1 154

Figure E5–2.—Digital photography front (impact side) face of panel 23–1 at 960 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 192. 
NASA/TM—2009-213642/REV1 155

Figure E5–3.—Digital photography of back face of panel 23–1 at 960 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 192. 
NASA/TM—2009-213642/REV1 156

Figure E6–1.—Digital photography of edges and faces of panel 26–3 at 1267 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 208. 
NASA/TM—2009-213642/REV1 157

Figure E6–2.—Digital photography front (impact side) face of panel 26–3 at 1267 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 208. 
NASA/TM—2009-213642/REV1 158

Figure E6–3.—Digital photography of back face of panel 26–3 at 1267 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 208. 
NASA/TM—2009-213642/REV1 159

Figure E7–1.—Digital photography of edges and faces of panel 27–2 at 1350 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 213. 
NASA/TM—2009-213642/REV1 160

Figure E7–2.—Digital photography front (impact side) face of panel 27–2 at 1350 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 213. 
NASA/TM—2009-213642/REV1 161

Figure E7–2.—Digital photography front (impact side) face of panel 27–2 at 1350 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 213. 
NASA/TM—2009-213642/REV1 162

Figure E8–1.—Digital photography of edges and faces of panel 27–1 at 1392 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 211. 
NASA/TM—2009-213642/REV1 163

Figure E8–2.—Digital photography front (impact side) face of panel 27–1 at 1392 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 211. 
NASA/TM—2009-213642/REV1 164

Figure E8–3.—Digital photography of back face of panel 27–1 at 1392 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 211. 
NASA/TM—2009-213642/REV1 165

Figure E9–1.—Digital photography of edges and faces of panel 26–2 at 1459 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 200. 
NASA/TM—2009-213642/REV1 166

Figure E9–2.—Digital photography front (impact side) face of panel 26–2 at 1459 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 200. 
NASA/TM—2009-213642/REV1 167

Figure E9–3.—Digital photography of back face of panel 26–2 at 1459 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 200. 
NASA/TM—2009-213642/REV1 168

Figure E10–1.—Digital photography of edges and faces of panel 24–1 at 1541 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 195. 
NASA/TM—2009-213642/REV1 169

Figure E10–2.—Digital photography front (impact side) face of panel 24–1 at 1541 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 195. 
NASA/TM—2009-213642/REV1 170

Figure E10–3.—Digital photography of back face of panel 24–1 at 1541 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 195. 
NASA/TM—2009-213642/REV1 171

Figure E11–1.—Digital photography of edges and faces of panel 26–1 at 1718 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 198. 
NASA/TM—2009-213642/REV1 172

Figure E11–2.—Digital photography front (impact side) face of panel 26–1 at 1718 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 198. 
NASA/TM—2009-213642/REV1 173

Figure E11–3.—Digital photography of back face of panel 26–1 at 1718 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 198. 
NASA/TM—2009-213642/REV1 174

Figure E12–1.—Digital photography of edges and faces of panel 24–2 at 1825 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 203. 
NASA/TM—2009-213642/REV1 175

Figure E12–2.—Digital photography front (impact side) face of panel 24–2 at 1825 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 203.
NASA/TM—2009-213642/REV1 176

Figure E12–3.—Digital photography of back face of panel 24–2 at 1825 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 90° impact angle. Test GRCC 203. 
NASA/TM—2009-213642/REV1 177

Appendix F.—Test Data 

PDL External Tank Foam Impact Testing at 45° Angle on 6- by 6-in. 

Reinforced Carbon-Carbon (RCC) Flat Panels 

Notable Observations From the Appendix F Test Series 

1. The RCC material used in this test series was tested in the as-manufactured condition. 

2. GRCC 230 ARAMIS data was obscured due to coating and substrate debris throughout the run. 

This can be seen in the data dropouts on the deformation contours and trace plots for this test. 

