delamination of a strong film from a ductile substrate during indentation unl...
Tests of Graphite Polyimide Sandwich Panels Camarda
1. - I I I!!! 111111111!!!!!!!!!
NASA Technical Paper 1785
Tests of Graphite/Polyimide
Panels in UniaxialEdgewise
Charles J. Camarda
DECEMBER 1980
NASA
TP
1785
c. 1 ~
----Sandwich
Compression
2. TECH LIBRARY KAFB, NM
0067732
NASA Technical Paper 1785
Tests of GraphitdPolyirnideSandwich
Panels in UniaxialEdgewiseCompression
Charles J. Camarda
Latzgley Research Center
Halnptm,Virginia
NASANational Aeronautics
and Space Administration
Scientific and Technical
Information Branch
1980
3. SUMMARY
An experimentalandanalyticalinvestigationhasbeen made ofthe local
andgeneralbucklingbehavior of graphite/polyimide(Gr/PI)sandwichpanels
simplysupportedalong a l l fouredgesand loaded inuniaxialedgewise com-
pression. Material properties ofsandwichpanelconstituents(adhesiveand
f a c i n g s ) were determinedfromflatwise-tensionandsandwich-beam-flexure
tests. Results from theflatwise-tension tests e s t a b l i s h e d a s u i t a b l e cure
c y c l ef o r FM-34l polyimidefilmadhesivewhich was theadhesiveused to f a b r i -
cate theflatwise-tension,sandwich-beam,andbucklingspecimens. A cell-edge
bondingtechniqueusing a l i q u i dv e r s i o no f FM-34 polyimideadhesive was
investigatedand results i n d i c a t e dt h a t a considerable mass savings may be pos-
s i b l eu s i n g a cell-edgeadhesive.Tensileandcompressivematerial properties
ofthefacings(quasi-isotropic, symmetric, laminates ( ( [0,+45,90,-45],)of
Celion2/PMR-1 5) were determined a t 116 K, room temperature,and 589 K (-250°F,
room temperature, and 600°F) usingthesandwich-beam-flexure test method. Buck-
lingspecimens were 30.5 by 33 c m (12 by 1 3i n . ) , had quasi-isotropic, symmetric
f a c i n g s ( [0,+45,90],),and a g l a s s / p l y i m i d e honeycomb core (HRH-3273-3/8-4).
Core thicknesses were varied(0.635,1.27,1.91,and2.54 cm (0.25,0.50,0.75,
and1.00in.))andthreepanelsofeachthickness were t e s t e d a t room temper-
a t u r e to i n v e s t i g a t ef a i l u r e modesandcorrespondingbucklingloads.Specimens
0.635 c m (0.25 i n . ) t h i c k f a i l e d by overallbuckling a t loads close to theana-
l y t i c a l l yp r e d i c t e db u c k l i n gl o a d ; a l l o t h e rp a n e l sf a i l e d by facewrinkling.
Resultsofwrinkling tests indicatedthatseveralbucklingformulas were uncon-
servativeandthereforenotsuitablefordesignpurposes; a recommended wrin-
klingequation is presented.
INTRODUCTION
Preliminarystructuralstudiesofadvanced space transportationsystems
usingadvancedcompositestructuralmaterialsofhigh-strengthfibersand
polyimideresin matrices i n d i c a t e t h a t a reductionof up t o 25 percentinvehi-
cle s t r u c t u r a l mass is obtained by thedirectreplacementof aluminum panels
withgraphite/plyimide(Gr/PI)panels(refs. 1 and2).Furthermore, prelimi-
narystudies of t h e a f t body f l a po ft h eS p a c eS h u t t l eO r b i t e r( r e f .3 )i n d i c a t e
thatcompressionloads are t h ep r i m a r yd e s i g nc o n d i t i o nf o rt h i ss t r u c t u r a l com-
ponentandbecause a b i a x i a l state of stress e x i s t si nt h ec o v e rp a n e l s a sand-
wichpanel was chosen. The presentstudyfocusesonGr/PIstructuralsandwich
panelswhich may haveapplication as coverskinsonlightlyloadedcomponents
s u c ha st h ea f t body flapoftheSpaceShuttleOrbiter.Basedonthe l o w magni-
FM-34 filmadhesive:manufactured by AmericanCyanamid Company,
Bloomingdale Division.
2Celion: registeredtrademark of CelaneseCorporation.
3HRH 327: r e g i s t e r e d trademark of HexcelProducts,Inc.
4. IllllIllIlllllllIll1I I lIlllIlllIIll I1 11111l1l1ll11l1l1l1
tude and biaxial natureof these loads, a minimum-gage, quasi-isotropic,SF-
metric Gr/PI laminate ( [ O f +45,901s ) was chosen for the facingsof these sand-
wich panels in the present study.
The purposesof the present study are to analytically and experimentally
investigate the local and general buckling behaviorof minimum-gage Gr/PI sand-
wich panels capable of use at temperatures ranging from 116 to 589K (-250° to
600°F), to verify the fabrication method used in manufactureof the panels, and
to determine the material properties of the [0,+45,901s Gr/PI sandwich panel
facings.
Buckling specimens 30.5 by 33.0cm (12 by 13 in.) were designed and fabri-
cated with various core thicknessesto study local and general instability
failure modes. The buckling specimens were tested in uniaxial edgewise compres-
sion at room temperature (R.T.) andwere simply supported along allfour edges.
Several analysis methods (refs. 4 to8) were used to determine upper and lower
bounds on critical stresses relating to intracellular buckling (dimpling), wrin-
kling, shear crimping, and general panel instability and are evaluated in this
study for their capability in predicting buckling loads and modes of Gr/PI
sandwich panels.
The panels were fabricated using a commercially available high-temperature
film adhesive, FM-34, to bond the core to the facings. Flatwise-tensile tests
were performed using the sandwich panel facing laminate orientation,core, and
adhesive to determine a suitable fabricationcure cycle and the tensile adhesive
bond strength in a core-to-facing bond situation. In addition, flatwise-tensile
tests were used to evaluate BR-34, a liquid version of the FM-34 film adhesive,
as a cell-edge adhesive.
Sandwich-beam-flexure tests were performed to determine modulus, strength,
and Poisson's ratio of the facings. The flatwise-tensile tests and sandwich-
beam-flexure tests were conducted at temperatures of 116K, R.T., and 589 K
(-250°F, R.T., and 600OF). Quality control standards for fabricationof all
specimens were high to minimize scatter in the data. Results of the tests are
presented in tabular and graphical form. Results of the beam tests were ana-
lyzed statistically and a best-fit third-order polynomial relating stress and
strain was fit through the data.
Certain commercial materials are identified in this paper in order to spe
ify adequately which materials were investigated in the research effort.In no
case does such identification imply recommendation or endorsement of the produc
by NASA, nor does it imply that the materials are necessarily the only ones or
the best ones available for the purpose.In many cases equivalent materials are
available and would probably produce equivalent results.
SYMBOLS
Values are given in bothSI and U.S. Customary Units. The measurements and
calculations weremade inU.S. Customary Units.
2
5. DQxI DQy
FC
Fcu
G
Nx Ny
stiffness matricesof sandwich panel
width of plate
coefficients of polynomials used in regression analysis
flexural stiffness of composite facings
transverse shear stiffnessof sandwich plate inx- and
y-directions, respectively
flexural stiffness of orthotropic sandwich plate in x- and
y-directions, respectively
twisting stiffnessof orthotropic sandwich plate
elastic modulus
modulus of core in z-direction
facing modulus
facing modulus inx- and y-directions, respectively
tangent modulus
average elastic moduliof laminate in x- and y-directions,
respectively
lower of flatwise core compressiveor tensile strengths,or
core-to-facing bond strength
compressive ultimate strength
shear modulus
core shear modulus in xz-plane
core shear modulus in yz-plane
facing shear modulus in xy-plane
total number of points in regression analysis
length of plate
number of half sine waves in x- and y-directions, respectively
resultant normal forcesin x- and y-directions, respectively
3
6. P
SCJ/E
S
Tg
t
t C
tf
t f l
tf2
th
Vf
VV
X I Y I Z
6
6
1-1
1-1xr1-1y
Dxy' pyx
"
P
0
ocr i m
*d i m
*wr
load
standarderror of estimate
honeycomb c e l l s i z e
glass transition temperature
thickness
corethickness
average facingthickness
thickness of facing 1
thickness of facing 2
total sandwich panelthickness
fiber volume fraction
void volume fraction
rectangularcoordinates
i n i t i a l panel waviness
strain
Poisson'sratio
Poisson's ratio of orthotropic plate associated w i t h bending of
plate i n x- and y-directions,respectively
average Poisson'sratios of orthotropicplateassociated w i t h
extension of plate i n x-and y-directions,respectively
density
stress
critical stress associated w i t h shearcrimping
critical stress associated w i t h dimpling
critical stress associated with wrinkling
Subscripts :
av average
4
7. cr critical
iindexofsummation
maxaximum
ultltimate
XIYcoordinatedirections
1,2directionsparallelandperpendiculartofiberdirection,respectively
TEST SPECIMENSI APPARATUS, AND PROCEDURE
Graphite/Polyimide Materials
This program was conducted as part of the NASA program, Composites for
Advanced Space Transportation (CASTS) (ref. 1). The CASTS effort focusedon
graphite/polyimide, and a significant part of the program included evaluating
and characterizing various fiber and resin materials. As a result of these
evaluations, the materials used in different phases ofthe present study varied
as improved systems were identified. Specifically, for flatwise-tensile tests,
the laminates were HTS4-l/PMR-15; for sandwich-beam-flexure tests, the laminates
were Celion6000/PMR-15; for buckling tests, the laminates wereCelion
3000/PMR-15. The primary purpose of the flatwise-tensile tests was to evaluate
adhesive tensile strengths and therefore the difference in facing materials was
not critical. The thinnest gage prepreg, Celion 3000/PMR-15, was chosen over
the Celion6000/PMR-15 to minimize mass of the sandwich panels, and the Celion
fiber was chosen over the HTS fiber because Celion exhibitsless material prop-
erty degradation than HTS at elevated temperatures.
Flatwise-Tensile Specimens
Forty-six 7.62 by 7.62-an (3 by 3-in.)specimens were fabricated using pre-
cured [0,+45,901s laminates of HTS-l/PMR-15 Gr/PI facings, glass/polyimide
honeycomb core (HRH-327-3/16-6 or 8) and the desired adhesive. A schematic dia-
gram of a typical specimenis shown in figurel. Details of fabrication proce-
dures and cure cycles are given in reference 9 and tableI, respectively.
