- II I!!! 111111 111 !!!!! 

NASA Technical Paper 1785 

Tests of Graphite/Polyimide 

Panels in Uniaxial Edgewise 

Charles J. Camarda 

DECEMBER 1980 

NASA 

TP 

1785 

c. 1 ~ 

- - 

Sandwich 

Compression 
TECH LIBRARY KAFB, NM 

0067732 

NASA Technical Paper 1785 

Tests of GraphitdPolyirnide Sandwich 

Panels in Uniaxial Edgewise Compression 

Charles J. Camarda 

Latzgley Research Center 

Halnptm, Virginia 

NASA 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1980 
SUMMARY 

An experimental and analytical investigation has been made of the local 

and general buckling behavior of graphite/polyimide (Gr/PI) sandwich panels 

simply supported along all four edges and loaded in uniaxial edgewise compression. Material properties of sandwich panel constituents (adhesive and 

facings) were determined from flatwise-tension and sandwich-beam-flexure 

tests. Results from the flatwise-tension tests established a suitable cure 

cycle for FM-34l polyimide film adhesive which was the adhesive used to fabricate the flatwise-tension, sandwich-beam, and buckling specimens. A cell-edge 

bonding technique using a liquid version of FM-34 polyimide adhesive was 

investigated and results indicated that a considerable mass savings may be possible using a cell-edge adhesive. Tensile and compressive material properties 

of the facings (quasi-isotropic, symmetric, laminates (( [0,+45,90,-45],) of 

Celion2/PMR-1 5) were determined at 11 6 K, room temperature, and 589 K (-250°F, 

room temperature, and 600°F) using the sandwich-beam-flexure test method. Buckling specimens were 30.5 by 33 cm (1 2 by 13 in.), had quasi-isotropic, symmetric 

facings ( [0,+45,90],), and a glass/plyimide honeycomb core (HRH-3273-3/8-4). 

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 overall buckling at loads close to the analytically predicted buckling load; all other panels failed by face wrinkling. 

Results of wrinkling tests indicated that several buckling formulas were unconservative and therefore not suitable for design purposes; a recommended wrinkling equation is presented. 

INTRODUCTION 

Preliminary structural studies of advanced space transportation systems 

using advanced composite structural materials of high-strength fibers and 

polyimide resin matrices indicate that a reduction of up to 25 percent in vehicle structural mass is obtained by the direct replacement of aluminum panels 

with graphite/plyimide (Gr/PI) panels (refs. 1 and 2). Furthermore, preliminary studies of the aft body flap of the Space Shuttle Orbiter (ref. 3) indicate 

that compression loads are the primary design condition for this structural component and because a biaxial state of stress exists in the cover panels a sandwich panel was chosen. The present study focuses on Gr/PI structural sandwich 

panels which may have application as cover skins on lightly loaded components 

such as the aft body flap of the Space Shuttle Orbiter. Based on the low magniFM-34 film adhesive: manufactured by American Cyanamid Company, 

Bloomingdale Division. 

2Celion: registered trademark of Celanese Corporation. 

3HRH 327: registered trademark of Hexcel Products, Inc. 
IllllIllIlllllll Ill1 I I lIlllIlllIIll I1 11111l1l1l11l1l1l1 

tude and biaxial nature of these loads, a minimum-gage, quasi-isotropic, SFmetric Gr/PI laminate ( [Of +45 , 901 s) was chosen for the facings of these sandwich panels in the present study. 

The purposes of the present study are to analytically and experimentally 

investigate the local and general buckling behavior of minimum-gage Gr/PI sandwich panels capable of use at temperatures ranging from 116 to 589 K (-250° to 

600°F), to verify the fabrication method used in manufacture of the panels, and 

to determine the material properties of the [0,+45,901s Gr/PI sandwich panel 

facings. 

Buckling specimens 30.5 by 33.0 cm (12 by 13 in.) were designed and fabricated with various core thicknesses to study local and general instability 

failure modes. The buckling specimens were tested in uniaxial edgewise compression at room temperature (R.T.) and were simply supported along all four edges. 

Several analysis methods (refs. 4 to 8) were used to determine upper and lower 

bounds on critical stresses relating to intracellular buckling (dimpling), wrinkling, 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 fabrication cure 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 sandwichbeam-flexure tests were conducted at temperatures of 116 K, R.T., and 589 K 

(-250°F, R.T., and 600OF). Quality control standards for fabrication of 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 analyzed 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 both SI and U.S. Customary Units. The measurements and 

calculations were made in U.S. Customary Units. 

2 
DQx I DQy 

FC 

Fcu 

G 

Nx Ny 

stiffness matrices of sandwich panel 

width of plate 

coefficients of polynomials used in regression analysis 

flexural stiffness of composite facings 

transverse shear stiffness of sandwich plate in x- and 

y-directions, respectively 

flexural stiffness of orthotropic sandwich plate in x- and 

y-directions, respectively 

twisting stiffness of orthotropic sandwich plate 

elastic modulus 

modulus of core in z-direction 

facing modulus 

facing modulus in x- and y-directions, respectively 

tangent modulus 

average elastic moduli of laminate in x- and y-directions, 

respectively 

lower of flatwise core compressive or 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 forces in x- and y-directions, respectively 

3 
P 

SCJ/E 

S 

Tg 

t 

tC 

t f 

tfl 

tf2 

th 

Vf 

VV 

XIYIZ 

6 

6 

1-1 

1-1xr1-1y 

Dxy' pyx 

" 

P 

0 

ocr im 

*d im 

*wr 

load 

standard error of estimate 

honeycomb cell size 

glass transition temperature 

thickness 

core thickness 

average facing thickness 

thickness of facing 1 

thickness of facing 2 

total sandwich panel thickness 

fiber volume fraction 

void volume fraction 

rectangular coordinates 

initial panel waviness 

strain 

Poisson's ratio 

Poisson's ratio of orthotropic plate associated with bending of 

plate in x- and y-directions, respectively 

average Poisson's ratios of orthotropic plate associated with 

extension of plate in x- and y-directions, respectively 

density 

stress 

critical stress associated with shear crimping 

critical stress associated with dimpling 

critical stress associated with wrinkling 

Subscripts : 

av average 

4 
cr critical 

i index of summation 

max aximum 

ult ltimate 

XIY coordinate dirctions 

1,2 directions parallel and perpendicular to fiber direction, respectively 

TEST SPECIMENS I 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 focused on 

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 of the 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 Celion 6000/PMR-15; for buckling tests, the laminates were Celion 

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 Celion 6000/PMR-15 to minimize mass of the sandwich panels, and the Celion 

fiber was chosen over the HTS fiber because Celion exhibits less material property 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 precured [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 diagram of a typical specimen is shown in figure l. Details of fabrication procedures and cure cycles are given in reference 9 and table I, 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 mode at 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 prior to 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 
Sandwich-Beam-Flexure Tests 

Specimens.- Sandwich-beam-flexure specimens consisted of Gr/PI facings and 

glass/polyimide honeycomb core as shown in figure 2. The honeycomb core was 

HRH 327-3/16-8 glass/polyimide and was cut into strips 2.54 cm (1.00 in.) wide 

by 55.88 cm (22.0 in.) long by 3.175 cm (1.25 in.) high. The test facing of 

the beam was a [O,+45,9O,-45Is laminate of Celion 6000/PMR-15 and the opposite facing had a laminate orientation of [02,+45,90,-451,. The additional Oo 

layers of the non-test facing insured failure of the test facing. These laminate orientations were chosen to avoid microcracks in the laminate which are 

believed to occur when adjacent layers are stacked at an angle greater than or 

equal to 90° with respect to one another. The honeycomb core was filled with 

BR-34 liquid adhesive and glass beads throughout the length of the beams, 

except for the 7.62-cm (3.00-in.) test section in the center of the beams, to 

prevent premature adhesive failure. Details of the fabrication of the sandwich beam specimens are presented in reference 9. 

Apparatus and instrumentation.- Each specimen was instrumented at the midspan of the beam with a high-temperature strain rosette (WK-03-060-WR-350) 

oriented at Oo, 45O, and 90° with the load axis and bonded to the test facing, 

and a single strain gage (WK-03-125-AD-350) oriented at Oo with the load axis 

and bonded to the non-test facing. These gages were manufactured by MicroMeasurements Division of Vishay Intertechnology, Inc. The strain gages were 

bonded to the outer surfaces of the beam using a polyimide adhesive (either 

"Bond 610 or PLD-700 available from Micro-Measurements and BLH Electronics, 

respectively) . 

