NASA Technical Memorandum 109095

38-z/

Preliminary Evaluation of HybridTitanium Composite Laminates

J. L. Miller, D. J. Progar, W. S. Johnson, and T. L. St. Clair

Langley Research Center, Hampton, Virginia

(NASA-TM-I09095) PRELIMINARYEVALUATION OF HYBRID TITANIUM

COMPOSITE LAMINATES (NASA. Langley

Research Center) 31 p

N94-29472

Unclas

April 1994

G3/24 0003821

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National Aeronautics and

Space AdministrationLangley Research CenterHampton, Virginia 23681-0001

https://ntrs.nasa.gov/search.jsp?R=19940024969 2020-06-14T07:40:35+00:00Z

PRELIMINARY EVALUATION OF HYBRID TITANIUM COMPOSITE

LAMINATES

J.L. Miller 1, D. J. Progar 2, W. S. Johnson 3 and T.L. St. Clair 2

Abstract: In this study, the mechanical response of hybrid titanium composite laminates (HTCL)

was evaluated at room and elevated temperatures. Also, the use of an elastic-plastic laminate

analysis program for predicting the tensile response from constituent properties was verified. The

improvement in mechanical properties achieved by the laminates was assessed by comparing the

results of static strength and constant amplitude fatigue tests to those for monolithic titanium sheet.

Two HTCL were fabricated with different fiber volume fractions, resin layer thicknesses and

resins. One panel was thicker and was poorly bonded in comparison to other. Consequently, the

former had a lower tensile strength, while fewer cracks grew in this panel and at a slower rate.

Both panels showed an improvement in fatigue life of almost two orders of magnitude. The model

predictions were also in good agreement with the experimental results for both HTCL panels.

Keywords: adhesive bonding, graphite fibers, isothermal fatigue, laminated titanium,crack growth.

INTRODUCTION

Both military and commercial aircraft are being designed to fly faster and to last longer than

ever before. This requires structural materials that are capable of operating at higher temperatures

yet exhibit improved fatigue and damage tolerance properties. One material system that merits

attention is a hybrid titanium and polymeric matrix composite laminate.

1National Research Council Resident Research Associate

2Senior Scientist

3Senior Research Engineer

In themid 1960'sKaufman[1] showedthatthefracturetoughnessof adhesively

laminatedaluminumplieswasimprovedin comparisonto thatof anequivalentmonolithicplate

dueto theindividualpliesfailing in aplanestressstate.In themid70'sJohnsonandcolleagues

[2, 3] showedthatsignificantimprovementsin fatigueandcrackgrowthresistancecouldalsobe

realizedby adhesivelylaminatingthinaluminumpliestogether.In theearly 80'sJohnson

extendedthiswork to showthatadhesivelylaminatedtitaniumpliesimprovedfracturetoughness

by almost40%,increasedfatiguelife by anorderof magnitude,andsloweddownthough-the-

thicknesscrackgrowthratesby 20%[4]. In themid 80'sresearchersatDelft Universityand

AlcoadevelopedARALL (AramidReinforcedAluminumLaminates)wherenotonly werethin

pliesof aluminumadhesivelylaminatedtogether,butaramidfiberswerealsoincludedin the

bondline[5,6]. Theinclusionof thefibersofferedevenmoreimprovementin mechanical

behaviorof thelaminate,primarilyduetoreductionsin crackgrowthratesresultingfrom afiber

bridgingphenomenon.Someof theARALL laminateswereevenprestrainedsothealuminum

plieswerein astateof residualcompression,furtherimprovingfatigueproperties.Similar

improvementshavebeenofferedbyGLARElaminates[7] thatareessentiallythesameconcept,

butglassfibersreplacethearamid.Thus,thehistoryof laminatedmetalshasshowndefinite

mechanicaladvantagesthatcantranslateto weightsavings.Theselaminatedaluminumpartsare

nowflying onseveralcommercialandmilitary aircraft.

