NASA/TM-- 1998-208806

Thermomechanical Fatigue Durability of

T650-35/PMR- 15 Sheet Molding Compound

Michael G. Castelli

Ohio Aerospace Institute, Cleveland, Ohio

James K. Sutter

Lewis Research Center, Cleveland, Ohio

Dianne Benson

ProTech Lab Corporation, Cincinnati, Ohio

November 1998

https://ntrs.nasa.gov/search.jsp?R=19990009286 2018-05-11T16:56:27+00:00Z
The NASA STI Program Office... in Profile

Since its founding, NASA has been dedicated to

the advancement of aeronautics and space

science. The NASA Scientific and Technical

Information (STI) Program Office plays a key part

in helping NASA maintain this important role.

The NASA STI Program Office is operated by

Langley Research Center, the Lead Center for

NASA's scientific and technical information. The

NASA STI Program Office provides access to the

NASA STI Database, the largest collection of

aeronautical and space science STI in the world.

The Program Office is also NASA's institutional

mechanism for disseminating the results of its

research and development activities. These results

are published by NASA in the NASA STI Report

Series, which includes the following report types:

TECHNICAL PUBLICATION. Reports of

completed research or a major significant

phase of research that present the results of

NASA programs and include extensive data

or theoretical analysis. Includes compilations

of significant scientific and technical data and

information deemed to be of continuing

reference value. NASA's counterpart of peerreviewed formal professional papers but

has less stringent limitations on manuscript

length and extent of graphic presentations.

TECHNICAL MEMORANDUM. Scientific

and technical findings that are preliminary or

of specialized interest, e.g., quick release

reports, working papers, and bibliographies

that contain minimal annotation. Does not

contain extensive analysis.

CONTRACTOR REPORT. Scientific and

technical findings by NASA-sponsored

contractors and grantees.

CONFERENCE PUBLICATION. Collected

papers from scientific and technical

conferences, symposia, seminars, or other

meetings sponsored or cosponsored by

NASA.

SPECIAL PUBLICATION. Scientific,

technical, or historical information from

NASA programs, projects, and missions,

often concerned with subjects having

substantial public interest.

TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific

and technical material pertinent to NASA's

mission.

Specialized services that complement the STI

Program Office's diverse offerings include

creating custom thesauri, building customized

data bases, organizing and publishing research

results.., even providing videos.

For more information about the NASA STI

Program Office, see the following:

• Access the NASA STI Program Home Page

at http://www.sti.nasa.gov

• E-mail your question via the Internet to

help@sti.nasa.gov

• Fax your question to the NASA Access

Help Desk at (301) 621-0134

• Telephone the NASAAccess Help Desk at

(301) 621-0390

Write to:

NASA Access Help Desk

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076
NASA/TM-- 1998-208806

Thermomechanical Fatigue Durability of

T650-35/PMR- 15 Sheet Molding Compound

Michael G. Castelli

Ohio Aerospace Institute, Cleveland, Ohio

James K. Sutter

Lewis Research Center, Cleveland, Ohio

Dianne Benson

ProTech Lab Corporation, Cincinnati, Ohio

Prepared for the

Symposium on Time-Dependent and Non-Linear Effects in Polymers and Composites

sponsored by the American Society for Testing and Materials

Atlanta, Georgia, May 4-5, 1998

National Aeronautics and

Space Administration

Lewis Research Center

November 1998
Acknowledgments

The authors would like to thank Mr. Chris Burke, Mr. Ralph Corner, Ms. Linda Inghram, Ms. Linda McCorkle,

and Mr. Dan Scheiman for their expert technical assistance in the various laboratory facilities at NASA LeRC;

Mr. Dennis Keller for assistance with the statistical planning aspects; Dr. Gary Roberts for helpful technical

discussions and supplying the neat PMR-15 resin, and Mr. Kevin Kannmacher and Mr. Wayne Maple

of AADC for fabricating and supplying the SMC panels.

Trade names or manufacturers' names are used in this report for

identification only. This usage does not constitute an official

endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
Thermomechanical Fatigue Durability of T650-35/PMR-15 Sheet Molding Compound

Michael G. Castelli

Ohio Aerospace Institute

NASA Lewis Research Center

Cleveland, Ohio

USA

James K. Sutter

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio

USA

Dianne Benson

ProTech Lab Corporation

Cincinnati, Ohio

USA

ABSTRACT: Although polyimide based composites have been used for many years in a wide variety of elevated

temperature applications, very little work has been done to examine the durability and damage behavior under

more prototypical thermomechanical fatigue (TMF) loadings. Synergistic effects resulting from simultaneous

temperature and load cycling can potentially lead to enhanced, if not unique, damage modes and contribute to a

number of nonlinear deformation responses. The goal of this research was to examine the effects of a TMF

loading spectrum, representative of a gas turbine engine compressor application, on a polyimide sheet molding

compound (SMC). High performance SMCs present alternatives to prepreg forms with great potential for low

cost component production through less labor intensive, more easily automated manufacturing. To examine the

issues involved with TMF, a detailed experimental investigation was conducted to characterize the durability of a

T650-35/PMR-15 SMC subjected to TMF mission cycle loadings. Fatigue damage progression was tracked

through macroscopic deformation and elastic stiffness. Additional properties, such as the glass transition

temperature (Tg) and dynamic mechanical properties were examined. The fiber distribution orientation was also

characterized through a detailed quantitative image analysis. Damage tolerance was quantified on the basis of

residual static tensile properties after a prescribed number of TMF missions. Detailed microstructural

examinations were conducted using optical and scanning electron microscopy to characterize the local damage.

The imposed baseline TMF missions had only a modest impact on inducing fatigue damage with no statistically

significant degradation occurring in the measured macroscopic properties. Microstructural damage was,

however, observed subsequent to 100 h of TMF cycling which consisted primarily of fiber debonding and

transverse cracking local to predominantly transverse fiber bundles. The TMF loadings did introduce creep

related effects (strain accumulation) which led to rupture in some of the more aggressive stress scenarios

examined. In some cases this creep behavior occurred at temperatures in excess of 150 °C below commonly cited

values for Tg. Thermomechanical exploratory creep tests revealed that the SMC was subject to time dependent

deformation at stress/temperature thresholds of 150 MPa/230 °C and 170 MPa/180 °C.

Introduction

High performance polymeric composites (PMCs) continue to be the focus of a number of research efforts

aimed at developing cost effective, light weight material alternatives for advanced aerospace and aeropropulsion

applications. These materials not only offer significant advantages in specific stiflhess and strength over their

current metal counterparts, but present the further advantage that structures can be designed and manufactured to

eliminate joints and fasteners by combining individual components into integral subassemblies, thus making them

extremely attractive for commercial applications. Of particular interest to elevated temperature applications, are

polyimide matrix based composite materials which exhibit outstanding thermal stability providing for short and

long term uses to 550 and 300 °C, respectively [1]. PMR-15 is one such polyimide which has seen considerable

NASA/TM--1998-208806 1
use in aeropropulsion applications due to its good thermo-oxidative stability, relatively low cost and availability

in a variety of forms [2].

With current emphasis on low cost manufacturing aspects of advanced composite structures, there is

heightened interest on high performance sheet molding compounds (SMCs). SMCs effectively serve to reduce

the costs associated with component production using prepregs, where variable costs are generally associated

with labor, secondary processes and scrap. Using compression molding, SMCs can be molded into complicated

shapes facilitating the use of simple charge patterns, part consolidation and molded-in inserts, which reduce

labor, equipment, and operation costs for preparatory and secondary processes [3]. Specific to the present study

is a carbon fiber reinforced PMR-15 SMC which has been used in a number of elevated temperature static aero

applications, including oil exposed helicopter gearboxes [2] and shrouds for gas turbine engine-inlet housings [4].