The sandwich beams were placed in a four-point bending test apparatus 

(fig. 3) which supported the beam on rollers 48.26 cm (19.00 in.) apart with 

flat sections 2.54 cm (1.00 in.) wide machined in them. Load was applied by 

a 222-kN (50-kip) capacity hydraulic testing machine which acted at two points 

on the top flange of the beam spaced 10.16 cm (4.00 in.) apart and symmetric 

about the beam's center. For testing at temperatures other than room temperature the specimen was instrumented with a thermocouple attached to the test 

facing and the test fixture and specimen were completely enclosed in an environmental chamber and either heated or cooled to the desired test temperature. 

Specimens were allowed to soak at the test temperature for 20 minutes to insure 

thermal equilibrium. A data handling system consisting of 40-channel scanner, 

digital voltmeter, plotter, printer, clock, and calculator was used to record 

and reduce data. 

Procedure.- The load signals from the load cell on the hydraulic testing 

machine were input to one channel of the scanner. Strain signals were input to 

selected scanner channels and initally set to zero using Wheatstone bridge balance (for non-room-temperature tests, initial strain signals were set to zero 

after thermal equilibrium). Strains were corrected for transverse sensitivity 

of the gages and nonlinearity of the bridge circuit. Thermocouples were connected to the scanner through a 273 K (32OF) cold-junction reference. 

Beams were tested to failure during the test, load was applied at a rate of 

80 N/sec (20 lbf/sec), data were recorded every 3 seconds, and a stress-strain 

6 
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 leastsquares sense, of a third-order polynomial relating stress and strain according 

to the polynomial equation: 

A more detailed explanation of the analysis is given in reference 10. 

To assess the magnitude of scatter of experimental points about the regression equation, the standard error of estimate So/,, which is a measure of the 

mean deviation of the sample points from the regression line, is determined as 

follows : 

\ j-4 

This method of statistical analysis is similar to that presented in reference 10. 

Buckling Tests 

Specimens.- Design considerations of the buckling specimens are given in 

appendix A. Specimens were 30.5 by 33.0 cm (12 by 13 in.) with core 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 a completed buckling specimen; details on specimen manufacture are given in reference 9; details of 

significant panel parameters are listed in table 11. 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 edges of the test 

panels was considered -to be a realistic representation of the boundary conditions that actual panels on the shuttle body flap will experience. A test 

fixture, similar to that of reference 4, was fabricated to simply support all 

four edges of the sandwich panel and allow adjustments to be made during loading which would align the specimen and thus maintain a uniform strain distribution across the panel. Details of the simple supports are shown in 

7 
figures 5(a) and (b). The stainless-steel alignment sheet embedded in each of 

the potted ends of the panel fit into stainless-steel knife edges which fit 

into steel V-groove load blocks as shown in figure 5 (a) . The load blocks fit 

into adjustable end loading heads which were attached to the hydraulic load 

machine. The end loading heads contained a flat stainless-steel bar which, 

together with the aligning screws, was used to obtain a uniform longitudinal 

strain across the specimen. Figure 6 shows the buckling specimen in the test 

fixture. The sides of the panel were simply supported by knife edges which 

were supported by Z-section steel beams as shown in figure 5(b). The side 

supports maintained a relatively snug fit against the panel because of the high 

degree of flatness of the panels. However, because of the raised scalloped 

doublers the side supports could not extend the complete length of the panel. 

The Z-section beams were braced so that motion of the side supports was 

restrained. The knife edges of the side supports were bolted snugly in place at 

two locations on two sides as shown in the schematic of figure 5(b) and as partially shown in figure 6. The side supports were positioned 1.27 cm (0.5 in.) 

from each side edge, making the test section width 30.5 cm (1 2 in.). 

A 222-kN (50-kip) hydraulic load machine was used to compress the panels, 

A mercury-vapor light source was used in conjunction with a photographic line 

grid having a pitch of 19.7 lines/cm (50 lines/in.) to determine out-of-plane 

displacements and mode shapes using the grid-shadow Moire/ technique as discussed in references 11 and 12. A camera was positioned perpendicular to the 

sandwich panel and the light source formed an angle of 30° with that 

perpendicular. 

Each panel was instrumented with 12 single, foil-type strain gages and 2 

45O strain rosettes, Micro-Measurements WK-03-125-AD-350 and WK-03-060-WR-350, 

respectively, as shown schematically in figure 7. The positioning of the 

gages allowed measurement of longitudinal strain distributions on each facing 

across the panel length and width. Back-to-back longitudinal strain gages were 

positioned at five points on the panel (four corner points and a centrally 

located one). The purpose of the back-to-back gages was to detect bending of 

the panel and to determine the general buckling load and possibly the wrinkling 

load. The data acquisition system used to reduce and store data is identical 

to that mentioned earlier for the sandwich beams. 

Procedure.- During each test, the hydraulic testing machine was operated 

in a displacement control mode at a rate of approximately 0.020 cm/sec 

(0.008 in/sec), strain gages were scanned approximately every 3 seconds, and 

the specimen was loaded to failure. Raw data were converted to engineering 

units, printed in real time, and stored on a disk. Stresses were calculated 

by dividing the load by the combined cross-sectional area of the two facings. 

Gages were balanced prior to testing using Wheatstone bridge circuits as discussed earlier. Prior to tesing, panels were loaded up to approximately 

50 percent of failure load and were aligned using the adjustable screws shown 

in figure 6. The panel was then unloaded and the Moire' grid positioned in 

front of the specimen. Strain gages were then zeroed and load was applied to 

the specimen until failure. 

8 
TEST RESULTS 

Flatwise Tensile Tests 

Preliminary tests indicated that significant improvements of bond strengths 

could be obtained by abrasively cleaning the edges of the honeycomb and by 

dipping the core in primer instead of brush or roller coating it on the core. 

(See ref. 9.) 

A series of flatwise tensile tests of specimens, bonded with FM-34 using 

various cure cycles, aided in the selection of a suitable cure cycle. Two 

specimens were tested at room temperature for each cure-cycle variation listed 

in table I. Specimen failures occurred by either adhesive bondline rupture or 

facing delamination; figures 8 and 9 show the two modes of failure. Results of 

those tests, listed in table 111, indicate that cure cycles 1 and 5 produced 

the strongest bonds, with failures occurring in the facing. Delamination of 

the facings also occurred with cure cycle 4 but because the bond cure temperature of 61 6 K (6500F) was greater than the facing cure temperature of 603 K 

(625OF) the interlaminar shear strength of the facing was degraded and failure 

loads were lower. Bonding one face of the specimen at a time, with the face 

to be bonded positioned under the core (cure cycle 3), provided good filleting 

between the face and core but did not enhance the strength of the bond. 

Instead, bond strengths were lower and failures occurred in the second of the 

two bonds. Six specimens were tested at 589 K (600OF) ; two specimens were fabricated at each of three cure cycles 1, 5, and cure cycle 1 with a higher cure 

temperature (603 K (625OF) ) . It was hoped that the higher cure temperature 

would improve the elevated temperature bond strength. 

Cure cycle 1 with the elevated cure temperature was chosen because of the 

higher bond strengths at elevated temperature and because maintaining a vacuum 

during cure would help eliminate volatiles produced during the cure of the FM-34 

adhesive. Although trapped volatiles did not degrade the strengths of the 7.62 

by 7.62 cm (3 by 3 in. ) specimens, it would be more difficult to vent the volatiles in larger panels. 

Sixteen flatwise-tension specimens were fabricated using cure cycle 1 with 

a cure temperature of 603 K (625OF). Test results from these specimens are presented in table IV. Flatwise-tensile strengths at room temperature and 116 K 

(-250OF) increased from 1 .6 MPa (230 psi) to an average value of 3.2 (470 psi) 

when the edges of the honeycomb were cleaned and primed as mentioned earlier. 

Failures at this stress level were usually by facing delamination as shown in 

figure 9. Flatwise-tensile strengths at 589 K (600OF) were generally higher 

than 1.4 MPa (200 psi) with failures occurring in the bondline, similar to the 

room-temperature test shown in figure 8. 

Flatwise-tensile test results at room temperature of specimens bonded using 

BR-34 as a cell-edge adhesive are presented in table IV(b) . Most of these specimens failed by facing delamination. However, for these specimens the facings 

9 
delaminated locally about each cell edge as shown in figure 10 and usually 

resulted in slightly lower strengths. When local facing delamination 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 reference 13. The mass of 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 result in a mass savings 

equivalent to 10 percent 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 adhesive can 

result in considerable mass savings without necessarily sacrificing bond 

strength and that further research in this area is 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 tests are presented in tables V 

and VI and in figures 11 to 16. As shown in table V, the scatter of test data, 

as determined by the standard error of estimate, was lowest for the room temperature 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 61 0 psi) as compared to respective average ultimate strengths of 567.7 MPa (82.34 ksi) and 

334.0 MPa (48.44 ksi) (see table VI). The average compression ultimate strain 

was 1.38 percent at room temperature and 0.657 percent at 589 K (600OF). 

