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NASA/TP--2000-209760
Oxidation of Reinforced Carbon-Carbon
Subjected to Hypervelocity Impact
Donald M. Curry
NASA Johnson Space CenterHouston, Texas 77058
Vuong T. PhamNASA Johnson Space Center
Houston, Texas 77058
Ignacio NormanBoeing North American, Inc.Houston, Texas 77058
Dennis C. Chao
Boeing North American, Inc'Houston, Texas 77058
March 2000
https://ntrs.nasa.gov/search.jsp?R=20000032537 2018-05-15T19:48:13+00:00Z
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NASA/TP--2000-209760
Oxidation of Reinforced Carbon-Carbon
Subjected to Hypervelocity Impact
Donald M. Curl 3,
NASA Johnson Space CenterHouston, Texas 77058
[htong T. Pham
NASA Johnson Space CenterHouston, Texas 77058
lgnacio NormanBoeing North American, Inc.Houston, Texas 77058
Dennis C. Chao
Boeing North American, lncHouston, Texas 77058
National Aeronautics and
Space Administration
Lyndon B. Johnson Space CenterHouston, Texas 77058
March 2000
Acknowledgments
The authors wish to express their appreciation to Eric Christensen and Frankel Lyons of the JSC
Hypervelocity hnpact Technology Facility and J. D. Milhoan of the NASA JSC Atmospheric
Reentry Materials and Structures Evaluation Facility for participating in and contributing to this
work.
Available from:
NASA Center for AeroSpace Infomlation7121 Standard Drive
Hanover, MD 21076-1320301-621-0390
National Technical Information Service
5285 Port Royal RoadSpringfield, VA 22161
703-605 -6000
This report is also available in electronic fi_rm at http://techreports.larc.nasa.gov/cgi-bin/NTRS
Contents
Page
Abstract ................................................................................................................................. 1
Introduction ........................................................................................................................... 1
RCC Test Articles ................................................................................................................. 2
Hypervelocity hnpact Tests .................................................................................................. 3
Arc Jet Testing .................................................................................................................. 7
Test Facility and Conditions ............................................................................................. 7
Results ................................................................................................................................... 9
Test Data Analysis and Correlation .................................................................................. 14
2500°F Tests ................................................................................................................ 14
2800°F Tests ................................................................................................................ 15
Application to Flight ......................................................................................................... 20
Concluding Remarks ............................................................................................................. 22
References ............................................................................................................................. 22
Appendix A, ARMSEF Orbital Debris Arc Jet Tests ........................................................... 25
Appendix B, Hole Growth Data, Linear Growth Rate Assumption ...................................... 41
Appendix C, Error Analysis .................................................................................................. 49
Tables
Table l.
Table 2.
Table 3.
Table 4.
Table 5.
Summary of Reinforced Carbon-Carbon Hypervelocity Test Results ................... 5
RCC Orbital Debris Arc Jet Test Program Hole Growth Summary ...................... 12
Substrate Mass Loss of Hypervelocity Impacted RCC for the 2500°F Tests ........ 15
An Impacted RCC Hole Growth Rate [Correlated From Test Data] ..................... 17
Correlations for an Impacted RCC Hole Growth ................................................... 20
Figures
Figure 1. Meteoroid/orbital debris relative risks for various components using
omnidirectional flux analysis ......................................................................................... 1
Figure 2. RCC test article dimensions (typical) ................................................................... 2
Figure 3. RCC hole size correlation ..................................................................................... 4
Figure 4. Typical posttest picture of a hypervelocity-impacted 6-in. x 6-in. RCC panel ..... 5
Figure 5. Hypervelocity impact energy vs RCC coating damage area ................................ 5
iii
Contents
(continued)
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Figure
Page
6, NASA model 116 ! front face, pre arc jet test 53 fl-lb impact energy .................. 6
7. NASA model i 152 front face, pre arc jet test 212 ft-lb impact energy ................ 6
8. NASA model 1161 back face, pre arc jet test 53 ft-lb impact energy .................. 6
9. NASA model i 152 back face, pre arc jet test 212 ft-ib impact energy ................ 6
10. Test specimen holder test configuration ................................................................. 8
11. Specimen in chamber during lest ........................................................................ 8
12. Hypervelocity impact damage pre and post arc jet test 2500°F/50 psf model # 1159 10
13. NASA/JSC model #1151 pre and post arc jet exposure 2800°F/100 psf ........... 1 !
14. Video hole growth data for a 2500°F/0.25-in.-dia. impacted hole ..................... 13
15. Video hole growth data for a 2800°F/0.25-in.-dia. impacted hole ..................... 13
16. Conical shape of hole growth at 2800°F ............................................................. 14
17. Front and back hole growth rate comparison ...................................................... 16
18. Comparison of test data and predictions for front hole growth rate (pressure) .. 18
Figure 19. Comparison of test data and predictions for front hole growth rate
(impact energy) ........................................................................................................................ ! 8
Figure 20. Comparison of test data and predictions for back hole growth rate (pressure) .. 19
Figure 21. Comparison of test data and predictions for back hole growth rate
(impact energy) ........................................................................................................................ 19
Figure 22. Temperature histories for the stagnation area of the wing leading edge
panels 4, 7, 9, 14, 16, and 19 for the ISS mission case R 233K, 57-deg inclination
entry trajectory ............................................................................................................... 2 !
Figure 23. Heat flux and pressure history for the stagnation area of the wing leading
edge panels 9 ISS mission case R 233K, 57-deg inclination entry trajectory ............... 21
Figure 24. RCC hole growth history for ISS case R, entry trajectory .................................. 22
iv
RCC
MOD
SiC
HVI
H1T-F
ARMSEF
IS&AG
ISS
Acronyms and Nomenclature
reinforced carbon-carbon
meteoroid/orbital debris
silicon-carbide
hypervelocity impact
Hypervelocity Impact Technology Facility
Atmospheric Reentry Materials and Structures Evaluation Facility
JSC Image Science and Analysis Group
International Space Station
Dh- hole diameter, cm
d - Impacting particle diameter, cm
Pp - density of impacting particle, g/cm 3
V - particle velocity, Km/s
0 - impact angle, deg
P = pressure, lbf/ft 2
Po = atmosphere pressure (2116.2 lbf/ft 2)
IE = impact energy, fl-lbf
D f = front coating hole diameter growth rate, in/sec
D b = back coating hole diameter growth rate, in/sec
X1 = function of pressure defined in table 5
X2 = function of impact energy defined in table 5
X3 = function of pressure defined in table 5
X4 = function of impact energy defined in table 5
Abstract
This paper presents results fl'om arc jet tests conducted at the NASA Johnson Space Center (JSC)
on reinforced carbon-carbon (RCC) samples subjected to hyperveiocity impact. The RCC test
specimens are representative of RCC components used on the Space Shuttle Orbiter. The arc jet
testing established the oxidation characteristics of RCC when hypervelocity projectiles, simulating
meteoroid/orbital debris (MOD), impact the RCC material. In addition to developing correlations
for use in trajectory simulations, we discuss analytical modeling of the increased material oxidation
in the impacted area using measured hole growth data. Entry flight simulations are useful in
assessing the increased Space Shuttle RCC component degradation as a result of impact damage
and the hot gas flow through an enlarging hole into the wing leading-edge cavity.
Introduction
During the early flights of the Space Shuttle, micrometeoroids and orbital debris were not
recognized as a significant hazard to the Orbiters. However, more recent models of the space
environment suggest that MOD pose a significant threat to the orbital vehicle. The RCC wing
leading edge and nose cap are especially vulnerable to this MOD environment and have been
identified as areas of the Orbiter that appear to pose the highest risk for critical failure. Figure 1
presents the relative risk of the RCC as compared to the other Orbiter components.
0.25
0.20',e"
re"D 0.15III
N,.-I
< 0.10
rr"0z 0.05
0.00
Figure 1. Meteoroid/orbital debris relative risks for various components
using omnidirectional flux analysis.
Because RCC is a structural composite and because of the brittle nature of the RCC silicon
carbide (SIC) coating and the requirement to maintain oxidation protection, the effect of impact
damage on thc mission life and structural performance of the RCC components is of concern.
Impact damage may occur from a variety of threats with a wide range of impact speeds.
Meteoroids and human-made orbital debris represent a source of high-velocity impact damage,
while other sources of damage include launch debris, runway debris from landing, and handling
and service damage. Orbital debris can impact orbiting spacecraft at velocities from less than 1
km/sec to over 14 km/sec, and meteoroids from 10 km/sec to over 70 km/sec.
The effects of hypervelocity impact damage on the oxidation perfonnance on the RCC material
has been investigated experimentally. Specifically, arc jet tests have been performed on RCC
samples impacted at hypervelocity to establish oxidation performance and develop corrections
for predicting the increased oxidation during a typical Orbiter entry. Hypervelocity impact
testing produced "complete penetration" or "through-holes" in the test specimens. The arc jet
tests were performed on RCC samples impacted at hypervelocity to establish the rate of
enlargement of the "through-holes" due to oxidation. The hole growth rate was used to develop
corrections for predicting the increased oxidation during a typical Orbiter entry. The typical
entry hole growth rates can be used to calculate the hot gas influx into the wing leading-edge
cavity through the resulting "burn through'" hole and assess structural vehicle component
damage. These hole growth correlations, along with the probabilities of impact damage, are used
to assess MOD impacts on the Orbiter RCC vehicle cornponents for the establishment of flight
rules and initial impact hole acceptability criteria. The objective of this paper is to describe these
tests, and present test results and correlations.