The primary objective of the present research was to evaluate the durability of PMR-15 SMC subjected to a

thermomechanical fatigue (TMF) loading spectrum. The TMF mission spectrum was representative of conditions

found at a mid-stage within a gas turbine engine compressor. Researchers at Allison Advanced Development

Company (AADC) and NASA Lewis Research Center investigated the use of PMR-15 SMC as the material

comprising a mid-stage inner vane endwall [5]. Such a component resides in the engine flow path and is

subjected to not only high airflow rates, but also elevated temperatures and pressures. Thus, the application

represents a much more aggressive use of the material than those cited previously and raises obvious concerns

related to the fatigue durability and damage tolerance. A survey of the literature on polyimide SMCs and their

various applications reveals that very little information is available which details the structural durability of such

materials, particularly in the light of more prototypical thermomechanical loading conditions. Therefore, one of

the first goals of the current research was to determine reasonable maximum stress and temperature parameters

for the representative engine mission cycle, so as to take full advantage of the SMC's capabilities.

Toward this end, a detailed experimental investigation was conducted to characterize the fatigue durability

and damage tolerance of a T650-35/PMR-15 SMC subjected to TMF mission cycle loadings. Fatigue damage

progression was tracked on the basis of longitudinal stiffness degradation and strain accumulation, while damage

tolerance was quantified by residual static tensile properties after a prescribed number of TMF missions. The

two parameters of stiflhess and static response were selected not only because of application design and

performance criteria, but due to their commonplace use in the area of damage mechanics and material life

modeling [e.g., see ref. 6-9]. Sufficient tests were conducted for all of the conditions investigated so that the

statistical significance of the results could be assessed. Additional properties, such as fiber distribution

orientation, glass transition temperature, Tg, and dynamic mechanical properties were examined. Detailed

inspections were conducted using optical and scanning electron microscopy to characterize the microstructural

damage. As the TMF cycle promoted damage associated with creep deformation, the creep behavior was

investigated through a series of unique thermomechanical exploratory tests and compared with the response of

neat PMR-15. Emphasis was placed on determining stress/temperature thresholds for time dependent

deformation.

Material Details

Composition and Properties

The chopped carbon fiber polyimide based sheet molding compound examined in this study was

T650-35/PMR-15 SMC [10] supplied by Quantum Composites, Inc., Midland, MI, (QCI 15C, lot # 092343;

comparable to HyComp 310). The carbon fiber, Amoco's T650-35 (3K tow, UC309 sized), was chopped to

25 mm lengths and sprinkled with a randomized orientation (2-D) onto the matrix layer [see ref. 3 for details].

The composite panels had a nominal dimension of 10 x 20 x 0.22 (2 ply) cm and were compression molded at

AADC, Indianapolis, IN, using the conditions presented in Table 1.

The Tg was measured on selected representative panels with an RMS 800 instrument (Rheometrics

ScientifiC') deforming the specimen in torsion at a frequency of 1 Hz, a temperature ramp rate of 5 °C/rain, and

a nominal specimen geometry of 40 x 5 mm (length x width). Tg was calculated by the intercept method using

the storage modulus curve (for more details see ref. [11] pg. 245). The results indicated a dry Tg in the range of

NASA/TM--1998-208806 2
284to 306°Cwhichisgenerallyconsideredtobeunacceptablyowforthismaterialandindicativeof

insufficientpostcuring.Therefore,allpanelsweresubjectedto asecondpostcureto raisetheTg. In an effort to

optimize the Tg and also prevent extensive thermo-oxidative degradation, the effects of three potential secondary

postcure cycles consisting of soaks at 316 °C in 1 atm air were evaluated. Specifically, test samples were

postcured at 316 °C for either 4, 8, or 12 h. Prior to the secondary postcure, all panels were vacuum dried at

140 °C/76 cm Hg for 48 h. The average Tg values after the 4, 8, and 12 h postcures were 334, 339, and 345 °C,

respectively. Given that the anticipated maximum test temperature would be approaching the target temperature

of 316 °C, and desiring to test at a maximum temperature of at least 28 °C (i.e., 50 °F) below the Tg, the

secondary postcure of 12 h at 316 °C was selected and implemented for all of the test panels.

The quality of all the postcured panels was evaluated by nondestructive analysis using ultrasonic C-scan.

Past research in this area has successfully established a correlation between signal attenuation and void content

[12]. Good consistency was found among the panels used in the study and C-scan results generally indicated

void volumes in the range of 0.5 to 1.5 %. Additionally, the void content and fiber volumes were measured by

the methods described in the Test Method for Void Content of Reinforced Plastics (ASTM D 2734) and Test

Method for Fiber Content of Resin-Matrix Composites by Matrix Digestion (ASTM D 3171). The void content

and fiber volume values ranged from 0.9 to 1.7 % and 54 to 61%, respectively. It is important to note that the

fiber volumes are higher than those recommended for this type of SMC, which is generally desired to be in the

range of 50 to 55 %. Fiber volume fractions above 55 % can lead to poor fiber wetting, reducing the overall

mechanical performance of the material 1 (as will be seen in the discussion to follow).

Fiber Orientation Distribution

Carbon fiber orientation distribution has a significant effect on mechanical properties for SMCs. Therefore,

a representative sample was subjected to a detailed examination to quantify the fiber orientation distribution at

various depths into the thickness of the sample. This analysis was performed at ProTech Lab Corp., Cincinnati,

OH. Fiber orientations at three depths (0.7, 1.0, and 1.5 ram) were examined through progressive polishing. The

2

second depth represented the SMC's midsection. The area scanned was 218 mm (approximately square) and

was deemed to be representative of the plane. Fifty-five photographs, each representing approximately

3.95 IIlIn 2, were used to compose the scanned area at each of the three depths. The detailed procedure for

measuring the fiber orientation distribution is given in Appendix A.

The image analysis was performed by analyzing the carbon fiber distribution in 10 degree increments from

0 to 180 degrees, where the 0 and 180 degree orientations coincide and are parallel to the longitudinal specimen

axis (i.e., the specimen loading axis). In theory, the fiber orientation in the plane being examined should be

random for sheet molding compound composites. However, achieving a completely random distribution of

carbon fibers during the prepreg process is very difficult. The image analysis results for depths 1 to 3 suggest

that the carbon fiber distributions are not random, but rather, bi-modal. The fiber distributions through each of

the three depths differ slightly, as might be statistically anticipated and depth 2 is only slightly bi-modal.

However, depths 1 and 3 suggest a strong bi-modal distribution as illustrated Figure 1, where depth 3 is shown in

histogram form. The bi-modal distribution tends toward fiber alignment along the longitudinal axis

(i.e., 0 or 180 +30 degrees), leaving a significantly reduced fiber volume fraction along the 90 +30 degree axis.

The tendency for such a trend is not entirely unanticipated by the manufacturers I given that the fibers are dropped

onto the matrix sheet as it moves along on a conveyor system [3]. That is, the fibers tend to align in the moving

direction, though a minimization of this tendency is desired and sought.

In addition to acid digestion, fiber volumes were determined by image analysis. The fiber volume percent at

each depth was determined by randomly selecting ten digitized photographs at a magnification of 200x and

assuming that the general bundle orientation had no effect on the fiber volume. The results from these fiber

volume analyses are given in Table 2. Note that the values (56 to 61%) correspond well to the fiber volumes

obtained from acid digestion (54 to 61%).

1Personal communication with Dr. Joseph Reardon, HyComp, Inc., Cleveland, OH, 44130.

NASA/TM--1998-208806 3
TestingDetails

All coupon specimens were cut using abrasive water-jet machining; each of the panels (10 x 20 cm) yielded

three samples. After cutting and prior to testing, specimens were dried for 48 h at 140 °C and 76 cm Hg vacuum

and then stored in a desiccator until immediately prior to testing. The specimen geometry was a reduced gage

section dogbone geometry, shown in Figure 2. The relatively large radius of 36.8 cm forming the transition

section has been used extensively in advanced metal matrix composite testing for a variety of laminates [13].

