March 2003

NASA/TM-2003-212171

Thermal/Mechanical Response of a Polymer

Matrix Composite at Cryogenic Temperatures

Karen S. Whitley and Thomas S. Gates

Langley Research Center, Hampton, Virginia
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March 2003

NASA/TM-2003-212171

Thermal/Mechanical Response of a Polymer

Matrix Composite at Cryogenic Temperatures

Karen S. Whitley and Thomas S. Gates

Langley Research Center, Hampton, Virginia
Available from:

NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)

7121 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161-2171

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1

Abstract

In order for polymeric-matrix composites to be considered for use as structural materials in

the next generation of space transportation systems, the mechanical behavior of these materials at

cryogenic temperatures must be investigated. This paper presents experimental data on the

residual mechanical properties of a carbon-fiber polymeric composite, IM7/PETI-5, both before

and after aging. Both tension and compression modulus and strength were measured at room

temperature, –196°C, and –269°C on five different laminate configurations. One set of

specimens was aged isothermally for 576 hours at –184°C in an unconstrained state. Another set

of corresponding specimens was aged under constant uniaxial strain for 576 hours at –184°C.

Based on the experimental data presented, it is shown that trends in stiffness and strength that

result from changes in temperature are not always smooth and consistent. Moreover, it is shown

that loading mode and direction are significant for both stiffness and strength, and aging at

cryogenic temperature while under load can alter the mechanical properties of pristine, un-aged

laminates made of IM7/PETI-5 material.

Introduction

The National Aeronautics and Space Administration (NASA) has recently initiated the Space

Launch Initiative (SLI) program that will, in part, advance some of the key technologies required

for the next generation of launch vehicles. This next generation of space transportation systems

may require both reusable launch vehicles (RLV’s) and expendable launch vehicles (ELV’s) to

satisfy mission requirements. One of the key technologies identified for RLV’s and ELV’s has

been the reduction in structural weight through the use of advanced materials and manufacturing

methods. This reduction in structural weight must be tempered against the increased demands on

performance, damage tolerance, and lifetime durability. One potential source for structural

weight reduction is the replacement of traditional metallic cryogenic-fuel tanks with polymericmatrix-composite (PMC) tanks. The interest in design of polymeric-composite, cryogenic-fuel

tanks for launch vehicles goes back several years and includes research associated with the

National Aerospace Plane (NASP), single-stage-to-orbit (SSTO) vehicles [1], [2] and other

launch vehicle applications [3]. For internal and external tanks, design of the tank may take the

form of externally stiffened shells of PMC material or thin-walled sandwich shells constructed

with a lightweight core and PMC facesheets. Regardless of the design, the PMC-based tanks

will be required to safely carry pressure and flight loads and operate over temperatures that may

range from –250°C to +120°C. From a durability perspective, the primary performance criteria

of the PMC material is to retain mechanical properties within allowable limits over the life time

of the tank, while minimizing loss of cryogenic fuel due to permeation or leakage through the

tank wall.

Aside from their use in space launch vehicles, there have been very few applications of

PMC’s as structural materials in cryogenic environments. Consequently, a review of the

available literature provides a limited amount of experimental data on the mechanical properties

of polymers or PMC’s operating at cryogenic temperatures. For amorphous and crystalline

polymeric materials, Perepechko [4] provides details on a number of non-mechanical properties

that include thermal expansion, thermal conductivity, as well as, viscoelastic or dynamical

mechanical properties. From these studies, it is clear that many of the polymer properties are not
2

linear with respect to temperature over a range from room temperature to approximately absolute

zero. In the work by Pannkoke [5], fatigue tests were performed on unidirectional thermoplastic

composites at –196°C. The fatigue strength at 106

 cycles was found to be only 60% of the static

strength. In this work, it was recognized that large thermal stresses degraded the fatigue

performance. In a series of articles by Ahlborn [6], [7], static tests, and thermally cycled, and

mechanically cycled tests were performed on unidirectional and cross-ply thermoplastic

composites. Isothermal tests were performed at 23°C, –196°C, and –269°C and cyclic thermal

tests were performed between –196°C and 23°C. Strength, damage, and fatigue life were

measured for all test conditions. He found that the matrix-dominated properties, shear and

transverse tension, were largely related to the effects of temperature on static strength. As the

temperature decreased, increases in matrix-dominated strength, could be offset by the

development of thermal-stress-induced cracks. As in previous studies, the fatigue life was

reduced as the test temperature was decreased.

More recently, a compilation of test data for several PMC material systems [8] indicated that

tensile modulus, tensile strength, and compressive strength all generally increased as the test

temperature was decreased from 23°C to –269°C. Once again, it was found that thermal stresses

had a large influence on behavior and that the sensitivity of matrix-dominated properties to

temperature can be used to help explain the stress-strain response of laminated composites. Of

these matrix-dominated properties, transverse ply strength appears to play a critical role in the

development and growth of residual stress-induced matrix cracks. Schoeppner [9] recently

investigated the range of thermomechanical loading over which steady-state matrix cracking is

likely to occur. Shimoda [10] has performed additional studies related to matrix crack

development in PMC’s due to cryogenic-temperature exposure.

It is recognized that a broad spectrum of factors influence the mechanical properties of PMC’s

including material selection, composite fabrication and handling, aging or preconditioning,

specimen preparation, laminate ply lay-up, test procedures, etc. The present study focuses on

test temperature, preconditioning methods, and laminate configuration as the primary test

variables. It is expected that by focusing on those three variables the results of this study will aid

in the development of future material qualification methods and design verification testing.

Toward that purpose, this paper provides experimental results and test methods developed from a

series of thermal/mechanical tests. The selected test conditions represent a range of exposure

times, loads and temperatures similar to those experienced during the lifetime of a liquidhydrogen fuel tank.

In particular, this paper first presents a description of the material system, IM7/PETI-5, and

the material characterization test-plan, including a description of the experimental methods that

were used for tension and compression testing at room temperature and cryogenic temperatures.