RCC Test Articles
Test articles consisted of SiC-coated RCC samples. Figure 2 shows the nominal dimensions of
the coating and substrate that make up the 0.25-in.-thick RCC test samples. These samples are
representative of RCC components on the Space Shuttle Orbiter, which include the nose cap,
wing leading edge panels, an area between the nose landing gear door and nose cap, and a small
area surrounding the forward attach fitting of the external tank to the Orbiter.
Silicon-Carbide coating _ _' ~.02 in
jr/ SilicCarbon-Carbon composite/
on- arbide coatin __, ,-.02 in
~.25 in
Figure 2. RCC test article dimensions (typical).
The test sample substrate is an all carbon composite laminate fabricated in a multiple pyrolysis
and densification process from a 19-ply phenolic prepreg graphite lay-up. The substrate has a
density of 90 ib/ft _ to 100 Ib/W and is typically 0.17 in. to 0.21 in. thick.
The oxidation-resistant SiC coating is formed in a diffusion reaction process. It is typically
0.02 in. to 0.04 in. thick. Further oxidation resistance is provided by impregnation with
tetraethyl-orthosilicate that, when cured, leaves a silicon dioxide residue throughout the coating
and substrate. Any surface porosity or micro cracks are filled by an application of a surface
sealant (sodium silicate/SiC mixture) in the final step of the fabrication process. More details on
the development and fabrication of the Orbiter RCC applications are given by Curry, et al. ',2
Hypervelocity Impact Tests
Hypervelocity impact (HVI) testing of RCC was conducted at the JSC Hyperveiocity Impact
Technology Facility (HIT-F) from August 27, 1997 to November 21, 19977 The HIT-F
supported the testing by providing 24 tests investigating the effects of hypervelocity impacts of
projectiles simulating MOD striking RCC.
Two types of targets were tested: ten flat plate 6-in. x 6-in. RCC target samples, and a curved
surface sample used for pretesting and a full-scale RCC wing leading panel. All samples had a
thickness of approximately 0.25 in. The primary objective of these tests was to impact each
sample twice, with each impact producing a separate and complete penetration. Impact holes of
•125 in., 0.25 in., or .375 in. diameter were produced. The samples had to be impacted by the H1T-
F such that two circular disks measuring 2.8 in. in diameter with the impact hole in the center could
be cut from each sample. This sample size was required to fit the specirnen holders for arc jet
testing. Data were collected in the HIT-F for correlation of projectile size to hole size. After each
sample was tested in the HIT-F, it was sent to the JSC Arc Jet Facility.
The objective for testing the full-scale wing leading edge panel was to collect data for correlation
of projectile size to the hole size through the top of the panel and the secondary damage that
would occur to the internal and external bottom of the panel. There was an interest in what kind
of secondary damage would occur to the wing span' insulation inside the wing leading edge panel.
These results were reported separately in reference 3.
Previous hypervelocity impact testing conducted at JSC had established the threshold for several
damage modes on coated RCC samples• Equations have been developed relating impact
conditions to RCC damage, such as penetration depth and hole size? ,5 An equation was
developed to predict hole size when complete penetration of RCC occurs. This equation was
used to predict the diameter/velocity required to create the three hole sizes used in this
investigation. Figure 3 provides a comparison between predicted hole size and HVI data.
3.00
2.50
E 2.000
v
?_ 1.50
o 1.00ZE
0.50
Dh = 2.20 d pl, I/_ (V cosO) 1/3- 0.36
Where: DI,- hole diameter, cm
d - Impacting particle diameter, cm
Pv - density' of impacting particle, g/cm 3V -particle velocity, km/s
0 - impact angle, deg
S --Regression El!__
0.00
0.00 0.50 1.00 1.50 2.00
d (V cos8)1/3 pl/3
Figure 3. RCC hole size correlation.
Table 1 presents a summary of the impact test results, and figure 4 shows a representative photo
of the target damage. Figure 4 shows the result of a 5/64-in.-dia. aluminum sphere impacting the
6-in. x 6-in. RCC sample at 6.63 km/sec normal to the surface. Because the SiC coating is
brittle, the coating is damaged in an area around the impact site over 10 times larger than the
projectile diameter. The hole through the specimen exposes the RCC substrate over an area 4
times larger than the projectile diameter. The graph in figure 5 illustrates the approximate front
and back coating damage diameter as a function of impact energy. The coating on the rear side
of the RCC is damaged in a wide area, larger than the front. Internal damage in the RCC
substrate can extend significantly beyond the external damage that is readily detected by visual
observation. Internal damage in the RCC consists of micro cracks and delaminations.
The hole diameters selected for producing the arc jet test specimens were of 0.125, 0.25, and
0.375 in., with kinetic energies of 50, 100, and 200 ft/lb, respectively.
The photographs presented in figures 6 and 7 represent the RCC front face impact damage as a
result of 53 and 212 fl-lb of impact energy. Figures 8 and 9 show a photograph of the back face
damage. These models are presented to illustrate the minimum and maximum energy levels used
in the subject test program and the resulting RCC damage.
4
Table 1. Summary of Reinforced Carbon-Carbon Hypervelocity Test Results
JSC Projectile Projectile Projectile Projectile Projectile
HIT-F Sample Diameter Mass Velocity Velocity KineticEnergy
Shot No. Used* (mg) (ft/s) (kin/s} (j)
A3085 10.1 5/64" 12.39 21,195 6.46
A3086 9. [ 5/64" 11.65 22,508 6.86
A3087 8.1 5/64" l 1.53 21,753 6.63
A3088 7.1 5/64" 11.62 22,639 6.90
A3089 6.1 5/64" 11.70 23,033 7.02
A3093 4.1 1/16" 5.86 22,278 6.79
A3095 5.1 1/16" 5.68 23,098 7.(14
A3096 3.1 1/16" 5.82 23,262 7.09
A3097 10.2 1/16" 5.67 22,803 6.95
A3098 9.2 1/16" 5.97 23,361 7.12
A3099 8.2 0.125 cm 2.8l 23,361 7.12
A3100 7.2 0.125 cm 2.92 23,033 7.02
A3101 6.2 0.125 cm 2.85 23,295 7.10
A3105 4.2 1/16" 5.28 23,197 7.07
A3103 2.1 1/16" 5.71 23,000 7.01
A3108 3.2 5/64" 11.61 23,000 7.01
*Commercially available sizes; projectile malerial AI2017-T4;
258.53
274.12
253.41
276.61
288.29
135.09
140.75
146.28
ProjectileKinetic
Energy
(fl-lb)
190.69
202.19
186.92
204.03
212.64
99.64
103.82
107.9
101.01
Through Hole
Damage
(Hole Size)
9x9 mm
9x9.5 trml
9.5x9 nml
9.5x9.5 nun
9x9.5 mm
6.5x6.5 nml
6.5x7 mm
6.5x7 mm
6.5x6.5 mm
Fronl
Dialneler of
Damage
(in.)
0.867
0.819
0.92
0.932
0.888
0.752
0.776
0.792
0.756
Back
Diameter
of Damage
(in.)
1.035
1.208
1.083
1.159
1.18
0.952
0.962
0.874
0.966136.94
151.32 111.61 6.5x6.5mm 0.801 0.965
71.23 52.54 3.5x3.5 nun 0.673 0.701
71.95 53.07 3.5x3.5 mm 0.649 0.675
71.83 52.98 4x4mm 0.685 0.725
131.96 97.33 0.803 1.142
103.48140.29
285.26
0.777 0.968
0.913 1.069
6.5x6mm
6.5x6mm
hole-9.5x9 mm210.41
projectile shape sphere; all impact angles are 0 deg (normal to the surface).
NASA JSC HIT'F Shot #A3087Target: Shuttle RCC Test Model #11Projectile Diameter: 5/64 in.Velocity: 6.63 km/sImpact Angle: 0
Figure 4. Typical posttest picture of a
hypervelocity-impacted 6-in. x 6-in. RCC panel.
• Front Diameter of Damage Inch
o Back Diameter of Damage Inch
1.4
i
d
1.2
= 0.8-,_
C,
L)L)
0.6
O
0 o0
0
0
Ooe•
o •
t0 50 100 150 200 250
Kinetic Energy fl-lb
Figure 5. Hypervelocity impact
energy vs RCC coating damage area.
Figure6. NASAmodel1161front face,prearcjet test53ft-lb impactenergy.
Figure7. NASAmodel1152front face,prearcjet test212ft-lb impactenergy.
Figure8. NASAmodel1161backface,prearcjet test53ft-lb impactenergy.
Figure9. NASAmodel1152backface,prearcjet test212ft-lb impactenergy.
Arc Jet Testing
The objective of the arc jet testing was to establish the oxidation characteristics of RCC with
hypervelocity-impacted damage. RCC specimens were exposed to constant heating conditions
for specific test points at temperature of 2500°F and 2800°F and pressures of 50 psf to 180 psf.
Test Facility and Conditions
This test program was conducted in JSC's Atmospheric Reentry Materials and Structures
Evaluation Facility (ARMSEF). This facility simulates atmospheric entry conditions by heating
test gas or gas mixtures electrically using a segmented, constricted, 10-megawatt direct current
arc heater. For this test program, the gas mixture was nitrogen and oxygen to simulate air. Once
the air was heated to a high-energy (high-enthalpy) level inside the arc heater, it was then
expanded supersonically through a conical nozzle, which has a 5-in. diameter at the exit plane.
The high-enthalpy and supersonic flow field is formed and then captured by a downstream
supersonic diffuser. Inside the chamber, a hydraulic model insertion system permits the test
articles to be located outside the flow field until the required test conditions have been
established. Model insertion and retraction are normally performed within 1 sec from the
command signal, thereby producing a step pulse heating profile on insertion and an immediate
cessation of oxidation upon retraction. During the test, the chamber static pressure was
maintained below 400 microns of Hg by a four-stage steam ejector system.