This geometry was successfully extended to high temperature PMC testing, facilitating gage section failures

while avoiding the use of tabs [14].

The mechanical testing system was a closed-loop, servo-hydraulic system manufactured by MTS TM with a

load capacity of 89 kN featuring hydraulic actuated, water cooled, diamond pattern serrated, wedge grips.

Longitudinal strain measurements were obtained using an MTS TM air cooled extensometer with a 1.27 cm gage

length mounted on the edge of the specimen. Specimens were heated using a quartz lamp system and actively

cooled with forced air enabling the rapid thermal cycling necessary for the TMF cycle to be discussed. One of

the formidable difficulties associated with tree TMF testing (i.e., stress and temperature simultaneously dynamic)

of PMCs is that of temperature measurement and control. Thus, a technique and control scheme were developed

specifically for use with quartz lamp heating and PMCs [14]. First, a SMC temperature calibration specimen was

manufactured with a series of internal K-type thermocouples (TCs) at known locations. This specimen was used

to optimize the axial temperature gradients over the gage section to +1% of the nominal desired temperature. A

K-type TC was then embedded in a block of neat PMR-15 and attached to the calibration specimen using a small

metal mounting clip as shown in Figure 3a. Before each test, an externally mounted TC block was calibrated

against the internal TCs of the SMC calibration specimen as illustrated in Figure 3b. The relationship between

the imbedded and external TCs was found to be linear. The block was then mounted on a test specimen in the

precise calibration location and used subsequently to measure and control the temperature. Repeatability of the

relationship between the block and the calibration specimen was verified by removing and installing the setup

several times and examining consistency; acceptable variations of _+2°C were observed.

All static tensile tests were conducted in accordance with the Standard Test Method for Tensile Properties of

Plastics (ASTM D 638) in displacement control with a loading rate of 0.5 ram/rain. The TMF mission cycle used

for this study is shown in Figure 4 (note that all loads are tensile). This generic cycle was determined by AADC

researchers to be representative of the gas turbine compressor mid-stage inner vane endwall application. The

mission consists of eight secondary segments, seven of which represent idle to maximum engine conditions, and

one which represents a redline engine condition where the stress level is prescribed to be a 5 % increase over

maximum. Note that the idle, maximum, and redline stresses and temperatures (_, cyMax,o-RL, and T I, TMax,TRL,

respectively) are not specified, since determining these parameters was part of the research objective. Here, the

goal was to determine a mission where the material was loaded as aggressively as possible, but survive a

minimum of 100 h of mission cycling. As shown, the TMF mission time is approximately 50 minutes; therefore,

100 h of mission cycling corresponded to approximately 120 missions.

Once the TMF mission was specified, the test matrix consisted of testing to the two conditioning states of

50 and 100 h, in addition to the 0 hr TMF state which represents the unconditioned material. Residual tensile

properties were then examined. A minimum of six tests was conducted at each of the TMF conditioning states

satisfying issues regarding statistical significance. The experiments were intentionally designed with respect to

i) order of tests and ii) specimen selection for condition. Given that each panel yielded three samples, one of

each of these was used to examine the three conditioning states. For example, panel-A yielded samples A1, A2

and A3; A1 was used in a 0 h TMF residual test, A2 in a 50 h TMF residual test, and A3 in a 100 h TMF residual

test. Thus, the design of experiments incorporated panel to panel variations for any one TMF state, but not

specimen to specimen variation within a panel. The panel to panel variation was considered more significant.

Several specimens were also tested to 50 and 100 TMF h states for purposes of destructive examination to detail

the state of microstructural damage using optical and scanning electron microscopy.

NASA/TM--1998-208806 4
Results and Discussion

Static Tensile Properties

One of the first issues to resolve on the PMR-15 SMC was the as-manufactured static properties. These

properties would in-turn assist in formulating the maximum and redline parameters specified for the TMF

mission cycle. Shown in Figure 5 are the room temperature (i.e., 22 °C) and 260 °C static tensile behaviors. The

22 and 260 °C tests yielded an average strength ((f_t) of 259 and 230 MPa and an average strain to failure (e f) of

1.0 and 0.47 %, respectively. Two points were worthy of noting. First, the cy_atvalues appeared to fall short of

those advertised by the manufacturers of the comparable SMC, HyComp 310, (Dexter/HyComp) [15] by

approximately 25 % at 22 °C and 17 % at 260 °C. This shortfall in tensile properties was likely due to the high

fiber volume fraction state noted earlier. Further, the properties given in [15] were generated using four and six

ply based materials, as opposed to two plies. The thicker materials will tend to show less scatter, as the through

thickness characteristics become more homogeneous.

Second, note the distinctive "reverse" curvature (concave upward) of the tests performed at 260 °C indicative

of a stiffening effect with increased loading. The curvature becomes markedly noticeable at an approximate

stress/strain of 170 MPa/0.4 %. This effect is likely due to fiber straightening and/or rotation of the fiber

segments into the loading direction, enabled by the viscous behavior of the matrix at 260 °C. This behavior was

somewhat surprising, given that the temperature is significantly below the Tg (~ 337 °C). However, it will be

discussed later in the exploratory creep deformation section that time dependent matrix deformation can be

initiated at markedly low temperature values.

Given the apparently "viscous" response at 260 °C noted in Figure 5, a slightly lower temperature was

examined for potential use as T Maxin the TMF cycle with the intent to avoid this type of deformation behavior as

being typical for the mission cycle. However, the 260 °C target was felt to be a minimum for representing the

TRI_condition. Shown in Figure 6 is the static tensile response at the slightly lower temperature of 232 °C.

Having established, within the first few tests, that the stiffening effect was not manifested at this temperature, a

full complement of tests was conducted (i.e., 6 repeats) to establish statistically meaningful static properties;

these properties are also shown in Figure 6. Note there is a relatively large deviation in properties as one might

expect with a SMC, particularly in view of the fiber orientation distribution information presented earlier. This

variation was especially significant with regards to ef, where the standard deviation (Y) was found to be

approximately 25 % of the mean value. At a minimum, this suggests that the panel-to-panel variation in

properties can be quite large.

Determination of TMF Capabilities and TMF Deformation

With T Maxspecified to be 232 °C and T RI_established at 260 °C, the remaining key parameter to be

determined was cyMax(recall that o-RI_is specified at 1.05cyMax). As an initial estimate, the maximum fatigue stress

was taken as ((f_t _ 3Z) where (f_t is taken at T Max. Thus, the first set of TMF mission parameters featured

cyMax= 162 MPa with the remaining parameters, as shown in Table 3, where the stress values correspond to the

column marked (orM"x= cy_t - 3Z). The specific idle parameters were selected as representative of the application

without regard to the maximum and redline values. Also note that a cyMaxof 162 MPa represents a value

corresponding to 72 % of cy_t, a seemingly modest level in view of the fact that T Maxwas more than 100 °C below

Tg. A representative "cycle 1" deformation response for this TMF mission cycle is given in Figure 7, where the

strain plotted is the total component (i.e., thermal and mechanical). As shown in Figure 7, the TMF cycle

induces a relatively complex deformation behavior. Each of the secondary segments (idle to maximum) prior to

the redline cycle is discernible as the material experienced a notable amount of creep deformation at maximum

conditions. This effect is seen to decrease with progressive secondary segments as the material approaches a

stabilized deformation response. When the redline condition was reached, the creep response was significantly

revived, and then followed by a nominally elastic response corresponding to the remaining secondary segments

of the mission cycle.