Steel load blocks were bonded to the facings of the specimens and each block h
a tapped hole for attaching a loading rod. Universal joints were attached
between the testing machine and the loading rods to assure proper alignment of
the fixture in the loading machine.The specimens were tested in a universal
testing machine operating in a displacement control modeat a constant rate of
0.13 cm/min (0.05 in/min). Test temperatures other than room temperature were
obtained using an environmental chamber positioned within the crossheads and
posts of the testing machine. Specimens were held at desired test temperatures
for 15 minutes priorto testing to insure thermal equilibrium. Maximum load
was recorded for each test and converted to a normal tensile stress.
. .
4HTS graphite fiber: product of Hercules Incorporated.
5
8. Sandwich-Beam-Flexure T e s t s
Specimens.-Sandwich-beam-flexurespecimensconsisted of Gr/PI facingsand
glass/polyimide honeycomb core as shown i n f i g u r e 2.Thehoneycomb core was
HRH 327-3/16-8 glass/polyimideand was c u t i n t o strips 2.54 cm (1.00in.)wide
by55.88 cm (22.0in.)longby3.175 cm (1.25in.)high. The test f a c i n go f
t h e beam was a [O,+45,9O,-45Islaminate of Celion 6000/PMR-15 andtheoppo-
s i t e facinghad a laminateorientationof[02,+45,90,-451,.Theadditional Oo
l a y e r s of t h en o n - t e s tf a c i n gi n s u r e df a i l u r e of t h e test facing.These lami-
n a t e o r i e n t a t i o n s were chosen to avoid microcracks inthelaminatewhich are
believed to occur when a d j a c e n t l a y e r s are stacked a t ananglegreaterthan or
equal to 90° w i t hr e s p e c t t o oneanother.The honeycomb core was f i l l e d w i t h
BR-34 liquidadhesiveandglassbeadsthroughoutthelength of t h e beams,
e x c e p tf o rt h e 7.62-cm (3.00-in.) test s e c t i o ni nt h ec e n t e r of t h e beams, t o
prevent premature a d h e s i v ef a i l u r e . Details ofthefabricationofthesand-
wich beam specimens are p r e s e n t e di nr e f e r e n c e 9.
Apparatusandinstrumentation.-Eachspecimen was i n s t r u m e n t e da tt h e mid-
spanofthe beam with a high-temperaturestrain rosette (WK-03-060-WR-350)
o r i e n t e d a t Oo, 45O, and 90° withtheloadaxisandbonded to t h e test facing,
and a s i n g l es t r a i ng a g e (WK-03-125-AD-350) o r i e n t e d a t Oo w i t ht h el o a da x i s
andbonded to thenon-testfacing.Thesegages were manufacturedby Micro-
MeasurementsDivisionofVishayIntertechnology,Inc. The s t r a i ng a g e s were
bonded to t h e outer surfacesofthe beam using a polyimideadhesive(either
"Bond 610 or PLD-700 availablefromMicro-Measurementsand BLH E l e c t r o n i c s ,
r e s p e c t i v e l y ) .
The sandwichbeams were p l a c e d i n a four-pointbending test apparatus
(fig.3)whichsupportedthe beam on rollers 48.26 cm (19.00in.)apartwith
f l a t s e c t i o n s 2.54 c m (1.00in.)widemachinedin them.Load was applied by
a 222-kN (50-kip)capacityhydraulictestingmachinewhichacted a t two p o i n t s
onthe top flangeofthe beam spaced10.16 cm (4.00in.)apartand symmetric
aboutthebeam'scenter.Fortesting a t temperaturesotherthan room temper-
aturethespecimen was instrumentedwith a thermocoupleattached t o t h e test
facingandthe test fixtureandspecimen were completelyenclosedinanenvi-
ronmental chamberand eitherheated or cooled t o t h ed e s i r e d test temperature.
Specimens were allowed t o soak a t t h e test temperature for 20 minutes to i n s u r e
thermalequilibrium. A datahandlingsystemconsistingof40-channelscanner,
d i g i t a lv o l t m e t e r , plotter, p r i n t e r , clock, andcalculator was used t o record
andreducedata.
Procedure.-Theloadsignalsfromtheload cell o n t h e h y d r a u l i c t e s t i n g
machine were i n p u t t o onechannelofthescanner.Strainsignals were i n p u t t o
selectedscannerchannelsandinitally set to zerousingWheatstonebridgebal-
ance(fornon-room-temperature tests, i n i t i a l s t r a i n s i g n a l s were set to z e r o
a f t e rt h e r m a le q u i l i b r i u m ) .S t r a i n s were corrected f o rt r a n s v e r s es e n s i t i v i t y
ofthegagesandnonlinearityofthebridgecircuit.Thermocouples were con-
nected to thescannerthrough a 273 K (32OF) cold-junctionreference.
Beams were t e s t e d t o f a i l u r e d u r i n g t h e test, load was applied a t a r a t e o f
80 N/sec (20lbf/sec),data were recordedevery 3 seconds,and a s t r e s s - s t r a i n
6
9. curve was plotted in real time. Quantities were stored in volts and engineering
units on magnetic tape and printed during each test.A data reduction program
used the longitudinal stresses and strains of replicate tests as input to a
regression analysis to determine the coefficients of a best fit, in the least-
squares sense, of a third-order polynomial relating stress and strain according
to the polynomial equation:
A more detailed explanationof the analysis is given in reference 10.
To assess the magnitudeof scatter of experimental points about the regres-
sion equation, the standard errorof estimate So/,, which is a measure of the
mean deviation of thesample points from the regression line, is determined as
follows:
j - 4
This method of statistical analysis is similar to that presented in reference10.
Buckling Tests
Specimens.- Design considerations of the buckling specimens are givenin
appendix A. Specimens were30.5 by 33.0 cm (12 by 13 in.) withcore thicknesses
of 0.635, 1.27, 1.91, and 2.54 cm (0.25, 0.50, 0.75, and 1.00 in.). Facings of
all sandwich panels were similar and were symmetric quasi-isotropic 8-ply lami
nates of Celion/PMR-15 [0,+45,901,. Figure 4 shows acompleted buckling speci-
men; detailson specimen manufacture are given in reference 9; detailsof
significant panel parametersare listed in table11. Quality-control standards
(refs. 1, 2, and 9) for fabrication of the panels were very high to minimize
scatter in experimental data.
Apparatus. "- ~~ and .instrumentation.- Simply supporting the edgesof the test
panels was considered-tobe a realistic representation of the boundary condi-
tions that actual panelson the shuttle bodyflap willexperience. A test
fixture, similar to that of reference4, was fabricated to simply support all
four edges of the sandwich panel and allow adjustmentsto be made during load-
ing which would align the specimen and thus maintain a uniform strain distri-
bution across the panel. Details ofthe simple supports are shown in
7
10. figures5(a)and ( b ) . The stainless-steelalignmentsheet embedded i ne a c h of
thepottedends of thepanel f i t i n t os t a i n l e s s - s t e e lk n i f ee d g e sw h i c h f i t
i n t o steel V-groove loadblocks as shown i n f i g u r e 5 (a). The load blocks fit
intoadjustableendloadingheadswhich were a t t a c h e d to t h eh y d r a u l i cl o a d
machine. The endloadingheadscontained a f l a ts t a i n l e s s - s t e e lb a rw h i c h ,
togetherwiththealigning screws, was used to o b t a i n a uniformlongitudinal
s t r a i n across thespecimen.Figure 6 shows thebucklingspecimeninthe test
f i x t u r e . The sides of thepanel were simply supported by knifeedgeswhich
were supported by Z-section steel beams as shown i nf i g u r e5 ( b ) . The side
supportsmaintained a r e l a t i v e l ys n u gf i ta g a i n s tt h ep a n e lb e c a u s e of thehigh
degree of f l a t n e s s of thepanels.However,because of t h e raised scalloped
doublersthe side supports could notextendthe complete lengthofthepanel.
The Z-section beams were braced so thatmotion of t h e side supports was
r e s t r a i n e d . The knifeedgesofthe side supports were boltedsnuglyinplace a t
t w o locationson two sides as shown intheschematicoffigure5(b)and as par-
t i a l l y shown i nf i g u r e6 . The s i d es u p p o r t s were positioned1.27 c m (0.5in.)
from eachsideedge,makingthe test sectionwidth30.5 c m (12 i n . ) .
A 222-kN (50-kip)hydraulicloadmachine was used to compress t h ep a n e l s ,
A mercury-vapor l i g h t source was usedinconjunctionwith a photographicline
gridhaving a p i t c h of 19.7lines/cm(50lines/in.) to determineout-of-plane
displacementsand mode shapesusingthegrid-shadow Moire/ technique as dis-
cussedinreferences 11 and12. A camera was positionedperpendicular to t h e
sandwichpanelandthelight source formed anangleof 30° w i t ht h a t
perpendicular.
Eachpanel was instrumentedwith12single,foil-typestraingagesand 2
45O s t r a i n rosettes, Micro-Measurements WK-03-125-AD-350 and WK-03-060-WR-350,
r e s p e c t i v e l y , as shown s c h e m a t i c a l l yi nf i g u r e 7. The p o s i t i o n i n g of t h e
gagesallowedmeasurementoflongitudinalstraindistributions on eachfacing
across thepanellengthandwidth.Back-to-backlongitudinalstraingages were
positioned a t fivepointsonthepanel(fourcornerpointsand a c e n t r a l l y
located one). The purposeoftheback-to-backgages was to detect bending of
thepaneland to determinethegeneralbuckling load andpossiblythewrinkling
load. The dataacquisitionsystemused to reduceand store d a t a is i d e n t i c a l
to thatmentioned earlier for thesandwich beams.
Procedure.-Duringeach test, thehydraulictestingmachine was operated
i n a displacementcontrol mode a t a rateofapproximately0.020 cm/sec
(0.008 i n / s e c ) ,s t r a i ng a g e s were scannedapproximatelyevery 3 seconds,and
thespecimen was loaded to f a i l u r e . Raw d a t a were converted to engineering
u n i t s ,p r i n t e di n real time, andstoredon a d i s k .S t r e s s e s were calculated
by dividingtheload by t h e combinedcross-sectional area ofthe two facings.
Gages were balanced prior to testingusingWheatstonebridge circuits as dis-
c u s s e de a r l i e r .P r i o r to tesing,panels were loaded up to approximately
50 percentoffailureloadand were a l i g n e du s i n gt h ea d j u s t a b l e screws shown
i nf i g u r e6 . The panel was thenunloadedandthe Moire' g r i dp o s i t i o n e di n
frontofthespecimen.Straingages were thenzeroedand load was applied to
thespecimenuntilfailure.
8
11. TESTRESULTS
Flatwise T e n s i l e T e s t s
Preliminary tests indicatedthatsignificantimprovementsof bond s t r e n g t h s
couldbeobtained by abrasivelycleaningtheedgesofthe honeycomband by
dippingthe core inprimerinsteadofbrush or roller coating it onthe core.
(Seeref.9.)