Also presented, are the techniques used to precondition the IM7/PETI-5 material, followed by

details on acquiring the residual property measurements. In addition, residual-stress analysis is

used to evaluate thermal stresses as a function of ply lay-up, test temperature, and preconditioning. Results of the residual stiffness and strength are also evaluated as a function of ply

lay-up, test temperature, and pre-conditioning. Finally, photographs of the fracture and edge

surfaces are shown, and discussed.
3

Material System

The PMC material used in this study, IM7/PETI-5, consisted of a continuous high-strength,

intermediate-modulus, carbon fiber in a thermoplastic polyimide matrix. All test materials were

laminated composites fabricated at the NASA Langley Research Center. These composite

panels consisted of [0]12 and [90]12 unidirectional laminates, [±25]3S and [±45]3S angle-ply

laminates, and a 13-ply laminate given by [45/903/-45/ 03]S. The bar notation over the 0

indicates that the 03 ply group spans the mid-plane of the laminate. Figure 1 provides a

schematic illustrating the 0° ply and the 90° fiber directions relative to the specimen dimensions.

These lay-ups were chosen to provide basic lamina-level material constants and for the case of

the [±25]3s and [45/903/-45/ 03]S laminates, to be representative of a composite wall in a typical

cryogenic-propellant tank. For a cryogenic tank, the orientation of the 0° in the [45/903/-45/ 03]S

laminate would be along the longitudinal axis of the tank.

All composite panels were fabricated by hand lay-up. The bagging and cure processes

employed are consistent with standard practices. Specifically, a full vacuum was applied to each

bagged panel during the entire cure cycle. The cycle started with a ramp to 260°C, at

5°C/minute. After a 1-hour hold at 260°C, the autoclave was pressurized to 1480 kPa at 170

kPa/minute while the temperature was increased to 370°C at 5°C/minute. Then, the temperature

was held at 370°C for 1 hour, and the pressure was maintained until the temperature cooled to

38°C. After curing, through-transmission, ultrasonic inspection of the laminates indicated that

there were no significant internal anomalies in any of the panels. The glass transition

temperature (Tg) of the as-received composite material was 267°C, as measured by the peak in

the Tan d curve of tests run in a dynamic mechanical analysis (DMA) test [11].

Material Characterization

The material characterization test plan consisted of two phases. The first phase addressed

material aging where the material was subjected to long- term exposure at cryogenic temperature

in either an unloaded (unconstrained) or statically loaded state. The second phase was residual

property characterization of pristine and aged specimens at three temperatures (23°, –196°, and

–269°C). This second phase provided fundamental material properties and an understanding of

material behavior as a function of prior aging conditions and test temperatures. The complete

test matrix, including the phase-one aging conditions and the phase-two residual property tests,

are shown in table 1.

All tests were conducted on coupon-type specimens that were cut from larger panels prior to

aging. Three replicates of each condition were used to generate the test data. A schematic of the

tension-test coupon is presented in figure 1. The 254-mm-long tension coupons were mounted in

the test set-up such that 63 mm of each end were gripped between two serrated face-plates,

leaving a 128-mm test section. The face-plates were clamped together by using six hex-head

bolts, which were torqued to 61 N-m with a hand-held torque wrench. The tensile coupons

tested at room temperature were 152.4 mm long, and 38.1 mm of each end were mounted in

hydraulic wedge grips, leaving a 76.2-mm test section.

The general set of guidelines used for the tension test included ASTM D3039-76 and the

Suppliers of Advanced Composite Materials (SACMA) SRM 4R-94 test method [12]. The

tensile modulus was calculated by using ASTM D3039-76 and a linear regression least-squares
4

curve fit of the stress-strain data in the linear region of the stress-strain curve. The linear region

of the stress-strain curve was most commonly defined as being between 1000μe and 3000μe for

all test conditions and specimen ply lay-ups.

Because an existing end-loaded compression fixture, developed by Northrup Grumman, had

been successfully used for previous compression testing of IM7/PETI-5, for the NASA High

Speed Research Program [13], it was selected as the fixture to be used for the compression

testing in this study. A schematic of the compression coupon is presented in figure 2. The

gauged coupon is placed in the center cut-out section of the compression test fixture, as shown in

figure 3(a). A side view of the assembled fixture is shown in figure 3(b). The compression tests

were performed by using the SACMA recommended method, SRM 1R-94 [14], and the ASTM

D695 references as guides. The compressive strength and modulus were calculated according to

the method described in ASTM D695, by calculating the slope of the tangent to the initial linear

portion of the stress-strain curve. In most cases, the linear region of the stress-strain curve from

the compression test was defined as being between 1000μe and 2000μe for all test conditions and

specimen ply lay-ups.

Test phase one - Material aging

Another unique test fixture, also developed under the NASA High Speed Research Program,

was used to produce a constant tensile-strain condition for half of the total number of specimens

used during the aging phase. Several fixtures, like the one shown in figure 4, were constructed

of Invar material (coefficient of thermal expansion (CTE) = 1.4 μ mm/mm – °C) and could

accommodate two rectangular specimens. The high stiffness of the fixture relative to the test

specimen and the low CTE of the Invar material ensured a constant-strain condition during the

entire aging period. The ability of the test-fixture to maintain a constant strain within a specimen

was verified by applying a pre-load to a gauged specimen mounted in the fixture, and monitoring

the strain while aging at –184°C. Each specimen was individually preloaded to the desired strain

level by compressing a series of spring-type washers that react against the frame to put the

specimen into tension. The washers were compressed by turning an adjacent nut with a handheld wrench, while the strain in the specimen was monitored with a longitudinal extensometer

mounted on the specimen. The preload strain levels, given in table 1, were selected to be

approximately 50% of the not-aged room-temperature failure strain. A corresponding set of

specimens was aged in an unconstrained condition. The constant-strain fixtures with specimens

and the corresponding unconstrained specimens were placed into a large cryo-chamber that

maintained a constant temperature of –184°C for an aging time of 576 hours.

Test phase two - Residual properties

In phase two of the test program, residual strength and stiffness of all three sets of specimens

(pristine, aged without load, aged with load) were measured at room temperature (24°C),

–196°C, and –269°C. All residual-property tests were performed with a servo-hydraulic test

machine, using a displacement rate of 1.27 mm/min. The isothermal cryogenic test conditions at

–196°C and –269°C were achieved by immersing the test specimen and load introduction

apparatus into liquid nitrogen or liquid helium, respectively. In order to reach thermal
5

equilibrium, the specimen stayed immersed in a constant level of the cryogen for at least 15

minutes prior to mechanical loading.