The facility data acquisition system obtained facility and test article data at a rate of 10 Hz.
Surface temperature was monitored by a 0.865-micron optical pyrometer aimed in such a way
that the entire field of view was covered by the specimen in the area between the hole and the
edge of the specimen. Once the hole starts to grow, the exposed and burning carbon substrate
will affect the measurements. Because of the potential erroneous surface temperature feedback,
the test runs were made with constant arc heater parameters as identified in the calibration runs.
Two super-VHS video cameras were mounted outside on the test chamber. One was used to
visually monitor and record the test articles' surfaces for transient hole diameter measurements.
The other was used to monitor the flow field and the overall interior of the test chamber.
The arc jet test specimens were fabricated from a 6-in. x 6-in. panel that was first impacted at the
appropriate energy level to produce a desirable hole diameter. The sample was then cut to
approximately 2.8 in. diameter with the impacted hole at the center to make an arc jet test
specimen. Test specimens were cut from an impacted panel using a high-power CO2 laser. One of
the benefits from the laser cutting was the melted SiC coating deposited on the exposed carbon
edge, which acted as another oxidation protection layer. Once a specimen is cut from a panel, four
small retention pin holes about 1/8 in. diameter and about 1/8 in. deep were drilled at 90 deg apart
on its exposed carbon edge. The entire exposed carbon edge was then repaired with a mixture of
Sermabond 487 and 1000 Grit SiC powder that was cured by heating to 300°F in air.
7
A sketch of the test specimen holder used during these tests is shown in figure 10. This figure
shows the 2.0-in.-diameter hole through which gases ingested through the impact holes can pass
without restriction. Ceramic rods support the holder such that a water-cooled conical interface
block may smoothly divert the flow. The outlet slot between the holder and interface block has
an area that is over 15 times that of the largest posttest impact hole area. Since the pressure just
outside the slot is less than 0.6 psf, sonic conditions are maintained at the impact hole with large
margin. In addition the stagnation pressure measurement at the forward end of the conical flow
diverter indicated that sonic conditions existed at the impact holes.
2.0"Dia.
0 °
0 °
Flow •
F Insulator / Holder Back Hole
/7[ _sFl°w_verter
/_ i>O
Interface Sting
Z2,8 " Dia _ 4.0" Dia. Siliconized Carbon Holder
RCC Disk
With Impact Standing Shock
Hole
Figure 10. Test specimen holder testconfiguration.
Figure 11. Specimen in chamber during test.
A view of one of the fixtures during the arc jet test is presented in figure 11. This photograph
shows the normal shock in front of the specimen and the flow of ingested gases out the rear of the
fixture. The arc jet nozzle, plasma flow field glow and the mirrors that are used to record the
specimen face during testing are visible.
Test conditions were established by measuring the temperature of an undamaged specimen with an
optical pyrometer that has repeatedly agreed to within 35°F with imbedded thermocouples in RCC
calibration specimens (during many test programs over the last five years). The test specimens
were exposed to constant heating conditions that were established with the undamaged specimens.
No attempt was made to adjust to target temperatures, since the presence of the impact damage was
expected to change the temperature measurement. Facility operating parameters such as current,
voltage, test gas flow rate, oxygen percentage, chamber pressure, arc column pressure, energy
balance enthalpy, and gas injection manifold pressure were recorded during the calibration test and
monitored for repeatability during subsequent tests. The arc jet test facility parameters used in this
test program were air mass flow rate, m = 0.2 - 0.7 lbm/sec; heater current, i = 530 - 1175 amps;
and bulk enthalpy, h = 2400 - 7200 BTU/lbm.
A videocameramountedoutsidethechamberviewedeachmodel indirectly through a window
with the use of a small mirror mounted to one side of the exit nozzle of the plasma source. Since
the models faced directly into the plasma flow field, they were seen obliquely in the mirror,
making the circular articles and the holder appear elliptically shaped. The test specimen surface
was videotaped before and after each test and during the tests by viewing through a first surface
mirror at an angle of approximately 45 deg. These recordings are stored in the facility and are
available for viewing. Posttest photographs with back lighting to emphasize the outlines of the
penetration holes were taken for each test.
Specimen weights were taken to help assess substrate mass loss. It should be understood that
erosion of the substrate and the SiC repair at the periphery and the erosion of SiC coating near
the impact hole could affect the interpretation of these data.
Results
Fourteen arc jet tests were performed and are summarized in table 2. This table identifies the test
conditions assigned to each specimen, impact energy sustained, the approximate level of visible
pretest damage, arc jet exposure test time and conditions, and post arc jet test dimensions of the
hole growth.
Figures 12A and !2B present the front and back pre arc jet exposure test photographs of
specimen NASA # 1159, typical test specimen. It can be noted of the coating damage, which
resulted from the hypervelocity impact and penetration, that the backside damage has a larger
spalled coating area than the front face coating spallation area. The specimen NASA #1159 was
exposed to 2500°F, 50 psf for 450 sec.
Photographs of this typical posttest specimen (figures 12A-E) show the oxidation and growth of
the impacted frontal area (12C) and of the backside area (I 2D) that occurred during the arc jet
test. The hole growth of the test specimen grew by the oxidation of the exposed carbon-carbon
substrate. Hole growth as a result of coating loss was not expected as there is negligible coating
loss for temperatures between 2800°F - 3000°F for the Orbiter leading edge RCC2
Figure 12E presents an oblique view of NASA specimen #1159, where it can be seen that there
was additional substrate loss that formed a dishing out of the carbon sandwiched between the
SiC front and back coating.
Figure 13A and 13B present the pre arc jet exposure test photograph of specimen NASA #115 I
that can be compared with the post arc jet exposure photograph shown in views 13C and 13D.
For this test condition (2800°F and 100 psf) there was significant front face coating erosion. In
general, the front face of the damaged region grew faster than the backside of the specimen
giving the hole a conical shape that is typical of all the 2800°F tests.
View 12A. Frontside pre arc jet. View 12B. Backside pre arc jet.
View 12C. Frontside post arc jet. View 12D. Backside post arc jet.
Figure 12. Hypervelocity impact damagepre and post arc jet test 2500°F ! 50 psf
model # 1159.
View 12E. Post arc jet test oblique.
lO
View 13A. Frontside. View 13B. Backside.
View 13C. Frontside. View 13D. Backside.
Figure 13. NASA/JSC model #1151 pre and post arc jet exposure 2800°F/100 psf.
11
Table 2. RCC Orbital Debris Arc Jet Test Program Hole Growth Summary
JSC TargetedArc Panel # Hole
Jet /Shot # DiameterID# [in.]
1161 8.2 0.125
1166 7.2 0.125
1165 6.2
1157 5.1
1151 4.1
1158 3.1
1160 10.2
1159 9.2
1145 10.1
1144 9.1
1143 8.1
1142 7.1
1152 6.1
1172 3.2
Test
Conditions
(°F/psf/s)
280011001450
2800/50/450
0.125 2800/180/450
0.250 2500/180/450
0.250 2800/100/450
0,250 2800/50/450
0.250 2800/180/450
0.250 2500/50/450
0.375 2500/50/450
0.375 2800/180/150
0.375 2500/180/450
0.375 2800/100/408
0.375 2800/180/304
0.375 2800_0/450
Impact
energy(fi-lb)
52.54
53.07
52.98
103.82
99.64
107.9
101.01
111.61
190.69
202.19
186.92
204.03
212.64
210.41
Pretest Measurements
AverageHole
Diameter
[in.l
Front
CoatingDiameter
[in.]
Back
CoatingDiameter
[in.1
Posttest Measurements
Front HoleDimensions
[in.l
Back HoleDimensions
[in.]
0.138 0.673 0.701 1.525-1.621 0.992-1.043
0.t38 0.649 0.675 1.003-1.056 0.642-0.659
0.157 0.685 0.725 0.652-0.744 0.390-0.491
0.266 0.776 0.962 0.552-0.653 0.680-0.769
0.256 0.752 0.952 1.890-1.909 1.216-1.233
0.266 0.792 0.874 1.249-1.309 0.853-0.862
0.256 0.756 0.966 0.948-1.136 0.670-0.764
0.256 0.801 0.965 0.476-0.721 0.607-0.778
0.354 0.867 1.035 0.691-0.762 0.799-0.809
0.364 0.819 1.208 1.010-1.064 0.677-0.680
0.364 0.92 1.083 0.741-0.811 0.846-0.885
0.374 0.932 1.159 1.941-1.978 1.240-1.257
0.364 0.888 !.18 1.350-1.503 1.091-1.120
0.364 0.913 1.069 1.423-1.437 0.989-1.002
The JSC Image Science and Analysis Group (IS&AG) used videotapes of each test to make an
assessment of the hole growth (Appendix A)? Dimensional measurements of the diameters of
both the larger front face of each hole (called the outer hole) and the back face of the each hole
(called the inner hole) were made for 14 tests. Appendix A contains photographs for time cuts
that are typical for the videotape images used by the IS&AG group to determine the raw data
inner and outer hole growths. The minimum uncertainty in the diameter measurements is
estimated to be 0.01 in. The minimum uncertainty applies only when the hole edge is well
defined, focus is good, and there is no residual SiC to interfere with the edge identification,
generally true for inner hole edge only. Outer edge uncertainty is .03 to .05 in. because of the
appearance of new oxidizing (burning) rings of SiC. As these new rings glow brightly their edge
blurs the image, giving a false image width. Typical hole diameter measurements as a function
of time that the JSC IS&AG made for the 14 test articles are shown in figures 14 (2500°F) and
15 (2800°F).