A series of f l a t w i s e t e n s i l e tests of specimens,bondedwith FM-34 using
various cure c y c l e s ,a i d e di nt h es e l e c t i o no f a s u i t a b l ec u r ec y c l e . Two
specimens were t e s t e d a t room temperature for e a c hc u r e - c y c l ev a r i a t i o nl i s t e d
i nt a b l e I. Specimen failuresoccurred by eitheradhesivebondlinerupture or
facingdelamination;figures 8 and 9 show t h e t w o modes offailure.Resultsof
those tests, l i s t e d i n t a b l e 111, i n d i c a t et h a t cure c y c l e s 1 and 5 produced
thestrongestbonds,withfailuresoccurringinthefacing.Delaminationof
thefacings also occurredwithcurecycle 4 butbecausethe bond curetempera-
ture of 61 6 K (6500F) was greaterthanthefacing cure temperatureof 603 K
(625OF) theinterlaminarshearstrengthofthefacing was degradedandfailure
loads were lower. Bondingonefaceofthespecimen a t a time, withtheface
to bebonded positionedunderthe core (curecycle3),providedgoodfilleting
betweenthefaceand core b u t didnotenhancethestrengthofthebond.
Instead, bond s t r e n g t h s were lower and f a i l u r e so c c u r r e di nt h es e c o n do ft h e
t w o bonds.Sixspecimens were t e s t e d a t 589 K (600OF) ; t w o specimens were fab-
r i c a t e d a t each of t h r e ec u r ec y c l e s 1 , 5,andcurecycle 1 with a highercure
temperature (603 K (625OF) ) . I t was hoped t h a tt h eh i g h e rc u r e temperature
wouldimprovetheelevatedtemperature bond strength.
Curecycle 1 withtheelevatedcuretemperature was chosen because ofthe
higher bond s t r e n g t h s a t e l e v a t e d temperature andbecausemaintaining a vacuum
during cure wouldhelpeliminatevolatilesproducedduringthecureofthe FM-34
adhesive.Althoughtrappedvolatilesdidnotdegradethestrengthsofthe7.62
by 7.62 c m (3 by 3 i n . ) specimens, it would be more d i f f i c u l t to ventthevola-
t i l e s inlargerpanels.
Sixteenflatwise-tensionspecimens were fabricatedusingcurecycle 1 with
a cure temperatureof603 K (625OF). T e s t results fromthesespecimens are pre-
s e n t e di nt a b l e IV. F l a t w i s e - t e n s i l es t r e n g t h s a t room temperatureand116 K
(-250OF) increasedfrom 1 .6 MPa (230 psi) to an averagevalueof3.2(470psi)
when theedgesofthe honeycomb were cleanedand primed as mentioned earlier.
F a i l u r e s a t t h i s stress l e v e l were usually by facingdelamination as shown i n
f i g u r e9 .F l a t w i s e - t e n s i l es t r e n g t h s a t 589 K (600OF) were generallyhigher
than1.4 MPa (200psi)withfailuresoccurringinthebondline, similarto t h e
room-temperature test shown i n f i g u r e 8.
F l a t w i s e - t e n s i l e test r e s u l t s a t room temperatureofspecimensbondedusing
BR-34 as a cell-edgeadhesive are p r e s e n t e di nt a b l eI V ( b ) . Most of t h e s e spec-
imens f a i l e d by facingdelamination. However, for thesespecimensthefacings
9
12. delaminated locally about eachcell edge as shown in figure 10 and usually
resulted in slightly lower strengths.When local facingdelamination did not
occur, strengths were similar to results of the FM-34 film adhesive. Flatwise
tensile strengths using BR-34 were much higher than results presented in refer-
ence 13. The massof the BR-34 adhesive was 0.244 kg/m2 (0.05 lbm/ft2) which is
a 59-percent reduction in mass compared with FM-34 film adhesive having a mas
of 0.586 kg/m2 (0.12 lbm/ft2) . The use of BR-34 would resultin a mass savings
equivalent to 10percent of the total sandwich panel mass for a panel consisting
of 8-pl Gr/PI facings and a 1.27-cm (0.50-in.) thick core having a density of
64 kg/m 3 (4 lbm/ft3) .
Results of the bond study indicate that a liquid cell-edge adhesivecan
result in considerablemass savings without necessarily sacrificingbond
strength and that further research in this areais warranted. However, since
flatwise-tensile strengths with BR-34 were not consistent, FM-34 film adhesive
was used to fabricate the sandwich beam and buckling specimens.
Sandwich-Beam-Flexure Tests
Results of the sandwich beam flexure testsare presented in tables V
and VI and in figures1 1 to 16. As shown in table V, the scatter of test data,
as determined by the standard error of estimate, waslowest for the room tem-
perature and 116 K (-250OF) tensile tests. Maximum scatter occurred for the
elevated and room temperature compression tests in which the standard errors
of estimates SolE were 10.67 MPa (1 547 psi) and 11.10 MPa (1 610 psi) as com-
pared to respective average ultimate strengths of 567.7 MPa (82.34 ksi) and
334.0 MPa (48.44 ksi) (see tableVI). The average compression ultimate strain
was 1.38 percent at room temperature and 0.657 percent at 589K (600OF).
Average ultimate strengths of the laminate were slightly higher in compression
than tension for each test temperature. Ultimate strengths of theCelion
6000/PMR-15 [0,+45,90,-451s laminates were higher than results for HTS/PMR-15
as reported in references 10 and14 except for tensile strength at 589K (600OF)
reported in reference 14. Average room temperature tensile and compressive
ultimate strengths for the HTS/PMR-15 laminateswere 450.6 and 532.4 MPa (65.36
and 77.23 ksi), respectively, as compared with 565.2 and 567.7 MPa (81.98 and
82.34 ksi) for Celion 6000/PMR-15. Average tensile ultimate strengths at 116 K
(-250OF) increased by 8.5 percent over roomtemperature values while strengths
at 589 K (60O0F) decreased by 43 percent. Average compressive ultimate
strengths at 116K (-250OF) and 589K (600°F) increased and decreased, respec-
tively, by 13.8 and 41.2percent from room-temperaturevalues.
Modulus values of the Celion 6000/PMR-15 laminates were higher for all test
temperatures than values reported in references 10 and 14 for HTS/PMR-15 lami-
nates. This difference is probably due to the higher fiber volume fraction,
72 percent for the Celion/PI laminatesof the present study compared with 43to
55 percent for the HTS/PMR-15 laminates of references 10 and 14. Modulus values
at 0.2 percent strain and 116K (-250OF) were about 10 percent higher than
values at room temperature. Modulus valuesat 589 K (60O0F) were about the
same as room-temperature values. Stress and tangent modulus as a functionof
strain for various temperatures are presented in figures11 to 16. Table V
10
13. lists t h e c o e f f i c i e n t s of theregressionequation,usedinthereduction of the
experimentaldata.Thedatapointsinthefiguresrepresentexperimentalvalues
a l l r e p l i c a t e tests; t h e s o l i d l i n e i n t h e f i g u r e s is t h eb e s t - f i tt h i r d - o r d e r
polynomialobtainedfromtheregressionanalysis.Tangentmodulus as a func-
t i o n of s t r a i n was c a l c u l a t e d by d i f f e r e n t i a t i o n of thethird-order,polynomial
(eq. (1)).Tensilemodulusvalues were f a i r l yl i n e a rt h r o u g h o u tt h eu s a b l e
s t r a i nr e g i o n (E I 0.35percent) as shownby figures11,13,and15. Compres-
sivemodulusvaluestended t o benonlinear a t room temperatureand became l i n e a r
a t 589 K (600OF) as shown,byfigures 14 and 16.
R e p r e s e n t a t i v et e n s i l ea n dc o m p r e s s i v ef a i l u r e s are shown i n f i g u r e s 17 and
18,respectively. Most compressivefailuresoccurredneartheedge of t h e
p o t t e d s e c t i o n o f t h e honeycomb next t o t h e l o a d tabs.
Buckling Tests
Two modes o f p a n e l f a i l u r e were discernablefromexperimental results:
wrinklingandoverallbuckling. Specimens with a core thickness tc ofapprox-
imately0.635 cm ( 0 . 2 5i n . )f a i l e d by overallbuckling,and a l l otherspecimens,
havingnominal core thicknessesof1.27,1.91,and2.54 c m (0.5,0.75, and
1.00 i n . ) ,f a i l e d bywrinkling. None of t h e panels tested failed bylaminate
yield,dimpling, or shear crimping. Theshadow Moire/ method was u s e f u l i n
determining mode shapesof t h e overallbucklingspecimensbut was not able to
determinewrinkling mode shapesbecauseofthehighstiffnessandbrittlenature
of t h e Gr/PI facingsand,hence,therelativelysmallout-of-planedisplace-
ments.Theuseof a f i n e r Moire/ l i n e g r i d wouldincrease the s e n s i t i v i t yo ft h e
opticaltechniqueandpossiblyenable t h e determinationof local buckling modes.
Wrinkling. -. ~ ~ - .. specimens.- R e s u l t s o f l o n g i t u d i n a l s t r a i n u n i f o r m i t yacross
specimen width a r ep r e s e n t e di nf i g u r e s1 9 ( a )a n d( b )f o r two values of applied
loadand t w o d i f f e r e n tp a n e l s . The a d j u s t a b l e test f i x t u r e was u s e f u li n eliw
i n a t i n gl a r g es t r a i nv a r i a t i o n sc a u s e d bymisalignment, similar t o test f i x t u r e s
used i nr e f e r e n c e 4. S t r a i n s were f a i r l yu n i f o r m across thewidth of thepanel
as shown i nf i g u r e s1 9 ( a )a n d( b ) . However, s l i g h t l yh i g h e rs t r a i n sa n ds t r a i n
v a r i a t i o n s do occur a t t h e edgesofthepanels as was n o t e di nr e f e r e n c e 4.
T r e n d s i n s t r a i n d i s t r i b u t i o n s a t t h e l o w loadlevel,44.48 kN (10 000 l b f ) ,
were similar to t r e n d s a t thehigherloadlevelof88.96 kN (20 000 l b f ) .T h e r e
were no c o n s i s t e n tt r e n d si ns t r a i nd i s t r i b u t i o n sf r o mp a n e l t o panel.However,
most ofthewrinklingspecimensdidfailneartheendofthe side simple
s u p p o r t s w h e r e s l i g h t l y h i g h e r s t r a i n s were recorded.