Stress in the test specimens was calculated by using load, as measured by the test-machine

load cell, divided by the original cross section of the specimen. Strain in the tensile-test

specimens was measured by using a combination of a cryogenic-rated axial extensometer (MTS

model 634.11F-21) and bonded electrical-resistance strain gages (Measurements Group WK-00-

250BG-350). Strain in the room-temperature compressive-test specimens was measured by using

strain gages, Measurements Group CEA-06-125WT-350, and the compression specimens tested

at cryogenic temperatures used Measurements Group WK-06-120WT-350 gages. Strain gage

selection was based on considerations discussed in the report [15], “Strain Gage Selection,” such

as material CTE, static-load test conditions, isothermal test temperatures, and gage availability.

The WK-series gages are self-temperature-compensating to correct for thermal drift. Placement

of these sensors on the tensile specimens is shown in figure 1, and placement on the compression

specimen is shown in figure 2. All the strain gages were bonded in a back-to-back configuration.

Prior to aging, representative samples were polished along one edge. Sections (2.6 x 1.3 cm)

were potted in a clear epoxy mold and were then circulated in an automated polishing wheel at

150 rpm’s with an applied weight of 5 pounds per specimen. Consecutive grades of grit paper

(600, 800, 1200 grit) were used for the polishing. The specimens received a final polish with a

solution of silica alumina suspension fluid. The polished specimen edges were examined for

damage (microcracks, delaminations) using optical microscopy, and photomicrographs were

taken to establish the baseline condition. After aging, but prior to the destructive residual tests,

visual examination of all lay-ups was again performed to determine if the exposure conditions

generated any microcracks, damage, or change in surface morphology. Photomicrographs

representative of the specimen’s edge before and after aging are shown in figure 5.

Residual-Stress Analysis

The previous studies cited herein indicate that thermal stresses alone may degrade laminate

performance. For thermal loading, stresses are induced at the ply level by mismatches in

expansion or contraction of ply constituents and by constraining effects caused by adjacent plies

that prevent a free relative expansion or contraction. The residual stresses in any ply can be

calculated by using classical laminated plate theory [16]. In general, for a unidirectional

laminate under plane stress, the constitutive relationship is given by

s

s

s

e

e

g

e

e

g

11

22

12

11 12

12 22

66

11

22

12

11

22

12

0

0

0 0

Ï

Ì

Ô

Ó

Ô

 ̧

 ̋

Ô

 ̨

Ô

È

Î

Í

Í

Í

 ̆

 ̊

 ̇

 ̇

 ̇

Ï

Ì

Ô

Ó

Ô

 ̧

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Ô

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Ô

Ï

Ì

Ô

Ó

Ô

 ̧

 ̋

Ô

 ̨

Ô

Ê

Ë

Á

Á

Á

ˆ

 ̄

 ̃

 ̃

 ̃

= -

Q Q

Q Q

Q

T

T

T

(1)

where sij are the normal stress components in the principal material directions, eij are the normal

strains in the principal material directions, and Qij are the reduced stiffness coefficients. In this

equation, the subscripts”1” and “2” denote the fiber direction and inplane direction perpendicular
6

to the fiber, respectively. The thermal strains, in the principal material directions, eij

T , are

calculated using

 

e

e

g

a

a

11

22

12

1

2

0

T

T

T

T

Ï

Ì

Ô

Ó

Ô

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Ì

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= D (2)

The thermal strains are proportional to the temperature change, DT , which is the temperature

difference between the test condition and the stress-free condition at cure. The constants of

proportionality are the coefficients of thermal expansion (CTE), a1 and a2. The thermal shear

strain, g 12

T , is zero because an orthotropic material only exhibits dilation along its principal

material coordinate directions, with a change in temperature. Clearly, from equation (2) it is

apparent that for large values of DT , which occur at cryogenic temperatures, the magnitude of

the thermally induced strains is highly dependent on the material CTE values. In particular,

nonlinear behavior of the CTE with respect to temperature, as illustrated in table 2 of reference

[17], may result in residual thermal stresses that also vary in a nonlinear manner with respect to

temperature. To illustrate this effect, CTE values and material properties given in table 3 were

used with classical laminated plate theory to predict the maximum transverse ply stress, s22, as a

function of laminate stacking sequence, test temperature, and laminate preconditioning. The

results of these calculations are given in table 4. It should be noted that CTE values of pristine

material were used for all conditions and the stress-free temperature was assumed to be the glass

transition temperature of IM7/PETI-5. With respect to temperature, table 4 indicates that the

trend for the not-aged condition is for the transverse stress to increase with a decrease in

temperature. However, both of the isothermal-aged conditions predict the largest stress (s22) to

occur at the intermediate temperature, –196°C. With respect to ply lay-up, the highly

constrained [45/903/-45/ 03]S laminate is predicted to have the highest stress at any given

temperature. However, the [±45]3S angle-ply laminate is predicted to have stress magnitudes

close to those predicted for the [45/903/-45/ 03]S laminate. With respect to preconditioning, no

clear trends are evident from a residual-stress analysis.

Results and Discussion

In this section of the present paper, elastic modulus and residual strength results are presented

for laminates with [0]12, [±25]3S, [45/903/-45/ 03]S, [90]12, and [±45]3S, ply stacking sequences.

Tables 3 and 5 contain the tensile and compressive elastic modulus, respectively, at the three

different test temperatures, for each lay-up, and for each aging condition. Each value in the

table represents the average of three replicates, along with the standard deviation. The lamina inplane moduli values, E1 and E2, were calculated directly from stress-strain behavior of the [0]12

and [90]12 unidirectional laminates, respectively. The laminate effective modulus value Ex of the

[±25]3S and [45/903/-45/ 03]S laminates was calculated from the laminate stress-strain behavior.

The lamina in-plane shear modulus G12 was calculated indirectly from the stress-strain behavior

of the [±45]3S laminates by using the equation,
7

 G Ex

xy

12 2 1 = ( +n ) (3)

where nxy is the laminate Poisson’s ratio and Ex is the laminate longitudinal modulus. The form

of equation (3) is customarily associated with isotropic materials, however for the special case of

a [±45]S laminate subjected to axial loading, the shear modulus can be calculated by using this

equation. (See reference [18]).