Average ultimate strengths of the laminate were slightly higher in compression 

than tension for each test temperature. Ultimate strengths of the Celion 

6000/PMR-15 [0,+45,90,-451s laminates were higher than results for HTS/PMR-15 

as reported in references 10 and 14 except for tensile strength at 589 K (600OF) 

reported in reference 14. Average room temperature tensile and compressive 

ultimate strengths for the HTS/PMR-15 laminates were 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 room temperature values while strengths 

at 589 K (60O0F) decreased by 43 percent. Average compressive ultimate 

strengths at 116 K (-250OF) and 589 K (600°F) increased and decreased, respectively, by 13.8 and 41.2 percent from room-temperature values. 

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 laminates. This difference is probably due to the higher fiber volume fraction, 

72 percent for the Celion/PI laminates of the present study compared with 43 to 

55 percent for the HTS/PMR-15 laminates of references 10 and 14. Modulus values 

at 0.2 percent strain and 116 K (-250OF) were about 10 percent higher than 

values at room temperature. Modulus values at 589 K (60O0F) were about the 

same as room-temperature values. Stress and tangent modulus as a function of

strain for various temperatures are presented in figures 11 to 16. Table V 

10 
lists the coefficients of the regression equation, used in the reduction of the 

experimental data. The data points in the figures represent experimental values 

all replicate tests; the solid line in the figures is the best-fit third-order 

polynomial obtained from the regression analysis. Tangent modulus as a function of strain was calculated by differentiation of the third-order ,polynomial 

(eq. (1)). Tensile modulus values were fairly linear throughout the usable 

strain region (E I 0.35 percent) as shown by figures 11, 13, and 15. Compressive modulus values tended to be nonlinear at room temperature and became linear 

at 589 K (600OF) as shown,by figures 14 and 16. 

Representative tensile and compressive failures are shown in figures 17 and 

18, respectively. Most compressive failures occurred near the edge of the 

potted section of the honeycomb next to the load tabs. 

Buckling Tests 

Two modes of panel failure were discernable from experimental results: 

wrinkling and overall buckling. Specimens with a core thickness tc of approximately 0.635 cm (0.25 in.) failed by overall buckling, and all other specimens, 

having nominal core thicknesses of 1.27, 1.91, and 2.54 cm (0.5, 0.75, and 

1.00 in.), failed by wrinkling. None of the panels tested failed by laminate 

yield, dimpling, or shear crimping. The shadow Moire/ method was useful in 

determining mode shapes of the overall buckling specimens but was not able to 

determine wrinkling mode shapes because of the high stiffness and brittle nature 

of the Gr/PI facings and, hence, the relatively small out-of-plane displacements. The use of a finer Moire/ line grid would increase the sensitivity of the 

optical technique and possibly enable the determination of local buckling modes. 

Wrinkling . -. ~ ~-. . specimens.- Results of longitudinal strain uniformity across 

specimen width are presented in figures 19(a) and (b) for two values of applied 

load and two different panels. The adjustable test fixture was useful in eliw 

inating large strain variations caused by misalignment, similar to test fixtures 

used in reference 4. Strains were fairly uniform across the width of the panel 

as shown in figures 19(a) and (b). However, slightly higher strains and strain 

variations do occur at the edges of the panels as was noted in reference 4. 

Trends in strain distributions at the low load level, 44.48 kN (10 000 lbf), 

were similar to trends at the higher load level of 88.96 kN (20 000 lbf). There 

were no consistent trends in strain distributions from panel to panel. However, 

most of the wrinkling specimens did fail near the end of the side simple 

supports where slightly higher strains were recorded. 

Longitudinal strains as a function of stress were calculated for each 

strain-gage position on the panel. Results of several tests (panel numbers BT-5 

and BT-6) are presented in figures 20(a) and (b). Back-to-back strain variation 

was usually lowest in the center of the panels (x = y = 0). Irregularities in 

slopes were noted in some specimens as shown in figure 20(b) for panel number 

BT-6. These irregularities in slope occur at too low a load to be an indication 

of wrinkling or some form of local instability as mentioned in reference 4. The 

irregularities in the present study were possibly caused by some interference or 

interaction of the test fixture. Material behavior was slightly nonlinear to 

failure, similar to results of the four-point flexure tests as,noted earlier. 

11 
Back-to-back stress-strain data did not predict the onset of local buckling 

(wrinkling) and the use of the force stiffness method of reference 15 to predict 

wrinkling was unsuccessful. All panel failures were abrupt with no indication 

of local instability. It would probably be necessary to extensively instrument 

both sides of a facing to calculate facing bending strains and predict local 

buckling. Modulus values at 0.2 percent strain, maximum back-to-back strain 

variation at 0.6 percent strain, theoretical wrinkling stress, and experimental 

ultimate stress and strain values of each panel are presented in table VII. 

Maximum back-to-back strain variation was fairly low, considering the size and 

complexity of the sandwich panels. Compressive modulus values at 0.2 percent 

strain of the sandwich panels which used Celion 3000 material were slightly 

higher than results of beam tests which used the Celion 6000 material; the average modulus of all wrinkling specimens is 53.9 GPa (7.82 x lo6 psi) as compared 

to 48.95 GPa (7.1 0 x 1 O6 psi) obtained using the four-point beam flexure test 

method. Since the fiber volume fraction Vf of the beam specimens was higher 

than that of the buckling specimens (72 percent compared with approximately 

61 percent) it appears that the thinner gage Celion 3000 material did not 

experience any degradation in modulus. Results of replicate tests indicate 

that scatter was low. Scatter in critical wrinkling stress ranged from a minimum of 7.6 MPa (1.1 ksi) for the 1.27 cm (0.5 in.) specimens to a maximum of 

89 MPa (1 3 ksi) for the 2.54 cm (1 .OO in.) specimens. This amounts to a range 

from minimum to maximum of 1.7 to 29 percent, respectively, when compared to 

average critical stress values. From tables I1 and VI1 some trends in results 

are evident: 

1. Average failure stresses of the wrinkling specimens decrease as core 

height tc increases. This is characteristic of a wrinkling or local buckling 

type of instability. Average failure stresses were 452, 354, and 311 MPa 

(65.6, 51 .4, and 45.1 ksi) for the 1 .27-, 1 .91-, and 2.54-cm (0.50-, 0.75-, and 

1.00-in.) thick cores, respectively. 

2. Average failure strains were 0.87, 0.71, and 0.63 percent for the 1.27-, 

1 .91-, and 2.54-cm (0.50-, 0.75-, and 1.00-in.) thick cores, respectively. 

3. Specimens with higher total facing thicknesses had higher failure loads; 

however, these specimens did not necessarily have higher failure stresses. 

This is because the thicker facings had a lower fiber volume fraction Vf 

because not enough resin was removed during the consolidation phase of laminate 

fabrication. 

4. Panels with the largest value of initial waviness had the lowest 

ultimate load. 

5. Ultimate strains of the wrinkling specimens were well below ultimate 

laminate strain results from the beam tests. 

As mentioned earlier, most of the wrinkling specimens failed close to the 

end of one of the side simple supports. Failure of a 1.27-cm (0.50-in.) panel 

is illustrated in figures 21 (a) and (b) ; the failure extends across the panel to 

the top of the left-side simple support. The failures were nearly perpendicular 

to the direction of load. Wrinkling failure was most noticeable in the 1.27-cm 

(0.50-in.) specimens in which the facings separated from the core due to a ten12 
sile failure of the adhesive. Failed panel BT-4 (fig. 22(a) ) illustrates the 

outward buckling of the facing; the panel was cut 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 symmetric orantisymmetric; however, laminate failures 

on either facing were similar which suggests that failures were symmetric. This 

agrees with results of references 4 and 16 which indicate that for honeycomb 

cores, where the modulus of the core in the direction of the load is much less 

than the modulus of the core in the direction perpendicular to the facings, sym 

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 fiber strain on the convex side 

of the buckled panel. The specimens exhibited a very short postbuckling 

region as evidenced by the experimental results of 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 reference 4, 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 perpendicular 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 sine waves in the length and width directions as expected. The outof-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-hand simple support. All other specimens failed in 

the center. Moirg fringe patterns of a 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 to be 

symmetric in the longitudinal direction; however, nonzero displacements appear 

to occur near the right-hand simple support. Displacements do occur at the 

corners of the panel since the simple supports do not extend the total panel 

length. As the panel approaches failure, mode shapes tend to be nonsymmetric 

(fig. 27(d)). As mentioned in reference 4, 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 wave in the length and width 

direction. 

Comparison of analytical a,nd experimental results.- The analysis assumes 

the r&m-temperature unidirectional properties and dimensions listed in 

table IX. 