Thispreliminarystudywill builduponpastexperienceandapply thelaminated/hybrid

technologyto highertemperaturemetalsandadhesivesin hopesof demonstratingthatthese

systemswill beusefulin highspeedaircraft. Two unidirectionalhybrid titaniumcomposite

laminates(HTCL) weremadewith differentfiber volumefractions,resinlayerthicknessesand

resins.Fromthelimitedsupplyof material,theHTCL weretestedstaticallyto failureand

cyclically in fatigue. Theresultswerecomparedto eachotherandto ansimilar thicknesspiece

of monolithictitaniumto assesslaminationaffects.A laminateanalysiscodewasusedto predict

thehybrid laminateresponsebaseduponconstituentpropertyinputandto verify is applicability

tomodelthemechanicalresponseof this typeof materialsystem.

MATERIALS

TheHTCL werefabricatedas7.6cm x 17.8cmpanelswithTi-6A1-4V alloy for the

metalportionof thelaminatedpanel.Prior to laminating,thetitaniumalloywasgiventhe

following Pasa-Jell107"surfacetreatment:

o

2.

.

4.5.6.7.8.9.10.

Washed with acetone

(Panel 1) Abraded surface with 150 grit sandpaper(Panel 2) Gritblasted with 120 aluminum oxide gritWashed with methyl alcoholBrushed Pasa-Jell 107 on surfaces, let set for 10 min

Repeat 4.Washed with hot running tap water, then cold tap waterPlaced in ultrasonic dimineralized water bath for 15 minRemoved and air dried in laboratory hood for 10 minDried in forced-air oven at 100°C for 10 minPrimed within two hours of treatment

Two HTCL were prepared using two different polyimide adhesive materials. The prepreg

(adhesive tape) was made in-house (Polymeric Materials Branch) on a multipurpose prepregging

machine.

Panel 1 was prepared using LARCrM-IA 4 mole % offset [8 mole % phthalic anhydride

(PA) end cap] in N-methyl-2-pyrrolidone (NMP) [8] and IM7 graphite fibers as a unidirectional

powder coated prepreg, approximately 0.33 mm thick. The LARCrU-IA was obtained from

Imitec. The primer, which was prepared in-house, was applied to the titanium as a 10.6 wt. %

LARC_-IA 3 mole % offset (6 mole % PA) amic acid solution in gamma-butyrolactone. The

solution was brushed on the titanium and staged in a forced-air oven for 15 min each at 50", 100 °

and 225°C.

Panel 2 was prepared with LARCrU-IAX [a version of LARCrU-IA in which the

backbone was modified slightly with 10 mole % of a more rigid diamine, p-phenylenediamine, to

improve solvent resistance] amic acid (again a 4 mole % offset as with LARCVM-IA), 30 wt. %

solids solution in NMP with a viscosity of 30,400 cp and IM7 fibers solution coated as a

unidirectional prepreg approximately 0.18 mm thick. The titanium sheets were primed with a 7.5

*Trade name for titanium surface treatment available from Semco, Glendale, CA, USA

wt % solution of LARCrU-IAX in NMP and air dried for one hour after which they were placed

in a forced-air oven and heated for one hour each at 150 ° and 230°C.

MANUFACTURING and PROCESSING

The laminate assembly was arranged in an open-ended matched-die mold by alternating

six layers of titanium sheets with five layers of adhesive tape. TX1040 release cloth was placed

on the bottom and top of the laminate assembly to prevent the laminate from sticking to the

mold. The mold had previously been coated with a mold release agent. A hydraulic press with

30.5 cm x 30.5 cm heated platens and load cell was used to process the laminates.

The initial attempt to process Panel 1, under 0.10 MPa at 343°C for one hour, did not

appear to provide adequate flow in the adhesive, as indicated by the lack of any excess adhesive

polymer outside the laminate and the thickness of the laminate. Therefore, the laminate was

further processed at a pressure of 13.8 MPa at 371 °C for one hour. The average adhesive

thickness per layer was determined to be 0.28 mm.