12
0.80
0.60
e-._
,..-0.40
E
0.20
0.00
• Outer Hole Diameter
_'lnner Hole Diameter
Surface Temp 2500 °F
Test Surface Pressure 50 psf
Impacted Hole .025 in. Dia.
Projectile Kinetic Energy I 12 ft-lb
NASA]JSC Model # 1159
0 100 200 300 400 500
Time, seconds
Figure 14. Video hole growth data for a2500°F/0.25-in.-dia. impacted hole.
e_Outer Hole Diameter
InnerHote Diameter2.0 l:7_/'_p_S_'¥ ........
Test Press I_) psf
Impacted Hole Size
.025 inch Dia
1.5 Projectile Kinetic •
._ I Energy I00 F|-lb"_ M<xlel # 1151
_'1.0¢D
E
0.5
0.0
0 100 200 300 400 500
Time, seconds
Figure 15. Video hole growth data for a
2800°F/O.25-in.-dia. impacted hole.
Hole growths for the 2500°F tests showed only loss of the RCC substrate and no visible SiC
coating loss; thus the 2500°F hole growth was circular on the exposed substrate with virtually no
difference in the outer and inner diameter.
Hole growths of the 2800°F test specimens developed a conical shape when viewed in cross
section. The widest hole cone was formed by the 100-psf test cases, and the narrowest cone
slopes were formed by the test cases tested at 180 psf. The conical hole shape phenomena is
illustrated in figure 16. The test data are grouped to show the conical shape (outer ring diameter
minus inner ring diameter).
The hole growth data derived from the assessmentsof the videotapes of arc jet testing, the test
measurements made at the Hypervelocity Lab, and the measurements performed at the ARMSEF
were used for final hole growth prediction methods.
13
NeD fi ib Impact Energy Hole Final Growth (Front - Back)
MOO fl lb_t Energy Hole Final Growth (Front - Back)
N200 fl ib Impact Energy Hole Final Growth (Front - Back)
0.8
09• 0.7
t-O
0.6-_-'- _
--05_ _ °
"6 ._ 0.4
o 0.3
o 0.1
, 0
5O lO0
Test Pressure,
Figure 16. Conical shape of hole growth at 2800°F.
Test Data Analysis and Correlation
2500°F Tests
Test data from the 2500o1 = series indicated that early in the test period the front hole growth rate
was much higher than the back hole. Approximately 200 sec into the test the back hole diameter
becomes equal to the front hole diameter and the hole growth stopped. The results of comparisons
between the pretest and posttest tesl specimens indicated the final hole size is either equal or less
than the initial size of the front coating damage area (the coating was either broken into pieces or
spalled) caused by the hypervelocity impact test. There is no coating oxidation at 2500°F; thus
the hole growth was limited to the size of the impact damaged area (i.e., exposed bare substrate,
figure 12).
14
Comparisons between these test data and the RCC substrate mass loss correlation (developed
from the test data of non-impacted and non-coated RCC 7) predictions show that the actual mass
loss rate is around two to four times higher than predicted. However, the mass loss rate has no
significant relationship with the level of the impact energy except that the final size of the hole is
proportional to the level of the impact energy. The higher impact energy causes more coating
damage, coating removal, and substrate exposure. The results of the data analysis for the 2500°F
test data group are shown in table 3.
Table 3. Substrate Mass Loss of Hypervelocity Impacted RCC for the 2500°F Tests
Test Pressure Impact Damage Hole Damage Hole *Ratio Final RCC Substrate Mass
Specimen psf Energy Initial Final Diameter Hole Loss
JSC ID ft-lb Diameter PTM Growth Post Test Measurement
(in) (in) lbm/sec-sq ft
1145 50 190 0.354 0.656 1.85 3.605E-03
1143 180 186 0.364 0.672 1.85 3.635E-03
1159 50 Ill 0.256 0.594 2.32 4.050E-03
!157 180 103 0.266 0.531 2.00 3.466E-03
*Ratio = final hole diameter / initial hole diameter
Test Pressure Impact Front Hole RCC RCC MLRp-rM MLRv-m
Specimen psf Energy Diameter Substrate Substrate /MLRpRD /MLRpRDJSC 1D ft-lb VTD Loss Rate Loss Rate
inches VTD PRD
lbm/sec-sq ft Ibm/sec-sq ft
1145 50 190 0.649 3.553E-03 8.784E-04 4.1 4.05
1143 180 186 0.600 3.058E-03 1.667E-03 2.18 1.83
1159 50 Iil 0.605 4.123E-03 8.784E-04 4.61 4.69
1157 180 103 0.529 3.449E-03 1.667E-03 2.08 2.07
MLR = RCC Substrate Mass Loss RatePTM = Posttest Measurement
PRD = Correlation Prediction (w/o an impacted hole)VTD = Video Test Data at 200 sec after Test
2800°F Tests
Test data from the 2800°F series showed that the characteristics of the holes are quite different
from the 2500°F series. The holes from the 2500°F test are irregular in shape but keep a shape
similar to one resulting from the impact test (figure 12). The holes from 2800°F are large, round
shapes with a significant difference in size between the front hole and back hole diameters
(figure 13).
Inspection of the test specimen after the impact test and before the arc jet thermal test indicates that
the through-hole diameter from the front to the back is about the same size with circumferential
damage/delamination of the hole wall surface. Upon insertion of the test specimen into the arc jet
stream, the diameter of the front face surface hole increases much more than the back surface.
This phenomenon can be seen from the video test data illustrated in Appendix A.
15
The2800°Ftestseriesincludedthree different levels of impact energy and three different
pressure environments. Based on these test data, correlations of an impacted RCC hole growth
as a function of the level of impact energy and pressure were developed. The correlations
assumed a constant hole growth rate with respect to time. Plots of the test data and the constant
hole growth rate liqe are shown in Appendix B. Since the front and back hole growth rates are
different, the development of the correlations for the front and back hole growths were
performed separately. A typical plot of the test data and the assumed constant hole growth rate is
shown in figure 17.
Development of an impacted RCC hole growth correlation proceeds in two phases. In phase
one, 18 constants of the hole diameter growth rates--9 for the front end of the hole and 9 for the
back of the hole--were established based on the test data of 9 different test conditions. These
growth rates are shown in table 4. With these growth rates, two correlations were developed:
one for the growth rate as a function of the level of impact energy and the other for the growth
rate as a function of pressure ratio (the test pressure divided by atmospheric pressure). These
two correlations were then used as the second-generation independent variables and a regression
analysis was performed. The final correlations are shown in table 5. As the correlations
indicate, the effect of the pressure environment on the hole growth rate is a second-order
polynomial curve and the effect of the impact energy on the hole growth rate is a power curve.
1.4
1.2
2: l.o
_0.8
0.6
--_0.4O
_Z
0.2
Figure 17.
0.0
Model # 1158
2800 deg F, 50 psf,
100 fl-lb
• Back ttole Test Data
O Front Hole Test Data
_Back Hole Growth Line
_Front Hole Growth Line
0 200 400
Time, sec
Front and back hole growth rate comparison.
Figures 18 to 21 show a comparison of the test data and predictions, using the correlations, for both
front and back coating hole growth rates, and illustrate the effects of impact energy and pressure on
hole growth. As expected with increasing impact energy, the hole growth rate increases as a result
of coating and substrate damage. As the arc jet test pressure was increased from 50 to 180 lb/fi 2,
hole growth increased as the pressure change from 50 to 100 lb/ft 2, but then decreased at a pressure
16
of 180 lb/ft 2. This reduction in hole growth is a result of a passive glassy layer (SiO2) being
formed on the carbon surface, since the carbon substrate has been impregnated with tetraethyl-
orthosilicate as part of the RCC oxidation protection system.
Table 4. An Impacted RCC Hole Growth Rate [Correlated From Test Data]
(A) Front Coating Hole
Impact
Energyfl-lb
50
50
50
I00
100
100
200
200
200
Pressure
psf
50
100
180
50
100
180
50
100
180
TemperatureoF
2800
2800
2800
2800
2800
2800
2800
2800
2800
Hole Diameter
Growth Rate
in./sec
1.68-03
3.04E-03
1.20E-03
1.98E-03
3.24E-03
1.80E-03
2.01E-03
3.20E-03
i .82E-03
(B) Back Coating Hole
Impact
Energyft-lb
50
50
50
100
100
100
200
200
200
Pressure
psf
50
100
180
50
100
180
50
100
180
TemperatureoF
2800
2800
2800
2800
2800
2800
2800
2800
2800
Hole DiameterGrowth Rate
in./sec
1.36E-03
1.96E-03
1.18E-03
1.44E-03
2.18E-03
1.73E-03
1.64E-03
2.36E-03
1.93E-03
17
Figure 18.
Figure 19.
[]
0.0035
rr 0.0030t-
O 0.0025
_ 0.0020E. __
o 0.0015
o"1-
0.0010
0
-- Front Hole Prediction, IE = 50 ft-lb
- - - Front Hole Prediction, IE = 100 ft-lb
Front Hole Prediction, IE = 200 ft-lb
Front Hole Test Data, IE = 50 ft-lb
Front Hole Test Data, IE = 100 ft-lb
Front Hole Test Data, IE = 200 ft-lb
.'rY \\',
#" _"
50 100 150 200
Pressure, psf
Comparison oftest data and predictions forfront hole growth rate
(pressure).