13 
- I 

From laminate theory Ex = Ey = Ef = 51 .97 GPa (7.538 x 1 O6 psi) and 

pxy = 0.3075. These results agree with experimental results from the 

sandwich beam flexure tests in which the average modulus, .Ex = 48.95 GPa 

(7.1 x lo6 psi) and bv = 0.347. Since laminates were quasi-isotropic, 

symmetric ([0,+45,90Is), A16 and A26 coupling terms of the sandwich were 

identically zero, and the Dl6 and D26 coupling terms were negligible. Analytical results, assuming a lamina thickness of 0.0076 cm (0.003 in.), are 

presented and compared with experimental results in tables VI1 and VI11 and 

in figure 28. The overall buckling analysis described in reference 7, which 

included the core shear flexibility, agreed well with experimental overall 

buckling results. Local and general buckling formulas used in the present 

analysis are presented in appendix B. 

- - 

The average experimental overall buckling stress was 264 MPa (38.3 ksi) 

and compared exactly with the analytically predicted overall buckling stress. 

From experimental wrinkling results it appears that equations (B5), (B6) and 

(B8) were unconservative and impractical to use from a design standpoint. 

Equation (B4), however, was conservative in its prediction of symmetric wrinkling loads and is useful for design purposes. Wrinkling results obtained by 

using equation (B6) and assuming 6max = 0.01 cm (0.004 in.) were 7, 26, and 

32 percent higher than experimental results for the 1.27-, 1 .91-, and 2.54-cm 

(0.50-, 0.75-, and 1.00-in.) thick cores, respectively. The equation for 

overall buckling is equation (B11) and, as explained earlier, when core shear 

flexibility is accounted for, the results using this equation agree exactly 

with average experimental values of Dcr. 

CONCLUDING REMARKS 

The purpose of the present study was to investigate the buckling behavior, 

local and general, of graphite/polyimide (Gr/PI) sandwich panels capable of use 

at temperatures ranging from 11 6 to 589 K (-250 to 600°F) as the sandwich skin 

of the space shuttle body flap. The adhesive and facing material properties 

were investigated and buckling formulas for predicting local and general sandwich panel instabilities were evaluated. Results of a bond study include a fabrication technique for adhesively bonding sandwich structures and a cure cycle 

for FM-34 film adhesive which produced flatwise tensile strengths in excess of 

3.4 MPa (500 psi) at 116 K and R.T. (-25O0F and R.T.) in 1 .4 MPa (200 psi) at 

589 K (600OF) . Results also indicate that a liquid cell-edge adhesive (BR-34) 

can result in considerable panel mass savings (10 percent) without necessarily 

sacrificing bond strength; however, further research is necessary since flatwisetensile strengths using BR-34 were not consistent. Material property tests of 

quasi-isotropic, symmetric laminates ([0,+45,90Is) of Celion 600.0/PMR-15 Gr/PI 

material indicate that it maintains suitable structural properties for shortterm use at temperatures from 11 6 to 589 K (-250° to 60O0F). 

Experimental results of flat rectangular honeycomb sandwich panels which 

were simply supported along all four edges and tested in uniaxial edgewise compression indicate that two modes of panel failure, wrinkling or overall buckling, can occur depending on the core thickness. As predicted analytically, 

specimens with a core thickness of 0.635 cm (0.25 in.) failed by overall buckling and all other specimens, having nominal core thicknesses of 1 .27, 1.91 , 

14 
and 2.54 cm (0.50 , 0.75, and 1.00 in.), failed by wrinkling. The shadow Moire/ 

method was useful in determining mode shapes of the overall buckling specimens 

but was not able to detect wrinkling. 

Results of the wrinkling tests indicated that several analytical methods 

were unconservative and therefore not suitable for design purposes. Most of 

the wrinkling specimens failed near the side simple supports. The failure mode 

appeared to be symmetric wrinkling with failures occurring because of tensile 

rupture of the adhesive. Some trends in wrinkling results are: 

1. Average failure stresses of the wrinkling specimens decrease as core 

thickness increases and are 452, 354, and 311 MPa (65.6, 51.4, and 45.1 ksi) 

for the 1.27-, 1.91-, and 2.54-cm (0.50-, 0.75, and 1.00-in. ) thick cores, 

respectively. 

2. Average failure strains were 0.87, 0.71, and 0.63 percent for the 1.27-, 

1.91-, and 2.54-cm (0.50-, 0.75-, and 1.00-in.) thick cores, respectively. 

3. Panels with the largest value of initial waviness had the lowest 

ultimate load. 

The average experimental overall buckling stress of the 0.635-cm (0.25-in.) 

thick specimens was 264 MPa (38.3 ksi) and compared exactly with the analytically predicted overall buckling stress. All the overall buckling specimens 

except one failed in the center on the concave facing by compression. 

Langley Research Center 

National Aeronautics and Space Administration 

Hampton, VA 23665 

December 2, 1980 

15 
APPENDIX A 

DESIGN CONSIDERATIONS FOR BUCKLING SPECIMENS 

Preliminary studies of structural loads on 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 laminates of [0,+45,90Is Gr/PI were chosen for the 

facings of the sandwich skin of the body'flap. Therefore, buckling specimens 

similar to the sandwich skin of the body flap are examined in the present study. 

Only symmetric laminates were considered in the present investigation to 

prevent laminate warpage during the cure cycle caused by 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 be sufficient 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 reference 4. 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 substantial mass savings over the thicker gage Celion 6000 material. Average 

thickness per ply of the Celion 3000 laminates were 0.007 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 589 K (600OF) is either Hexcel HRH-327-3/16-4 or HRH-327-3/8-4 glass/PI 

which has a density of 64 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 chosen at test conditions because they closely represent conditions 

actual shuttle body 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 references 9, 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 equations of reference 8 were used to predict local instability modes and associated loads. The local and general buckling equations are 

also presented in appendix B. Buckling loads were computed for various ply 

thicknesses, core thicknesses, and operating temperatures. 

Unidirectional laminate material properties used in the design of the buckling specimens were obtained from references 19 and 20. Honeycomb core material 

properties were obtained from reference 5. Some of the material properties used 

in the present analysis are presented in table IX. Various cores and core 

16 
__ 

APPENDIX A 

thicknesses (0.635 to 2.54 cm (0.25 to 1.00 in.)) and panel lengths and widths 

(1 0.2 to 122 cm (4.0 to 48.0 in.) ) were analytically investigated at various 

temperatures (room temperature to 589 K (6000F)), and design envelopes, 

typified by figure 28, were determined. Since the laminate orientation of the 

facings is quasi-isotropic and symmetric, the average elastic modulus Ex or 

Ey was used for the facing-modulus Ef equations in appendix B. Results of 

critical stress as a function of core thickness for an assumed ply thickness of 

0.0076 cm (0.003 in.) are shown in figure 28. 

- 

- 

The design envelope curves in figure 28 indicate that either overall buckling, dimpling, laminate strength, or wrinkling could be critical failure modes 

depending on scatter in material properties and different analysis techniques. 

Since it is desirable to verify as many analytical predictions for various 

failure modes as possible, the honeycomb core with the larger cell size (0.95 cm 

(3/8 in.) ) was chosen because it lowers the dimpling stress to values closer to 

the other critical stresses. A panel size of 30.5 by 30.5 cm (12 by 12 in.) was 

adequate to investigate several failure modes. 

The honeycomb core near the loaded ends of the specimens was potted with 

BR-34 liquid polyimide adhesive, and tapered end tabs of [+451 glass/PI were 

bonded at each end to prevent local end failures such as core crushing or end 

brooming; scalloped doublers were bonded beneath the end tabs to enhance load 

diffusion into the panel and help reduce stress concentrations. A stainlesssteel sheet was embedded in the BR-34 potting at each end to align the specimens in the knife edges. Laminates were bonded to the core and end tabs and 

doublers were secondary bonded using FM-34 film adhesive. Significant panel 

parameters, related to the fabrication and quality of the wrinkling and overall buckling specimens such as fiber and void volume fractions and the glass 

transition temperature, are presented in table 11. Facing and total sandwich 

panel thickness measurements were made at various panel locations and initial 

panel waviness 6 was measured as explained in reference 9. Because of good 

fabrication and quality control procedures the panels were consistent in 

dimensional and material properties. Average thickness per ply of all wrinkling specimens was 0.0079 cm (0.0028 in.) with maximum variations in total 

laminate (8 plies) thicknesses averaging only 0.00451 cm (0.001 78 in.) ; average 

variation in total sandwich panel thicknesses was only 0.0059 cm (0.0023 in.). 

Maximum panel waviness 6max averaged only 0.0097 cm (0.0038 in.). 