L o n g i t u d i n a ls t r a i n s as a function of stress were c a l c u l a t e df o re a c h
strain-gagepositiononthepanel. R e s u l t s of s e v e r a l tests (panelnumbers BT-5
and BT-6) are p r e s e n t e di nf i g u r e s2 0 ( a )a n d( b ) . Back-to-back s t r a i nv a r i a t i o n
was u s u a l l y lowest i nt h ec e n t e r of thepanels(x = y = 0 ) . I r r e g u l a r i t i e s i n
slopes were n o t e di n some specimens as shown i nf i g u r e2 0 ( b ) for panel number
BT-6. T h e s ei r r e g u l a r i t i e si ns l o p eo c c u r a t too l o w a load to beanindication
ofwrinkling or some form of local i n s t a b i l i t y as mentionedinreference 4. The
i r r e g u l a r i t i e s i n t h e p r e s e n t s t u d y were possiblycausedby some i n t e r f e r e n c e or
i n t e r a c t i o n of the test f i x t u r e . Material behavior was s l i g h t l yn o n l i n e a r t o
f a i l u r e , similar to r e s u l t s of t h ef o u r - p o i n tf l e x u r e tests a s , n o t e d earlier.
11
14. Back-to-back s t r e s s - s t r a i n d a t a d i d n o t predict t h eo n s e t of localbuckling
(wrinkling)andtheuse of t h e force s t i f f n e s s methodofreference 15 to p r e d i c t
wrinkling was unsuccessful. A l l p a n e lf a i l u r e s were abruptwithnoindication
of local i n s t a b i l i t y . It wouldprobablybenecessary to extensivelyinstrument
bothsides of a facing to calculatefacingbendingstrainsandpredict local
buckling.Modulus values a t 0.2 p e r c e n ts t r a i n , maximum back-to-back s t r a i n
v a r i a t i o n a t 0.6 p e r c e n ts t r a i n ,t h e o r e t i c a lw r i n k l i n g stress, andexperimental
ultimate stress andstrainvalues of eachpanel are p r e s e n t e di nt a b l e V I I .
Maximum back-to-back s t r a i n v a r i a t i o n was f a i r l y low, c o n s i d e r i n gt h es i z ea n d
complexity of thesandwichpanels.Compressivemodulusvalues a t 0.2 percent
strainofthesandwichpanelswhichusedCelion 3000 material were s l i g h t l y
higherthan results of beam tests whichusedtheCelion 6000 material; theaver-
agemodulus of a l l wrinklingspecimens is 53.9 GPa (7.82 x lo6 psi)as compared
to 48.95 GPa (7.1 0 x 1 O6 psi) obtainedusingthefour-point beam f l e x u r e test
method.Sincethe fiber volume f r a c t i o n Vf of the beam specimens was higher
thanthatofthebucklingspecimens (72 percentcomparedwithapproximately
61 percent) it appearsthatthethinnergageCelion 3000 m a t e r i a ld i dn o t
experienceanydegradationinmodulus. R e s u l t s of replicate tests i n d i c a t e
t h a t scatter was low. S c a t t e ri n c r i t i c a l wrinkling stress ranged from a mini-
mum of 7.6 MPa (1.1 k s i ) f o rt h e 1.27 cm (0.5 in.)specimens to a maximum of
89 MPa (13 k s i ) for t h e 2.54 c m (1.OO in.)specimens.Thisamounts to a range
from minimum to maximum of 1.7 to 29 percent,respectively, when compared to
a v e r a g ec r i t i c a l stress values. From t a b l e s I1 and V I 1 some t r e n d si n results
areevident:
1. Averagefailure stresses ofthewrinklingspecimensdecreaseas core
height tc increases.This is c h a r a c t e r i s t i c of a wrinkling or local buckling
typeofinstability.Averagefailure stresses were 452, 354, and 311 MPa
(65.6, 51 .4, and 45.1 k s i ) f o rt h e 1 .27-, 1 .91-, and 2.54-cm (0.50-, 0.75-, and
1.00-in.)thick cores, respectively.
2. Average f a i l u r e s t r a i n s were 0.87, 0.71, and 0.63 percent for t h e 1.27-,
1 .91-, and 2.54-cm (0.50-, 0.75-, and1.00-in.)thick cores, respectively.
3. Specimenswithhigher total facingthicknesses had h i g h e rf a i l u r e loads;
however,thesespecimensdidnotnecessarilyhavehigherfailure stresses.
This is becausethethickerfacings had a lower f i b e r volume f r a c t i o n Vf
becausenotenoughresin was removed duringtheconsolidationphaseoflaminate
f a b r i c a t i o n .
4. Panelswiththelargestvalue of i n i t i a lw a v i n e s s had t h e lowest
ultimate load.
5. Ultimate strainsofthewrinklingspecimens were well below ultimate
l a m i n a t es t r a i n results from t h e beam tests.
A s mentioned earlier, most ofthewrinklingspecimensfailed close t o the
end of oneoftheside simple supports. F a i l u r e of a 1.27-cm (0.50-in.)panel
is i l l u s t r a t e d i n f i g u r e s 21 (a) and(b) ; t h ef a i l u r ee x t e n d s across thepanel to
t h e top o ft h el e f t - s i d e simple support. The f a i l u r e s were nearlyperpendicular
to t h ed i r e c t i o n of load.Wrinklingfailure was most n o t i c e a b l ei nt h e 1.27-cm
(0.50-in.)specimensinwhichthefacingsseparated from t h e core due to a ten-
12
15. sile failureof the adhesive. Failed panel BT-4 (fig. 22(a)) illustrates the
outward buckling of the facing; the panel wascut along the dashed line in that
figure to further illustrate the tensile failure of the adhesive which was pr
cipitated by wrinkling (fig, 22(b)). Figure 23 is a side view of two different
panels (tc = 1.27 cm (0.50 in.)). It is not conclusive from the side views
whether the failures were symmetricor antisymmetric; however, laminate failures
on either facing were similar which suggests that failures were symmetric.This
agrees with results of references4 and 16 which indicate that for honeycomb
cores, where the modulusof the core in the directionof the load is muchless
than the modulusof the core in the direction perpendicular to the facings,s y m
metric wrinkling will occur at a lower load than that for antisymmetric
wrinkling.
Overall buckling specimens.- Experimental results of overall buckling
specimens are presented in table VI11 and figures 24(a), (b), and (c). The
experimental critical overall buckling stress was determined from the applied
load associated with the reversal of extreme fiberstrain on the convex side
of the buckled panel. The specimens exhibited a very short postbuckling
region as evidenced by the experimental resultsof PC, and Pult as shown
in table VIII. Average values of Pcrl P,lt, Ucr, and Uult are 101.9 kN
(22 903 lbf),106.3 kN (23 897 lbf), 264.1 MPa (38.3 ksi), and 275.5 MPa
(39.96 ksi), respectively, and corresponding scatter is 21, 20, 21, and
20 percent.
Similar to results of reference4 , all the overall buckling specimens
failed on the concave side of the specimen in a typical compressive failure
mode. Most of the specimens failed in the center, all the failures were per-
pendicular to the direction of load as shown in figure 25.The MoirL method
was useful in visualizing the deflected mode shapes of the specimens and
determining the effectiveness of the simple supports. Panel number BT-2 was
the only specimen which failed near a simple support. Photographs of Moir4
fringe patterns of panel BT-2 indicated that it did not deform symmetrically
in half sinewaves in the length and width directions as expected. The out-
of-plane deformation of panel BT-2 with increasing load is illustrated in
figures 26(a), (b), and (c). As shown, the peak out-of-plane deformation
occurs in the upper right-hand portion of the specimen.This panel eventually
failed near the lower left-handsimple support. All other specimens failed in
the center. Moirg fringe patterns ofa typical buckling specimen are shown in
figures 27(a) to (d) for increasing load. As shown, the maximum out-of-plane
displacement does occur in the center of the panel. Displacements seem tobe
symmetric in the longitudinal direction; however, nonzero displacements appear
to occur near the right-hand simple support. Displacementsdo occur at the
corners ofthe panel since the simple supportsdo not extend the total panel
length. As the panel approaches failure, mode shapes tend to be nonsymmetric
(fig. 27(d)). As mentioned in reference4, it is very difficult to simulate
true simply supported boundaries when the buckled mode shape occurs at
m = n = 1 or the buckled shape is a half sine wavein the length and width
direction.
Comparison of analytical a,ndexperimental results.- The analysis assumes
the r&m-temperature unidirectional properties and dimensions listed in
table IX.
13
16. - I
From laminatetheory Ex = Ey = Ef = 51 .97 GPa (7.538 x 1O6 psi) and
pxy = 0.3075. These results agreewithexperimental results fromthe
sandwich beam f l e x u r e tests i n whichtheaveragemodulus, .Ex = 48.95 GPa
(7.1 x lo6 psi) and bv = 0.347. Sincelaminates were quasi-isotropic,
symmetric ([0,+45,90Is), A16 and A26 coupling terms of thesandwich were
identicallyzero,andthe Dl6 and D26 coupling terms were n e g l i g i b l e . Ana-
l y t i c a lr e s u l t s ,a s s u m i n g a laminathicknessof 0.0076 cm (0.003 i n . ) , are
presentedandcomparedwithexperimentalresultsintables VI1 and VI11 and
i nf i g u r e 28. The o v e r a l lb u c k l i n ga n a l y s i sd e s c r i b e di nr e f e r e n c e 7, which
includedthe core s h e a rf l e x i b i l i t y ,a g r e e d well withexperimentaloverall
bucklingresults. Local andgeneralbucklingformulas used i nt h ep r e s e n t
a n a l y s i s are presentedinappendix B.
-
-
The averageexperimentaloverallbuckling stress was 264 MPa (38.3 k s i )
andcompared exactlywiththeanalyticallypredictedoverallbuckling stress.
From experimentalwrinkling results it appearsthatequations (B5), (B6) and
(B8) were unconservativeand impractical to u s e from a designstandpoint.
Equation (B4), however, was conservativein its predictionof symmetric wrin-
klingloadsand is usefulfordesignpurposes.Wrinkling results obtained by
usingequation (B6) andassuming 6max = 0.01 cm (0.004 in.) were 7, 26, and
32 percenthigherthanexperimental results for t h e 1.27-, 1 .91-, and 2.54-cm
(0.50-, 0.75-, and1.00-in.)thick cores, respectively. The equation for
overallbuckling is equation (B11) and, as explained earlier, when core shear
f l e x i b i l i t y is accounted for, t h er e s u l t su s i n gt h i se q u a t i o na g r e ee x a c t l y
withaverageexperimentalvalues of Dcr.