Front Hole Prediction P=50 psf
- - - Front Hole Prediction
...... Front Hole Prediction
• Front Hole Test Data
Front Hole Test Data
[] Front Hole Test Data
0.0035_Y
rr 0.0030¢-
_ 0.0025
o 0.0020E._
O 0.0015
oI
P=100 psf
P=180 psf
P=50 psf
P=100 psf
P=t80 psf
,..,.1.- w
[email protected]"_" @
[]0.0010 u u , u
0 50 100 150 200 250
Impact Energy (IE), ft-lb
Comparison of test data and predictions for front hole growth rate
(impact energy).
18
Figure20.
--Back Hole Prediction, IE = 50 ft-lb
- - - Back Hole Prediction, IE = 100 ft-lb
...... Back Hole Prediction, IE = 200 ft-lb
• Back Hole Test Data, IE = 50 ft-lb
i Back Hole Test Data, IE = 100 ft-lb
[] Back Hole Test Data, IE ---200 ft-lb
0.0025
t"¢ ,* _jl__ _ _, ",
r- I' ..J \ "
o.oo o ..',I,,,--,,"".• %,0
"/ \\i
E.-.__ 0.0015E3
o •-P
0.0010 I i I
0 50 100 150 200
Pressure, psf
Comparison of test data and predictions for back hole growth rate
(pressure).
Q=¢-
E
o
¢-
o
--Back Hole Prediction, P=50 psf
- - - Back Hole Prediction, P=100 psf
...... Back Hole Prediction, P=180 psf
• Back Hole Test Data, P= 50 psf
Back Hole Test Data, P=100 psf
[] Back Hole Test Data, P=180 psf
0.0025
0.0020
0.0015
J
. !,. w'''°°
13
0.0010 i I I I
0 50 100 150 200 250
Impact Energy (IE), ft-lb
Figure 21. Comparison of test data and predictions for back hole growth rate
(impact energy).
19
Table 5. Correlations for an Impacted RCC Hole Growth
Correlation for the Front Coating Hole Growth Rate:
D f= -0.00168398 + X I + 0.794886 (X2)
Where D i= Front Coating Hole Diameter Growth Rate, in/sec
XI =-1.54763 (t'), + 0.163622 (I')-1.11310E-3Po Po
X2 = 1.02886E-3 (IE) 0.Is6
P = Pressure, psf
Po = latm (2116.2 psf)IE = Impact Energy, fl-lb
Correlation for the Back Coating Hole Growth Rate:
D b = -0.00164586 + (X3) + 0.953631 (X4)
Where D h = Back Coating Hole Diameter Growth Rate, in/sec
X3 =-7.05562E-! (--_-P)" + 7.88547E-2 (-_,) + !.22376E-5Po
X4 = 6.50434E-4 (I12) 0.210._69
Application to Flight
The correlations developed from this test program only cover two temperature points, 2500°F
and 2800°F. To perform a flight simulation requires a range of temperatures, from 500°F to
3000°F. Therefore, the following assumptions were made:
1. The mass loss correlation developed from the current 2500°F test data for substrate mass loss
will hold for the temperature range of 500°F to less than 2800°F (this correlation is in a form
of factors of the RCC substrate mass loss correlations developed from the test data of
nonimpacted and noncoated RCC).
2. There is no coating mass loss in the temperature range of 500°F to less than 2800°F (this was
concluded from a previous RCC over-temperature test programT). Therefore, during the
flight while the temperature is below 2800°F, the hole will only grow up to the size of the
impacted coating damage area.
3. At the temperature range of 2800°F to 3000°F, the coating mass loss rate remains a constant
with respect to the temperature (again, this was concluded from a previous RCC over-
temperature test program_), and the mass loss/hole growth rate correlations developed from
the current 2800°F test data will hold throughout this temperature range. Based on these
assumptions, plus the new correlations, a flight simulation program was developed for the
flight application.
20
Assessmentof anentryenvironmentusingthemethodologydevelopedbeginswith a thermal -
entry analysis for the entire RCC leading edge subsystem. A typical International Space Station
(ISS) Shuttle mission trajectory case R, 233K forward cg with 57 deg inclination was chosen.
Figure 22 presents the temperature histories that represent the entire wing leading edge. For this
assessment of the RCC wing leading edge areas where temperatures are predicted to be below
2800°F, the hole growth is not expected to expand beyond the coating damage surrounding the
impacted hole. An examination of RCC temperatures reveal that only panels 8, 9, and 10
stagnation area would experience entry temperatures above 2800°F where major coating erosion
would be predicted. The entry environment for panel 9 stagnation area is presented in figure 23.
The assessment of the impact damage area of the wing leading edge panel requires the prediction
of the expected hole growth history. This hole growth history is used to evaluate the
consequences of the potential damage effects to the Orbiter structure. For this example a 0.25-
in.-dia, hole resulting from MOD impact was assumed. Figure 24 presents the predicted hole
growth history for the ISS Case R trajectory. The back hole growth rate is used to calculate the
hot gas influx into the wing leading edge cavity for assessing wing spar insulation and structural
damage. Final hole diameter of 1.76 and 1.415 in., respectively, for the front and back surfaces
is predicted.
3OO0
2500
500
[] P(4) 2033 deg FO P(7) 2552 degF
_P(9) 2903 deg F
A P(14) 2319degF
• P(16) 2328 deg F
4- P(19) 2330degF
0
0 500 1000 1500 2000
Entry Time, Seconds
•""-"--- Heat Flux
- - Surface Pressure, psf
40 -- -- 1600
o 1400
<_ 30 1200
g 1ooo-_ 20 800 _"
600IT ___,, 10 400
"r 200
0 0
0 500 1000 1500 2000
Entry Time, seconds
Figure 22. Temperature histories for thestagnation area of the wing leading edge
panels 4, 7, 9, 14, 16, and 19 for the ISSmission case R 233K, 57-deg inclination
entry trajectory.
Figure 23. Heat flux and pressure history forthe stagnation area of the wing leading edge
panels 9 ISS mission case R 233K, 57-deginclination entry trajectory.
21
_--"'--Surfac_e Temperature, °F . Back Coating Hole Diameter--- --front Coating Ho_euiameter
3000 .... 30
2800 °F ---------_
2500 _" Coating & 2.5
L
Substrate
2000 ,, Oxidation _ 20
= 1500 i 1.5
. I Oxidation.. E
° !/jJl ooo 1.o i5
500 05
0 ' 00
0 500 1000 1500 2000Entry Time, seconds
Figure 24. RCC hole growth history for ISS case R, entry trajectory.
Concluding Remarks
This report presented results from arc jet tests conducted on hypervelocity impacted RCC test
specimens and hole growth rates measured with video techniques and posttest examinations.
Included were correlations developed for both front face and back face mass loss as a function of
hypervelocity impact energy, pressures, and two test temperatures, 2500°F and 2800°F. An
assessment of a typical entry with an assumed impact damage to estimate hole growth was
included. These correlations, along with the probabilities of impact damage, can be used to
assess MOD impacts on the Orbiter RCC and hot gas flux though an enlarging hole with
resultant impingement on the internal insulation components.
References
I. Curry, D. M., Scott, H. C., and Webster, C. N., Material Characteristics of Space Shuttle Reinforced
Carbon-Carbon, Proceedings of 24 'h National SAMPE Symposium, Volume 24, Book 2, 1979, pp.
1524-1531.
2. Dotts, R. L., Curry, D. M., and Tillian, D. J., Orbiter Thermal Protection System, Space Shuttle
Technical Conference, NASA CP-2342, Part 2, 1983, pp. 1062-1081.
3. Lyons, Frankel, Hypervelocity Impact Testing of Reinforced Carbon-Carbon (RCC), NASA 3SC
28398, May 1998 (internal publication; contact author for documentation).
22
, Christiansen, E. L., Curry, D. M., Kerr, J. H., Cykowski, E., and Crews, J. L., Evaluation of the lmpact
ResistaJlce qfRe#!forced Carbon-Carbon, Proceedings of the Ninth International Conference of
Composite Materials, Madrid_ Spain, 1993.
5. Christiansen, E. L. and Friesen, L. Penetration Equations for Thermal Protection Materials,
International Journel of Impact Engineering, Vol. 20, 1997, pp. 153-t64.
6. Gaunce, M., Assessment of Orbital Debris Test Hole Growth, NASA Memo SN3-98-009, March 24,
1998.
7. Williams, S. D., Curry, Donald M., Chao, Dennis C., and Pham, Vuong T., Ablation Analysis of the
Shuttle Orbiter Oxidation Protected Reitforced Carhou-Carhon AIAA Journal of Thermophysics and
Heat Transfer, Volume 9, No.3, July-September, 1995, pp. 478-485.
23
Appendix A
ARMSEF Orbital Debris Arc Jet Tests
Hole Growth Measurements
JSC Image Science and Analysis Group
25
Figure A-I. Photo of Model 1142.
2.00
1.80
1.60
1.40
1,20
l.00
0.80
_" 0.60
0.40
0.20
0.00
ARMSEF RCC Orbital Debris
Model 1142
Hole Diameters vs. Time
ilL-iN
mill II
m m
mmn imm!
mmm m ee _mmmm =e*
mmmmm .+e °+"
iN _eOO *.O
"=mmm _000$ .
_'ee'ee m _:t_r_l_post-Rasma Measure
---- OuterHolePo_-RasmaMeasure
50 100 150 200 250 300 350 400 450
Time (sec.)
Figure A-2. Graph of hole growth for Model 1142.
500
26
Figure A-3. Photo of Model 1143.