Panel 2 was processed slightly differently than Panel 1. The laminate assembly was

arranged in the bottom part of the die mold, placed in a forced-air oven without the top part of

the die mold, then staged to remove some of the volatiles prior to the bonding procedure. The

laminate assembly was heated for one hour each at 150 °, 175 ° and 230°C. The laminate

assembly with the top part of the die mold was then placed in the press where a pressure of 13.8

MPa was applied while heating the assembly to 350°C in approximately one hour. The pressure

and temperature was then held for one hour. The average adhesive thickness for Panel 2 was

0.12 mm. Schematic diagrams indicating the thickness variations are shown in Figure 1.

ANALYTICAL METHODS

The AGLPLY code [9], developed originally to analyze metal matrix composites, was

used to evaluate its potential for predicting the HTCL's response based on constituent input. The

The use of trade names in this paper does not constitute endorsement, either expressed or implied, by the NationalAeronautics and Space Administration.

4

programperformsanelastc-plasficanalysisof symmetriccompositelaminatedplatesunderin-

planemechanicalloads.Thelaminapropertiesarecalculatedvia thevanishingfiberdiameter

(VFD) modelwhichassumesa ruleof mixturescontributionof thefiber to themodulusin the

longitudinaldirectionwhile offeringno transverseconstraintby thefiber [9]. AGLPLY

computes the overall laminate elastic moduli, the local fiber and matrix stresses and strains in

each ply as well as the overall laminate strains for the entire elastic-plastic loading regime. This

program has proven successful in predicting the mechanical response of metal matrix composites

in numerous studies [ 10-12]. Table 1 displays the constituent properties used to predict the

composite properties.

EXPERIMENTAL PROCEDURE

Tests were conducted on both monolithic Ti-6A1-4V and the two HTCL. From the

limited supply of material manufactured, six HTCL specimens were fabricated, three from each

panel. One specimen from each panel was tested statically to failure to examine the tensile

behavior and two were tested cyclically in fatigue, one at room temperature and the other at 177 °

C. To assure a valid comparison between the monolithic material and the titanium in the HTCL,

Rockwell A hardness tests (RHA) were conducted on both materials. Since the lamination

process involves elevated temperatures, it was of interest to determine whether the titanium was

affected by this processing. The titanium laminae and the monolithic sheet were both 68 RHA;

thus it is assumed that the titanium laminae are the same as the monolithic material, only varying

in thickness.

The fiber volume fraction in the laminates were determined by digital image analysis of

polished (2.5 _tm finish) cross sections using an image processing system. A series of digital

image scans were made over the adhesive layers and the average volume fraction of fibers was

determined from the area fraction of the fibers measured in each scan. Twenty scans were made

for Panel 1 and ten scans were made for Panel 2. Greater variations in fiber distribution in Panel

1 necessitated more scans.

Theroomtemperaturetensileresponsewasevaluatedusing19mmwidestraight-sided

specimens.Thespecimensweretestedto failure in laboratoryair ata loadingrateof 150N/sec

ona250kN servo-hydraulictestframe,equippedwith hydraulicgrips; theappliedgripping

pressurewas14MPa. Axial strainwasmeasuredon thesurfaceof thespecimenusingastrain

gageextensometerwith a25mm gagesection.ResultswererecordedonanX-Y plotter.

Constantamplitudefatiguetestswereconductedonstraight-sidedspecimens,19mm

wide(W), containinga6.4mmdiameter(d) centerhole,rendingd/W--0.33.Theholesin the

monolithictitaniumwereproducedbyelectronicdischargemachining,while theHTCL were

drilled. Theinteriorof theholesin bothmaterialshadthesamesmoothsurfacefinish. The

surfaceof thespecimenswerepolishedalongthegagesectionto a3 lim finish. All testswere

conductedona250kN servo-hydraulictestframeatacyclic frequencyof 10Hz andanR=0.1.