NASA/TM--1998-208806 5
Oneobviouseffectwastheamountof strainrecoveryexperiencedwithinthecycle.Forthetestshownin

Figure7, atotalcreepstrainof 0.14%accumulatedatmaximumandredlineconditions.However,after

unloadingfromthefinalsecondarysegment,only0.7%(i.e.,half)remained.Further,if timepermitted,more

strainwouldlikelyhavebeenrecovered.Unfortunately,theredlineconditionof(yMax= 170MPaat260°C

provedtobetoodemandingto allowthefull missiondeformationresponseto stabilize.NotethattheSMCmean

ef(seeFigure6)at232°Cisonlyslightlygreaterthanthatexperiencedafterjustonemissioncycle.With

progressivemissions,acyclicstrainaccumulationeffect(comparabletocreepratchettinginmetallicmaterials

[16])ensuedleadingtocompletefractureof thesample.Althoughthistestisrevealingregardingmaterial

responseundersevereTMFconditions,theresultpreventedanimportantresearchobjective:toestablishresidual

propertiesubsequent to 100h of TMFmissioncycling.Therefore,TMFcyclerevisionswerenecessary.

Shownin Figure8is thefinalmodifiedTMFmissioncyclewhichwasusedfor all residualproperties.This

cyclewasbaseduponstressparametersdefinedby(yMax= (f_t_4Z. Thevaluesfor(yMa_andO"xkwerereducedto

143and150MPa,respectively,withall otherparametersheldconstant(seeTable3). Thiso-Ma_valuerepresents

64%of(f_tat232°C.Representativedeformationresponsesof theSMCtothismissioncyclearegiveninFigure

9,wherebothcycles1and120areshown.Severaldramaticchangesin thedeformationbehaviorarenotedwhen

comparedtothatobservedinFigure7. First,thestress/strainresponseis seentobeessentiallylinearelastic,with

onlymodesttimedependent effectsobservablewithinagivencycle.Further,theviscouseffectsappeartobe

wellstabilizedwithinagivenmissioncyclewiththeredlineconditionhavingaseeminglyinconsequential effect

ontheoveralldeformationbehavior.Notethatthematerialcontinuesto experienceaminordegreeofpermanent

strainratchettingontheorderof 0.05to0.1%strain,whichtendsto accumulateduringtheearlycycles.Having

determinedthattheT650-35/PMR-15SMCwascapableof sustainingthenewmissioncycleloadingfor 100h,

thefull complement of testswasconductedtothe50and100hstatesof conditioningandthencheckedfor

residualproperties.

Residual properties

As indicated earlier, progressive fatigue damage accumulation was tracked through monitoring elastic

stiffness degradation and damage tolerance was quantified on the basis of the static tensile property retention.

Shown in Figure 10 are the elastic stiflhess values measured as a function of accumulated TMF missions. These

values were measured isothermally at 232 °C (i.e., T Ma_)by applying a small elastic load (35 MPa) immediately

after each mission cycle. Given the relatively large deviation in elastic stiflhess (see Figure 6), to facilitate

specimen-to-specimen comparison, this property was normalized with respect to the original value measured

prior to testing. The one major conclusion drawn from this data is that the TMF mission cycling had little to no

effect on elastic stiffness. Of the 12 tests shown, six were cycled to 60 and 6 were cycled to 120 mission cycles,

the data divide essentially equally above and below the original value, indicating the lack of a clear or overriding

trend. Further, the data from Figure 10 indicates that the modest changes observed are incurred early in the

cyclic conditioning (prior to ~ 20 TMF missions) with only minimal exceptions. Subsequent to these early

changes, the material tends to be cyclically neutral with respect to stiffness changes. The early changes may be

indicative of a slight degree of fiber straightening, corresponding to the observed strain ratchetting and/or

substructural damage (discussed later). Also note, in Figure 10, identical symbols represent specimens taken

from the same original panel: one tested to 60 missions and the other tested to 120 missions. No panel-specific

patterns corresponding to elastic stiffness degradation were observed.

Shown in Figure 11 are residual static tensile properties at 232 °C after 50 and 100 h of mission cycling

(six tests for each condition); the 0 hour results from Figure 6 are also shown for comparison. Though the figure

is relatively crowded and difficult to distinguish single tests, it serves to show the significant spread and/or

grouping associated with the data. Also given in Figure 11 are the means ( X ) and standard deviations (Z), for

ultimate strength, strain to failure, and elastic stiffness. By comparing the X and Z values for the two populations

(i.e., post 50 and 100 hour TMF conditions) through an analysis of variance to the 0 TMF h material behavior, it was

determined that there were no significant differences at the 95 % confidence level in any of the three properties. The

only property which exhibits a potential difference is the Z of elastic stiffness for the 100 hour TMF condition,

NASA/TM--1998-208806 6
whichwassignificantat92.5%confidencelevel,butnotatthe95%level.ThisincreasedZoverthatexhibited

priortoTMFcyclingisconsistentwiththetrendrevealedinFigure10,wheresomeof thematerialtendstostiffen

slightlyandsometendstobecomemorecompliant,buttheoverallmeanremainsconstant.Similartothe

progressivestiffnesschangesdiscussedabove,panel-specifictrendsassociatedwithresidualstatictensileproperties

wereexaminedandnonewereevident.

Anotheresidualpropertyexaminedwas Tg -- to check for potential aging. It is well known that elevated

temperature exposures can advance the cure states of polyimides, potentially introducing changes in mechanical

properties. The result indicated no change in Tg after 100 h of TMF mission cycling, which was anticipated given

that the total time spent at 232 and 260 °C was 35 and 5 h, respectively. This observations is consistent with studies

conducted on PMR-15 composites where volume changes were not found to occur after 100 h at 316 °C [17] and Tg

was found to be stable up to 2000 h of exposure at 260 °C [18]. Thus, given the modest exposure times in the

present study, it was assumed that the material did not incur any noteworthy aging.

Microstructural Examination

Microstructural examinations were conducted on several specimens subjected to TMF mission cycling.

Although there was only minimal indication of damage corresponding to the macroscopic property degradation,

there was clear evidence of highly localized damage at the microstructural level which was not evident in the

untested control samples. Damage was predominantly associated with fiber/matrix interface de-bonding at fibers

oriented at angles of 90 _+40 degrees, where the specimen longitudinal axis coincides with 0 or 180 degrees, that is,

the loading direction. At points where such de-bonded fibers intersected with other fibers or fiber bundles having a

different orientation, the cracks appear to have propagated from one bundle to the next, but generally remained

confined to fibers oriented in the range specified above. This common pattern is well illustrated in Figure 12a where

a surface-visible crack is seen in the presence of at least two fiber orientations (upper portion of picture). The crack

likely initiated at the fiber interface oriented at approximately 105 degrees, then connected with the fibers oriented at

approximately 60 degrees and proceeded to cause a localized interface de-bond along this orientation, albeit, in

much less aggressive fashion. As would be expected, whenever a crack at a fiber intersection such as this is noted,

the more dominant crack is generally associated with the fiber orientation closest to 90 degrees. Note that there is a

third dominant fiber bundle orientation visible in the lower portion of the photograph where a starting crack is also

associated; the orientation is approximately 135 degrees.

In general, it appears that when the crack front encountered a bundle with a predominant longitudinal (PL)

orientation, the fiber bundle effectively bridged the crack, leading to cases where the crack propagated "around" the

bundle to another having a predominantly transverse (PT) orientation. This "around" crack path was seen to be

associated either with the cut end or the outer diameter of the PL bundle. An example of the first is shown in

Figure 12b looking into the specimen thickness, where PT oriented fibers appear with a near circular cross-section,

and PL oriented fibers appear elliptical. The crack shown here, connected to the surface, appears to propagate

transverse to the loading direction through the PT bundles, but takes a path around the cut end of the PL bundle.