200
RCC Orbital Debris
Model 1143
Hole Diameters vs. lime
1.80
1.60
,- 1.40o.=_
1.20
1.00E
O.8O
o 0,60
0,40
0,20
0.00
• Inner Hole
• Outer HOe
-- Inner Hale Post-Plasma IV_qsure
--- - Outer Hde Post-Plasma Measure
.... • .. • . ... . .=.=r--
..,.l,nllll,,l,llll,llll**l,=lIl*l*,,,,---
1_ 1_ 2_ 2_ 300 3_ 4_ 4_
_me (sec.)
Figure A-4. Graph of hole growth for Model 1143.
50o
27
Figure A-5. Photo of Model 1144.
2.00
ARMSEF RCC Orbital Debris
Model 1144
Hole Diameters vs Time
1.80
1.60
1.40i'mU
,_E 1.20
_1.00 ira--"
0.80 =tin
=oool v." ...-
0.40 _1_ _O0,20
0.00
Inner Hole
II Ou[er Hole
_lnner Hole Post-Plasma Measure
-- -- -- Outer Hole Pos[-Plasma Measure
0 50 100 150 200 250 300 350 400 450
Time (sec.)
Figure A-6. Graph of hole growth for Model 1144.
5OO
28
Figure A-7. Photo of Model 1145.
2.00
ARMSEFRCC Orbital DebrisModel 1145
Hole Diameters vs. Time
1.80
1.60
1.40
g 1.20
1.00
;_ 0.80
0.60
040
0.20
000
4, Inner Hole
• Outer Hole
Inner Hole Post-Plasma Measuro
-- - -- Outer Hole Post-Plasma Measure
,,lUUlnannamnUnunu• nulu.! llllllliUUillllllllliliilul"-"--
Iil_Jl_ O'OOv
50 100 150 200 250 300 350 400 450
Time (sec.)
Figure A-8. Graph of hole growth for Model 1145.
5OO
29
Figure A-9. Photo of Model 1151.
200
1.80
1,60
ARMSEFRCC Orbital Debris
Model 1151
Hole Diameters vs. Time
mmmmm mmmm
mmmmmm,-. 1,40
1.00 illll II|l|= _****
-r- 0.60 : 00_0 0 00_'''• 0'¢' ..................
, =OO • • Inner Ho!eO4O lilm"_* °" • outer Hole
Z._@" Inner Hole Post-Rasma Measure
0.20 J" Outer Hole Post-Plasma Measure
/0.00
0 50 100 150 200 250 300 350 400 450
_me(sec.)
Figure A-10. Graph of hole growth for Model 1151.
500
3O
Figure A- 11. Photo of Model 1152.
2,00
1.80
1.60
,_. 1.40
"_ 1.20
_) 1.00
0.80
0.601
0.40
0,20
000
ARIVISEFRCC Orbital Debris
Model 1152
Hole Diameters vs. Time
ill Illl
•
iiiii nl _O, t_
,.111111_I i i IIII O(,I $$
m 0_000v
II_ , 4''" 4_ . Inner-H_e-
41N_ O • Outer Hole
P4ql_ _ Inner Hole Post-Rasma Measure
- -- -- Outer Hde Post-Plasma Measure
50 100 150 200 250 300 350 400 450
_me (_c.)
Figure A-12. Graph of hole growth for Model 1152.
5O0
31
FigureA-13. Photo of Model 1157.
2,00
ARMSEF RCC Orbital Debris
Model 1157
Hole Diameters vs. Time
1.80
1.60
A 1.40
(3
c= 1.20
1,00
_5 080
OZ
0.60
0.40
0.20
0.00
• Inner Hole
• Outer Hole
--Inner Hole Post-Plasma Measure
-- -- -- Outer Hole Post-Plasma Measure
0 50 100 150 200 250 300 350 400 450
Time (sec.)
Figure A-14. Graph of hole growth for Model 1157.
5OO
32
Figure A-15. Photo of Model 1158.
] .40
AI_Bff RCC Orl_tal Debris
Modd 1158
Hole Diameters vs. Time
1.20
tOO
i-
t-
_" 0.80
$Ei_ 0.60
0.40
0.20
0.00
i _u_w
lillll Ill
nlllllm ll
Illll II
Illl A 'Ill IIOOIf-----ml .=i
i III- iiOilliO i
j,.," o.O°°
_oo,lo* * • out_ _eIrr_rl-lOePast-RasrnaMeasure
- - - C_er Fide Pcst-Plasma Measure
1_ 1_ 2_ 2_ 300 3_ 4_ 4_
_me(sec.)
Figure A-16. Graph of hole growth for Model 1158.
33
FigureA-17.Photoof Model 1159.
0.80
0.70
0,60
O50
0.40
030
0,20
0.10
0.00
ARMSEF RCC Orbital Debris
Model 1159
Hole Diameters vs. ]_me
mmmmmmmmmmml||l|l|mi|ii I • "v
i u O_e 04
O I_m_ummmmmummmumi ***
[] ,•
F ---
• _1, Inner Hole•e
e_ • Out_ HOe[ _lnn_ Hole Post-Plasma Moasure,
- - - Outer Hde Po_-Piasma Measure
1_ 1_ 2_ 2_ 3_ 3_ 4_ 4_
_me (sec.)
Figure A-18. Graph of hole growth for Model 1159.
500
34
Figure A-19. Photo of Model 1160.
ARMSEF RCC Orbital Debris
Model 1160
Hole Diameters vs. Time
1 2G
l_OG
mmmmmm L ....• l
mm u )
I_@400 @ @0@0@@
• m@
mmmmm mmmmmmmmml m)*
4,4'
0
• Inner Hole
• Outer Ho_e
-- -- -- Outer Hole Post-Plasma Measure
50 100 150 200 250 300 350 400 450 500
lime (sec.)
Figure A-20. Graph of hole growth for Model 1160.
35
Figure A-21. Photo of Model 1161.
1.80
ARNE_ RCC Orbital Debris
Model 1161
Hole Diameters vs. lime
1.60
1.40
0.40
50 100 150 200 2,50 300 350 400 450
"nine (sec.)
Figure A-22. Graph of hole growth for Model 1161.
500
36
FigureA-23.Photoof Model1165.
200
180
1.60
1.40I,,-U
'- 1.20
100
_5 O80
" 0.60
0.400.20
0.00
RCC Orbital Del3fis
Model 1165
Hole DialT_ers v& ]]me
• Inner Hole
• Outer Hole
_lnner 14o_ePc6bPla_qa Measure
- - -- Outer Hole Post-Plasma Meastxe
,,,, ,, ,mmmmmmmm nn unqll|l,
um ,%%,I,****
OOOt4_ 0_
50 100 150 200 250 300 350 400 450
Time(._c.)
Figure A-24. Graph of hole growth for Model 1165.
5OO
37
FigureA-25.PhotoofModel1166.
200
ARMSEF RCC Orbital Debris
Model 1166
Hole Diameters vs. lime
180
1.60
1.40
120
1.00
0,80
0.60
0,40
0.20
0.00
• Inner Hole
• Outer Holo
Inner Hole Post-Plasma Measuro
- - - Outer Hole Post-Plasma Measure
ullnUlllllllllllllll[ -__
llmnUli nl llli$t
mmm immm m _+ll lllll
_m m" _i$,i*$1*i"
ll_@$111@Iolli$@$
0 50 100 t 50 200 250 300 350 400 450 500
Figure A-26. Graph of hole growth for Model 1166.
38
200
1.80
1160
_1,40
1.20
1.00
_ 0.80
"_ 0.60
0.40
0.20
0.00
0
Figure A-27. Photo of Model 1172.
_ RCC Offal Debris
Model 1172
Hole Diameters vs. ]]me
_, Inner Hole
• O.Jt_ Hde
Inner Fide Post-Plasma l'_,::a_e
--- Cute-t-ldePo_-Iqasrr_e _ _l ItI I
illllll BIll
. iililnlllllllllllllI OllOtlOOOOO4r'-_
Jiiq..i IllI 14,,_II*14'4'
T .,==OO_OOOtOOOO
i_i,_ **9'vv
50 100 150 200 250 300 350 400 450
"lime(s__...)
Figure A-28. Graph of hole growth for Model 1172.
50O
39
Appendix B
Hole Growth Data
Linear Growth Rate Assumption
41
09
08
07
-_ 0.6_E
05-$
E 04
01
00
Test Specimen: JSC1159Test Condition: 2500 °F, 50 psf, & 11161 ft-lb
tt
..:......-- -
_°,oO4 °'
al_n,,otD_• Back Hole Test Data
_ebelP . Front Hole Test Data- - - Back Hole Growth Line
Front Hole Growth Line
n l i l | I u | I I I I I I J | | I " n ,
100 200 300 400
Time, secFigure B-1 Test Data of An Impacted RCC Hole Growth
i |
5OO
09
08
07
Test Specimen: JSC1157Test Condition: 2500 °F, 180 psf, & 103.82 ft-lb
06
b'* ...Z.....-..--.-E 0"4_
.o;;-;.-0.3 Or, ,, " "
0
0.1
0.0
• Back Hole Test Data
Front Hole Test Data
- - - Back Hole Growth LineFront Hole Growth Line
50 100 150 200 250 300 350 400 450 500
Time, sec
Figure B-2 Test Data of An Impacted RCC Hole Growth
42
0.9
0.8
Test Specimen: JSC1145Test Condition: 2500 °F, 50 psf, & 190.69 ft-lb
0.7
(,3E
0.5
(1)E
0.4
a
-5 0.3"1-
0.2
0.1
0.0
$$e eolso888oSlOlo$15°
• Back Hole Test Data• Front Hole Test Data
- - - Back Hole Growth LineFront Hole Growth Line
! la l l | | " * " J " ' = • m | u | s | a m I • " • • • ' I = , , | I | n | • " • • • " • ....