TheHTCL wereevaluatedfor their fatigueresistancebyapplyingequivalentload-to-weight-

ratiosascomparedto themonolithicmaterial.Thedensitiesof thematerialswere2.9,3.4and

4.4g/ccfor Panel1,Panel2 andthemonolithicsheet,respectively.Variationsin thethickness

of theadhesivelayersresultedin differing laminatedensities.Theresultingappliedloadsfor the

laminateswere0.67and0.78timestheappliedloadin themonolithicmaterial. Elevated

temperaturefatiguetestswereperformedat 177°Cin aconvectionovenmountedon thetest

stand.A K-typethermocouplepositionedonthesurfaceof thespecimen,awayfrom thehole,

indicatedthetesttemperature.

Crackgrowthmeasurementsweremadeduringtheroomtemperaturetestingby

replicatingthecrackedregionneartheholewith anacetatefilm. A mirror imageof thecrackis

createdin thefilm whichcanthenbeexaminedmicrospcopicallyandmeasured.Boththefront

andbacksurfacesof thespecimenwerereplicated.This techniquewasnotapplicablefor the

elevatedtemperaturetests.Duringthesetests,cracklengthsweremeasuredwith a longfocal

lengthmicroscopeandvideoimagingsystemby viewingthespecimensurfacethroughawindow

in theoven.Thus,crackmeasurementsweremadeon thefront surfaceonly for theelevated

temperaturetests.

6

RESULTSAND DISCUSSION

Average fiber volume fractions were determined for each panel. Panel 1, the thicker

laminate, had 44.0!-_8.6% fibers while Panel 2 had 63.8:t:5.0% fibers in the adhesive layers.

Optical micrographs of the polished cross sections (Figure 2) show significant variations in the

fiber distribution in the two laminates: the fibers were much more uniformly distributed

throughout the adhesives layers in Panel 2, whereas many resin rich areas existed in Panel 1.

This non-uniformity contributes to the larger variation in the measured average volume fraction

for Panel 1.

The results of room temperature tension tests on the monolithic titanium and the HTCL

are shown in Figure 3. Due to the limited supply of HTCL, elevated temperature tensile

response was not evaluated. Panel 1 performed poorly in tension, failing at a lower strain than

the monolithic titanium and showing no improvement in elastic modulus. Panel 2 showed a

significant improvement in tensile strength, approximately doubling the strength of the titanium

alone, as well as an improved elastic modulus. The fracture modes of the two HTCL varied

greatly as well (Figure 4). Panel 1 suffered severe delamination during the test, a result of the

poor bonding between the adhesive and the titanium layers. In Panel 2, the titanium plies

showed more plastic deformation and the amount of delamination was less. In Panel 1 the

delaminations and fractured plies extended throughout the gage length of the specimen, while in

Panel 2, the damage was more localized.

The elastic-plastic laminate analysis model AGLPLY was used to predict the stress-strain

response of the various materials used in this study. Prior to applying the model to the HTCL,

the applicability of the model to predict behavior of the individual lamina was verified. Figure 5

displays the model predictions bounding the experimental data [13] for the tensile response of the

IM7/LARC TM-IA polymer composite over the applicable range of fiber volume fractions for the

composites tested. For the HTCL the predictions over the range of measured fiber volume

fractions are compared directly with the experimental data for Panels 1 and 2, in Figures 6 and 7,

respectively. Overall the predictions were in good agreement with the experimental results. For

bothpanels,thepredictionusingthelowestfiber volumefractiongivesthebestrepresentationof

theexperimentalresultsin eachcase:boththeelasticmodulusandthetensilestrengtharein

excellentagreementwhileyielding in thetitaniumpliesoccursasindicatedby thechangein

slope,atapproximately820and900MPalaminatestressfor Panel1andPanel2, respectively.