Thus, in general, a through-thickness view of the SMC such as that shown in Figure 12c revealed the vast majority

of transverse cracks in PT bundles. Further, there is an indication of damage progression from the 50 to the

100 hour TMF state, associated with higher transverse crack densities in the PT bundles, though the number of

samples viewed (two for each case) was not sufficient to make a quantitative assessment. These microstructures do

suggest, however, that given sufficient TMF cycling (significantly more than 100 h), the damage processes would

eventually cause degradations in the macroscopic properties.

Fracture surfaces of residual strength specimens for all conditions (0, 50, and 100 h of TMF) appeared similar

and are well represented by the fractography in Figure 13. The fracture surface was generally transverse to the

loading direction and revealed features dominated by pull-out of PL fibers/fiber bundles and separating of PT fiber

bundles. The pull-out did not appear to be influenced by ply to ply interfaces, which were generally

indistinguishable, both on the fracture surfaces and on the polished mounts. Thus, there was no indication of

ply-to-ply delamination. A number of bundles having PL orientations reveal some degree of fiber fractures, but this

NASA/TM--1998-208806 7
effectappearedsecondarytothepull-outfeature.Thegeneralfeatureswereconsistentwiththosecitedby

BeaumontandSchultz[19]for aroomtemperaturefatiguefailureofSMC-65.

Creep Response

During the process of determining the SMC's structural capabilities under TMF mission cycle loading, it

became obvious that the more aggressive parameters investigated, which led to "premature" specimen fracture,

induced a primary failure mode associated with creep deformation. This was facilitated by the fact that the TMF

mission cycle featured a series of elevated temperature stress holds (2.5 min each), allowing for creep. This

"excessive" creep was unanticipated given that the mission maximum and redline temperatures (232 and 260 °C,

respectively) were significantly below the average Tg of approximately 340 °C, raising concerns associated with the

time dependent deformation response.

Recently, some emphasis has been placed on characterizing the creep response of carbon fiber/polyimide PMCs

because of long-term durability issues [e.g., refs. 17,20,21], however, the work has been dominated by examining

the effects of thermal aging at temperatures much closer to the material's Tg. Further, such investigations have dealt

with continuous fiber reintorced materials, the much more common application. For PMR-15 SMC, the following

questions needed to be addressed: At what stress/temperature threshold levels does creep occur? Further, do

behaviors determined from routine dynamic mechanical loadings (e.g., storage and loss modulus, Tg .... etc.) give

insight into thresholds for time dependent responses? A concise exploratory examination was conducted to address

these critical issues.

Thermomechanical Creep Initiation -- A novel thermomechanical creep deformation test was conducted

where the temperature was ramped at 5 °C/min to the redline temperature with an applied static load. The

temperature ramp rate was felt to be sufficiently slow so as to allow for a quasi static thermal equilibrium through

the thickness of the sample. A preliminary test was conducted under zero load to assess the free thermal expansion

of the SMC during the temperature ramp. The specimen was then loaded at the redline stress level and subjected to

the same temperature excursion. The goal of this test was to identify the temperature at which creep initiated. The

results of this test are shown in Figure 14 for two different conditions consisting of the two redline levels discussed

previously. The creep strain plotted was reduced by subtracting off the time independent elastic and thermal strain

components. As seen in Figure 14, with the nominal o-_ of 150 MPa applied, the creep threshold was determined to

be approximately 230 °C. Further, the experiment revealed that the creep response at the redline temperature of

260 °C, remains relatively modest. Note that 230 °C essentially corresponds to the TMF mission cycle T Max,at

which point the stress is slightly less. These results correspond well to the deformation behaviors observed during

the (-42) TMF mission cycling, indicating that the creep experienced at maximum and redline conditions should be

minimal. The creep threshold temperature drops quickly, however, when the creep stress level is set to the more

demanding o-_ of 170 MPa. Under these conditions creep deformation is found to initiate at approximately 180 °C

which is well below both the T M_'and T RL. Again, the results here are consistent with the excessive creep observed

during the exploratory phase of establishing TMF capabilities; recall Figure 7. The data suggest that if a o-RLof

170 MPa is to be tolerated, then the T_ need be restricted to approximately 230 °C. It is important to note that the

creep deformations experienced were strictly associated with a viscoelastic response (no microstmctural damage)

and verified to be fully reversible/recoverable.

To gain further insight into the creep threshold temperature of the SMC, a comparable test was conducted on the

neat PMR-15 resin with an identical post-cure cycle. The results are given in Figure 15. The thermomechanical

threshold stress for temperatures below 260 °C was found to be ~21 MPa, which is seen to initiate creep at 240 °C.

By examining the tensile strength of the neat PMR-15 it was determined that this stress level corresponds to ~56 %

of the material (f_t at 260 °C. A stress/temperature viscoelastic threshold "lower-bound" was estimated for neat

PMR-15 by Kamvouris, et al. [11] to be approximately 204 °C, with stress levels up to 50 % of clot. This was

concluded because no creep strains were observed over time periods approaching 720 h. The data presented in

Figure 15 agrees and further defines this stress/temperature creep threshold, suggesting that for a temperature of

204 °C as proposed by Kamvouris et al., a stress level approaching 70 % of (f_t would be needed to induce time

NASA/TM--1998-208806 8
dependent deformation.Havingestablishedthisthreshold,astresslevelcorrespondingto64%of the260°CG_t

(i.e.,24MPa)wasappliedtoexamineamacroscopicstresslevel"comparable"tothatbeingappliedtotheSMC

(i.e.,150MPa).TheresultsindicatedthattheneatPMR-15resinexperiencedcreepinitiationatapproximately

215°C. It shouldbenotedthatthe"comparable"stresscomparisonbetweentheneatresinandSMCmaterialsisnot

tosuggestanidenticalstressstate.Unliketheneatresin,thelocalizedmatrixintheSMCissubjectedtoahighly

complexmultiaxialstressstate.Further,useof therespectived atvalues as the normalizing factors reflects only

modest mechanistic relevancy, as these parameters are dictated by significantly different failure mechanisms for the

two classes of materials. However, it remains noteworthy that the PMR-15 resin behavior clearly indicates that at

mid-stress levels (50 to 70 % dat), time dependent deformation is likely to occur at temperatures as low as 200 °C:

a temperature which is nominally 150 °C below commonly cited Tg values. The creep response measured on the

neat matrix confirms the low temperature creep deformation exhibited by the SMC, where the characteristic creep

behavior is dictated by the properties of the PMR-15. Also, if the possibility of interface damage associated with PT

fibers is considered, both the creep threshold and creep rates may be detrimentally affected, giving rise to properties

that are even less desirable than those of the matrix alone. Such an effect of "structural weakening" is relatively

common in cases of transversely reinforced systems, and has been noted with specific reference to elevated

temperature creep behavior of polyimide based composites [22].

Dynamic Mechanical Response -- A final aim of the exploratory SMC creep investigation was to relate the

thermomechanical creep thresholds to dynamic mechanical characterizations routinely performed on polymers and

their composites to determine Tg. Specifically, does the dynamic mechanical response provide quantitative insight

to the thermomechanical creep threshold? It is well known in the polymer science field that temperature dependent

dynamically measured properties (e.g., storage modulus, loss modulus ,Tg .... etc.) are not time independent

properties, but rather vary as functions of both temperature rate and loading rate. This fact, however, is often

overlooked or at least minimized by mechanics researchers, who generally tend to treat and report material

properties, such as Tg, as unique, time independent values (like melting points). Such issues of time dependency

are central to the consideration of a thermomechanical creep threshold.

A series of dynamic mechanical characterization curves for the SMC and neat PMR-15 are shown in

Figures 16 and 17, respectively. The loading rate dependency was examined over the range from 0.1 to 10 rad/sec

using the RMS 800 (RheometricsTM). The temperature sweep rate was maintained constant at 5 °C/rain for all of

the tests and the Tg values were determined by the intercept method from the storage modulus as before. All of

the samples used for this examination were taken from the same panel with the goal of minimizing material

variability, while highlighting the effects of time dependency. The Tgs of both the SMC and neat PMR-15 were

found to vary considerably with loading frequency. The shear loss modulus, G", representative of the imaginary

part of the complex modulus, is usually discussed in the context of its maximum at the c_ transition, by which the

Tg is often defined. However, this "out-of-phase" component is also indicative of sub-Tg transitions, or more

subtle viscous effects pertinent to the discussion of time dependency.