3. Lot 3, (22–36) in the following test series table refers to Lockheed Set 3A, processing Batch 22–

36, made specifically for the Return to Flight Program. 
NASA/TM—2009-213642/REV1 178

 

Appendix F Test Series
NASA/TM—2009-213642/REV1 179

Figure F1–1.—Pulse thermography and ultrasound post impact pretest and posttest images of reinforced carboncarbon 6- by 6-in. flat panels impacted with PDL foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 180

Figure F2–1.—ARAMIS out-of-plane deformation contours across centerline of 6- by 6-in. reinforced carbon-carbon 

flat panels measured at 37.1-s increments undergoing impact with PDL foam cylinders (nominally 1.5 in. in diameter 

by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 181

Figure F3–1.—ARAMIS centerpoint out-of-plane deformation vs. time of 6- by 6-in. reinforced carbon-carbon flat 

panels impacted with PDL foam cylinders (nominally 1.5 in. in diameter by 3 in.) at a 45° angle. 
NASA/TM—2009-213642/REV1 182

Figure F4–1.—ARAMIS color fringe plots depicting maximum deformation prior to material failure of 6- by 6-in.

reinforced carbon-carbon flat panels as they undergo impact with PDL foam cylinders (nominally 1.5 in. in diameter

by 3 in.) at a 45° angle.
NASA/TM—2009-213642/REV1 183

Figure F5–1.—Digital photography of edges and faces of panel 28–1 at 1122 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 217. 
NASA/TM—2009-213642/REV1 184

Figure F5–2.—Digital photography front (impact side) face of panel 28–1 at 1122 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 217. 
NASA/TM—2009-213642/REV1 185

Figure F5–3.—Digital photography of back face of panel 28–1 at 1122 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 217. 
NASA/TM—2009-213642/REV1 186

Figure F6–1.—Digital photography of edges and faces of panel 28–2 at 1266 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 229. 
NASA/TM—2009-213642/REV1 187

Figure F6–2.—Digital photography front (impact side) face of panel 28–2 at 1266 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 229. 
NASA/TM—2009-213642/REV1 188

Figure F6–3.—Digital photography of back face of panel 28–2 at 1266 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 229. 
NASA/TM—2009-213642/REV1 189

Figure F7–1.—Digital photography of edges and faces of panel 29–3 at 1296 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 223. 
NASA/TM—2009-213642/REV1 190

Figure F7–2.—Digital photography front (impact side) face of panel 29–3 at 1296 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 223. 
NASA/TM—2009-213642/REV1 191

Figure F7–3.—Digital photography of back face of panel 29–3 at 1296 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 223. 
NASA/TM—2009-213642/REV1 192

Figure F8–1.—Digital photography of edges and faces of panel 29–1 at 1495 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 220. 
NASA/TM—2009-213642/REV1 193

Figure F8–2.—Digital photography front (impact side) face of panel 29–1 at 1495 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 220. 
NASA/TM—2009-213642/REV1 194

Figure F8–3.—Digital photography of back face of panel 29–1 at 1495 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 220. 
NASA/TM—2009-213642/REV1 195

Figure F9–1.—Digital photography of edges and faces of panel 29–2 at 1723 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 222. 
NASA/TM—2009-213642/REV1 196

Figure F9–2.—Digital photography front (impact side) face of panel 29–2 at 1723 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 222. 
NASA/TM—2009-213642/REV1 197

Figure F9–3.—Digital photography of back face of panel 29–2 at 1723 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 222. 
NASA/TM—2009-213642/REV1 198

Figure F10–1.—Digital photography of edges and faces of panel 36–1 at 1920 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 225. 
NASA/TM—2009-213642/REV1 199

Figure F10–2.—Digital photography front (impact side) face of panel 36–1 at 1920 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 225. 
NASA/TM—2009-213642/REV1 200

Figure F10–3.—Digital photography of back face of panel 36–1 at 1920 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 225. 
NASA/TM—2009-213642/REV1 201

Figure F11–1.—Digital photography of edges and faces of panel 36–3 at 1998 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 230. 
NASA/TM—2009-213642/REV1 202

Figure F11–2.—Digital photography front (impact side) face of panel 36–3 at 1998 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 230. 
NASA/TM—2009-213642/REV1 203

Figure F11–3.—Digital photography of back face of panel 36–3 at 1998 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 230. 
NASA/TM—2009-213642/REV1 204

Figure F12–1.—Digital photography of edges and faces of panel 36–2 at 2105 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 227. 
NASA/TM—2009-213642/REV1 205

Figure F12–2.—Digital photography front (impact side) face of panel 36–2 at 2105 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 227. 
NASA/TM—2009-213642/REV1 206

Figure F12–3.—Digital photography of back face of panel 36–2 at 2105 ft/s with a PDL foam cylinder 

(nominally 1.5 in. in diameter by 3 in.) at a 45° impact angle. Test GRCC 227. 
NASA/TM—2009-213642/REV1 207

References 

1. Gehman, Howard W., et al.: Columbia Accident Investigation Board Report. Vol. 1, 

NASA00194, U.S. Government Printing Office, Washington, DC, Aug. 2003. 