Figure8comparestheexperimentalresultsto theAGLPLY predictionsfor themonolithic

titaniumandtheHTCL.Theshapeof thepredictioncurvescloselyresembletheexperimental

results,with thechangein sloperepresentingtheyieldingthatoccurswhenthestressin the

titaniumreaches820MPa. Thepredictionsalsopredicttheelasticmodulifairly well asshown

in Figure8.

Resultsof theconstantamplitudefatiguetestsoncenterholespecimensof themonolithic

titaniumandtheHTCL areshownin Figure9. For themonolithictitanium,increasingthe

temperaturereducestheoverallfatiguelife of thetitanium. FortheHTCL, all thespecimens

testedwereableto withstand106cycleswithoutfracturingthelaminatecompletely,bothat

roomandelevatedtemperature.However,thepanelsweretestedat anequivalentload-toweight

ratiowhichresultsin differentappliedstressesfor theindividualpanelsdueto differencesin

thicknessanddensity.Figure10showstheresultsin termsof theequivalentload-to-weightratio

tested.TheHTCL panelsshowedamarkedimprovementin fatiguelife of almosttwoordersof

magnitudeatbothroomandelevatedtemperatures.Furthercomparingtheresultson thebasisof

equivalentstress-to-weightratios,asshownin Figure11, illustratesthebenefitof the lower

densityof theHTCL: thethinner,lighterweightPanel2, hasagreaterstressperunit weightthan

themonolithictitanium,yet hasa longerfatiguelife. Similarresultsareshownfor Panel1.

Crackpropagationwasmonitoredonall HTCL specimens.Cracksinitiatedat thehole,

growingthroughtheexteriortitaniumpliesandthenlinking with theinteriorplies. Figure12

showstheedgereplicasof aspecimenfromPanel2, takenat differentstagesduringthetest,

illustratingthis pattern.Crackshadpropagatedthroughall thetitaniumpliesat 890,000cycles,

yet thespecimenremainedwholeandcontinuedtocarry loaduntil thetestwasstoppedat a

million cycles. Oncethetitaniumwascompletelycracked,thefiberscarriedall theload. The

8

stressin thefibers(basedon theaveragevolumefraction)with only thefiberscarryingload,was

calculatedto be445MPafor Panel1and495MPafor Panel2. Thesefiber stressesaremuch

lessthanthefiberultimatestrength(5310MPa),andtherefore,thefiberswouldcontinueto carry

theappliedloadindefinitelywithoutfracturing. If thesameloadswereappliedto themonolithic

titanium,failurewouldbeexpectedsincethesestressesaregreaterthantheendurancelimit of

thealloy (138MPa)asshownin Figure9.

Typically in metallicmaterials,thecrackpropagationrateincreaseswith crackgrowth

exponentially. For laminatedsheets,thecrackgrowthis fasterinitially, butslowsasthecrack

encountersthe interfacesbetweenlayers,yieldinganoveralllongerfatiguelife, asshownby

Johnson[4].However,for theHTCL testedin thisstudy,thecrackgrowthrateremainsconstant,

asindicatedby thelinearcurvesin Figures13-15.Sincecrackgrowthwasnotmonitoredin the

monolithictitanium,predictionsweremadeusingFASTRAN-II, a fatiguecrackgrowth

structuralanalysisprogram[14]. This program has been used extensively to predict fatigue

crack growth from available crack growth data on specific alloys. An assumed initial crack size

of 1 mm was used for these predictions to illustrate the extent of the reduction in the crack

growth rate in the HTCL as compared to monolithic titanium. These results are also shown in

Figures 13 and 14. The predictions indicate an increasing rate rather than the constant rate of

growth shown by the HTCL. The reduced crack growth in the HTCL is attributed to fiber

bridging, where the fibers in the adhesive layers act to bridge the gap across the cracks in the

titanium. The result is a reduced stress intensity at the crack tip and slower crack growth in the

HTCL than in the monolithic material. The phenomena of fiber bridging is active in composite

material systems [ 15-16] and contributes to the improved damage tolerance of these materials.