As the temperature progresses from ambient conditions, G" experiences a local minimum and then increases

as the c_ transition is approached. The significant observation in Figures 16 and 17 is that the approximate

temperature at which the increase in G" occurs, decreases with decreasing loading rate, showing a typical

positive strain rate dependence. For clarification, each of the local increase points are designated with an 'T' on

the plots. Though this observation is not uncommon, such behavior is generally not discussed in the context of

defining the onset of time dependent behavior. Note that as the loading rate is decreased, the state begins to

approximate a static load condition, facilitating a comparison with the static load thermomechanical creep

threshold tests. The G" trends indicate that it is not unreasonable to expect time dependent deformation at

temperatures as low as approximately 200 °C, which is in the range of 150 °C below commonly cited Tg values

for PMR-15 and its composites. This result compares well with the findings from the exploratory

thermomechanical creep threshold experiments. The fact that time dependent behavior can be experienced at

such modest temperatures needs to be highlighted when considering the long term use of these materials in

deformation critical applications.

NASA/TM--1998-208806 9
Summary/Conclusions

A detailedexperimental investigationwasconductedtocharacterizethethermomechanical fatigue(TMF)

durabilityanddamagebehaviorof thecarbonfiber/polyimidesheetmoldingcompound(SMC),

T650-35/PMR-15.HighperformanceSMCspresentalternativesto prepregformswithgreatpotentialforlow

costcomponentproduction.TheTMFloadingspectrumwasproposedbyAllisonAdvancedDevelopment

Companytoberepresentativeof agasturbineenginecompressorapplicationwheretheSMCwill beusedfor an

innervaneendwall.Thefiberdistributionorientationwascharacterizedthroughdetailedquantitativeimage

analyses, revealinganon-randomizedfiberdistributionorientation.Mechanicaldamageprogressionwastracked

macroscopicallyonthebasisof propertychanges,in additiontoexaminingotherpropertiesuchasglass

transitiontemperature, Tg, and dynamic mechanical properties. Damage tolerance was quantified through

residual static tensile properties after a prescribed number of TMF missions. Detailed microstmctural

examinations were conducted using optical and scanning electron microscopy to characterize the local damage.

The SMC was further evaluated through a series of exploratory thermomechanical creep tests designed to

determine stress/temperature creep thresholds. Similar thermomechanical tests on the neat PMR-15 resin

permitted a comparison of the creep characteristics to the PMR-15 SMC.

The imposed TMF missions were found to have only a modest impact on inducing fatigue-type damage when

limiting the exposure to 100 h. Microstructural damage was observed after 50 and 100 h of TMF cycling which

consisted primarily of fiber debonding and transverse cracking local to various fiber bundles. However, no

statistically significant degradations occurred in the elastic stiffness, or the residual properties of strain to failure

and ultimate tensile strength. The TMF loadings did, however, promote creep damage and excessive strain

accumulation which led to rapture in more aggressive stress scenarios. The creep behavior was found to occur in

some cases at temperatures more than 150 °C below glass transition temperatures commonly cited for PMR-15

composites. Thermomechanical exploratory creep tests revealed that below 260 °C (the redline temperature

investigated) the SMC undergoes time-dependent deformation at the stress/temperature threshold level of

150 MPa/240 °C. Stress increases above this level allowed for lower temperature thresholds in the range of

180°C. Finally, upon examining relatively slow loading rates during dynamic mechanical testing, trends revealed

by the loss modulus (G") were found to serve as good indicators of the creep threshold temperatures.

NASA/TM--1998-208806 10
APPENDIXA

Procedurefor Quantitative Image Analysis of the Fiber Orientation Distribution

. Metallographic images were captured in gray scale from a video monitor/computer which is

connected to a camera in the Nikon metallograph. A magnification of 50x was used to capture

images across the entire cross section. The images were then imported to NIH Image Analysis

program where the images were "seen as" 320 x 240 pixels (video squares) each with a level of gray

between 0 and 255 (0=white, 255=black).

2. The images were calibrated against a known standard so that measurements could be made in

millimeters.

. The angle tool within the image analysis program was used to measure the predominant orientation

of fibers within a fiber bundle region. The freehand tool was then used to outline this region to

measure its geometric area. A MS Excel spreadsheet was used to record all measurements.

. All fiber bundle regions at each cross section were analyzed. The total area of the fiber bundle regions

within each orientation category was calculated. The fiber volume percent within the fiber bundles was taken

into account to calculate the overall fiber volume orientation distribution.

NASA/TM--1998-208806 11
References

[1] Wilson, D., "Polyimides as Resin Matrices for Advanced Composites," Polyimides, Wilson, D. Stenzenberger, H.D.

and Hergenrother, P.M., Eds. Chapman and Hall, 1990, pp. 187-226.

[2] Stevens, T. "PMR-15 is A-OK," Materials Engineering, Oct. 1990, pp. 34-38.

[3] Allen, P. and Childs, B., "SMC: A Cost Effective Alternative to Prepreg Technology, 38 thInternational SAMPE

Symposium, 1993, pp. 533-46.

[4] Hoft, S.M., "Applying Advanced Materials to Turboshaft Engines," Aerospace Engineering, Vol. 15,

No. 2, 1995, pp. 27-30.

[5] Investigation of Low Cost High Temperature PMC Components, NASA Contract NAS3-97015, 1997.

[6] Hahn, H.T. and Kim, R.Y., "Fatigue Behavior of Composite Laminate," Journal of Composite Materials,

Vol. 10, 1976, pp. 156-80.

[7] Talreja, R., "Stiffness Based Fatigue Damage Characterization of Fibrous Composites" Fatigue of Composite

Materials," Technomic Publishing Company, Lancaster, Pennsylvania, 1987, pp. 73-81.

[8] Yang, J.N. and Liu, M.D., "Residual Strength Degradation Model and Theory of Periodic Proof Tests for

Graphic/Epoxy Laminates," Journal of Composite Materials, Vol. 11, 1977, pp. 176-203.

[9] Reifsnider, K.L. and Stinchcomb, W.W., "A Critical Element Model of the Residual Strength and Life of

Fatigue-Loaded Composite Coupons," Composite Materials: Fatigue and Fracture, ASTM STP 907,

H.T. Hahn, Ed., ASTM, Philadelphia, 1986, pp. 298-303.

[10] Reardon, J.P. and Thorpe, J.D., U.S. Patent No. 5126085 920630, 30 Jun. 1992.

[11] Kamvouris, J.E., Roberts, G.D., Pereira, J.M. and Rabzak, C., "Physical and Chemical Aging Effects in PMR-15 Neat

Resin," High Temperature and Environmental Effects on Polymeric Composites: 2 nd Volume, ASTM STP 1302,

Thomas S. Gates and Abdul-Hamid Zureick, Eds. ASTM 1997, pp. 243-58.

[12] Roth, D.J., Baaklini, G.Y., Sutter, J.K., Bodis, J.R.,, Leonhardt, T. and Crane, E.A., "NDE Methods Necessary for

Accurate Characterization of Polymer Matrix Composite Uniformity, Advanced High Temperature Engine Materials

Technology Program, NASA CP 10146, 1994, paper 11.

[13] Castelli, M.G., "A Summary of Damage Mechanisms and Mechanical Property Degradation in Titanium Matrix

Composites Subjected to TMF Loadings," Thermal-Mechanical Fatigue of Aircraft Engine Materials, AGARD Conference

Proceedings 569, Mar., 1996, pp. 12:1-12.