Measured laminate tensile and compressive strength values are listed in tables 6 and 7,

respectively, along with the associated standard deviation. Failure was defined as the point of

complete loss of load-carrying capability during the test. Due to the nonlinear nature of the

stress-strain behavior in the [±45]3S laminates, their tensile strength was defined as the initial

point of deviation from the nearly horizontal portion of the stress-strain curve. Normalized

tensile modulus and strength of the [0]12, [±45]3S, and [45/903/-45/ 03]S laminates are presented in

figures 6-11 and normalized compressive modulus and strength are presented in figures 12-17.

Both modulus and strength values have been normalized against the corresponding results for the

not-aged condition tested at room temperature for each laminate.

Effect of Temperature on Modulus and Strength

By examining the results from specimens subjected to the not-aged test condition, the effects

of temperature on tensile modulus and strength were found. The fiber-dominated [0]12 laminate

experienced a decrease in both the modulus and strength due to testing at cryogenic

temperatures, with up to a 35% decrease in strength for the –196°C case. Conversely, the shear

modulus and longitudinal strength of the matrix-dominated [±45] 3S laminate increased as the

temperature decreased. The shear modulus increased by as much as 35% and the strength by as

much as 50% when tested at –269°C. The transverse modulus (E2) showed a slight decline at

cryogenic temperatures while the transverse strength dropped by approximately 70% when the

temperature was reduced to –269°C. The tensile modulus and strength of the not-aged [45/903/-

45/ 03]S laminates decreased by nearly 20% at –196°C, with some reverse in this decline as the

temperature was lowered to –269°C. The [±25]3S laminates showed little sensitivity in modulus

or strength to cryogenic temperature.

The compressive modulus and strength increased for all the laminates that were tested at

cryogenic temperatures. In most cases, the highest modulus and strength occurred at the lowest

test temperature (–269°C). Of all the laminates tested, the [90] 12 laminates experienced the

greatest increase (67%) in strength while the [±45]3s laminates experienced the greatest increase

(57%) in modulus at cryogenic temperatures.

Effect of Aging on Modulus and Strength

To determine the effects of aging, the panels subjected to the not-aged condition were

compared to the corresponding panels subjected to the aged-without-load and the aged-with-load

conditions. The results show that, in general, aging had very little effect on the tensile modulus

of the fiber-dominated [0]12 laminate. Aging did cause a slight increase in tensile strength for the

[0]12 laminate, with the most significant change occurring when tested at –269°C after aging with
8

load. In contrast, aging did not significantly affect the shear modulus or the longitudinal strength

of the matrix-dominated [±45]3S laminate.

For the [45/903/-45/ 03]S laminate, aging increased the modulus. The [45/903/-45/ 03]S laminate

that was aged with load, experienced the largest increase (18%) in modulus, which occurred at

–269°C. Aging also increased the strength of the [45/903/-45/ 03]S laminate, with the most

significant increase occurring for the aged-with-load condition, where the strength increased

relative to the baseline strength by 21% at –196°C and 11% at –269°C. In general, aging

decreased the modulus of the [±25]3S laminate and had little effect on the laminate strength.

From the limited data for the [90]12 laminate, it appears that aging had little effect on the

transverse modulus and the strength.

As expected, the standard deviation in strength measurements was more significant than the

standard deviation associated with modulus measurements. In part, large standard deviations in

the strength values can be attributed the fact that strength is defined by a single value, whereas

modulus is an averaged value, given by the slope of the stress-strain curve. Strength in

laminated composites is also more sensitive to processing, handling, and test-parameter

variability.

The compressive modulus and strength of the [0]12 laminate were not greatly influenced by

aging. The matrix-dominated [±45]3S laminate showed a decrease in modulus and strength due to

aging in most cases, and in general, the modulus decreased more than the strength. The

compressive modulus of the [45/903/-45/ 03]S laminate tended to decrease after aging. However,

the compressive strength of the [45/903/-45/ 03]S laminate increased due to aging. The greatest

increase (88%) occurred at –269°C after aging with load, however, the standard deviation

associated with those strength values was quite high. The room-temperature modulus and

strength of the [±25]3S laminate were unaffected by aging. However, at –196°C both the

modulus and the strength increased due to aging. The [90]12 laminate was not significantly

affected by aging.

Fracture Surfaces

The fracture surfaces of each failed specimen were examined for any distinguishing features

that could be attributed to test temperature or aging conditions. It was observed that aging did

not produce any distinguishing characteristics in the fracture surfaces of the tensile specimens.

The greatest distinction in the appearance was observed when comparing the fracture surfaces of

the room-temperature-test specimens with those of the cryogenic-temperature-test specimens.

However, there were no discernable differences between the specimens tested at –196°C and the

specimens tested at–269°C.

The fracture surface of the [45/903/-45/ 03]S laminates tested at room temperature and –196°C

are shown in figures 18 and 19, respectively. The fractured edge of the [45/903/-45/ 03]S

laminates tested at room temperature appeared more jagged, such that the ±45° laminae were

more apparent. The [45/903/-45/ 03]S laminates tested at –196°C failed such that the fracture

surface was nearly perpendicular to the direction of the applied load. The [±25]3S laminates

tested at room temperature (figure 20) displayed irregular fracture surfaces, whereas the [±25]3S

laminates tested at –196°C displayed a very clean-cut failure at a 25° angle that is shown in

figure 21. All the [0]12 laminates failed by splitting along the 0° direction. The [0]12 laminates

tested at room temperature displayed a catastrophic splintering as shown in figure 22, whereas
9

the [0]12 laminates (figure 23) tested at –196°C, were characterized by only a few splits through

the thickness. In general, the failure of the [45/903/-45/ 03]S, [±25]3S, and the [0]12 laminates at

cryogenic temperatures could be described as “clean breaks,” as opposed to the irregular

fragmented breaks that occurred at room temperature. The clean breaks indicate that a more

brittle fracture occurred at cryogenic temperatures, and the irregular breaks indicate that a more

ductile fracture occurred at room temperature. There was no difference in the fracture surface of

the [90]12 laminates, which all failed straight across the width of the specimen in the 90°

direction, regardless of the test temperature. Due to the nonlinear stress-strain behavior of the

[±45]3S laminate, loading to the point of fracture did not always occur.