17 
APPENDIX B 

BUCKLING FORMULAS USED IN THE PRESENT STUDY 

There are several instability modes which cancause failure of a sandwich 

structure; as shown in figure 29 they are: intracellular buckling (face dimpling), face wrinkling (either symmetric or antisymmetric), and shear crimping. 

Intracellular buckling is a localized mode of instability which occurs only when 

the core is not continuous, as in the case of honeycomb or corrugated cores. As 

shown in figure 29(a), the facings buckle in a platelike fashion directly above 

core cells, with cell edges acting asedge 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 thickness of the core. Depending on the nature of the 

material properties of the core the facings can buckle symmetrically or anti

metrically. For the honeycomb cores, in which the elastic modulus parallel to 

the facings is very low compared with the modulus in the direction perpendicula

to the facings, failure is usually b 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 very short due to a low transverse shear modulus of the core. This mode occurs suddenly and 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 method of 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 formulas for local buckling of a sandwich panel subject to 

uniaxial compression and appropriate references are given as follows: 

Dimpling : 

From references 8 and 22 to 24, for isotropic facings 

18 
From reference 4 

APPENDIX B 

From reference 4, assuming orthotropic faces 

odim = 0.825 - (B3) 

3 4 (1 - iixyiiyx) 

where Ef, and Efy are the facing moduli n the x- and y-direction, respectively, and is the facing shear modulus in the xy-plane. Gfxy 

For isotropic faces, equation (B3) reduces to 

Facing wrinkling ( symmetric) : 

From references 11 and 23, the lower bound on wrinkling stress is 

and the upper bound is 

where is the modulus of the core in the direction normal to the facings 

and tc is the thickness of the core, 

From reference 23, accounting for initial facing imperfections 

19 
APPENDIX B 

where Fc is the flatwise sandwich strength and 6 is the amplitude of initial 

waviness in the facings. 

tC 

t f 

From reference 8, for - 50 

Uwr = 0.5 ( Gc XZECZEf) 'I3 

tC 

t f 

and for - > 50 

From reference 4 

+ 

Shear crimping: 

From reference 8 

and from reference 22, 

20 

. . . . ". . I 
APPENDIX B 

Overall buckling: 

From reference 7 

where 

w13 = 

3 Dl2 

+- 

2 

w22 = DQy 

Ex(?) 

and 
APPENDIX B 

The buckling load of the sandwich is obtained from equations (B1 1 ) to (B14) 

by minimizing with respect to m and n, the number of half-waves in the 

buckle pattern in the length and width directions of the plate, respectively. 

The smallest n consistent with the assumption of simply supported plates is 

n = 1. 

22 
REFERENCES 

1. Davis, John G., Jr., compiler: Composites for Advanced Space Transportation 

Systems - (CASTS) . NASA TM-80038, 1979. 

2. Dexter, H. Benson; and Davis, John G., Jr., eds.: Graphite/Polyimide Composites. NASA CP-2079, 1979. 

3. Design and Test Requirements for the Application of 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, James P.: 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 Analysis of Sandwich Plates and Shells. NASA CR-1457, 1 969. 

9. Camarda, Charles J.: Experimental Investigation of Graphite/Polyimide Sandwich Panels in Edgewise Compression. NASA TM-89895, 1980. 

10. Raju, B. Basava; Camarda, Charles J.; and Cooper, Paul A.: ElevatedTemperature Application of the IITRI Compression Test Fixture for 

Graphite/Polyimide Filamentary Composites. NASA TP-1496, 1979. 

11. Dykes, B. C.: Analysis of Displacements in Large Plates by the Grid-Shadow 

Moire'Technique. Experimental Stress Analysis and Its Influence on 

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, Jon G.: Development of Lightweight Graphite/Polyimide Sandwich 

Panels. Non-Metallic Materials, Volume 4 of National SAMPE Technical 

Conference Series, SOC. Aerosp. Mater. & Process Eng., c.1972, 

pp. 605-61 4. 

14. Shuart, Mark J.; and Herakovich, Carl T.: An Evaluation of the Sandwich 

Beam in Four-Point Bending as a Compressive Test Method for Composites. 

NASA TM-78783, 1 978. 

23 
15. Jones, Robert E.; and Greene, Bruce E.: The Force/Stiffness Technique for 

Nondestructive Buckling Testing. A Collection of Technical Papers - 

AIAA/ASME/SAE 15th Structures, Structural Dynamics and Materials Conference, Apr. 1974. (Available as AIAA Paper 74-351.) 

16. Norris, Charles B.; Ericksen, Wilhelm S.; March, H. W.; Smith, C. B.; and 

Boller, Kenneth H.: 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, Robert M.: Mechanics of Composite Materials. McGraw-Hill Book Co., 

c. 1975. 

18. Ashton, J. E.; Halpin, J. C.; and Petit, P. H.: Primer on Composite Materials: Analysis. Technomic Pub. Co., Inc., c.1969. 

19. Hanson, Morgan P.; and Chamis, Christos C.: Graphite-Polyimide Composite 

for Application to Aircraft Engines. NASA TN 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, Howard G.: Analysis and Design of Structural Sandwich Panels. 

Pergamon Press, 1969. 

22. Sandwich Construction for Aircraft. ANC-23, Pts. I and 11, Air Force-NavyCivil Subcommittee on 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 
TABLE I.- CURE: CYCLES OF FLATWISE-TENSILE SPECIMENS 

.. ". .. - ~~ 

Description ~. ~~ 

Vacuum + 0.34 MPa (50 psi) at R.T. 

Cure to 589 K ( 600°F) at 5 K/min (g°F/min) and hold for 2 hours 

No post cure 

Vacuum + 0.34 MPa (50 psi) at R.T. 

Cure to 450 K (350OF) at 5 K/min (g°F/min) and hold for 2 hours 

Post cure at 589 K (600OF) and hold at temperature for 2 hours 

with clamps 

._ - 

Same as cure cycle 1 but bond top and bottom facings separately 

.. " 

with facings to be bonded on bottom 

~ . . " - . . . .. - -. . . . . . . . . - . . . . . .. . - - 

Vacuum + 0.34 MPa (50 psi) at R.T. 

Cure to 61 6 K ( 650°F) at 5 K/min (g0F/min) and hold for 1 .5 hours 

Same as cure cycle 1 but don't apply vacuum 

~~ 

.. . .. 

, . . . . -. . ~~ 

25 
BT-1 61 .2 

BT-2 60.7 

BT-3 60.3 

(61 .1) 

TABLE 11.- SIGNIFICANT PANEL PARAMETERS 

[Values in parentheses are averages of replicate specimens] 

I 

(a) SI Units 

percent 

.78260 

.78402 I " 'a 63.6 y? 1.3984 

BT-5 58.4 2.4 589 1.4087 

BT-6 63.9 1.9 606 1.4052 

(62) (2.27) (599) (1.4041) 

t C 

cm 

0.66632 

.66576 

.66751 

(. 66726) 

1 .2923 

1 -2969 

1 .2972 

(1 .2954) 

cm cm 

0.05955 

(. 05729) (.11456) 

.05786 .05870 

.05814 .05870 

0.05757 

0.05362 0.05249 

.05701 -05475 

.05390 .05419 

(. 05433) (. 05433) 

0.11 71 2 

.11684 , 

.1166 

(. 0940) 

0.1 061 2 

.11176 

.lo809 

( .10866) 

- 

0.00732 0.00940 0.0025 0.003 

.00732 .00813 ~ .0051 ~ .005 

.00729 .0109 ' .0025 .005 

(. 0071 6) (. 00940) (. 00381 ) (. 004) 

0.00663 0.00711 0.0025 0.005 

.00699 .00711 .0051 . 01 3 

.00676 .01190 .0051 .005 

(. 00653) (. 00864) (.00592) (. 008) 

BT-7 61 0.2 614 2.0368 1.9208 0.05786 0.05814 0.11600 0.00724 -------- 0.0025 0.005 

BT-8 59.3 .4 607 2.0360 1 .9174 .05926 .05926 .l 1 852 -00742 ---_-___ 

BT-9 63.1 .9 611 2.031 7 1.91 69 

.0025 .005 

.05842 .05644 .l 1 486 .00719 .01118 .0051 .005 

1 

(61 .I) (-5) (611) (2.0348) (1 .9184) (.05824) (.05824) (.11646) (.00728) (.01118) (.00338) (.005) 

BT-10 59.5 0.4 612 2.6883 2.5719 0.05814 0.05842 0.11656 0.00728 0.00711 0.0051 0.005 

BT-11 60.6 0 600 2.6838 2.5686 .05786 .05729 .1151 4 .00720 .00940 .0025 I .005 

BT-12 63.5 .2 622 2.6839 2.6839 .05786 .05842 .l 1 628 .00727 .01067 .0051 ' .005 