2. LSDYNA Keyword User's Manual. Version 970, Livermore Software Technology 

Corporation, Livermore, CA, 2003. 

3. Impact Testing on Reinforced Carbon-Carbon Flat Panels With Ice for the Space Shuttle 

Return to Flight Program, Final Test Report. NASA/TM—2007-213641, 2007. 

4. Orbiter RCC Flat Panel Impact Testing Plan. 

5. Vision Research. High Speed Digital Imaging Systems, Vision Research, Inc., Wayne, NJ, 

2007. www.visionresearch.com Accessed July 11, 2007. 

6. Photron. Precision High-Speed Imaging Systems, Photron USA, Inc., San Diego, CA, 2007. 

www.photron.com Accessed July 11, 2007. 

7. GOM. Optical Measuring Techniques, GOM mbH. Braunschweig, Germany, 2007. 

www.gom.com Accessed July 11, 2007. 

8. Trilion. Optical Quality Systems. Trilion Quality Systems, West Conshohocken, PA, 2007. 

www.trilion.com. Accessed July 11, 2007. 

9. Schmidt, T.; Tyson, J.; and Galanulis, K.: Full-Field Dynamic Displacement and Strain 

Measurement Using Advanced 3D Image Correlation Photogrammetry: Part I. Exp. Tech., 

vol. 27, no. 3, 2003, pp. 4750. 

10. Lee, Mikyong, et al.: Application of 3D Measurement System With CCD Camera in 

Microelectronics. Advanced Packaging, 2003, pp. 3334. 

11. Koenig, J. : Silicon Carbide Coating Thickness Measurements of Various RCC panels and 

Plates, Final Report to NASA MSFC, George C. Marshall Space Flight Center, Huntsville, 

AL 35812. Contract No. NAS8-00159, December 2005. SRI–ENG–05–23–10359.14. 

12. Koenig, J.: Correlation of RCC Substrate Properties, Final Report to NASA MSFC, George 

C. Marshall Space Flight Center, Huntsville, AL 35812. Contract No. NAS8-00159, April 

2005. SRI–ENG–06–22–10359.14 

13. Curry, Donald M.; Wong, Kenneth A.; and Baccus, Ronald K.: Mechanical Property 

Characterization of Reinforced Carbon-Carbon (RCC) for the Space Shuttle Return to Flight 

Effort. NASA JSC63036, 2006. 

14. Gilat A.; and Goldberg R.K.: Experimental Study of the High Strain Rate Tensile and Shear 

Response of Reinforced Carbon Carbon (RCC) Composites. NASA TM213644 (to be 

published), 2007. 

15. Carney, Kelly, et al.: Material Modeling of Space Shuttle Leading Edge and External Tank 

Materials for Use in the Columbia Accident Investigation. Proceedings of the 8th 

International LSDYNA Users Conference, Dearborn, MI, 2004. 

16. Goldberg, Robert K.; and Carney, Kelly S.: Modeling the Nonlinear, Strain Rate Dependent 

Deformation of Shuttle Leading Edge Materials With Hydrostatic Stress Effects Included. 

Proceedings of the 8th International LSDYNA Users Conference, Dearborn, MI, 2004. 

17. Lyle, Karen H., et al.: Application of Non-Deterministic Methods to Assess Modeling 

Uncertainties for Reinforced Carbon-Carbon Debris Impacts. Proceedings of the 8th 

International LSDYNA Users Conference, Dearborn, MI, 2004. 

18. Fasanella, Edwin L., et al.: Test and Analysis Correlation of Form Impact Onto Space Shuttle 

Wing Leading Edge RCC Panel 8. Proceedings of the 8th International LSDYNA Users 

Conference, Dearborn, MI, 2004. 

19. Gabrys, Jonathan, et al.: The Use of LSDYNA in the Columbia Accident Investigation and 

Return to Flight Activities. Proceedings of the 8th International LSDYNA Users 

Conference, Dearborn, MI, 2004. 
NASA/TM—2009-213642/REV1 208

20. Fasanella, Edwin L.; Boitnott, Richard L.; and Kellas, Sotiris: Dynamic Crush 

Characterization of Ice. NASA/TM2006-214278, 2006. 