The fracture surfaces of the specimens tested in compression did not show any differences

due to test temperature or aging conditions, possibly due to the fact that the test fixture enclosed

most of the compression coupon. However, each set of laminate configurations displayed unique

failure characteristics, as expected. In all cases, failure occurred contiguously above or below

the strain-gage section. The [0]12-laminate compression failures did not split along the length, as

in the tension tests. Instead, the fracture surface could be described as having a short broom-like

appearance along the failed edge (figure 24). The [90]12 laminates failed with a clean-cut,

beveled edge across the width (figure 25). The [±25]3S laminates failed sharply along a ±25°

angle ply and failure may have propagated from any one of the four corners of the specimen

(figure 26). The [45/903/-45/ 03]S laminate, failed straight across the width of the specimen, with

the fracture surface having a crushed appearance (figure 27).

Edge Surface Morphology

Optical examination of surface morphology was intended to provide data on the initiation and

growth of any damage that occurred during specimen aging. Photomicrographs of the

specimen’s polished edges taken during this study are shown in figure 5. The photos in the first

column are for the not-aged material, the second column are specimens that have been aged

without load, and the third column are specimens that have been aged with load. Very few

microcracks or similar damage were observed before or after aging in any of the lay-ups.

However, the surface morphology shows some definite degradation along the exposed edges

after aging without load, and it appears that further surface degradation occured after aging with

load. This degradation occurred in all lay-ups and can be described as pitting in the matrix

regions.

Summary and Conclusions

Five different laminates made of IM7/PETI-5 were evaluated for tensile and compressive

strength and stiffness at room temperature and at two cryogenic temperatures. The effects of

laminate configuration, test temperature, and preconditioning or aging were investigated. Aging

consisted of 576 hours of exposure in a –184°C environment, both with and without constantstrain mechanical load. Specimens were also examined for evidence of damage or

microcracking before and after aging.

Examination of the basic laminate properties such as strength and stiffness, in the longitudinal

and transverse directions, indicates that cryogenic temperatures can have an appreciable
10

influence on behavior, as expected, but unexpected trends can occur. For example, the tensiontest data show that the longitudinal and transverse stiffness and strength decreased as the test

temperature decreased. Conversely, the tensile shear modulus and strength increased as the

temperature decreased. For tension loading, some of the laminate stacking sequences, such as

[45/903/-45/ 03]S and [±25]3S, were influenced less by a reduction in temperature, than the other

laminates. However, the [45/903/-45/ 03]S laminate was more sensitive to cryogenic temperatures

than the [±25]3S laminate. For compression loads, exposure to cryogenic temperatures produced

an increase in both the modulus and strength of all the laminates. The greatest increase in

compressive strength occurred for the [90]12 laminates and the greatest increase in modulus

occurred for the [±45]3S laminates.

Based on lamination theory, residual-stress calculations indicate that the transverse tensile ply

stresses, associated with matrix cracking, can be quite high at cryogenic test temperatures. The

[±25]3S and [45/903/-45/ 03]S laminates were predicted to have the highest values of residual

stress, which is important because these transverse residual stresses may reduce the strength and

stiffness of the laminates by accelerating the initiation, growth, and accumulation of transverse

microcracks in a ply.

Aging the material at a constant cryogenic temperature caused changes in the strength and

stiffness, compared to the corresponding values for the not-aged or as-received condition. In

general, it appears that this type of aging will increase both the tensile strength and stiffness,

particularly in the [±45]3s, [90]12, and [45/903/-45/ 03]S laminates. Of particular interest, is the fact

that isothermal-cryogenic aging while at a constant strain condition produced some additional

increase in strength and stiffness, as compared to aging in an unconstrained condition. The aging

tends to decrease the compressive strength and stiffness in the [±45]3s, [90]12, and the [0]12

laminates. For the IM7/PETI-5 material, aging at cryogenic temperatures also produces

degradation in the matrix material along the exposed edges. This degradation may have an

adverse influence on residual strength due to an increase in possible failure initiation sites.

The aging did not influence the fracture surface characteristics. Cryogenic temperatures

affected the fracture surfaces of only the tensile specimens. In general, the fracture surfaces

from the room-temperature tensile tests appeared catastrophic, having irregular breaks, whereas

the specimens tested at cryogenic temperatures exhibited clean breaks with a more regular

appearance. The differences in the fracture surfaces of the tensile specimens were attributed to

the ductility of the matrix material at room temperature verses the brittle nature of the matrix at

low temperatures.

Overall, the results of the present study show that the trends for changes in stiffness and

strength with changes in temperature are not always as expected, indicating that tests should be

run at several temperature levels. In addition, the loading mode and direction are significant for

both stiffness and strength characterization. Likewise, aging or long-term exposure at cryogenic

temperature while under load can alter the mechanical properties of the as-received laminate,

implying that design values should take into account thermomechanical exposure over the lifetime of the structure.
11

Acknowledgements

The authors would like to express their thanks to Lisa Hawks, Everitt Brown, and Ed Townsley

of NASA Langley Research Center for their technical support.

References

1. Morino, Y., Ishikawa, T., Aoki, T., Kumaza, H., Hayashi, Y., "Feasibility Study of

CFRP Material Application to the Cryogenic Propellant Tank of Reusable Launch

Vehicles in," 6th Japan International SAMPE Symposium, Japan, 1999, pp. 1127-1130.

2. Robinson, M. J., "Composite Structures on the DC-XA Reusable Launch Vehicle," 28th

International SAMPE Technical Conference, Seattle, Washington, 1996.

3. Callaghan, M. T., "Use of resin composites for cryogenic tankage," Cryogenics, Vol. 31,

No. 4, 1991, pp. 282-287.

4. Perepechko, I., Low-Temperature Properties of Polymers, Mir Publishers, Moscow,

1997.

5. Pannkoke, K., Wagner, H. J., "Fatigue properties of unidirectional carbon fiber

composites at cryogenic temperatures," Cryogenics, Vol. 31, No. 4, 1991, pp. 248-251.