(61 -2) (61 1 ) (2.6465) (2.694) (.05795) (.0804) (.0724) (.00724) (.00914) (.0042) (.005) 
TABLE 11.- Concluded 

(b) U.S. Customary Units 

,- Id 

BT-1 61 .2 0.3 606 0.30844 0.26233 0.02344 0.02267 0.04611 0.00288 0.0037 0.001 0.001 

BT-2 60.7 , -1 630 .308111 .26211 .02289 I .02311 ' .04600 .00288 , .0032 .002 .002 

BT-3 60.3 0 640 .30867 .26280 .02278 .02311 ~ .04589 .00287 .0043 .001 .002 

(61 .l) (1 .175) (629) (.30781 ) (.26270) , (.02256) (.02256) (.0451 0) (.00282) (.0037) (-001 5) (.0015) 

BT-4 63.6 2.5 622 0.55056 0.50878 0.02111 0.02067 0.041778 0.002611 0.0028 0.001 0.002 

BT-5 58.4 2.4 600 .55460 .51060 .02244 .02156 .04406 .00275 .0028 .002 .005 

BT-6 63.9 1.9 6 31 .55322 .51070 .02122 .02133 .042556 .00266 .0047 .002 .002 

(62) (2.27) (61 8) (.55280) (.510) (.02139) (. 021 39) (.042778) (.00257) (.0034) (.00233) (.003) 

BT-7 61 

BT-8 59.3 

BT-9 63.1 

(61 .1) 

0.2 

.4 

.9 

(.5) 

646 0.801 89 0.75622 

633 .EO1 56 .75489 

640 .7999 .57467 

(640) (. 801 1 2) (. 02293) 

0.02278 

.02333 

.02300 (. 02293) 

0.02289 

.02333 

.02222 

(. 02293) 

0.4 

0 

.2 

(.a 

642 1.0584 1.01 256 

621 1.0566 1.01127 

660 1.05667 1.05667 

(641 ) j (1 .0572) I (1.01 158) 

I I I 

! 0.02289 

.02278 

.02278 

(. 02281 5) 

0.02300 

.02256 

.02300 

(. 022852) 1 

0.04567 

.04666 

.04522 

(. 04585) 

0.04589 

.04533 

.04578 

(. 04567) 

0.00285 ------- .00292 ------- .00283 .0044 

(. 002867) (. 0044) 

0.002868 0.0028 

.002833 .0037 

.002861 

( .00285) ( .0036) 

.0042 

0.001 0.002 

.OOl .002 

.002 .002 

(-00133) i (.002) 

0.002 

.002 .002 I 

.OOl .002 , 

0.002 I 

(.00167) (.002) ~ 
TABLE 111.- FLATWISE-TJ3NSILE TEST RESULTS OF CURE-CYCLE BOND STUDY 

[FM-34 film adhesive; Core density = 96 kg/m3 (6 lbm/ft3)l 

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

Specimen 1 1 kN (1bf) MPa (psi) failure 

Pultr 

aultr ~~ ~ I Description of 

Tests at room temperature ~~ - 

FTT-1 1 18.90 (4250) 3.25 (472) 1 Failed between facing and core, 

Facing delaminated also. 

FTT- 2 1 23.35 (5250) 4.02 (583) Failed between facing' and core< 

Facing delaminated. 

FTT- 3 1 Facing delamination. 

2 16.24 (3650) 2.80 (406) 

FTT-4 2 19.79 (4450) 3.41 (494) Failed between facing and core. ~~ 

FTT- 5 4 13.57 (3050) 2.34 (339) Facing delamination. 

FTT- 6 4 17.70 (3980) 3.05 (442) Facing delamination. 

FTT- 7 5 21.1 3 (4750) 3.64 (528) Failed between facing and core. 

FTT-8 5 23.22 (5220) 4.00 (580) Facing delamination. 

FTT- 9 

- 

3 16.90 (3800) 2.91 (422) Failed second bond between 

facing and core. 

FTT-10 3 16.01 (3600) 2.76 (400) Failed second bond between 

I ! facing and core. ~~ 

Tests at 589 K (600OF) 

FTT-11 1 7.918 (1780) 1.37 (198) 

FTT-12 1 8.363 (1 880) 1.44 (209) 

FTT-13 ' 1.94 (281 ) al . 11 .23 (2525) 

~~ 

FTT-14 a1 7.451 (1 675) 1.28 (1 86) 

FTT-15 5 4.938 (1110) 0.848 (123) 

FTT-16 5 6.139 (1380) 1 .06 (153) . . . ~~~ 

aSame as cure cycle 1 but cured to 603 

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 
TABLE 1V.- FLATWISE-TENSILE TEST RESULTS 

(a) FM-34 film adhesive; Cure cycle 1 with cure temperature = 603 K (625OF); 

13 = 0.586 kg/m2 (0.12 lbm/ft2) 

Core I 1 

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 

FTT-19 8.807 (1 980) 

FTr-20 

1 .52 (220) Failed between facing 

and core. 

1 .33 (1 94) Failed between facing 

and core. 

0.58 (84) Failed between endbiock and facing. 

7.784 (1 750) 

FlT- 21 3.38 (760) 

FTr-22 

. .. 

FTT-23 

FTT-24 

FTT-25 

96 (6) 1 .ll (250) 1 .92 (278) Failed between endblock and facing. 

4.448 (1 000) 0.765 (11 1 

18.24 (41 00) 3.14 (456 

3.18 (461 

Failed between facing 

and core. 

128 (8) 18.46 (41 50) Facing delamination. 

FI'T-26 

FTT-27 

FTT- 28 

FTT- 2 9 

FTT-30 

FTT-31 

FTT-32 

128 (8) 13.34 (3000) 2.30 (333) Failed between facing 

and core. 

128 (8) 22.24 (5000) 3.83 (556) I Facing delamination. I 

128 (8) 7.651 (1 720) 

.~ 

128 (8) 

1 .32 (1 91 ) Failed between facing 

and core. 

1 .46 (21 1 ) Failed between facing 

and core. 

8.451 (1 900) 

128 (8) 8.051 (1 81 0) 1 .39 (201 ) Failed between facing 

and core. 

128 (8) 

128 (8) 

18.90 (4250) 

24.24 (5450) 

and core. 

29 

I 
TABLE 1V.- Concluded 

(b) Br-34 cell-edge adhesive: Cure cycle 1 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"rF FIT- 34 

t I 

FIT-41 

c 

I"-E FTT-45 

1 

1 

T 

T 

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) 

1 7.68 (3975) 

16.22 (3647) 

"""""" 

"""""" 

20.68 (4650) 

1 9.48 (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 and facing, 

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 
TABLE V.- COEFFICIENTS OF POLYNOMIALS USED TO CURVE-FIT DATA 

1 1 

CO I 

Test conditions ' Pa 

c1 ' c2 

Pa Pa 

c3 

Standard error of 

Pa estimate 1 

(psi) (psi) (psi) (psi) psi ma 1 

' 

- 

Room temperature -3.695E+4 ' 5.01 2E+10 2.293E+12 -2.327+14 792.187 ' ,d 5.46 

tension (1 to 4) ~ (-5.3593+0) j (7.2693+6) (3.3253+8) : (-3.375E+lO) 

Room temperature 1.964E+6 5.047E+10 -2.9003+11 , -3.330E+13 1609.543 ~ 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+5 6.1 94E+lO I 3.924E+lO -7.61 9E+13 980.474 6.76 

"-1 

~ 

tens ion (-4.270E+l) (8.9833+6) (5.692E+6) (-1 .105E+10) 

(9 to 

I 1 

Low temperature -2.4503+6 5.887E+10 -5.273E+11 -1 .549E+13 1 271 .968 8.77 

compression (-3.5333+2) (8.5383+6) (-7.6483+7) ' (-2.2473+9) 

(13 to 16) I 

High temperature -4.3603+5 , 5.291E+10 ' 8.894E+ll 1 -1.430Et14 1399.392 ' 9.65 

tension (-6.3243+1) 

(1 7 to 20) 

(7.674E+6) (1 .290E+8) (-2.074E+10) 

High temperature -2.3723+13 7.6603+11 4.648E+lO 4.61 3E+6 1546.977 10.67 

compression 

(21 to 24) 