0 50 100 150 200 250 300 350 400 450 500
Time, secFigure B-3 Test Data of An Impacted RCC Hole Growth
0.9
0.8
0.7
Test Specimen: JSC1143Test Condition: 2500 °F, 180 psf, &
0.1
• Back Hole Test DataFront Hole Test Data
- - -, Back Hole Growth LineFront Hole Growth Line
0.0 • i • n I • I I I i I I l I i l t • • i • • , • n , • • i i • i • , l , i , • i n n n • n ....
50 100 150 200 250 300 350 400 450 500
Time, secFigure B-4 Test Data of An Impacted RCC Hole Growth
43
2.0
Test Specimen: JSC1166/53.07 ft-lb Impact Energy
1.8 Test Conditions: 2800 dig F/50 psf
1.6• Back Hole Test Data
1.4 4 Front Hole Test Data
- - - Back Hole Growth line
_ Front Hole Growth Line1,2
I4>OO4• 'q_'q__ r o°°°°' -/- '_5 0.8
0.6 _•_'_ _ --" "._''-- I
0.2
0.0
0 50 100 150 200 250 300 350 400 450
Time, sec
Figure B-5 Test Data of An Impacted RCC Hole Growth
2.o
I Test Specimen: JSC1161/52.54 ft-lb Impact Energy1.8 Test Conditions: 2800 deg F/100 psf
Back Hole Test Datao
'°f • Front Hole Test Data
1.4 [ .... Back Hole Growth Line _,,,_•9 I
_ I----FrontHoleGrowthLine -o_ • " 1
_ 1.2 ....
g
I ;i_.......:.,.,,..,,T_ 0.8
0.6 I_ • __,Jl_r .- :;,;,;,"- I
0.4
0.2
0.0
0 50 100 150 200 250 300 350 400 450
Time, sec
Figure B-6 Test Data of An Impacted RCC Hole Growth
44
2.0TestSpecimen:JSC1165/52.98ft-lbImpactEnergy
1.8 TestConditions:2800degF/180psf
1.6 • BackHoleTestData• FrontHoleTestData
1.4 .... BackHoleGrowthLine. m. FrontHoleGrowthLine
_ 1.2 ......
g_ 1.0
o 0.8 _ ,_,,,. ,o,tt..,.,,.,__ . • _;_._--oo ....... """"" ""_:-:-.-H-.'_-_- -o-- °" °°" I
-. • ""_--_-z_r" -'- - "'''" I0.4
0.2
0.0
0 50 100 150 200 250 300 350 400 450
Time, sec
Figure B-7 Test Data of An Impacted RCC Hole Growth
2.0
Test Specimen: JSC1158/107.9 ff-lb Impact Energy
1.8 Test Conditions: 2800 deg F/50 psf
1.6 • Back Hole Test Data
• Front Hole Test Data
_, 1.4 .... Back Hole Growth Line
- -- - Front Hole Growth Line
- _g 1.2 ...... .
1.0 .. • • •_._"_ I
• •9"v t
0.6 * ,t.__ _ I,..'t "14"°'°'-° I
0.4
0.2
0.0
0 50 100 150 200 250 300 350 400 450
lime, sec
Figure B-8 Test Data of An Impacted RCC Hole Growth
45
2.0 f Tes t Specimen: JSCl151/99.64 ft-lb Impact Energy /
18 _ Test Conditions: 2800 deg F/100 psf . _'_t'
f I1.6 _ I . " Front Hole Test Data / _'_ I
[- I ' " " Back Hole Growth Line / _ I
-_ 14 I --Fr°ntH°leGr°wthLine ]_,&_12 ..t. "e'e°t
__ 0.8
0.6
0.4
0.2
0.0
0 50 100 150 200 250 300 350 400 450
Time, sec
Figure B-9 Test Data of An Impacted RCC Hole Growth
t-Ot'-
.w
Eom
_3
O"1-
20
18
16
14
12
10
0.8
06
04
02
00
Test Specimen: JSC1160/101 0 ft-lb Impact EnergyTest Conditions: 2800 deg F/180 psf
• Back Hole Test DataFront Hole Test Data
•..- - Back Hole Growth Line--Front Hole Growth Line
0 50 100 150 200 250 300 350
Time, sec
Figure B-10 Test Data of An Impacted RCC Hole Growth
400 450
46
2.0TestSpecimen:JSC1172/210.4ft-lbImpactEnergy
1.8 TestConditions:2800degF/50psf
1.6 • Back Hole Test DataFront Hole Test Data
- - - Back Hole Growth Line1.4 --------Front Hole Growth Line
.E 1.2
E 1.0 +_ ++ _o"•'° d_o'•'o 1m 0.8 d__o_
_. +'+ . - II o'°"°Tp_ ,.''+*j, b._ .,;ooO0.6, _ o..o,o o ato"O _ e
1O+ I0.0 ............................
0 50 100 150 200 250 300 350 400 450
Time, sec
Figure B-11 Test Data of An Impacted RCC Hole Growth
1.8
o.-...0.8 . .-_- ..;;o.-
0 50 100
°
. - " "o" •" • Back Hole Test Data0.6 ° -..'_o • • • • Front Hole Test Data
,o,_'_ • " - - - - Back Hole Growth Line
_I_ 4 " -- "Fr°nt H,ole Growtl Line
0.4 ........
0.2 Test Specimen: JSC1142/204.0 ft-lb Impact Energy
0.0 , Test ,Conditions: 2800 deg, F/100 psI ......
400 450150 200 250 300 350
Time, sec
Test Data of An Impacted RCC Hole GrowthFigure B-12
47
2.0
1.8
1.6
.._ 1,4
._c
1.2
E 1.0
/5m 0.8
0.6
0.4
0.2
0.0
0
Test Specimen: JSC 1152/212.6 ft-lb Impact Energy
Test Conditions: 2800 deg F/180 psf
• Back Hole Test Data
• Front Hole Test Data •
.... Back Hole Growth Line • • •
Front Hole Growth Line • • •
41_ _••v • __"- O O . o . . -e ee ee _ o- .-"
_ °
. n . _ I n n I 1 i n l _. i I n I u n J I n u u I * * . t | , _ _ . I i • • i ! , . •
50 100 150 200 250 300 350 400
Time, sec
Figure B-13 Test Data of An Impacted RCC Hole Growth
450
48
Appendix C
Error Analysis
49
Appendix C - Error Analysis
This error analysis is based on the hole growth or mass loss rate test data. The hole growth test
data was obtained by two different methods: pretest and posttest measurements from test
specimens and measurement from the test video images. The hole growth measurements made
by the JSC IS&AG indicated a minimum uncertainty of 0.01-in.-dia. measurements for the inner
hole edge and 0.03 to 0.05 in. for the outer edge diameter. These uncertainties represent an error
of approximately 1% to 7% in the final hole diameter as measured from the videotapes. Surface
temperature measurement was obtained using calibration models to correlate the surface
temperatures measured by laser pyrometer with thermocouples. Test data indicated that the laser
pyrometer's readings agreed with the thermocouple within -1.2% to +0.5% of its readings (at
temperatures of 2875°F to 3260°F). The presence of an impact damage in the test specimen
results in an increased temperature in the local impact area, however, no change in the test
conditions were made and all tests were conducted using the calibrated test conditions of a non-
impacted test specimen. The comparison of the steady-state readings between laser pyrometer
and thermocouples on RCC calibration models can be seen on the data fax from Ron Lewis
(NASA/JSC ARMSEF test facility) to I. Norman (Boeing North American) that is attached as
the last page of this appendix. Previous analysis of instrumented RCC test specimens indicated
that the recorded Iaser pyrometer surface temperature is within 30°F of the thermal math model
predicted surface temperature
Error Analysis for an Impacted RCC Hole Growth at 2500°F Tests
There were only four test data points available for 2500°F test condition, two each at 50 psf and
180 psf pressure environment. Information from pretest and posttest measurements for both test
specimens and recorded real-time video images were analyzed to determine the front and back
hole growth, and their growth rates. Both methods of obtaining the test data have their
advantages. The video image measurements provided very important transient, time-dependent,
history information. It showed that at 200 sec after test initiation the front hole growth stopped,
and its measurement agreed very well with the measurement of the posttest specimen's front hole
diameter. The basic test data for 2500°F test series is shown in table 3.
Measurements from both test specimens and recorded video images were used to perform a
regression analysis. The resulting correlation is in terms of a ratio of the RCC substrate mass
loss rate with and without an impacted hole. The regression line derived from these data along
with +10% error bands are shown in figure C-1. As can be seen, this provides an excellent
correlation of the data and was used in the flight simulation.
5O
Error Analysis for an Impacted RCC Hole Growth at 2800 °F Tests
For the 2800°F test series, ten test specimens were used fox" nine test points (nine test conditions
with three different pressures and three different levels of the impact energy). The last test point
at 180 psf pressure and 200 fl-lb of impact energy was tested twice. The first test was aborted at
150 sec after test start; the second was aborted at 14 sec after test start, then retested and aborted
again at 290 sec into the test. This was the only test point that did not have a perfect test run.
To eliminate the time dependency of the hole growth rate, test data were translated from hole
diameter growth history to hole diameter growth rate assuming an average constaqt growth rate
for each test condition. Because a significant difference appeared between the front and the back
hole growth rates, separated correlations were developed for the front and back hole growth.
Comparisons of the regression lines with test data along with +10% error bands are shown in
figures C-2 through C-7. These test data are in excellent agreement with the regression lines,
except two data points for the front hole growth rate (figure C-4) and one point for the back hole
growth rate are out of 10% error band (figure C-7).
