Post on 07-Jun-2020
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
_"
L
z\
\
\
\\
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.
14,
COc_
oO
c_
GO •
c_O'_
oO
CID r
cot<
T B
O0 _ C_
T"
T B
00
aiQ
O4O
°,F,,_
o
o
0
.i0
15
16
-T-_
¢.q
oi,,,_
0
"0
2
.S¢1)
0
0w,,_
0
o
e4
|
\
'\I s
\
Illll
|
"\i
llllJllll
mmm
_ .,-= C_lm m m
¢0 ¢- e=
l I|
\ ' :'Il o |
Is
i'ii[ii'i[ i'l'i[iilij [
0 0 0 00 0 0 00 _ 0L_I T- T-
_mQa.
_N0
0
If)
- 0
- 0
0- 0
0
E
.m
,_
Jo
Eo
E
i
°_
oc-o
o
o
N
NN8
17
.......i
E
_4
0
O
.4
18
%
\
j i 04
I 0 III
,,0 _-
0
o
o
o
o_
oo
-__-_
o
19
\
\
\
\
\
\
i
\
IIIIIIillllillllllllllllmaZil
I I; 0
O O OO O O
O
o
_uE
o
o
_U
o
o
_E
_d
2O
m_
od
E
E
wm
m
o
o
0
o
o
o.
_z
"i°°_!I
0
U)
0
_'_
g!e_
_
M,-..ON
oo
--, o
_'_ __._
._,_
22
0
, ,,,,,, I,, ,I ...... __ZZZIZ_ZZ_ZIZIZZZZZZZZ_ZZIZZZZZZZIZZZZZZZZZ.ZZIZZZZZZII _
...............ZZIIIilII0 ZZZI!ZZZZZZZIZZII_ZZZZA_ 0 [] _ _
.........................._.... ........ _ ......................._ ......................
........................... _........... _ ....................... _.......................... _
...........................i_o.... [] _ .......=_ _ ................
_!!!!!!!!!!!!!?i!?!!2 - _ !!!!!!!!!!!??i!!!!!!!!!?!?!!!!?!!!!!!!!!!!!!!!_!!!?!__..........................._.... __ .............!..........................._ ..................
.....................,__i/ I..............................................................................
.............................................................................................*............................ il............................ )......................... ).........................._... _i 0 _0_ ___
...........................i.........................•............................_........... _ ........_.......................... _
................::::::::::::::::::::::::::::::::::::::::::::::::::::.................... _
........................................................ i ............................ ) ............................ ) ..........................
...... i , , , i , , , I , , ,I I I
0 0 0 0 0 0
0
_3
E
, 1 , , .... i .... I .... _ "_--.!-:.-'-'...1..._...!-!.--:-t-'.--:-.-!-..!...!-..'-'-.'...'...!...................l ...................!................... ,,-
...................i--71 o IIIIIiZZIIZZIIIi=ZZZIZZiZIZZ:;Z_ _
...................!............e,. ........ _ .........._ ............._..................
...................!................... ": . ...........FI .............!...................
tt _ L] ..............._.........F..................r..........._ o [] <1......."= ..........._l- i
7::.................. ; ............ -- ......... ._....................................... i ................... "
...................i............< _ _ .........i................:1::::::............._...................
..................._............_. c ,- .........i.................._....................i...................o o.-- t_ t_ _ .... 1."................ _................
...................i............_- _ _ .........i..............l i......i _
.................. ; ................... _ ................... , .................... i ................. 4.................... i ................... _ t_
................... ;................... *................... i.................... i................. +................... i................... >'to
...................;..................._...................._...................._................. , ..................._...................
...................;...................4....................l.................i...................D................._............................ •................... _.................... _................. :.._................ _......... "_......;...................
................... i................... _...................... i................... ,i-
...................i..............................................................i......................................,..........._:::::: ............._.____._...................-_
............._...................................!......._, ._...................:...................i..__._.__.Z.=======================================================.......... 'F................... + .............. I .................... i ................... @......................................
............_"_ ............. T.............................................................. =
.... I .... I .... l, ............ 1 .... _0 _'_ "
1 I I I _ __
I
24
• II
_,,...._......._ ...............i.................i.................i.....................i.................i.................i.................i.................J................
i
.... i.... I.... i.... I.... i.... I .... ] ....I I I I I I I
r:
ffl
O
O
q.
O
o
N
O
U
o
EZ
O
O
L..
o_8_
"T" o
O
25
e',
m
U
U
000
COCr)CO
• i,,,_
°l,-t
o
0on(
c.)
c.)
e_J
26
,,, ,I .... l .... I .... I .... I .... I ....E EEE_ , • , _o. _o _l--
_, ,_ll ii _ i I ) _ ) 0 _
17 _'0"01- m m _1%3
N **
!•- g
.... I .... I,,,,I,,,,I,,,,1 ,,,,I,,,,I I I I I I
"E_EU
eU
§
°i1
Z
o
0
0
0
i ,,°e_,_
e _a
_.. __
_°_._
_.-
27
.... i,,,,i,,,,t .... t .... i .... i ....
E P
L ,--@._
,,_ / _ / ,,,, , ,,,'" II _ II -- ,,, _.l I._ u_ i._
_. / _:-a"al- / ' ' ',
• ¢_
"X 0
--._
_._- u ___> _ _
_ ,-- ,_ /,zOID 1_, 1 / ml
_i_ _ _ I .... I , , , , I .... I .... I .... 1 ....I I I I I I
,E
C
E0
.o._E"_ _ ._
c_
_ ¢"
E=
Z _=
_ .__#
L, ["_ e'_
o _._
o _I_
,4
28
I I I I_ I Ii i l i i i i e i l i i a I I i i l i l i a i i a e i i
• _- IJ=
?I
_=®
,,,ul,, ,,I,,,,1,,,,I,,,,I ,,,,1_,,,I I I I I I
e-
0
0
oo
-o° _,
i °Z
:..._"_
o
_'_8go
_'_
_ o
N'-
_._
29
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