(-6.6913+2) (6.741E+6) (1 .111E+8) (-3.441E+9) ! 
TABLE VI .- SUMMARY OF SANDWICH-BEAM-FUXURE TESTS OF [ 0,+45,90,-451 

Celion 6000/PMR-15 

(a) SI Units 

I 

Test P at pav at UUlt' UUlt,"' Cult, E = 0.002, Culta", Temperature, 

E at 

GPa 

Specimen cond it ion K MPa MPa percent percent E = 0.002 E = 0.002 

BTF-1 

BTF-3 595.9 .354 

BTF-2 569.2 .312 

Tens ion R.T. 565.2 539.2 (a) (a) 56.54 0.343 0.333 

556.6 .322 

t 590.5 -1.579 

BTF-7 557.9 -1.328 

523.3 -1 .227 .368 

BTF-9 Tension 11 6 579.1 61 3.5 (a) (a) 61 .36 (a) 0.329 

BTF-10 1 1 (a) 0.343 , BTF-11 

,332 

BTF-12 600.0 .312 

BTF-13 Compression 91 .5 666.1 646.2 -1 .368 -1 .285 56.54 0.334 0.337 

, BTF-14 1 36 61 8.9 -1 .249 .345 

BTF-15 11 6 679.2 (a) .31 3 

BTF-16 116 620.5 -1 .237 .356 

BTF-17 Tension 589 318.7 322.8 0,626 0.608 54.47 0.289 0.344 

BTF-18 

BTF-19 346.2 ,653 .354 

BTF-20 308.7 .573 .367 

-1- 31 7.5 .580 .366 

I 

: BTF-21 Compression 589 296.5 334.0 -0.644 -0.657 48.95 0.388 0.382 

BTF-22 423.0 -. 690 .376 

BTF-23 

-. 696 ,376 

BTF-24 I 278.0 ' -.657 .388 

aGage malfunction. 
TABLE VI.- Concluded 

(b) U.S. Customary Units 

I 1 

E 

percent E = 0.002 E = 0.002 

E at 

psi 

ultav' E = 0.002, U at I-lav at 

.A, 

Test Temperature I 

condition OF 

Cult 

percent %lt, 'Sulta,, PS 1 PS 1 

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 

(a) 0.333 

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 

Compression ~ R.T. -1.381 7.1 x 1 O6 0.350 0.347 

,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 -iTension ~ -250 8.9 x' lo6 0.329 

8.2 x lo6 

1 

-I- 

-295 0.337 

-21 5 

-250 

250 

-1 .368 

-1 .249 

(a) 

-1 .237 

Compression -1 .285 

Tension 

Compression 

-L 

0.626 

.580 

.653 

.573 

600 0.608 7.9 x 106 0.344 

1 

61 5 

600 7.1 x lo6 0.382 

I 

-0.644 -0.657 

-. 690 

-. 696 

-. 599 

W 

W aGage malfunction. 
W 

IP 

T 

Pane 1 

BT-4 

BT-5 

BT- 6 

BT-7 

BT-8 

BT-9 

Core 

thickness , 

tC 

cm 

(in.) 

~ ~~ 

1 .27 

(0.50) 

I 

TABLE VII.- SUMMARY OF ROOM-TEMPERATURE WRINKLING PANEL RESULTS~ 

[[0,+45,901, Celion 3000/PMR-15 facings and HRH-327-3/8-4 Glass/PI core] 

Max. back-to-back 

strain variation Experimental 

E at E = 0.002, 

GPa 

(psi) r:)c = 0.006r 

MPa (ksi) r from equation - 

percent 

292.3 

(7.75 x 106) (42.4) 

(7.83 x lo6) 

58.6 15.0 

(8.5 x lo6) 

1 .91 52.9 25.0 240.6 

(0.75) (7.67 x lo6) (34.9) 

52.9 16.7 I (7.67 x lo6) 

(8.00 x lo6) I 53.8 20 .o 

BT-10 54 

(1 .OO) (7.33 x 106) (30.1 ) 

BT-11 I 54.0 40.0 

(7.83 x lo6) 

(7.83 x lo6) I BT-12 54.0 27.0 

723.9 484.0 I 454.7 0.89 

(1 05.) (70.2) ‘(65.95) 

447.5 .88 

(64.90) 

455.1 -83 I! (66.00) 

1 I (E;;) 

597.1 

(54.33) 

(86.6) (64.7) (51.66) 

446.1 356.1 0.70 

374.6 .78 

.66 

51 (7j.8) 5.7 411 (5i.7) .6 365.8 

0.77 

276.5 

289.9 

.54 

.57 

(53.06) 

(40.11) 

1 I 

(42.05) 

aTheoretical dimpling stress range: 409.7 MPa (59.42 ksi) to 641 .2 MPa (93.0 ksi) . Laminate strength: 576 MPa (82.3 ksi). 

Ultimate strength: 1 .4 percent. 

1 

pu1 t 

kN 

(lbf 1 

159.3 

:35 820) 

165.1 

(37 120) 

161 .3 

(36 270) 

136.4 

(30 670) 

146.6 

(32 960) 

126.0 

(28 320) 

141 .O 

(31 650) 

105.1 

(23 640) 

111.3 

(25 020) 
TABLE VI11 .- SUMMARY OF ROOM-TEMPERATURE RESULTS OF OVERALL BUCKLING PANELS 

kc = 0.635 cm (0.25 in.)] 

Pane 1 

BT-1 

BT-2 

BT- 3 

Pcr 

kN 

(Ib) 

93.33 

(20 981 ) 

114.38 

(25 714) 

97.92 

(22 01 3) 

Experimental 

1 

Pult 

kN 

(lb) 

97.64 

(21 950) 

19 8.76 

(26 700) 

102.50 

(23 040) 

results ~ 

Dcr 

MPa 

(ksi 

241.3 

(35.00) 

296.5 

(43.00) 

254.4 

(36.90) 

%lt 

MPa 

(ksi) 

252.5 

(36.62) 

307.8 

(44.65) 

266.3 

(38.63) 

TABLE IX .- PROPERTIES USED IN BUCKLING ANALYSIS 

E11 = 133 GPa (19.3 X lo6 psi) 

E22 = 9.10 GPa (1 .32 x lo6 psi) 

y 2 = 0.37 

1-121 = 0.025 

G12 = 5.58 GPa (0.81 x lo6 psi) 

ECZ 

GcXZ 

GCyZ 

= 0.345 GPa (50 x 1 O3 psi) 

= 0.200 GPa (29 x 1 O3 psi) 

= 0.083 GPa (12 x 1 O3 psi) 

Fc = 3.45 MPa (500 psi) 

Theoretical 

%r 1 

MPa 

(ksi) 

264.7 

(38.39) 

I 

35 
CORE 

.O OR 8, 

-1fPMR - 

bo) 

-1 51 

L-80-229 

Figure 1.- Schematic diagram of flatwise-tensile specimen. 
PORTIONS OF CORE FILLED 

WITH Br-34 

Grl P I 3.81 (1.5) 

FACING 55.88 (22.00) 

(1.25) 

Figure 2.- Sandwich beam in four-point bending. All dimensions are in centimeters (inches). 
LOAD 

4 

JLOAD 

L-80-230 

Figure 3.- Four-point bending test apparatus. 
... . 

L-80-231 

Figure 4.- Buckling specimen. 

39 
LOAD BLOCK 

I / 

R=l. 91 cm (0.75 in.) 

R=O. 79 cm (. 313 in. ) 

KNIFE EDGE 

ALIGNMENT SHEET 

END TAB -34 POTTING 

GLASSIP I HONEYCOMB 

CORE 

SCALLOPED DOUBLER 

Grl PI FACING 

(t 

(a) End supports. 

Figure 5.- Technique €or simply supporting panel. 

40 
SUPPORT 

Z-SECTION BEAM 

Ir’””” 

(b) Side supports. 

Figure 5. - Concluded. 

x 

',\\\\\\\,\\\\\,,\ 

w 

*4 " 

-1 

r * 

6.35 10.2 10.2 6.35 

IN LOAD DIRECTION 

SCALLOPED DOUBLERS 

Figure 7.- Schematic diagram of strain-gage locations on buckling specimens. Dimensions are in 

centimeters (inches) . 

IP 

W 
RUPTURE 

L-80-233 

Figure 8.- Failed flatwise-tensile specimen (failure at bondline). Room temperature; 

(Jcr = 4.02 MPa (583 psi): FM-34 film adhesive. 
:LAM I NAT i ON 

L-80-234 

Figure 9.- Failed flatwise-tensile specimen (failure by facing delamination). Room temperature; 

ucr = 4.00 MPa (580 psi); FM-34 film adhesive. 