CONCLUDING REMARKS
The purposeofthepresentstudy was to investigatethebucklingbehavior,
l o c a l andgeneral,ofgraphite/polyimide(Gr/PI)sandwichpanelscapableofuse
a t temperaturesranging from 1 1 6 to 589 K (-250 to 600°F) as thesandwichskin
ofthespaceshuttle body f l a p . The adhesiveandfacingmaterialproperties
wereinvestigatedandbucklingformulasforpredicting local andgeneralsand-
w i c hp a n e li n s t a b i l i t i e s were evaluated. R e s u l t s of a bond studyinclude a fab-
ricationtechniqueforadhesivelybondingsandwichstructuresand a cure c y c l e
f o r FM-34 filmadhesivewhichproduced flatwise t e n s i l es t r e n g t h si ne x c e s so f
3.4 MPa (500 p s i ) a t 116 K and R.T. (-25O0F and R.T.) i n 1 .4 MPa (200 p s i ) a t
589 K (600OF). R e s u l t s also i n d i c a t et h a t a liquidcell-edgeadhesive (BR-34)
c a nr e s u l ti nc o n s i d e r a b l ep a n e l mass savings (10 percent)withoutnecessarily
s a c r i f i c i n g bond strength;however,furtherresearch is necessarysinceflatwise-
t e n s i l es t r e n g t h su s i n g BR-34 werenotconsistent.Materialproperty tests of
quasi-isotropic, symmetric laminates ([0,+45,90Is) ofCelion 600.0/PMR-15 Gr/PI
material i n d i c a t et h a t it m a i n t a i n ss u i t a b l es t r u c t u r a lp r o p e r t i e sf o rs h o r t -
term use a t temperatures from 1 1 6 to 589 K (-250° t o 60O0F).
Experimentalresultsofflatrectangular honeycombsandwichpanelswhich
were simplysupportedalong a l l fouredgesand tested inuniaxialedgewise com-
p r e s s i o ni n d i c a t et h a t two modes ofpanelfailure,wrinkling or o v e r a l l buck-
ling,can occur dependingonthe core thickness. A s p r e d i c t e da n a l y t i c a l l y ,
specimenswith a core thicknessof 0.635 cm (0.25 i n . )f a i l e d by o v e r a l l buck-
lingand a l l otherspecimens,havingnominal core thicknessesof 1 .27, 1.91 ,
1 4
17. and2.54 c m (0.50 , 0.75,and 1.00 i n . ) , failed by wrinkling.Theshadow Moire/
method was u s e f u l i n determining mode shapesoftheoverallbucklingspecimens
but was n o t able to d e t e c t w r i n k l i n g .
R e s u l t s of thewrinkling tests i n d i c a t e d t h a t s e v e r a l a n a l y t i c a l m e t h o d s
were unconservativeandthereforenotsuitable for designpurposes. Most of
thewrinklingspecimensfailedneartheside simple supports.Thefailure mode
appeared t o be symmetric wrinkling with failures occurring because of tensile
rupture of theadhesive. Some t r e n d si nw r i n k l i n g results are:
1. Averagefailure stresses of thewrinklingspecimensdecrease as core
thicknessincreasesand are 452,354,and 311 MPa (65.6,51.4,and45.1 k s i )
for the1.27-,1.91-,and 2.54-cm (0.50-, 0.75,and1.00-in. ) t h i c k cores,
respectively.
2. A v e r a g ef a i l u r es t r a i n s were 0.87, 0.71,and0.63percent for t h e 1.27-,
1.91-,and 2.54-cm (0.50-,0.75-,and1.00-in.) t h i c k cores, respectively.
3. Panelswiththelargestvalueofinitialwavinesshadthe lowest
ultimateload.
The averageexperimentaloverallbuckling stress ofthe 0.635-cm (0.25-in.)
t h i c k specimens was 264 MPa (38.3 k s i ) andcomparedexactlywiththeanalyt-
i c a l l yp r e d i c t e do v e r a l lb u c k l i n g stress. A l l theoverallbucklingspecimens
exceptone failed inthecenterontheconcavefacing by compression.
LangleyResearchCenter
NationalAeronauticsandSpaceAdministration
Hampton, VA 23665
December2,1980
15
18. APPENDIX A
DESIGN CONSIDERATIONSFOR BUCKLING SPECIMENS
Preliminary studiesof structural loadson the shuttle body flap (ref.3)
indicate that compression loads are the primary design condition and that a
biaxial state of stress is present. For this reason a sandwich panel design was
chosen. Furthermore, based on the low magnitude and biaxial nature of stresses,
minimum-gage symmetric laminatesof [0,+45,90Is Gr/PI werechosen for the
facings of the sandwich skin of thebody'flap. Therefore, buckling specimens
similar to the sandwichskin ofthe body flap areexamined in the present study.
Only symmetric laminates were consideredin the present investigationto
prevent laminate warpage during the cure cycle causedby bending-stretching
coupling terms (nonzero [B] matrix of the material). If nonsymmetric laminates
such as [ 0,+45,901 could be fabricated and forced flat and bonded symmetrically
with respect to the center line of the core, this would reduce the mass of the
panel and may besufficient to accommodate the low loads predicted for the body
flap. However, analysis techniques would have to be generalized to include
anisotropic facings as was done in reference4. Because of these fabrication
uncertainties, however, nonsymmetric laminates were not considered for the
experimental study.
Thin-gage Celion 3000 material was chosen because it would present a sub-
stantial mass savings over the thicker gage Celion 6000 material. Average
thickness per ply of the Celion3000 laminates were0.007 cm (0.0028 in.) as
compared to 0.0165 cm (0.0065 in.) for Celion 6000.
The lowest density commercially available core which can function structur-
ally at 589K (600OF) is either Hexcel HRH-327-3/16-4 or HRH-327-3/8-4 glass/PI
which has a density of64 kg/m3 (4 lbm/ft3) and either a 0.5 cm (3/16 in. ) or
a 0.95 cm (3/8 in.) cell size,respectively. Both of these cases were examined
in the analytical investigation.
Simply supported boundary conditions and uniaxial edgewise compression
loading were chosenat test conditions because they closely represent conditions
actual shuttlebody flap panels will experience. Both overall and local panel
buckling modes were considered in the analysis. Elements of the[AI and [Dl
matrices were calculated for the quasi-isotropic, symmetric Gr/PI sandwich based
on laminate theory presented in references9, 17,and 18. Overall buckling
equations (ref. 7) were minimized with respect to m and n, to predict overall
panel buckling load (assuming both infinite and finite core shear stiffness);
the local instability equationsof reference 8 were used to predict local insta-
bility modes and associated loads. The local and general buckling equations are
also presented in appendixB. Buckling loads were computed for various ply
thicknesses, core thicknesses,and operating temperatures.
Unidirectional laminate material propertiesused in the designof the buck-
ling specimens were obtained from references 19 and 20. Honeycomb core material
properties were obtained from reference5. Some of the material properties used
in the present analysis are presented in table IX. Various cores and core
16
19. __
APPENDIX A
thicknesses(0.635 to 2.54 c m (0.25 to 1.00in.))andpanellengthsandwidths
(10.2 to 122 cm ( 4 . 0 to 48.0 in.) ) were a n a l y t i c a l l y i n v e s t i g a t e d a t various
temperatures (room temperature to 589 K (6000F)),anddesignenvelopes,
t y p i f i e d by f i g u r e 28, were determined.Sincethe laminate o r i e n t a t i o no ft h e
f a c i n g s is quasi-isotropicand symmetric, theaverage elastic modulus Ex or
Ey was used for thefacing-modulus Ef equationsinappendix B. Resultsof
critical stress as a functionof core thicknessforanassumedplythicknessof
0.0076 c m (0.003in.) are shown i nf i g u r e 28.
-
-
The designenvelopecurvesinfigure 28 i n d i c a t et h a te i t h e ro v e r a l l buck-
ling,dimpling,laminatestrength, or wrinkling could be c r i t i c a l f a i l u r e modes
dependingon scatter i nm a t e r i a lp r o p e r t i e sa n dd i f f e r e n ta n a l y s i st e c h n i q u e s .
Since it is d e s i r a b l e to v e r i f y as many a n a l y t i c a lp r e d i c t i o n s for various
f a i l u r e modes as p o s s i b l e ,t h e honeycomb core withthelarger cell size(0.95 cm
(3/8in.) ) was chosen because it lowers thedimpling stress to values closer to
t h eo t h e r critical stresses. A p a n e ls i z e of 30.5by30.5 cm (12 by 12in.) was
adequate to i n v e s t i g a t e s e v e r a l f a i l u r e modes.
Thehoneycomb core neartheloadedendsofthespecimens was potted with
BR-34 liquidpolyimideadhesive,andtaperedendtabs of [+451glass/PIwere
bonded a t eachend to prevent local end f a i l u r e ss u c ha s core crushing or end
brooming; scalloped doublers were bondedbeneaththeend tabs to enhanceload
diffusionintothepanelandhelpreduce stress concentrations. A s t a i n l e s s -
steel sheet was embedded i n t h e BR-34 p o t t i n g a t each end to a l i g nt h es p e c i -
mens intheknifeedges.Laminates were bonded t o t h e core andendtabsand
doublers were secondarybondedusing FM-34 film adhesive.Significantpanel
parameters, related to thefabricationandqualityofthewrinklingandover-
a l l bucklingspecimenssuch as fiber andvoidvolumefractionsandtheglass
transitiontemperature,arepresentedin table 11. Facingandtotalsandwich
panelthicknessmeasurements were made a t v a r i o u sp a n e ll o c a t i o n sa n di n i t i a l
panelwaviness 6 was measured as explainedinreference 9. Because of good
f a b r i c a t i o n and qualitycontrolproceduresthepanelswereconsistentin
dimensionalandmaterialproperties.Averagethicknessperply of a l l wrin-
klingspecimens was 0.0079 c m (0.0028in.)with maximum v a r i a t i o n si n t o t a l
laminate(8plies)thicknessesaveragingonly 0.00451 c m (0.00178 i n . ) ; average
v a r i a t i o ni n total sandwichpanelthicknesses was only 0.0059 cm (0.0023in.).
Maximum panelwaviness 6max averagedonly0.0097 c m (0.0038in.).
17
20. APPENDIX B
BUCKLING FORMULAS USED INTHE PRESENT STUDY
There are several instability modes whichcan cause failureof a sandwich
structure; as shown in figure 29 they are: intracellular buckling (face dim-
pling), face wrinkling (either symmetricor antisymmetric), and shear crimping.
Intracellular buckling is a localized mode of instability which occursonly when
the core is not continuous, as in the caseof honeycomb or corrugated cores. As
shown in figure 29(a), the facings buckle in a platelike fashion directlyabove
core cells, with cell edges actingas edge supports. These buckles can deform
sufficiently to cause permanent, plastic deformations and can eventually lea
the face wrinkling instability mode (fig. 29(b)).The face wrinkling mode is a
localized buckling of the facings in which the wavelength of the buckles is o
the same order as the thicknessof the core. Depending on the nature of the
material properties of the core the facings can buckle symmetrically or antis
metrically. For the honeycomb cores, in which the elastic modulus parallel to
the facings is verylow compared with the modulus in the direction perpendicula
to the facings, failure is usuallyby symmetric wrinkling (ref. 16).
Shear crimping (fig. 29(c)) is considered to be a special form of general
instability for which the buckle wavelength is veryshort dueto a low trans-
verse shear modulus of the core.This mode occurs suddenlyand usually causes
the core to fail in shear; however, it may also cause a shear failure in the
core- to-f acing bond.