Error Analysis for Flight Simulation Application
The temperature range of the test data from this hypervelocity impacted RCC hole growth test
program was limited: two temperature points, 2500°F and 2800°F. The flight simulation was
based on previous test data correlations (reference 7) and the correlations developed from this
current test program.
The following assumptions were made:
. The 2500°F test series had two different pressures and two different levels of impact energy.
Therefore, the substrate mass loss rate correlation for 2500°F condition was developed as a
linear function of pressure. No significant effect of impact energy on the substrate mass loss
appeared in the 2500°F test data; therefore, it was not included in the 2500°F correlation.
, Because there is no impact/thermal test data available for temperature between above 2500°F
and below 2800°F, an assumption was made that the RCC substrate mass loss correlation
developed from 2500°F test data of the current test program, in conjunction with the bare
RCC mass loss correlations developed from the previous RCC test programs, will hold for
that temperature range (reference 7).
. Because of the limited test data and to simplify the correlation formulations, it was assumed
that the RCC mass loss rate was constant rate with respect to time. The correlations
developed in this program were
• For 2500°F condition, the mass loss rate was an average over 200 sec test time.
• For 2800°F condition, the mass loss rate was an average over 450 sec test time.
51
Flight simulations made using the above assumptions and correlations provide a good
engineering prediction (+10%) of expected hole growth knowing the initial impact damage (i.e.,
penetration diameter). Spatial hole growth for a typical Space Station return entry, with an initial
hole penetration of 0.025, is shown in figure 24. The correlations are in good agreement with the
limited test data. The error analysis indicates that the correlations can be used with confidence in
predicting the growth of impacted holes in RCC for flight conditions.
0
rr¢J)t._0
....J
t._
(1)
¢J)
03
5.0
4.5
4.0
3.5
3.0
2.5
2.0
1.5
1.0
Test Temperature: 2500 deg F
Front Surface Hole Growth
I I Io Post-Test Measurement• Video Test Data
= = Correlation minus 10%
Correlation plus 10%
] [ I ....Mass Loss Ratio = Mass Loss Rate (With Hole) / Mass Loss Rate (Without Hole)
J L I 1
0 25 50 75 100 125 150 175Pressure, psf
2O0
Figure C-1 RCC Substrate Mass Loss Ratio of
With and Without An Impacted Hole
52
0.0030
or-
.,_, t-
E----
.w
D
O-r-
fie
fl) O
E.E
C)
O"-I-
0.0025
0.0020
0.0015
0.0010
0.0040
0.0035
0.0030
0.0025
0.0020
@ Front Hole
Front Hole
Test Data w/P = 50 psf
Correlation w/P = 50 psf
- ," - Correlation minus 10%
•-'-- - Correlation plus 10%
Jw
,J,wI
Test Tern 3erature • 2800°F
!y
0 25 50 75 100 125 150 175 200
Impact Energy (IE), ft-lb
Figure C-2 An Impacted RCC Front Hole Growth
225
0
I I I ITest Temperature • 2800°F
,jJ
F
I
m m •_q
I
m m
I
41, Front Hole Test Data w/P = 100 psf
_Front Hole Correlation w/P = 100 psf
- - - Correlation minus 10%
- Correlation plus 10%
I25 50 75 100 125 150 175 200
Impact Energy (IE), ft-lb
Figure C-3 An Impacted RCC Front Hole Growth
25O
250
53
cr
t-
e3
E.-
n
fl)o
-1-
0.0O3O
0.0025
0.0O2O
0.0015
I I I ITest Temperature • 2800°F
I I " Front Hole Test Data w/P = 180 3sf
Front Hole Correlation w/P = 180
,- - - Correlation minus 10%
m . Correlation plus 10%
I
2222_ f _J i _ _ w _ m
f _ maW
o
psf
n-
¢-
r-
E.-
£3
o"1-
0.0010
0.0030
0.0025
0.0020
0.0015
0.0010
0 25 50 75 100 125 150 175 200
Impact Energy (IE), ft-lb
Figure C-4 An Impacted RCC Front Hole Growth
Test Temperature • 2800°F
225 250
f
4, Back Hole Test Data w/P = 50 psf
Back Hole Correlation w/P = 50 psf
•, - - Correlation minus 10%
,.-- - Correlation plus 10%
oo
0 25 50 75 100 125 150 175 200
Impact Energy (IE), ft-lb
Figure C-5 An Impacted RCC Back Hole Growth
225 25O
.54
e5cr¢-
¢-U
E.---.w
tm
o--i-
0.0030
0.0025
0.0020
0.0015
0.0010
0.0030
I ITest Temperature • 2800°F
i
J JI
m
I
i
L
w f
• Back Hole Test Data w/P -- 100 psf
_Back Hole Correlation w/P = 100 psf
- - - Correlation minus 10%
- Correlation plus 10%
I [ I0 25 50 75 100 125 150 175 200 225 250
Impact Energy (IE), ft-lb
Figure C-6 An Impacted RCC Back Hole Growth
cr
(.9 _
E._
E3
o"-1-
0.0025
0.0020
0.0015
0.0010
• Back Hole Test Data w/P = 180 psf
Back Hole Correlation w/P = 180 )sf
- - - Correlation minus 10%
Correlation plus 10%
J
J,m
I
Test Temperature • 2800°F
25 50 75 100 125 150 175 200
Impact Energy (IE), ft-lb
Figure C-7 An Impacted RCC Back Hole Growth
225 250
.55
NOV 3 799 12:45 FR NASA/JSC B222/EHTS TO 9-2818531525 P.!
COMPARISON OF
STEADY-STATE READINGS
BETWEEN LASER PYRO
AND
THERMOCOUPLES ON RCC CALIBRATION MODELS
The tables below were copied directly from a page in the Test Director's logbook and show
differences in the laser pyro and thermocouple readings that have been typical for all RCC tests
performed in this facility. TC ! is a surface thermocouple that was located in the center of the
calibration puck; TC3 is another surface thermocouple that was located about halfway to the
edge of the puck. Run 2-1726-7 used calibration model 2B24 that was used on a number of
previous test programs, and run 2-1728-7 used a new RCC calibration model that ran a slightly
higher temperature. The response from TC3 during run 1728 was about 350°F low because the
TC was pulled back.
Run 2-1726-7: 10/27/97
Dual calibration model (pressure and heating rate) on left arm. Old RCC model 2B24 on right
arm with centered and off-centered T/Cs (#1 and #3).
m 1 p Q h-bulk t/c1 t/c 3 pym(Ib/s) (A) ,(psf) (BTU_2sec) BTU/Ib) (°F) (°F) (°F)0.4 460 196 109 3195 2875 2850 2875
0.6 570 285 141 3084 3140 3100 31350.6 600 289 149 3199 3175 3150 3190
0.6 660 160 3441 3220 3275300 3260,=
Run 2-1728-7: 10/29/97
Dual calibration model (pressure and heating rate) on left arm. New RCC model 1010 on right
arm, with centered and off-centered T/Cs (#1 and #3).
m
(Ib/s)
I
(A)p Q
(psf) (BTU/ft2sec)
t/c1
(°F)t/c 3
(°F)pyro(OF)
0.4 460 196 109 2944 2580 2910
56
Form ApprovedREPORT DOCUMENTATION PAGE OMB No. 0704-0188
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I" AGENCY USE ONLY (Loave BIank) 12"REPORTDATEMarch2000 I 3 REPORT TYPE AND DATESCOVEREDNASATechnical Paper
4. TITLE AND SUBTITLE 5. FUNDING NUMBERS
Oxidation of Reinforced Carbon-Carbon Subjected to Hypervelocity Impact
6. AUTHOR(S)
Donald M. Curry, Vuong T. Pham, Ignacio Norman*, Dennis C. Chao*
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
Lyndon B. Johnson Space Center
Houston, Texas 77058
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
National Aeronautics and Space Administration
Washington, DC 20546-0001
8. PERFORMING ORGANIZATIONREPORT NUMBERS
S-852
10. SPONSORING/MONITORINGAGENCY REPORT NUMBER
TP-2000-209760
11. SUPPLEMENTARY NOTES
* Boeing North American, Inc., Houston, Texas 77058
12a. DISTRIBUTION/AVAILABILITY STATEMENT
Available from the NASA Center for AeroSpace Information (CASI)7121 Standard
Hanover, MD 21076-1320 Subject Category: 27
12b. DISTRIBUTION CODE
13. ABSTRACT (Maximum 200 words)This paper presents results from arc jet tests conducted at the NASA Johnson Space Center on reinforced carbon-carbon (RCC)
samples subjected to hype_'elocity impact. The RCC test specimens are representative of RCC components used on the Space Shuttle
Orbiter. The arc jet testing established the oxidation characteristics of RCC when hypervelocity projectiles, simulating
meteoroid/orbital debris, impact the RCC material. In addition to developing correlations for use in trajectory simulations, we discuss
analytical modeling of the increased material oxidation in the impacted area using measured hole growth data. Entry flight simulations
are useful in assessing the increased Space Shuttle RCC component degradation as a result of impact damage and the hot gas flow
through an enlarging hole into the wing leading-edge cavity.
14. SUBJECT TERMS
arc jet engines; arc discharges; impact damage; hypervelocity impact; impact prediction;
computerized simulation; impact tests; debris; space debris; oxidation
17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATIONOF REPORT OF THIS PAGE
Unclassified Unclassified
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Prescribed by ANSI Std. 239-18298-102
15. NUMBER OFPAGES
63
16. PRICE CODE
19. SECURITY CLASSIFICATIONOF ABSTRACT
Unclassified
20. LIMITATION OF ABSTRACT
Unlimited