6. Ahlborn, K., "Durability of carbon fibre reinforced plastics with thermoplastic matrices

under cyclic mechanical and cyclic thermal loads at cryogenic temperatures,"

Cryogenics, Vol. 31, No. 4, 1991, pp. 257-260.

7. Ahlborn, K., "Cryogenic mechanical response of carbon fibre reinforced plastics with

thermoplastic matrices to quasi-static loads," Cryogenics, Vol. 31, No. 4, 1991, pp. 252-

256.

8. Schutz, J. B., "Properties of composite materials for cryogenic applications," Cryogenics,

Vol. 38, No. 1, 1998, pp. 3-12.

9. Schoeppner, G. A., Kim, R., Donaldson, S. L., "Steady State Cracking of PMC's at

Cryogenic Temperatures," 42nd AIAA/ASME/ASCE/AHS/ASC Structural Dynamics

and Materials Conference, Seattle, WA, AIAA-2001-1216, 2001.

10. Shimoda, T., "Study of CFRP Application to the Cryogenic Propellant Tank of Reusable

Launch Vehicle," 42nd AIAA/ASME/ASCE/AHS/ASC Structural Dynamics and

Materials Conference, Seattle, WA, AIAA-2001-1598, 2001.

11. Kampf, G., Characterization of Plastics by Physical Methods, Experimental Techniques,

and Practical Application, Hanser, Munich, 1986.
12

12. SACMA, “SACMA Recommended Test Method for Tensile Properties of Oriented

Fiber-Resin Composites” SRM 4R-94 (Composites Fabricators Association, 1999).

13. National Research Council, Ed., U.S. Supersonic Commercial Aircraft, National

Academy Press, Washington, D.C., 1997.

14. SACMA, “SACMA Recommended Test Method for Compressive Properties of Oriented

Fiber-Resin Composites” SRM 1R-94 (Composites Fabricators Association, 1999).

15. Groups, V. M., “Strain Gage Selection” TN 505 (Vishay Intertechnology, Inc., 1999).

16. Jones, R. M., Mechanics of Composite Materials, Scripta Book Company, Washington,

D.C., 1975.

17. Johnson, T. F., Gates, T. S., "High Temperature Polyimide Materials in Extreme

Temperature Environments," 42nd AIAA/ASME/ASCE/AHS/ASC Structural Dynamics,

and Materials Conference, Seattle, WA, AIAA-2001-1214, 2001.

18. Herakovich, C. T., Mechanics of Fibrous Composites, John Wiley & Sons, Inc., New

York, 1998.
13

Table 1. Test matrix illustrating the residual test temperature and aging condition for each

type of laminate

Aging condition (–184°C) Test

temperature

(°C)

No load Static load

(3000 me)

Static load

(4000 me)

24

[0]12, [90]12,

[±45]3s, [±25]3s,

[45/903/-45/ 03]s

[±45]12,

[0]12

[±25]3s,

[45/903/-45/ 03]s

-196

[0]12, [90]12,

[±45]3s, [±25]3s,

[45/903/-45/ 03]s

[±45]12,

[0]12

[±25]3s,

[45/903/-45/ 03]s

-269

[0]12, [90]12,

[±45]3s, [±25]3s,

[45/903/-45/ 03]s

[±45]12,

[0]12

[±25]3s,

[45/903/-45/ 03]s

Table 2. Coefficients of thermal expansion for not-aged IM7/PETI-5 material system

Temp μ mm/mm – °C

(°C) a1 a2

24 –1.30 19.45

–196 –1.49 20.05

–273 –2.77 18.46
14

Table 3. Measured tensile elastic modulus values

Specimen ply

lay - up

Temp

°C

Material

property

Not aged

modulus (GPa)

± Std dev.

Aged without load

modulus (GPa)

± Std dev.

Aged with load

modulus (GPa)

± Std dev.

[0]12 24 E1 157.6 ± 9.3 153.6 ± 5.9 147.2 ± 2.6

[0]12 -196 E1 151.6 ± 2.2 146.3 ± 11.5 149.2 ± 1.9

[0]12 -269 E1 145.1 ± 7.4 143.0 ± 8.8 158.6 ± 20.8

[±25]3S 24 Ex 71.9 ± 1.0 70.3 ± 1.6 59.8 ± 2.6

[±25]3S -196 Ex 75.7 ± 0.8 63.8 ± 8.5 71.5 ± 3.0

[±25]3S -269 Ex 73.1 ± 3.7 69.4* 65.6 ± 12.2

[45/903/-45/ 03]S 24 Ex 50.7 ± 1.0 46.7 ± 4.3 51.0 ± 1.1

[45/903/-45/ 03]S -196 Ex 40.6 ± 5.9 48.0 ± 1.8 48.6 ± 3.0

[45/903/-45/ 03]S -269 Ex 43.8 ± 1.9 44.3 ± 3.9 52.1 ± 1.3

[90]12 24 E2 8.7 ± 0.1 8.6 ± 0.1 NA

[90]12 -196 E2 7.5 ± 0.1 9.5 ± 2.0 9.6 ± 0.1

[90]12 -269 E2 7.8 ± 1.2 7.1 ± 0.8 NA

[±45] 3S 24 G12 4.6 ± 0.4 5.0 ± 0.1 5.8 ± 0.2

[±45] 3S -196 G12 5.8 ± 0.1 6.2 ± 0.1 6.0 ± 0.2

[±45] 3S -269 G12 6.2 ± 0.2 5.5 ± 0.9 6.1 ± 0.3

(NA = not available)

* Only one specimen tested.
15

Table 4. Calculated maximum transverse ply stress, s22, for thermal loading only

Maximum transverse ply stress (MPa)

Precondition Temp

(°C)

[0]12 [±25]3s [45/903/

-45/ 03]S

[±45]3s

24 0.0 18.2 41.3 40.6

Not aged –196 0.0 37.7 70.9 69.7

–269 0.0 45.7 83.7 82.3

24 0.0 18.9 40.8 40.1

Iso-aged without –196 0.0 43.6 88.3 86.5

–269 0.0 40.8 76.5 75.3

24 0.0 20.4 40.7 39.9

Iso-aged with load –196 0.0 43.1 89.3 87.5

–269 0.0 42.9 76.9 75.8

Table 5. Measured compressive elastic modulus values

Specimen ply

lay - up

Temp

°C

Material

property

Not aged

modulus (GPa)

± Std dev.