[14] Gyekenyesi, A.L., Castelli, M.G., Ellis, J.R., and Burke, C.B., "A Study of Elevated Temperature Testing Techniques for

the Fatigue Behavior of PMCs: Application to T650-35/AMB -21," NASA TM- 106927, July 1995.

[15] Dexter Composites Division Data Sheet for HyComp M-300 Series Sheet Molding Compound,

Dexter Corporation (now HyComp, Inc.), Cleveland, Ohio.

[16] Skrzypek, J.J., Plasticity and Creep: Theory, Examples, and Problems, R.B. Hetnarski, Ed, CRC Press, Inc.,

Boca Raton, FL, 1993, p. 128.

[17] Skontorp, A. and Wang., S.S., "High-Temperature Aging, and Associated Microstructural and Property Changes in

Carbon -Fiber Reinforced Polyimide Composites, Proceedings of the 9 th technical Conference of the American Society

of Composites, Technomic Pub., 1994, pp. 1203-12.

[18] Bowles, K.J., Roberts, G.D. and Kamvouris, J.E., "Long-Term Isothermal Aging Effects on Carbon Fabric-Reinforced

PMR-15 Composites: Compression Strength," High Temperature and Environmental Effects on Polymeric Composites:

2ndVolume, ASTM STP 1302, Thomas S. Gates and Abdul-Hamid Zureick, Eds. ASTM 1997, pp. 175-90.

[19] Beaumont, P.W.R. and Schultz, J.M., "Fractography," Failure Analysis of Composite Materials: Delaware Composites

Design Encyclopedia, Vol. 4, 1990, pp. 134-35.

[20] Pasricha, A., Dillard, D.A. and Tuttle, M.E., "Effect of Physical Aging and Variable Stress History on the Creep

Response of Polymeric Composites," Mechanics of Plastics and Plastic Composites, ASME,

MD-Vol. 68/AMD-Vol. 215, 1995, pp. 283-99.

[21] Brinson, L.C. and Gates, T.S., "Effects of Physical Aging on Long-Term Creep of Polymers and Polymer Matrix

Composites, International Journal of Solids and Structures, Vol. 32, No. 6, 1995, pp. 827-46.

[22] Rodeffer, C.D., Maybach, A.P. and Ogale, A.A., "Influence of Thermal Aging on the Transverse Tensile Creep

Response of a Carbon Fiber/Thermoplastic Polyimide Composite," Journal of Advanced Materials,

Jan. 1996, pp. 46-51.

NASA/TM--1998-208806 12
TABLE 1--T650-35/PMR-15 SMC panel processing details.

Imidization Cycle Cure Cycle Postcure Cycle

• Room temperature to 121 °C over

30 min

• Hold at 121 °C for 30 min

• Ramp to 204 °C over 30 min

• Hold at 204 °C for 1 hr

• Preheat mold to 260 °C

• Load imidized part into mold

• Close mold, apply no pressure and

wait 2 minutes

• Apply 4.1 MPa and heat mold to

316 °C

• Hold pressure and heat for 60 min

• Cool while maintaining 4.1 MPa

until mold temp is 204 °C and

unload part

• Heat from room temperature to 249

°C over 3 hr

• Hold at 249 °C for 3 hr

• Ramp to 288 °C over 2 hr

• Hold at 288 °C for 3 hr

• Ramp to 316 °C over 2 hr

• Hold at 316 °C for 12 hr

• Cool to room temperature over 6 hr

TABLE 2--Fiber volume percentages determined by quantitative image analysis.

Polishing Depth Fiber Volume, %

1 60.6

2 56.4

3 55.7

TABLE 3 --Temperature and stress values for TMF mission cycle.

Engine Condition Temperature, °C Stress, MPa Stress, MPa

(O "MAX = G ttlt -3Z) (G MAX = G ttlt -4Z)

Shutdown 26 0 0

Idle 123 36 36

Maximum 232 162 143

Redline 260 170 150

NASA/TM--1998-208806 13
@

6

_" 5

d

E

= 4

i

_, 3

oJ,,_

II

Orientation From Longitudinal Axix, degrees

Figure 1--Fiber orientation distribution found at depth 3 (1.5 mm into thickness)

revealing strong bi-modal trend.

Parallel --

_+0.02

_ _ S 368"3Radius

_ _ 75"00h E_: _

I \ / \1 , I ._......__44.5.________

...... - ...... .........

19"00_-_+0"05

178.0 Ref.

3_

+0.00

- 21.6 -0.20

Figure 2--Specimen geometry for tensile and TMF tests.

NASA/TM--1998-208806 14
PMR-15

Block

Five Internal

Thermocouples

Internal

Thermocouple

Thermocouple

Wire

o

J

"_. / Mounting _"

/ _ Clip / =

SMC .=

Specimen

300

200

100

/

/

/

/

/

/

/

/

/

/

/

/

CUT-A-WAY 0 t ' t ' t t

VIEW 0 50 100 150 200 250

Block Temperature, °C

(a) (b)

Figure 3--Temperature measurement and control scheme showing (a) the internal and external

block thermocouple configuration and (b) representative calibration data.

i

Redline

Maximum

+

f..

.=

f..

_. Idle

[-

Shutdown

Gas Turbine Engine Condition

I

I

I I I

I I' I

_-9

I

I

I I

I I

II

lq'

I

&

' I ' I ' I ' I

0 10 20 30 40 50

Time, min

Redline

Maximum

+

f..

Idle

Shutdown

Figure 4--TMF mission cycle representative of a gas turbine engine compressor

inner vane endwall application.

NASA/TM--1998-208806 15
200

100

0

0.000

238 246

260 °C X

." . 22 °C

I

273

Advertised 22 °C (_ult = 324 MPa

Advertised 260 °C (_ult = 269 MPa

0.002 0.004 0.006 0.008 0.010 0.012 0.014

Strain, m/m

Figure 5--Static tensile behavior of as-manufactured T650-35/PMR-15 SMC at 22 and 260 °C.

cJ)

300

200

100

I I I

0 I I I

0.000 0.004 0.008 0.012

Strain, m/m

Statistics

Tensile Strength, MPa

Mean: 225

Std. Dev.: 21

95% Conf.: 22

Strain To Failure, 10.2

Mean: 0.923

Std. Dev.: 0.229

95% Conf.: 0.240

Elastic Modulus, GPa

Mean: 31.2

Std. Dev.: 3.7

95% Conf.: 3.8

Figure 6--Static tensile behavior and property statistics of as-manufactured

T650-35/PMR-15 SMC at 232 °C.

NASA/TM--1998-208806 16
200

150

100

50

' ' ' I ' ' ' I ' ' ' I ' ' ' I

Creep deformation 1 1

Cycle 1- Maximum --'/'-_ _ 7 Redline "-]

I

Z | | |

0 I I I I

0.000 0.002 0.004 0.006 0.008

Total Strain, m/m

Figure 7--Stress/strain deformation behavior of T650-35/PMR-15 SMC with TMF

mission cycle parameters established by c_Max = (yult _ 3_.

Redline --

250 --

Maximim--

200

o

d

t.

" 150

t.

Idle--=-

E

100 -

50

' I ' I ' I ' I '

2L

----O--- Temperature ----O--- Stress

- ' i ' i ' i ' i

0 10 20 30 40 50

Time, min

200

150 Redline

--Maximim

100

<D

t.

50

-- Idle

Figure 8--Finalized TMF mission cycle; parameters are based upon (yMax = (yult _ 4Z.

NASA/TM--1998-208806 17
150

100

5O

0

0.000

' ' ' I ' ' ' I ' ' ' I '

Cycle 1 _aximum --J<dli_ne -

Idle __ __

t t t t t t t t t t

0.002 0.004 0.006

Total Strain, m/m

Figure 9--Stress/strain deformation behavior of T650-35/PMR-15 SMC with TMF

mission cycle parameters established by o Max = (yult _ 4_.