Overall buckling was calculated using the methodof reference 7 which
assumes simply supported boundary conditions and includes consideration of
core shear flexibility.
There are many references concerning the analysis and prediction of local
instability modes of failure of sandwich structures (refs. 4,6, 8, and 21 to
24). Formulas for predicting local instability vary among references and for
that reason several methods were used to predict local failure loads. An upper
and a lower bound were calculated for various failure modes and sandwich pa
thicknesses. The formulasfor local bucklingof a sandwich panel subject to
uniaxial compression and appropriate references are given as follows:
Dimpling :
From references8 and 22 to 24, for isotropic facings
18
21. From reference 4
APPENDIX B
From reference 4 , assumingorthotropicfaces
odim = 0.825 - (B3)
3 4 (1 - iixyiiyx)
where Ef, and Efy are thefacingmoduliinthe x- andy-direction, respec-
tively,and is thefacingshearmodulusinthexy-plane.
Gfxy
Forisotropicfaces,equation(B3) reduces to
Facing wrinkling ( symmetric) :
From references 11 and23,the lower boundonwrinklingstress is
and t h e upperbound is
where is themodulusofthe core i nt h ed i r e c t i o nn o r m a l to t h ef a c i n g s
and tc is thethicknessofthe core,
From r e f e r e n c e2 3 ,a c c o u n t i n gf o ri n i t i a lf a c i n gi m p e r f e c t i o n s
19
22. APPENDIX B
whereFc is theflatwisesandwichstrengthand 6 is t h ea m p l i t u d eo fi n i t i a l
wavinessinthefacings.
t C
tf
From r e f e r e n c e8 ,f o r - 50
Uwr = 0.5 (Gc XZECZEf)'I3
t C
tf
and for - > 50
From reference 4
+
Shearcrimping:
From reference 8
and from reference22,
20
.. ..".. I
24. APPENDIX B
The buckling loadof the sandwich is obtained from equations(B11 ) to (B14)
by minimizing with respect to m and n, the number of half-waves in the
buckle pattern in the length and width directionsof the plate, respectively.
The smallest n consistent with the assumptionof simply supported plates is
n = 1.
22
25. REFERENCES
1. Davis, JohnG., Jr., compiler: Composites for Advanced Space Transportation
Systems - (CASTS). NASA TM-80038, 1979.
2. Dexter, H. Benson; and Davis, JohnG., Jr., eds.: Graphite/Polyimide Com-
posites. NASA CP-2079, 1979.
3. Design and Test Requirements for the Applicationof Composites to Space
Shuttle Orbiter. Volume I. Summary. SD 75-SA-0178-1, Space Div.,
Rockwell International Corp., Dec. 3, 1975.
4. Pearce, T. R. A.: The Stability of Simply-Supported Sandwich Panels With
Fibre Reinforced Faceplates. Ph. D. Thesis, Univ. of Bristol, 1973.
5. Mechanical Properties of Hexcel Honeycomb Materials. TSB 120, Hexcel Corp.,
c.1975.
6. Plantema, Frederik J.: Sandwich Construction. The Bending and Buckling of
Sandwich Beams, Plates, and Shells. John Wiley& Sons, Inca,c.1966.
7. Peterson, JamesP.: Plastic Buckling of Plates and Shells Under Biaxial
Loading. NASA TN D-4706, 1968.
8. Sullins, R. T.; Smith, G. W.; and Spier, E. E.: Manual for Structural
Stability Analysisof Sandwich Platesand Shells. NASA CR-1457,1 969.
9. Camarda, Charles J.: Experimental Investigation of Graphite/Polyimide Sand-
wich Panels in Edgewise Compression. NASA TM-89895, 1980.
10. Raju, B. Basava; Camarda, CharlesJ.; and Cooper, PaulA.: Elevated-
Temperature Application of the IITRI Compression Test Fixturefor
Graphite/Polyimide Filamentary Composites. NASA TP-1496, 1979.
11. Dykes, B. C.: Analysis of Displacements in Large Platesby the Grid-Shadow
Moire'Technique. Experimental Stress Analysisand Its Influenceon
Design, M. L. Meyer, ed., Inst. Mech. Eng., c.1971, pp. 125-1 34.
12. Chiang, Fu-Pen: Moire'Methods of Strain Analysis. Exp. Mech., vol. 19,
no. 8, Aug. 1979, pp. 290-308.
13. Poesch, JonG.: Development of Lightweight Graphite/Polyimide Sandwich
Panels. Non-Metallic Materials, Volume 4of National SAMPE Technical
Conference Series, SOC. Aerosp. Mater. & Process Eng., c.1972,
pp. 605-61 4.
1 4 . Shuart, Mark J.; and Herakovich, CarlT.: An Evaluation ofthe Sandwich
Beam in Four-Point Bending as a Compressive Test Method for Composites.
NASA TM-78783,1 978.
23
26. 15. Jones, Robert E.; and Greene, BruceE.: The Force/Stiffness Technique for
Nondestructive Buckling Testing. A Collectionof Technical Papers-
AIAA/ASME/SAE 15th Structures, Structural Dynamics and Materials Confer-
ence, Apr. 1974. (Available as AIAA Paper 74-351.)
16. Norris, Charles B.; Ericksen, Wilhelm S.; March, H. W.; Smith, C. B.; and
Boller, KennethH.: Wrinkling of the Facings of Sandwich Constructions
Subjected to Edgewise Compression. Rep. No. 1810, Forest Prod. Lab.,
U.S. Dep. Agr., Nov. 1949.
17. Jones, RobertM.: Mechanics of Composite Materials. McGraw-Hill Book Co.,
c. 1975.
18. Ashton, J. E.; Halpin, J. C.; and Petit, P. H.: Primer on Composite Mate-
rials: Analysis. Technomic Pub. Co., Inc., c.1969.
19. Hanson, Morgan P.; and Chamis, ChristosC.: Graphite-Polyimide Composite
for Application to Aircraft Engines. NASATN D-7698, 1974.
20. Advanced Composites Design Guide. Volume IV - Materials. Third ed.,
Air Force Materials Lab., U.S. Air Force, Jan. 1973. (Available from
DTIC as AD 916 6821;.1
21. Allen, HowardG.: Analysis and Design of Structural Sandwich Panels.
Pergamon Press, 1969.
22. Sandwich Construction for Aircraft. ANC-23, Pts. I and11, Air Force-Navy-
Civil Subcommitteeon Aircraft Design Criteria,1951.
Part I- Fabrication, Inspection, Durability and Repair.
Part I1- Materials Properties and Design Criteria.
23. Honeycomb and Prepreg in Sandwich Construction. TSB 100, Hexcel Corp.,
c. 1974.
24. Military Standardization Handbook- Structural Sandwich Composites.
MIL-HDBK-23A, Dec. 30, 1968.
24
27. TABLE I.- CURE:CYCLES OF FLATWISE-TENSILESPECIMENS
.. ". .. - ~~
Description
~~
~.
Vacuum + 0.34 MPa (50 psi) a t R.T.
Cure to 589 K ( 600°F) a t 5 K/min (g°F/min)andhold f o r 2 hours
No post c u r e
Vacuum + 0.34 MPa (50 p s i ) a t R.T.
Cure to 450 K (350OF) a t 5 K/min (g°F/min) andhold for 2 hours
P o s tc u r e a t 589 K (600OF)andhold a t temperature for 2 hours
with clamps
._-
Same as c u r e cycle 1 butbond top andbottomfacings separately
. . "
w i t hf a c i n g s to bebondedon bottom
~ . . " - .. . .. - -.. . .. . . . .- . . . . .. . .--
Vacuum + 0.34 MPa (50 psi) a t R.T.
Cure to 61 6 K ( 650°F) a t 5 K/min (g0F/min) andhold for 1 .5 hours
Same as c u r e cycle 1 b u td o n ' t apply vacuum
~~
. . . . .
, . . . . -.. ~~
25
30. TABLE 111.- FLATWISE-TJ3NSILE TEST RESULTS OF CURE-CYCLE BONDSTUDY
[FM-34 film adhesive; Core density= 96 kg/m3 (6 lbm/ft3)l
. .. . . . . ... . . -
Specimen
1 1 kN(1bf) MPa(psi) failure
Pultr
aultr I Descriptionof~~ ~
Tests at room temperature
~~
-
FTT-1 3.25 (472) 1 Failed between facing and core,18.90 (4250)1
Facing delaminated also.
FTT-2 Failed between facing' andcore<4.02 (583)23.35 (5250)1
Facing delaminated.
FTT-3
1Facing delamination.
2.80 (406)16.24 (3650)2
FTT-4 Failed between facing and core.3.41 (494)19.79 (4450)2
~~
FTT-5 Facing delamination.2.34 (339)13.57 (3050)4
FTT- 6 Facing delamination.3.05 (442)17.70 (3980)4
FTT-7 Failed between facing and core.3.64 (528)21.1 3 (4750)5
FTT-8 Facing delamination.4.00 (580)23.22 (5220)5
FTT- 9
-
3 Failed second bond between2.91 (422)16.90 (3800)
facing and core.
FTT-10 Failed second bond between2.76 (400)16.01 (3600)3
I !facing and core.
~~
Tests at 589 K (600OF)
FTT-11 1.37 (198)7.918 (1780)1
FTT-12 1.44 (209)8.363 (1880)1
FTT-13 ' 1.94 (281 ). 1 1 .23 (2525)al
~~
FTT-14 1.28 (186)7.451 (1 675)a1
FTT-15 0.848 (123)4.938 (1110)5
FTT-16 1 .06 (153)6.139 (1380)5
. . . ~~~
aSame as cure cycle1 but cured to603
I
I
I
!
E
~~ ~
~~ __
Failed between facing and core.
Failed between facing and core.
. .
Failed between facing and core.
Failed between facing and core.
Failed between facing and core.
~~
~~
-
Failed between facing and core.
C (625OF).
28
31. TABLE 1V.- FLATWISE-TENSILE TEST RESULTS
(a) FM-34 film adhesive; Cure cycle1 with cure temperature= 603 K (625OF);
13 = 0.586 kg/m2 (0.12 lbm/ft2)
I 1Core
density,
(lbm/ft3)
kg/m3
Temperature,
' K
(OF)
R.T.
R.T.
589 (600)
589 (600)
589 (600)
116 (-250)
116 (-250)
R.T.
116 (-250)
"
R.T.
R.T.
589 (600)
589 (600)
589 (600)
116 (-250)
-
116 (-250)
kN (lbm)
Pult Uult, Description of
MPa (psi) failure
Spec imen
FTr-17 13.12 (2950) 2.26 (328) I Facing delamination. I
FTT-18 21 .80 (4900) 3.75 (544) I Facing delamination.
I
I
8.807 (1 980)FTT-19
FTr-20
1 .52 (220) Failed between facing
and core.
1 .33 (1 94) Failed between facing
and core.