Aged without load

modulus (GPa)

± Std dev.

Aged with load

modulus (GPa)

± Std dev.

[0]12 24 E1 135.2 ± 1.1 139.0 ± 3 129.9 ± 4.8

[0]12 -196 E1 142.0 ± 15.0 154.0 ± 5.9 147.9 ± 5.5

[0]12 -269 E1 146.9 ± 11.1 143.1 ± 6.0 139.9 ± 11.4

[±25]3S 24 Ex 63.2 ± 5.7 60.0 ± 4.1 62.8 ± 3.1

[±25]3S -196 Ex 64.1 ± 7.5 72.1 ± 8.0 72.6 ± 2.5

[±25]3S -269 Ex 77.5 ± 17.5 77.3 ± 9.4 69.5 ± 1.7

[45/903/-45/ 03]S 24 Ex 44.3 ± 5.1 46.3 ± 0.9 40.7 ± 7.6

[45/903/-45/ 03]S -196 Ex 60.6 ± 7.3 50.8 ± 4.8 52.2 ± 7.0

[45/903/-45/ 03]S -269 Ex 51.2 ± 4.2 47.6 ± 1.1 53.7 ± 5.0

[90]12 24 E2 11.6 ± 1.7 12.1 ± 0.5 NA

[90]12 -196 E2 11.9 ± 1.1 11.1 ± 0.5 NA

[90]12 -269 E2 12.8 ± 1.1 12.2 ± 01.6 NA

[±45] 3S 24 G12 4.8 ± 2.5 5.3 ± 0.5 4.3 ± 0.1

[±45] 3S -196 G12 6.7 ± 0.8 5.4 ± 0.5 6.0 ± 1.2

[±45] 3S -269 G12 7.6 ± 3.0 5.8 ± 0.3 7.0 ± 2.0

(NA = not available)
16

Table 6. Measured tensile strength values

Specimen ply

lay- up

Temp

°C

Not aged

strength (MPa)

± Std dev.

Aged without load

strength (MPa)

± Std dev.

Aged with load

strength (MPa)

± Std dev.

[0]12 24 1939.8 ± 276.3 1864.8 ± 42.42 1847.8 ± 49.4

[0]12 -196 1277.9 ± 207.5 1567.7 ± 39.0 1536.1 ± 125.8

[0]12 -269 1495.2 ± 326.2 1451.0 ± 275.1 1771.4 ± 186.3

[±25]3S 24 1132.2 ± 76.8 1265.9 ± 57.8 954.9 ± 83.8

[±25]3S -196 911.6 ± 49.5 902.4 ± 208.1 1119.5 ± 51.4

[±25]3S -269 1130.7 ± 15.0 1113.6* 1024.4 ± 209.2

[45/903/-45/ 03]S 24 711.0 ± 14.0 699.9 ± 67.0 842.6 ± 37.9

[45/903/-45/ 03]S -196 585.4 ± 118.7 747.9 ± 57.1 872.4 ± 39.7

[45/903/-45/ 03]S -269 656.5 ± 47.6 731.8 ± 64.4 806.1 ± 2.1

[90]12 24 46.7 ± 1.9 46.6 ± 3.0 NA

[90]12 -196 21.0 ± 3.5 23.5 ± 30.1 63.1 ± 4.9

[90]12 -269 17.5 ± 3.9 11.3 ± 4.7 NA

[±45] 3S 24 162.7 ± 2.3 165.0 ± 0.5 182.0 ± 4.3

[±45] 3S -196 241.6 ± 3.7 249.0 ± 3.0 245.4 ± 7.7

[±45] 3S -269 255.9 ± 8.2 257.0 ± 11.9 257.0 ± 4.7

(NA = not available) Note: Tensile strength of [±45]3S was defined as the initial point of deviation from the

tangent of the nearly horizontal slope of the stress strain curve.

*Only one specimen tested.
17

Table 7. Measured compressive strength values

Specimen ply

lay- up

Temp

°C

Not aged

strength (MPa)

± Std dev.

Aged without load

strength (MPa)

± Std dev.

Aged with load

strength (MPa)

± Std dev.

[0]12 24 857.5 ± 52.5 856.2 ± 102.4 817.4 ± 74.2

[0]12 -196 1124.8 ± 127.6 1167.2 ± 60.9 1115.2 ± 66.2

[0]12 -269 1193.5 ± 100.8 1234.7 ± 44.6 1156.1 ± 61.0

[±25]3S 24 502.9 ± 65.7 467.7 ± 37.2 524.8 ± 43.4

[±25]3S -196 604.2 ± 9.6 709.5 ± 0.3 685.6 ± 87.2

[±25]3S -269 714.8 ± 132.7 749.4 ± 27.2 571.2 ± 61.5

[45/903/-45/ 03]S 24 473.3 ± 121.8 432.6 ± 10.2 530.6 ± 107.5

[45/903/-45/ 03]S -196 375.7 ± 130.1 554.3 ± 29.8 709.5 ± 89.2

[45/903/-45/ 03]S -269 572.2 ± 28.6 619.1 ± 11.9 723.0 ± 113.4

[90]12 24 220.0 ± 10.0 280.0 ± 11.5 NA

[90]12 -196 293.7 ± 16.1 273.7 ± 52.7 NA

[90]12 -269 367.6 ± 35.1 301.9 ± 21.8 NA

[±45] 3S 24 426.7 ± 27.7 398.3 ± 26.4 364.8 ± 12.2

[±45] 3S -196 536.8 ± 73.3 511.0 ± 57.1 530.5 ± 21.4

[±45] 3S -269 578.1 ± 13.8 569.5 ± 10.38 454.6 ± 98.0

(NA = not available)
18

L

w

90°

0°

CL

t

Extensometer

Strain

 gage

CL

Test Temp. w (mm) L (mm)

 24°C 25.4 152.4

-196°C 19.0 254.0

-269°C 19.0 254.0

Figure 1. Tensile specimen schematic. Axial strain gages are mounted back-to-back on

specimen.

Figure 2. Compression specimen schematic. Stacked strain gages are mounted back-to-back on

specimen.

25.4 mm

76.2 mm

Strain

 gages

CL

CL

t
19

a) open view b) side view

Figure 3. Compression-test fixture.