Cracks propagated significantly faster in Panel 2 than in Panel l, at room and at elevated

temperature. Cracks propagated at much slower rates in both HTCL than predicted for the

monolithic material. Crack growth rates for the two laminates were calculated for each

individual crack, as shown in Table 2. In the monolithic titanium, the predicted crack growth

rate varies greatly depending on the length of the crack. For the monolithic titanium at an

9

appliedstress130MPa(sameasin Panel1),to extenda 1mmcrackto 1.5mm, a3 mmto 3.5

mm,anda 5 mmto 5.5mm,thecrackgrowthrateswouldbe9.72E-5mm/cycle,3.29E-4

mm/cycleand3.91E-3mm/cycle,respectively.In Panel1,thehighestrateis 1.67E-5mm/cycie,

over200timesslowerthanthegrowthin themonolithictitaniumatacracklengthof 5mm.

Similarly, for monolithicmaterialatanappliedstress211MPa(sameasin Panel2),to extenda

crackfrom4.5mm to 5 mmtheaveragecrackgrowthratewouldbe1.62E-2mrrdcycle,whereas

in Panel2, thehighestcrackgrowthratewas9.80E-5mm/cycle,over150timesslower. In

addition,thepredictionsindicatethatin themonolithicmaterialcrackgrowthbecomesunstable

andrapidfractureensues,atcracklengthsgreaterthan5 mm.

Thepoorbondingof Panel1mayhavereducedtheamountof loadtransferfrom the

fibersin theadhesivepliesto thetitaniumplies. Therefore,aweakerbondmayallow thefibers

to bethedominantloadcarryingcomponent.In fatigue,aweakbondis anadvantage,butat

compromiseis madein ultimatestrength,asisevidentby thebehaviorof Panel1(weakbond)

andPanel2(strongbond).

CONCLUSIONS

This investigationevaluatedthetensileresponseandthefatigueresistanceof HTCL at

roomtemperature.Theeffectof elevatedtemperatureon fatigueresistancewasalsoexamined.

ThelaminateanalysiscodeAGLPLY wasusedtopredicttheHTCL responsebasedupon

constituentpropertyinput. Resultsof thesetestswerecomparedto thosefor monolithictitanium

sheetsto assesstheimprovementin mechanicalbehaviorattainedby theHTCL. Two laminates

werefabricatedfrom differentresinsandof differingfiber volumefractionsandresinlayer

thicknesses,thentestedstaticallyin tensionandunderconstantamplitudefatigueloading. The

crackgrowthwasmonitoredandrecordedviaacetatefilm replication.

Theexperimentsshowedthatthequalityof theadhesivebondgreatlyinfluencesthe

mechanicalpropertiesachievablethroughlamination.A strongbondallowsloadtransfer

betweentheplieswithoutdelaminationoccurring,producinga higherstrengthmaterial,asin the

10

caseof Panel2. However,if fatigueresistanceis themajorconcern,a weakerbond,asin Panel

1, reducestheamountof loadtransfer,allowingthefiber strengthto dominate.BothHTCL

havelongerfatiguelives thanthemonolithictitanium,lastingoveramillion cyclesatboth room

andelevatedtemperatures,atappliedstressesof 130MPaand211MPafor Panel1andPanel2,

respectively.Crackingin thetitaniumpliesdid occur,but thespecimensdid not fracture. Fiber

bridgingwasassumedto dominatethecrackgrowthprocess.Onceall thetitaniumplieswere

fractured,thefiberscontinuedto carrytheappliedload. Theelastic-plasticlaminateanalysis

modelAGLPLY wasshownto predictthe laminates'tensileresponsefairly accurately,verifying

its applicabilityto modelingthemechanicalresponseof theHTCL. Overall,theHTCL providea

stronger,stifferandmoredamagetolerantalternativefor highertemperaturesto monolithic

materialswhileaddingareductionin weight,propertiesnecessaryfor applicationsto futurehigh

speedaircraft.