21. Fasanella, Edwin L., et al.: Test and Analysis Correlation of Foam Impact Onto Space Shuttle 

Wing Leading Edge RCC Panel 8. NASA/TM2005-213774 (ARLTR3490), 2005. 

22. Pereira, J. Michael, et al.: Forces Generated by High Velocity Impact of Ice on a Rigid 

Structure. NASA/TM2006-214263, 2006. 

23. Pereira, J. Michael; Melis, Matthew E., and Revilock, Duane M.: A Summary of the Space 

Shuttle Columbia Tragedy and the Use of Digital High Speed Photography in the Accident 

Investigation and NASA’s Return-to-Flight Effort. Proceedings of the SPIE 26th International 

Congress on High-Speed Photography and Photonics, vol. 5580, Alexandria, VA, 2004. 

24. Goldberg, Robert K.; Baccus, Ronald K.; and Carney, Kelly S.: Analysis of the Nonlinear 

Deformation Response of Shuttle Leading Edge Materials Including Coating Effects. 

NASA/TM2005-213842, 2005. 

25. Fasanella, Edwin L., et al.: Dynamic Impact Tolerance of Shuttle RCC Leading Edge Panels 

Using LSDYNA; Final Report. AIAA Paper 20053631, 2005. 

26. Schulson, Erland M.; Iliescu, Daniel; and Fortt, Andrew: Characterization of Ice for Returnto-Flight of the Space Shuttle. Part 1Hard Ice. NASA/CR2005-213643/PART1, 2005. 

27. Schulson, Erland M.; and Iliescu, Daniel: Characterization of Ice for Return-to-Flight of the 

Space Shuttle. Part 2Soft Ice. NASA/CR2005-213643/PART2, 2005. 

28. Shazly, Mostafa; Prakash, Vikas; and Lerch, Bradley A.: High-Strain-Rate Compression 

Testing of Ice. NASA/TM2006-213966, 2006. 

29. Carney, Kelly S., et al.: A Phenomenological High Strain Rate Model With Failure for Ice. 

Int. J. Solids Struct., vol. 43, no. 2526, 2006, pp. 78207839. 
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4. TITLE AND SUBTITLE

Impact Testing on Reinforced Carbon-Carbon Flat Panels With BX-265 and PDL-1034 

External Tank Foam for the Space Shuttle Return to Flight Program 

5a. CONTRACT NUMBER

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6. AUTHOR(S)

Melis, Matthew, E.; Revilock, Duane, M.; Pereira, Michael, J.; Lyle, Karen, H. 

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WBS 377816.06.03.02.04 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-15130 

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Subject Categories: 24, 15, and 27 

Available electronically at http://gltrs.grc.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Following the tragedy of the Orbiter Columbia (STS-107) on February 1, 2003, a major effort commenced to develop a better understanding 

of debris impacts and their effect on the space shuttle subsystems. An initiative to develop and validate physics-based computer models to 

predict damage from such impacts was a fundamental component of this effort. To develop the models it was necessary to physically 

characterize reinforced carbon-carbon (RCC) along with ice and foam debris materials, which could shed on ascent and impact the orbiter 

RCC leading edges. The validated models enabled the launch system community to use the impact analysis software LS-DYNA (Livermore 

Software Technology Corp.) to predict damage by potential and actual impact events on the orbiter leading edge and nose cap thermal 

protection systems. Validation of the material models was done through a three-level approach: Level 1-fundamental tests to obtain 

independent static and dynamic constitutive model properties of materials of interest, Level 2-subcomponent impact tests to provide highly 

controlled impact test data for the correlation and validation of the models, and Level 3-full-scale orbiter leading-edge impact tests to 

establish the final level of confidence for the analysis methodology. This report discusses the Level 2 test program conducted in the NASA 

Glenn Research Center (GRC) Ballistic Impact Laboratory with external tank foam impact tests on flat RCC panels, and presents the data 

observed. The Level 2 testing consisted of 54 impact tests in the NASA GRC Ballistic Impact Laboratory on 6- by 6-in. and 6- by 12-in. flat 

plates of RCC and evaluated two types of debris projectiles: BX-265 and PDL-1034 external tank foam. These impact tests helped 

determine the level of damage generated in the RCC flat plates by each projectile and validated the use of the foam and RCC models for use 

in LS-DYNA. 

15. SUBJECT TERMS

Space shuttle; Debris; Impact; Reinforced Carbon Carbon; BX-265; PDL-1043; External tank; Foam; Thermal protection system 

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