Figure 4. Constant-strain fixture.

Specimen

Grip

413 mm

178 mm

Spring

washers

76.2 mm

Compression

specimen

Bolt holes

Strain gages
20

 

(a) [0]12 laminate

 

(b) [±45]3S laminate

 

(c) [45/903/-45/ 03]S laminate

Figure 5. Photomicrographs of specimen’s edge prior to aging, after aging without load, and

after aging with load.

Not aged Aged without load Aged with load

[0]12

[±45]3s

[45/903/-45/ 03]S

[0]12 [0]12

[±45]3s [±45]3s

[45/903/-45/ 03]S [45/903/-45/ 03]S

1 mm

1 mm

1 mm

0°
21

0

0.2

0.4

0.6

0.8

1

1.2

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile

modulus

[0]1 2

Figure 6. Tensile modulus of [0]12 specimens normalized against the not-aged condition tested at

room temperature.

0

0.2

0.4

0.6

0.8

1

1.2

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile

strength

[0]1 2

Figure 7. Tensile strength of [0]12 specimens normalized against the not-aged condition tested at

room temperature.

[0]12

[0]12
22

0

0.2

0.4

0.6

0.8

1

1.2

1.4

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile shear

modulus

[ ± 45]3 s

Figure 8. Shear tensile modulus of [±45]3S specimens normalized against the not-aged

condition tested at room temperature.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile

strength

[ ± 45]3 S

Figure 9. Tensile strength of [±45]3S specimens normalized against the not-aged condition

tested at room temperature.

[±45]3S

[±45]3S
23

0

0.2

0.4

0.6

0.8

1

1.2

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile 

modulus

Figure 10. Tensile modulus of [45/903/-45/ 03]S specimens normalized against the not-aged

condition tested at room temperature.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

24 ̊C

-196 ̊C

-269 ̊C

Not aged Aged

without load

Aged

with load

Normalized

tensile

strength

Figure 11. Tensile strength of [45/903/-45/ 03]S specimens normalized against the not-aged

condition tested at room temperature.

[45/903/-45/ 03]S

[45/903/-45/ 03]S
24

0

0.2

0.4

0.6

0.8

1

1.2

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

modulus

[0]1 2

Figure 12. Compressive modulus of [0]12 specimens normalized against the not-aged condition

tested at room temperature.

0

0.5

1

1.5

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

strength

[0]1 2

Figure 13. Compressive strength of [0]12 specimens normalized against the not-aged condition

tested at room temperature.

[0]12

[0]12
25

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

modulus

[ ±45 ]3 s

Figure 14. Compressive modulus of [±45]3S specimens normalized against the not-aged

condition tested at room temperature.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

strength

[ ±45 ]3 s

Figure 15. Compressive strength of [±45]3S specimens normalized against the not-aged condition

tested at room temperature.

[±45]3S

[±45]3S
26

0

0.2

0.4

0.6

0.8

1

1.2

1.4

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

modulus

Figure 16. Compressive modulus of [45/903/-45/ 03]S specimens normalized against the not-aged

condition tested at room temperature.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

24°C

-196°C

-269°C

Not aged Aged

without load

Aged

with load

Normalized

compressive

strength

Figure 17. Compressive strength of [45/903/-45/ 03]S specimens normalized against the not-aged

condition tested at room temperature.

[45/903/-45/ 03]S

[45/903/-45/ 03]S
27

Figure 18. Fracture surface of a [45/903/-45/ 03]S laminate after tensile test at room temperature.

Figure 19. Fracture surface of a [45/903/-45/ 03]S laminate after tensile test at –196°C.

Figure 20. Fracture surface of a [±25]3S laminate after tensile test at room temperature.

Figure 21. Fracture surface of a [±25]3S laminate after tensile test at –196°C.

25.4 mm

19.0 mm

25.4 mm

19.0 mm
28

Figure 22. Fracture surface of a [0]12 laminate after tensile test at room temperature.

Figure 23. Fracture surface of a [0]12 laminate after tensile test at –196°C.

Figure 24. Fracture surface of [0]12 laminate after compression test at room temperature.

25.4 mm

19.0 mm
29

Figure 25. Fracture surface of a [±25]3S laminate after compression test at room temperature.

Figure 26. Fracture surface of a [90]12 laminate after compression test at room temperature.

Figure 27. Fracture surface of a [45/903/-45/ 03]S laminate after compression test at room

temperature.
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4. TITLE AND SUBTITLE

Thermal/Mechanical Response of a Polymer Matrix Composite at

Cryogenic Temperatures

5. FUNDING NUMBERS

721-21-10-01

6. AUTHOR(S)

Karen S. Whitley and Thomas S. Gates 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

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 L-18269

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11. SUPPLEMENTARY NOTES

This paper was presented at the 43rd AIAA/ASME/ASCE/AHS/ASC Structural Dynamics, and Materials

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13. ABSTRACT (Maximum 200 words)

In order for polymeric-matrix composites to be considered for use as structural materials in the next generation

of space transportation systems, the mechanical behavior of these materials at cryogenic temperatures must be

investigated. This paper presents experimental data on the residual mechanical properties of a carbon-fiber

polymeric composite, IM7/PETI-5, both before and after aging. Both tension and compression modulus and

strength were measured at room temperature, –196°C, and –269°C on five different laminate configurations.

One set of specimens was aged isothermally for 576 hours at –184°C in an unconstrained state. Another set of

corresponding specimens was aged under constant uniaxial strain for 576 hours at –184°C. Based on the

experimental data presented, it is shown that trends in stiffness and strength that result from changes in

temperature are not always smooth and consistent. Moreover, it is shown that loading mode and direction are

significant for both stiffness and strength, and aging at cryogenic temperature while under load can alter the

mechanical properties of pristine, un-aged laminates made of IM7/PETI-5 material.

14. SUBJECT TERMS

Polymer matrix composite; Cryogenic; Mechanical properties; Tensile properties;

15. NUMBER OF PAGES

34

Compressive properties; Aging; IM7/PETI-5 16. PRICE CODE

 17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

 OF ABSTRACT

Unclassified

20. LIMITATION

 OF ABSTRACT

 UL

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z-39-18

298-102