0.58 (84) Failed between end-
biock and facing.
7.784 (1 750)
3.38 (760)FlT-21
FTr-22
. . .
FTT-23
FTT-24
FTT-25
96 (6) 1 .ll (250) 1 .92 (278) Failed between end-
block and facing.
4.448 (1 000) 0.765 (11 1
18.24 (4100) 3.14 (456
3.18 (461
Failed between facing
and core.
Facing delamination.128 (8) 18.46 (41 50)
FI'T-26
FTT-27
FTT-28
FTT- 2 9
FTT-30
FTT-31
FTT-32
128 (8) 13.34 (3000) 2.30(333)Failedbetweenfacing
and core.
128 (8) 3.83 (556) I Facing delamination. I22.24 (5000)
7.651 (1 720)128 (8)
.~
128 ( 8 )
1 .32 (191) Failed between facing
and core.
1 .46 (211 ) Failed between facing
and core.
8.451 (1 900)
128 (8) 8.051 (1 810) 1 .39 (201) Failed between facing
and core.
128 (8)
128 (8)
18.90 (4250)
24.24 (5450)
and core.
29
I
32. TABLE 1V.- Concluded
(b) Br-34 cell-edge adhesive: Cure cycle1 with cure temperature= 603 K
(625OF); R.T.; p = 0.244 kg/m2 (0.05 lbm/ft2)
1
Core
density,
(lbm/ft3:
Specimen
kg/m3
I'
l"rFFIT- 34
t I
FIT-41
c
I " - EFTT-45
11
TT
i
T
T
-r
I
kN (lbm)
Pult
9.1 86 (2065)
5.627 (1 265)
5.783 ( 1 300)
- ~~~
11 .30 (2540)
3.09 (695)
11 .23 (2525)
12.41 (2790)
1 2.86 (2890)
17.68 (3975)
16.22 (3647)
""""""
""""""
20.68 (4650)
19.48 (4380)
MPa (psi)
Uult
1 .58 (229)
0.97 (141)
0.99 (1 44)
I .94 (282)
3.53 (77)
I .94 (281)
2.14 (310)
!.21 (321 )
1.57 (51 7)
1.31 (487)
-
Description of
failure
~
Failed between block and facing
__
Failed between block and facing
Failed between block andfacing,
Facing delamination (localized
__
around cell edges).
Facing delamination (localized
around cell edges).
- .
Facing delamination (1c::alized
around cell edges).
_ _ _ _ ~
Facing delamination (localized
."
around cell edges).
~-.
Facing delamination (localized
around cell edges.)
Failed between facing and core.
?acing delamination (not
. .
localized).
?ailed immediately at very low
load between block and facing.
-
?ailed immediately at very low
load between block and facing.
Pacing delamination.
Pacing delamination.
30
33. TABLE V.- COEFFICIENTS OF POLYNOMIALS USED TO CURVE-FIT DATA
1 1
COI
Testconditions ' Pa
c1' c2
PaPa
c3
Standard errorof
Pa estimate 1
(psi)(psi)(psi)(psi)psi m a 1
- ,d
' Room temperature-3.695E+4 ' 5.012E+102.293E+12-2.327+14792.187 ' 5.46
tension (1 to 4) ~ (-5.3593+0) j (7.2693+6) (3.3253+8) : (-3.375E+lO)
Roomtemperature1.964E+65.047E+10-2.9003+11 , -3.330E+131609.543 ~ 11 .10 :
compression(2.8483+2) ' (7.320E+6)(-4.200E+7) I (-4.8303+9)
(5 to 8)
I I I
Low temperature ' -2.9443+56.194E+lO I 3.924E+lO-7.619E+13 980.4746.76
"-1
~
tension(-4.270E+l)(8.9833+6)(5.692E+6) (-1 .105E+10)
(9to I
1
Low temperature-2.4503+65.887E+10-5.273E+11 -1 .549E+13 1 271.9688.77
compression(-3.5333+2)(8.5383+6)(-7.6483+7) ' (-2.2473+9)
(13 to 16) I
Hightemperature-4.3603+5 , 5.291E+10 ' 8.894E+ll 1 -1.430Et141399.392 ' 9.65
tension(-6.3243+1)
(1 7 to 20)
(7.674E+6)(1.290E+8)(-2.074E+10)
High temperature 1546.97710.67-2.3723+13 7.6603+11 4.648E+lO 4.61 3E+6
compression
(21 to 24)
(-3.441E+9)(1.111E+8)(6.741E+6)(-6.6913+2)
!
34. TABLE VI .-SUMMARY OF SANDWICH-BEAM-FUXURE TESTS OF[ 0,+45,90,-451
Celion 6000/PMR-15
(a) SI Units
I
Test pav atP atE = 0.002, Culta",Cult,UUlt,"'UUlt'Temperature,
E at
GPa
Specimen
condition percentpercentMPaMPaK E = 0.002 E = 0.002
BTF-1
BTF-3 595.9 .354
BTF-2 569.2 .312
0.3330.34356.54(a)(a)565.2 539.2R.T.Tens ion
556.6 .322
t 590.5-1.579
BTF-7 557.9-1.328
523.3 -1 .227.368
BTF-9Tension 1 1 6 579.1613.5 (a) (a) 61.36 (a) 0.329
BTF-10
1 1
(a) 0.343
, BTF-11 ,332
BTF-12 600.0 .312
BTF-13Compression 91.5666.1646.2 -1 .368 -1 .28556.540.3340.337
, BTF-14 1 36 618.9 -1 .249.345
BTF-15 1 1 6679.2 (a) .31 3
BTF-16 116 620.5 -1 .237.356
BTF-17Tension589 318.7322.8 0,626 0.608 54.470.2890.344
BTF-18
BTF-19 346.2,653 .354
BTF-20 308.7.573.367
-1- 317.5 .580.366
I
: BTF-21 Compression589 296.5334.0-0.644-0.65748.950.3880.382
BTF-22 423.0 -.690 .376
BTF-23 -.696,376
BTF-24 278.0 ' -.657.388IaGage malfunction.
35. TABLE VI.- Concluded
(b) U.S. Customary Units
1I
E
percent E = 0.002 E = 0.002
E at
psi
ultav' E = 0.002, U at I-lav at
.A,
TestTemperature I
condition OF
Cult
percent
%lt, 'Sulta,,
PS1 PS1
78.20
82.55
81 .98
86.42
80.73
86.87 82.34
85.65
80.92
75.90
83.99 88.97
95.91
87.02
96.60 93.72
89.77
98.51
90.00
46.22 46.81
46.05
50.21
44.77
43.00 48.44
61 .35
49.08
40.33
(a)
Specimen
4
Tension R.T. (a) 8.2 x 106 0.343
.312
.354
.322
0.333(a)
I
I
-1 .392
-1 .579
-1 .328
-1 .227
(a')
BTF-1
BTF-2
BTF-3
BTF-4
BTF-5
BTF-6
BTF-7
BTF-8
BTF-9
BTF-10
BTF-11
BTF-12
BTF-13
BTF-14
BTF-15
BTF-16
BTF-17
BTF-18
BTF-19
BTF-20
BTF-21
BTF-22
BTF-23
BTF-24
-c -L
0.347Compression ~ R.T. -1.3817.1 x 1 O6 0.350
,313
.356
.368
(a)
0.343
.332
.31 2
0.334
.345
,313
.356
0.289
,366
,354
.367
0.388
.376
,376
.388
I c -i-
0.329Tension ~ -250 8.9 x' lo6
8.2 x lo6
1 -I-
0.337-295
-21 5
-250
250
-1 .368
-1 .249
(a)
-1 .237
-1 .285Compression
Tension
Compression
-L
0.626
.580
.653
.573
0.608 7.9 x 106 0.344600
161 5
0.3827.1 x lo6600
I
-0.657-0.644
-.690
-.696
-.599
W
W aGage malfunction.
36. W
I P
T
Pane1
BT-4
BT-5
BT- 6
BT-7
BT-8
BT-9
Core
thickness ,
t C
cm
( i n . )
~ ~~
1 .27
(0.50)
I
TABLE VII.- SUMMARY OF ROOM-TEMPERATURE WRINKLING PANEL RESULTS~
[ [ 0 , + 4 5 , 9 0 1 , Celion 3000/PMR-15 facings and HRH-327-3/8-4 Glass/PIcore]
Max. back-to-back
strainvariation
Experimental
E a t E = 0.002,
GPa
(psi) r:)c = 0.006r
MPa ( k s i ) r from equation -
percent
292.3
( 7 . 7 5 x 1 0 6 )( 4 2 . 4 )
(7.83 x l o 6 )
5 8 . 61 5 . 0
(8.5 x l o 6 )
1.91 52.9 25.0240.6
(0.75)(7.67 x l o 6 ) (34.9)
52.9 1 6 . 7
I (7.67 x l o 6 )
(8.00 x l o 6 ) I53.8 20 .o
BT-10.54
(1 .OO) (7.33 x 106)(30.1 )
BT-11
I
54.040.0
( 7 . 8 3 x l o 6 )
( 7 . 8 3 x l o 6 ) IBT-12 54.027.0
723.9484.0 I 454.7 0.89
(1 0 5 . )( 7 0 . 2 )‘ ( 6 5 . 9 5 )
447.5 .88
(64.90)
455.1 -83
(66.00)
I !
1 I (E;;)
597.1
(54.33)
(64.7)(51.66)(86.6)
0.70446.1356.1
374.6 .78
.66
51( 7 j . 8 ) 5 . 7411( 5 i . 7 ) . 6 3 6 5 . 8
0.77
276.5
289.9
.54
.57
(53.06)
(40.11)
(42.05)
1 I
aTheoreticaldimplingstressrange: 409.7 MPa (59.42 k s i ) to 641 . 2 MPa (93.0 k s i ) .
Laminate strength: 576 MPa ( 8 2 . 3 k s i ) .
Ultimatestrength: 1 . 4 percent.
1
pu1t
kN
(lbf1
1 5 9 . 3
:35820)
165.1
( 3 71 2 0 )
161 .3
(36270)
136.4
(30670)
146.6
(32960)
1 2 6 . 0
(28320)
141 .O
(31650)
105.1
(23640)
111.3
(25020)
37. TABLE V I 1 1 .- SUMMARY OF ROOM-TEMPERATURE RESULTS OF OVERALL BUCKLINGPANELS
kc = 0.635 c m (0.25 in.)]