REFERENCES

1. Kaufman, J. G., "Fracture Toughness of 7075-T6 and -T651 Sheet, Plate and MultilayeredAdhesive-Bonded Panels," Journal of Basic Engineering, Vol. 89, 1967, pp. 503-507.

. Johnson, W. S. and Stratton, J. M., "Effective Remote Stresses and Stress Intensity Factorsfor an Adhesive Bonded Multi-Ply Laminate," Engineering Fracture Mechanics, Vol. 9,1977, pp. 411-421.

. Johnson, W. S., Rister, W. C. and Spamer, T., "Spectrum Crack Growth in AdhesivelyBonded Structure," Journal of Engineering Materials and Technology, Vol. 100, 1978,pp. 57-63.

4. Johnson, W. S., "Damage Tolerance Evaluation of Adhesively Laminated Titanium,"Journal of Engineering Materials and Technology, Vol. 105, 1983, pp. 182-187.

. Vogelesang, L. B. and Gunnink, J. W., "ARALL, A Material for the Next Generation ofAircraft. A State of the Art," Report LR-400, Dept. of Aerospace Engineering, DelftUniversity of Technology, The Netherlands (1983).

, Bucci, R. J., Mueller, L. N., Schultz, R. W., and Prohaska, J. L., "ARALL-Laminates--

Results from a Cooperative Test Program," Advanced Materials Technology '87, Society forthe Advancement of Material and Process Engineering, Vol. 32, 1987, pp. 902-916.

7. Schijve, J., "Development of Fibre-Metal Laminates, ARALL and GLARE, New FatigueResistant Materials," FATIGUE '93, May 3-7, 1993, pp. 3-20.

8. Progar, D. J. and St. Clair, T. L., "A New Flexible Backbone Polyimide Adhesive," Journalof Adhesion Science and Technology, Vol. 4, No. 7, 1990, pp. 527-549.

11

o

10.

11.

12.

13.

14.

15.

16.

Bahei-E1-Din, Y. A. and Dvorak, G. J., "Plasticity Analysis of Laminated CompositePlates," Journal of Applied Mechanics, Vol. 49, 1982, pp. 740-746.

Johnson, W. S., "Modeling Stiffness Loss in Boron/Aluminum Laminates Below the FatigueLimit," Long Term Behavior of Composites, ASTM STP 813, T. K. O'Brien, Ed., AmericanSociety of Testing and Materials, Philadelphia, 1983, pp. 160-176.

Johnson, W. S., "Fatigue Behavior of Continuous-Fiber Silicon Carbide/AluminumComposites," Composite Materials: Fatigue and Fracture, ASTM STP 907, H. T. Hahn,Ed., American Society for Testing and Materials, Philadelphia, 1986, pp. 161-175.

Johnson, W. S., Lubowinski, S. J. and Highsmith, A. L., "Mechanical Characterization ofUnnotched SCS-6/Ti-15-3 Metal Matrix Composites at Room Temperature," Thermal andMechanical Behavior of Metal Matrix and Ceramic Matrix Composites, ASTM STP 1080,

J° M. Kennedy, H. H. Moeller, and W. S. Johnson, Eds., American Society for Testing andMaterials, Philadelphia, 1990, pp. 193-218.

Hou, T. H., Johnston, N. J. and St. Clair, T. L., "Processing and Properties of

IM7/LARCrM-IA Polyimide Composites," to be published in the Proceedings of the 39thInternational SAMPE Symposium and Exhibition, Anaheim, CA, April 11-14, 1994.

Newman, J. C., Jr. "FASTRAN-II--A Fatigue Crack Growth Structural Analysis Program,"NASA Technical Memorandum 104159, February 1992.

Shaw, M. C., Marshall, D. B. and Evans, A. G., "Bridging Process in Metal-ReinforcedCeramics," Materials Research Society Symposium Proceedings, Vol. 170, 1990, pp. 25-31.