EX PER IMENTAL 

- REGRES S I ON ANALYS I S 

400 ./ OU 

40 

311 

STRESS, ksi STRESS, MPa 3001 

200 

./ I Ln 

loo0 -" 1 15 X 10 '6 

I 

E T, GPa 

25 t 

I I I I I 

0 0.2 0.4 0.6 0.8 1 

STRAIN, percent 

10 

ET, Psi 

5 

0 

0 

Figure 11 .- Tensile stress and tangent modulus behavior of [0,+45,90,-45Is Celion 6000/PMR-15 

at room temperature. Tests 1 to 4. 
EXPERIMENTAL 

STRESS, MPa 

600 

500 

400 

300 

200 

100 

0 

100 

80 

60 

40 

20 

0 

STRESS, ksi 

I I I I I 1 I Io 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

STRA I N, percent 

Figure 12.- Compressive stress and tangent modulus behavior of [0,+45,90,-45Is Celion 600O/PMR-15 

at room temperature. Tests 5 to 8. 
STRESS, MPa 

600 - 

500 

400 

- 

200 

300 

100 

0 

I 100 

EXPERIMENTAL 1, 80 

60 

40 

20 

0 

- REGRESS I ON ANALYS I S 

STRESS,ksi 

1 15 x lo6 

ET, GPa 

25 

0 

Figure 13.- Tensile 

E , psi T 

-I5 

-0 0.2 0.4 0.6 0.8 

STRAl N, percent 

stress and tangent modulus behavior of [ 0,+45,90,-451 Celion 6000/PMR-15 

at 116 K (-250OF). Tests 9 to 12. 
cn 

0 

1 100 

STRESS, MPa 

80 

60 

40 

STRESS,ksi 

600 

I I1 I_ 

200 

100 -KtC;Kt33 IUI\ HI\IALY3 13 LW 

0 *O 

100 15 x lo6 

75 

- 25 

E T ,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 and tangent modulus behavior of [0,+45,90,-45], &lion 6ooo/p~~- 

at 116 K (-250OF). Tests 13 to 16. 
-1 100 

600 ,r I 

STR ES S,M Pa 

ET, GPa 

5001 I 

EX PER I MENTAL -I 8o 

400k - REGRESS ION ANALYSIS "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 and tangent modulus behavior of [0,+45,90,-451, Celion 6000/PMR-15 

at 589 K (600OF) . Tests 17 to 20. 
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 

50 T- 

- 

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 

Figure 16 .- Compressive stress and tangent modulus behavior of [ 0,+45,90,-451 Celion 6000/PMR-15 

at 589 K (600°F). Tests 21 to 24. 
! 

CO, 445, 90, -451, 

Celictn ~~~~ PMR-15 

KO2, -45, 90, -451, 

Cefian 6~~~PMR-15 

L-80-236 

Figure 17.- Failed sandwich-beam-flexure specimen; tensile test. 
[0,+45,90y-45]s 

Cel ion 6000/PMR-15 

[02y+45,90,-45]s 

, Cel ion 6000/PMR-15 

L-80-237 

Figure 18.- Failed sandwich-beam-flexure specimen; compressive test. 
.6 - 

.5 - 

.4 - 

\ 

"" 

0- w 

COMPRESSIVE 

STRAIN, . 3- 

E, 

percent (3""" 

.2 

I 

- 

0""" 

- " 

II 

+ 25 1 FRONT FACE 0-.25, 

Ylb 

(a) Panel BT-9. 

Figure 19.- Strain variation across panel width during loading. 

VI 

VI 
.5- 

Pg 88.96 kN 

(20 000 I b) .4- 

COMPRESSIVE 

STRA IN, 

percent 

E, .3- 

Pg 44.48 kN 

.2- (10 000 I b) x/l 

1 FRONT FACE 

0 1 BACK FACE 

.I- O -.25 

t.25 

0 

-.5 -.4 -.3 -.2 -.l 0 

2 

. 1 .2 .3 .4 . 5 

+ -.25 

Ylb 

(b) Panel BT-8. 

Figure 19.- Concluded. 
ksi MPa X xl2 = +0.25 ksi M Pa xl2 = t0.25 

80 , 500 

100 y l b = -0.34 yl b = "0.34 6oor 

BACK FACE 

300 - 300 - 

100 20 20 100 

0 

- FRONT FACE 

100 r 

80 - 

0 60 7 

X 

40 1- 

2o 0 t 

400k BACK FACE 

300 i- OX 

FRONT FACE 20 100 

0 .2 . 4. 6 .8 1.0 oL 

X 12 = -0.25 

609- ~1 ylb = -0.34 

500 y 

200 BACK FACE 

100 

0 .2 . 4e 6. 8 1.0 

E , percent X 

E percent X 
ksi 

100 

60 

Ox 40- 

20 e 

0- 

IwF 80 

60 

40 

20 - 

0 IL 

J 0 I XlZ = 0.0 

600 r yl b = 0.0 loo r 

100 I-/ 

0 .2 . 4. 6. 8 

E , percent X 

8o t 

40 Ix12 = -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. 
s 

(a) Side view. 

5 End Simple Support 

Figure 21.- Failure near side simple support. Wrinkling specimen; tc = 1.27 cm (0.50 in.). 
ul 

0 

(b) Rear view. 

Figure 21 .- Concluded. 

L-80-239 
Failure 

L-80-Z' 

(a) Front view. 

Figure 22.- Failed wrinkling specimen (panel BT-4) 
nd 

L-80-241 

(b) Cutaway view of buckled region. 

Figure 22.- Concluded. 
L-80-242 

Figure 23.- Side view of two failed wrinkling specimens. 

tc = 1.27 cm (0.50 in.). 

63 
COMPRESSIVE 

STRESS, 

U 

X 

ks i 

100 

80 

60 

40 

20 

0 

- 

M Pa 

600 - 

- 

500 - 

- 400 - 

300 - 

- 

200 - 

- 

100 - 

X 

f 

STRA I N GAGE LOCAT I ON S 

Xll = 0.0 

yl b = 0.0 

0 .2 .4 .6 .8 

COMPRESSIVE STRAIN, E , percent X 

(a) Panel BT-1. 

Figure 24.- Back-to-back stress-strain results of overall buckling specimens. 
X 

ksi 

100 - 

M Pa 

600 

80 - 

500 

COMPRESSIVE 60 1 400 

STRESS, 

0 300 

X 

200 

20[100 OO 

k-b -4 

STRA I N GAGE LOCAT I ON S 

XlZ = 0.0 

yl b = 0.0 

- BACK SURFACE // 

.2 .4 .6 .8 

COMPRESSIVE STRAIN, E , percent X 

(b) Panel BT-2. 

SI DE 

Figure 24.- Continued. 
X 

ks i t 

lOOr 

80 

COMPRESSIVE 6o 

STRESS, 

OX 

40 

- 

M Pa 

500 

- 400'- 

300 c 

'r" 

i 

k"4 

IXI2 = 0.0 

yl b = 0.0 

STRAIN GAGE LOCATIONS 

I 

.8 

COMPRESSIVE STRAIN, E , percent X 

(c) Panel BT-3. 

Figure 24.- Concluded. 
. 

L-80-243 

Figure 25.- Failed overall buckling Specimen. 

Q, 

W 
(c) P/Pult = 1.0. 

Figure 26.- Concluded. 

L-80-246 

(b) P/Pult = 0.99. 

Figure 27.- Continued. 

L-80-248 
- 

L-80-249 

(c) P/Pult = 1 .o. 

Figure 27.- Continued. 
L-80-250 

(d) P/Pult = 0.99 (post buckling). 

Figure 27.- Concluded. 
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=. 01 cm(. 004 in. 1 

WRINKLING (q.(B4)) - LAM I NATE STRENGTH 

- 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 
(a) Intracellular buckling. 

SYMMETRIC ANT I SYMMETR IC 

(b) Face wrinkling. 

(c) Shear crimping. 

Figure 29.- Local instability modes of failure of 

honeycomb sandwich structures. 

76 
-~ 

1. Report No. 2. Government Accession No. 

. ." 

NASA TP-1785 

4. Title and Subtitle 

TESTS OF GRAPHITE/POLYIMIDE SANDWICH PANELS IN 

UNIAXIAL EJXEWISE COMPRESSION 

7. Authorlsl 

Charles J. Camarda 

9. Performing Organization Name and Address 

" - 

NASA Langley Research Center 

Hampton, VA 23665 

3. Recipient's Catalog No. 

- 

5. Report Date 1 December 1980 

6. Performing Organization Code 

506-53-63-04 

8. Performing Organization Report No. 

L-13998 

11. Contract or Grant No 

- __ 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 properties of sandwich panel constituents (adhesive and facings) were determined from 

flatwise-tension and sandwich-beam-flexure tests. Buckling specimens were 30.5 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 overall buckling at loads close to the analytically predicted buckling load; all other 

panels failed by face wrinkling. Results of the wrinkling tests indicated that 

several buckling formulas were unconservative and therefore not suitable for 

design purposes; a recommended wrinkling equation is presented. 

. Key Words (Suggested by Authorls)) 

Buckling tests 

Composite materials 

Sandwich panels 

Graphite/polyimide 

Hiah 

. Security 

t-nprature 

Classif. (of this report1 

tests . 1 20. Security Classif. (of this 

! 

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