Pane1
BT-1
BT-2
BT- 3
Pcr
kN
(Ib)
93.33
(20 981 )
114.38
(25 714)
97.92
(22 01 3)
Experimental
1
P u l t
kN
(lb)
97.64
(21 950)
19 8.76
(26 700)
102.50
(23 040)
results
~
Dcr
MPa
( k s i
241.3
(35.00)
296.5
(43.00)
254.4
(36.90)
%lt
MPa
( k s i )
252.5
(36.62)
307.8
(44.65)
266.3
(38.63)
TABLE I X .- PROPERTIES USED I N BUCKLINGANALYSIS
E11 = 133 GPa (19.3 X lo6 p s i )
E22 = 9.10 GPa (1 .32 x lo6 p s i )
y 2 = 0.37
1-121 = 0.025
G12 = 5.58 G P a (0.81 x lo6 p s i )
ECZ
G c X Z
GCyZ
= 0.345 GPa (50 x 1 O3 p s i )
= 0.200 G P a (29 x 1 O3 p s i )
= 0.083 GPa (12 x 1 O3 p s i )
Fc = 3.45 MPa (500 p s i )
Theoretical
%r1
MPa
( k s i )
264.7
(38.39)
I
35
38. CORE
.OOR 8,
-1fPMR-
b o )
-151
L-80-229
Figure 1.- Schematic diagram of flatwise-tensile specimen.
39. PORTIONS OF CORE FILLED
WITH Br-34
GrlPI 3.81(1.5)
FACING
55.88(22.00)
(1.25)
Figure 2.- Sandwich beam in four-point bending. A l l dimensions are in centimeters (inches).
45. x
',,,,
w
* 4
- 1
" r *
6.35 10.2 10.2 6.35
IN LOAD DIRECTION
SCALLOPEDDOUBLERS
Figure 7.- Schematic diagram of strain-gage locations on buckling specimens. Dimensions are in
centimeters (inches).
I P
W
49. EXPERIMENTAL
- REGRESS ION ANALYS I S
400 . / OU
40
311
STRESS, ksi
STRESS, MPa 3001
200
./ I L n
loo 0 -"
1 15 X 10
' 6
I
E T,GPa
25 tI I I I I
0 0.2 0.4 0.6 0.8 1
STRAIN, percent
10
ET, Psi
5
0
0
Figure 1 1 .-Tensile stressand tangent modulusbehavior of [0,+45,90,-45Is Celion 6000/PMR-15
at room temperature. Tests 1 to 4.
51. STRESS, MPa
600 -
500
400
-
200
300
100
0
I 100
EXPERIMENTAL 1,80
60
40
20
0
- REGRESSION ANALYS IS
STRESS,ksi
1 15 x lo6
ET, GPa
25
0
Figure 13.- Tensile
E ,psiT
- I 5
-0 0.2 0.4 0.6 0.8
STRAlN, percent
stress andtangentmodulusbehavior of [ 0,+45,90,-451 Celion 6000/PMR-15
a t 116 K (-250OF). Tests 9 to 12.
52. cn
0
1100
STRESS, MPa
80
60
40
STRESS,ksi
600
I I 1 I _
200
100 -KtC;Kt33 IUI HIIALY3 13 L W
0 * O
100 15 x lo6
75
-25
ET ,psi
-50
-
- 10
ET, GPa i
- 5
1 I I I I I
0
0
0.2 0.4 0.6 0.8 1.0 1.2 1.4
STRAIN, percent
Figure 14.- Compressive stress andtangentmodulusbehavior of [0,+45,90,-45],&lion 6 o o o / p ~ ~ -
a t 116 K (-250OF). Tests 13 to 16.
53. -1 100
600,r I
STRESS,M Pa
ET, GPa
5001I
EXPERIMENTAL -I 8o
400k - REGRESS IONANALYSIS "I
300- -
-
60
40
20
0
STRESS, ksi
100
-50
"75
,-
- 15 x lo6
1
- 10
ET, Psi
- 5
25 -
1 I I 1 I 1 I 0
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8
STRA IN, percent
Figure 15.- Tensile stress andtangentmodulusbehavior of [0,+45,90,-451, Celion 6000/PMR-15
a t 589 K (600OF). Tests 17 to 20.
54. 100
80
600 -
500 -
400 - 60 STRESS, ksi
300 -
40
200 -
100 - 20
0 0
STRESS, MPa
EXPERIMENTAL
-REGRESS ION ANALYS IS
100 15 x lo6
75
T-50
-
ET, GPa 5 ET, Psi
-
-10
25 -
I
0
I I I I
0.2 0.4 0.6 0.8 1.0 1. 2
10
STRAIN, percent
Figure16 .-Compressive stress andtangentmodulusbehavior of [ 0,+45,90,-451 Celion 6000/PMR-15
a t 589 K (600°F). T e s t s 21 to 24.
57. .6 -
.5 -
.4 -
""
0-
w
COMPRESSIVE
STRAIN, .3 -
E,
percent
(3"""
.2 - -I 0"""
"
I-
I
+ 25 1FRONT FACE
0 - . 2 5 ,
Ylb
(a) Panel BT-9.
Figure 19.- Strain variation across panel width during loading.
VI
VI
58. .5-
P g 88.96 kN
(20 000 Ib)
.4-
COMPRESSIVE
STRA IN,
percent
E, .3-
P g 44.48 kN
.2- (10 000 Ib) x/l
1FRONT FACE
0 1BACK FACE
.I-
O -.25
t.25
0
-.5 -.4 -.3 -.2 -.l 0 .1 .2 . 3 .4 .5
2+ -.25
Ylb
(b) Panel BT-8.
Figure 19.- Concluded.
59. ksi MPa X x l 2 = +0.25 ksi MPa xl2 = t0.25
80 , 500
100
yl b = -0.34 y l b = "0.346oor
BACK FACE
300 - 300-
10020 20 100
0
- FRONT FACE
100r
80 -
0 60 7X
40 1-
2o0 t
400k BACK FACE
300i- OX
FRONT FACE 20
100
0 .2 .4 . 6 .8 1.0 oL
X 12 = -0.25
609- ~1 y l b = -0.34
500y
200 BACK FACE
100
0 .2 .4 e 6 . 8 1.0
E ,percent
X
E percent
X
60. ksi
100
60
Ox 40-
20e
0 -
IwF80
60
40
20 -
0I-
L
J 0 I X l Z = 0.0
600r y l b = 0.0 loor
100I-/
0 .2 .4 . 6 . 8
E ,percent
X
8o t
40 I-
x12 = -1-0.25
500
0 .2 .4 . 6 . 8 1.0
0 .2 .4 . 6 . 8 1.0
E ,percent
X
(b) Panel BT-6.
Figure 20.- Concluded.
61. s
(a) Side view.
5 End Simple Support
Figure 21.- Failure near side simple support. Wrinkling specimen; tc = 1.27 cm (0.50 in.).
66. COMPRESSIVE
STRESS,
U
X
ksi
100
80
60
40
20
0
-
MPa
600-
-
500 -
- 400 -
300 -
-
200 -
-
100-
X
f
STRA IN GAGE LOCATIONS
Xll = 0.0
y l b = 0.0
0 .2 .4 .6 .8
COMPRESSIVESTRAIN, E ,percent
X
(a) Panel BT-1.
Figure 24.- Back-to-back stress-strain results of overall buckling specimens.
67. X
ksi
100 -
MPa
600
80 -
500
COMPRESSIVE
60 1400
STRESS,
0 300
X
200
20[100O O
k-b -4
STRAIN GAGE LOCATIONS
X l Z = 0.0
y lb = 0.0
- BACK SURFACE //
.2 .4 .6 .8
COMPRESSIVESTRAIN, E ,percent
X
(b) Panel BT-2.
SIDE
Figure 24.- Continued.
77. 1
u'Fc u
OVERALL BUCKLING (FLEXIBLE CORE)
OVERALL BUCKLING
(R lGlD CORE)
SHEAR CRIMP ING
WRINKLING
PLY THlCKNESS=0.0076 cm (0.003 in. 1
FACING LAMINATE:[0,+45,90Js
- . ~ " _ CORE: HRH - 327 - 318 - 4
DIMPLING (eq.(B2))
DIMPLING (eq.(Bl))
WRINKLING (eq.(BG))
6=. 01cm(. 004 in. 1
WRINKLING (q.(B4))
- LAMINATESTRENGTH
- Fcu = 565 MPa (82 ksil
0 EXPER IMENTAL RESULTS
. l L- "" "1"~ ~. -1 . . - 1 - 1 1
0
I
. 5 1.0 1. 5 2.0 2. 5 3.0 cm
0 . 2 . 4 .6 . 8 1.0 in.
.~ ~ 1- -~ -I 1 I
CORE THICKNESS, tc
Figure 28.- Comparison of analytical and experimental results.
75
78. (a) Intracellularbuckling.
SYMMETRIC ANT ISYMMETR IC
(b) Facewrinkling.
(c) Shearcrimping.
Figure 29.- L o c a l i n s t a b i l i t y modes o f f a i l u r e of
honeycombsandwich s t r u c t u r e s .
76
79. -~
1. Report No. 2. Government Accession No.
. ."
NASA TP-1785
4. Title and Subtitle
TESTS OF GRAPHITE/POLYIMIDE SANDWICH PANELSIN
UNIAXIAL EJXEWISE COMPRESSION
7. Authorlsl
Charles J. Camarda
9. Performing Organization Name and Address
" -
NASA Langley Research Center
Hampton, VA 23665
3. Recipient's C a t a l o g No.
-
5. Report Date
1 December 1980
6. PerformingOrganization Code
506-53-63-04
8. Performing OrganizationReport No.
L-13998
11. ContractorGrantNo
- _ _
2. Sponsoring Agency Name and Address
National Aeronautics and Space Administration
Washington, DC 20546
~~ ~
5. Supplementary Notes
..
1 13. Type of Repon and Period Covered
6. Abstract
An experimental and analytical investigation has been made of the local and genera
buckling behavior of graphite/polimide (Gr/PI sandwich panels simply supported
along all four edges and loaded in uniaxial edgewise compression. Material proper-
ties of sandwich panel constituents (adhesive and facings) were determined from
flatwise-tension and sandwich-beam-flexure tests. Buckling specimens were30.5 by
33 cm (12 by 13 in.), had quasi-isotropic, symmetric facings([O,+45,90Is), and a
glass/polyimide honeycomb core (HRH-327 -3/8-4). Core thicknesses were varied@
(0.635, 1.27, 1.91, and 2.54 cm (0.25, 0.50, 0.75, and 1.00 in.)) and three panels
of each thickness were tested at room temperature to investigate failure modes and
corresponding buckling loads. Specimens 0.635 cm (0.25 in.) thick failed by over-
all bucklingat loads close to the analytically predicted buckling load; all other
panels failed byface wrinkling. Results of the wrinkling tests indicated that
several buckling formulas were unconservative and thereforenot suitable for
design purposes; a recommended wrinkling equation is presented.
. Key Words (Suggested by Authorls))
Buckling tests
Composite materials
Sandwich panels
Graphite/polyimide
Hiah t-nprature tests . !. Security Classif. (of this report1 1 20. Security Classif. (of this pagel 1 21. NO. of Pages 22. Price
I
18. DistributionStatement
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