Bakuckas, J. G., Jr. and Johnson, W. S., "Application of Fiber Bridging Models to FatigueCrack Growth in Unidirectional Titanium Matrix Composites," Journal of CompositesTechnology and Research, Vol. 15, No. 3, 1993, pp. 242-255.

12

Table1. Constituentmaterialpropertiesof titaniumhybridcomposite.

Ti-6A1-4Va LARCTM -IA b

Yield Strength 827.4 MPa 71.7 MPa

Ultimate Strength 1068.7 MPa 121.9 MPa

Longitudinal Modulus 118.6 GPa 3.34 GPa

Transverse Modulus

Shear Modulus 45.62 GPa 1.26 GPa

% Elongation 5.8 6.0

Poisson's Ratio 0.33 0.33

IM7

N/A

5310.0 MPa c

275.8 GPa

13.79 GPa

200.0 GPa

1.8 c

0.25

a data from the current studyb data from [8]

c data from manufacturer

13

Table2. Crackgrowthratescalculatedfor variouscracklocationsin titaniumhybrid laminates.

PanelID

Applied Stress Temperature Locationof Cracka CrackGrowthRateMPa °C mm/cycle

1 130 25 Front-Left 1.67E-5

1 130 25 Back-Right 1.37E-5

2 211 25 Back-Right 6.84E-5

2 211 25 Front-Right 9.80E-52 211 25 Front-Left 4.30E-52 211 25 Back-Left 6.01E-5

2 208 177 Front-Left 8.25E-5

2 208 177 Front-Right 1.60E-4

a Locationassumesviewing from thefront.

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REPORT DOCUMENTATION PAGEForm_:WovedOMB No. 0704-0188

VA 22202-4302,1mcltothoOffk_o(Manag4¢n_and BucJl_l, Paloer_¢_ R_KlUCt_e Pmj=¢l (O'/04"01U_,W_m, _ _3.

1. A(_ENCYUSEONLY (Leave b/ank) 2. REPORT DATE 13. REPORT TYPE AND DATES COVERED

April 1994 J Technical Memorandum4. TITLE AND SUBTITLE S. FUNDING NUMBERS

Preliminary Evaluation of Hybrid Titanium Composite Laminates WU 505-63-50-04

e. ALrrHon(s)J. L Miller, D. J. Pro<jar,W. S. Johnson, and T. L. St. Clair

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

NASA Langley Research CenterHampton, VA 23681-0001

gl SPONSORING / MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

8. PERFORMING ORGANIZATIONREPORT NUMBER

10. SPONSORING / MONITORINGAGENCY REPORT NUMBER

NASA TM 109095

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION / AVAILABILITY STATEMENT

Unclassified- Unlimited

Subject Category - 24

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Inthis study, the mechanical response of hybridtitanium composite laminates (HTCL) was evaluated at room andelevated temperatures. Also, the use of an elastic-plastic laminate analysis program for predicting the tensileresponse from constituent properties was verified. The improvement in mechanical properties achieved by thelaminates was assessed by comparing the results of static strength and constant amplitude fatigue tests to thosefor monolithic titanium sheet. Two HTCL were fabricated with differentfiber volume fractions, resin layerthicknesses and resins. One panel was thicker and was poorly bonded in comparison to other. Consequently,the former had a lower tensile strength, while fewer cracks grew inthis panel and at a slower rate. Both panelsshowed an improvement in fatigue lifeof almost two orders of magnitude. The model predictions were also ingood agreement with the experimental results for both HTCL panels.

14. SUBJECTTERMS

Adhesive bonding; Graphite fibers; Isothermal fatigue; Laminated titanium;

Crack growth

17. SECURITY CLASSIFICATIONOF REPORT

Unclassified

18. SECURITY CLASSIFICATIONOF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATIONOF ABSTRACT

15. NUMBER OF PAGES

30

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)NSN 7540-01-280-5500 Prescribed by ANSI Std. Z3g-18

_g_102