O

eq

ee)

eq

1.2

1.1

!

._ 1.0 1

0.9

0.8

tO

z

' ' ' I ' ' ' I ' ' ' I '

0 0 0 0

0 0 0 0 0 0

0 0 0 0

_oo-oooooo_o []

" togo2

same symbol => specimens from same original panel

' ' I ' ' ' I ' ' ' I '

40 80 120

TMF Mission Cycle

Figure 10--Normalized elastic stiffness measurements taken isothermally at 232 °C

during TMF mission cycling.

NASA/TM--1998-208806 18
¢D

300

200

100

0

0.000

I I I

• j

f-./ /----'-f t..

/./

..- . _/t'- -- 0 hours TMF

/z- -- -- 50 hours TMF

...... 100 hours TMF

I I I

0.004 0.008 0.012

Strain, m/m

Statistics

Tensile Strength_ MPa

0 50

Mean: 230 243

E: 21 21

-2

Strain to Failure, 10

100

234

28

0 50 100

Mean: 0.923 0.940 0.830

Z: 0.229 0.244 0.265

Elastic Modulus_ GPa

0 50 100

Mean: 31.2 29.0 31.5

E: 3.7 7.5 8.8

Figure 11--Residual static tensile properties at 232 °C subsequent to

50 and 100 hours of TMF mission cycling.

NASA/TM--1998-208806 19
iiiiiiiiiiiiiiiiii

(a)

iiiiiiiiiiil!iiii_iii!i!iiii

iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii

(Y ( > (Y (b)

(Y < > (_ (c)

Figure 12--SEM (a) and optical (b) & (c) microscopy revealing TMF damage in the SMC.

(a) surface cracking associated with de-bonds; (b) view into specimen thickness revealing a

transverse crack proceeding around the end of a predominantly longitudinal bundle; (c) view into

specimen thickness showing transverse cracking associated with transversely oriented bundles.

NASA/TM--1998-208806 20
Figure13--Typical fracturesurfaceof T650-35/PMR-15SMCrevealingextensive

fiber andfiberbundlepull-out.

400 ' ' I .... I .... I .... I .... I ....

T = 5 °C/min

300 _aL = 150 MPa

"_ 200

,_ (_RL = 170 MPa

r_

z. 100

Creep initiation points _ l

, , I .... I .... I , , ; , I .... I ....

50 100 150 200 250 300

Temperature, °C

Figure 14--Thermomechanical creep threshold of T650-35/PMR-15 SMC at stress levels

of 170 MPa ((yult _ 3Z) and 150 MPa ((yult _ 42).

NASA/TM--1998-208806 21
6OO

:_. 400

°m

I.

= 5 °C/min

200

0

O

-200 , I

300

ff = 24 MPa O

(~ 64 % of I_ ult)

9:

Creep initiation points _ T

, , I , , , , I , , , , I , , , , I , , , , I , , ,

50 100 150 200 250

Temperature, °C

Figure 15--Thermomechanical creep threshold of neat PMR-15 at stress levels

of 21 and 24 MPa, N56 and 64 % of the 260°C (yult, respectively.

m

m

1000 =

100

10

1

0

__\ ...... 10 rad/sec, "'. -- - 1.0 rad/sec,

".. -- 0.1 rad/sec,

_"'.o ."

/

Loss Modulus, G" ./ /.." \ \'5. /

-_-_ ........ .f • _..

• / %-/"/

T = 5 °C/rain

local Increase points

.... I .... I .... I .... I .... I

100 200 300 400 500

Loading Rate

T = 346 °C

g

T = 344 °C

g

T = 334 °C

g

Temperature, °C

Figure 16--Loading rate dependent dynamic mechanical response of T650-35/PMR-15

SMC indicating dramatic shifts in Tg and G" trends suggesting rate dependency at

temperatures near 180 °C.

NASA/TM--1998-208806 22
100

10

0.1

Loading Rate

10 rad/sec, Tg = 322 °C

B _ 1.0 rad/sec, Tg = 316 °C

0.1 rad/sec, Tg = 305 °C

Loss Modulus, G "

ti

_ _ .I'" • . . . .... • •

/

= 5 °C/min local Increase points

.... I .... I .... I .... I .... I

0 100 200 300 400 500

Temperature, °C

Figure 17--Loading rate dependent dynamic mechanical response of neat PMR-15 corroborating

dramatic shifts in Yg and relatively low temperatures for rate dependency.

NASA/TM--1998-208806 23
Form Approved REPORT DOCUMENTATION PAGE

OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

November 1998 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Thermomechanical Fatigue Durability of T650-35/PMR- 15 Sheet

Molding Compound

6. AUTHOR(S)

Michael G. Castelli, James K. Sutter, and Dianne Benson

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

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135- 3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546- 0001

WU-523-21-13-00

8. PERFORMING ORGANIZATION

REPORTNUMBER

E-11390

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1998-208806

11. SUPPLEMENTARY NOTES

Prepared for the Symposium oi] Time Dependent and Non Linear Effects in Polymers and Composites sponsored by the American

Society for Testing and Materials, Atlanta, Georgia, May 4 5, 1998. Michael G. Castelli, Ohio Aerospace Institute, 22800 Cedar Point

Road, Cleveland, Ohio 44142; James K. Sutter, NASA Lewis Research Center; Dianne Benson, ProTech Lab Corporation, Cincinnati,

Ohio. Responsible person, Michael G. Castelli, organization code 5920, (216) 433 8464.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 24 Distribution: Nonstandard

This publication is available fiom the NASA Center for AeroSpace Information, (301) 6214)390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Although polyimide based composites have been used for many yem5 in a wide variety of elevated temperature applications, very little work has

been done to examine the durability and damage behavior under more prototypical thennomechanical fatigue (TMF) loadings. Synergistic effects

resulting fi'om simultaneous temperature and load cycling can potentially lead to enhanced, if not unique, damage modes and conuibute to a

number of nonlinear deformation responses. The goal of this research was to examine the effects of a TMF loading spectrum, representative of a

gas turbine engine compressor application, on a polyhnide sheet molding compound (SMC). High performance SMCs present alternatives to

prepreg forms with great potential for low cost component production through less labor intensive, more easily automated manufactming. To

examine the issues involved with TMF, a detailed experimental investigation was conducted to characterize the durability of a T650 35/PMR 15

SMC subjected to TMF mission cycle loadings. Fatigue damage progression was tracked ttu'ough macroscopic deformation and elastic stiffness.

Additional properties, such as the glass transition temperature (Tg) and dynamic mechanical properties were examined. The fiber disuibution

orientation was also characterized ttu'ough a detailed quantitative hnage analysis. Damage tolerance was quantified on the basis of residual static

tensile properties after a prescribed number of TMF missions. Detailed microstructural examinations were conducted using optical and scmming

electron microscopy to characterize the local damage. The hnposed baseline TMF missions had only a modest impact on inducing fatigue damage

with no statistically significant degradation occurring in the measured macroscopic properties. Microstructural damage was, however, observed

subsequent to 100 h of TMF cycling which consisted primarily of fiber debonding and transverse cracking local to predominantly transverse fiber

bundles. The TMF loadings did introduce creep related effects (strain accumulation) which led to rupture in some of the more aggressive stress

scenarios examined, ha some cases this creep behavior occurred at temperatures in excess of 150 °C below commonly cited values for Tg.

Thennomechanical exploratory creep tests revealed that the SMC was subject to time dependent deformation at stress/temperature ttu'esholds of

150 MPa/230 °C and 170 MPa/180 °C.

14. SUBJECT TERMS

Polymer matrix composite; Fatigue; Damage; Residual strength; Tensile properties

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

29

16. PRICE CODE

A03

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

Standard Form 298 (Flev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102