Sharma, A. V.

English

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(NASA-CIR-153961) EFFECT OF TEMPERATURE ON N77-83149
 
COAPOSITE SANDWICE STRUCTURES SUBJECTED TO
 
LOW VELOCITY IMPACT Annual Report, May 1976
 
- Aug. 1977 4North Carolina Agricultural and Unclas
 
Technical) 25 p 00/24 42060
 
ANNUJAL REPORT
 
EFmCT OF TEMPERATURE ON COMPOSITE SANDWICH STRUCTURES 
SUBJECTBD TO LOW VELOCITY INPACT 
BY 
A. VISHNU SiARNA 
AUGUST4 1977 
NSG - 1296 1976077-
Prepared for 
Langley Research Center
 
National Aeronautics I Space Administration
 
-Period Covered: May 1976 - August 19770
 
PNATIONA TECHNICAL 
INFORMATION SERVICE 
,DEPARIMEWP OF COMMERCE -V. 22161SPRINGFIEtD: 
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INTRODUCTION 
The use of granhite/epoxy (G/E) composites and composites fab­
ricated out of similar fibrous materials in the design of secondary
 
components of aircraft is gaining momentum in recent years. These
 
secondary structural components are generally designed withhoney­
comb core sandwiched between composite facings. In normal operation­
al mode, these components may be exposed to foreign object damage
 
such as the dropping of hand tools, runway debris, etc. Consequently,
 
it is of interest to study impact related damage caused by foreign
 
objects and develop design criteria in the use of composite sandwich
 
structures. Researchers such as Slepetz [11*, et al., Rhodes [2,3,
 
4], Awerbuch and Hahn [5] have conducted several studies on composites.
 
These composites, besides G/E, include graphite/S-glass, boron/aluminum,
 
borsic/titanium and other combinations. Different test procedures and
 
specimen geometries have been used in the above studies. In particular,
 
the studies of Rhodes are of interest and the findings of the present
 
report are related to the earlier studies by Rhodes.
 
Rhodes in his studies has performed impact tests using G/E and
 
Kevlar-49/epoxy sandwich specimens at room temperature. Some of
 
these studies were published [3] and other results have been made avail­
able to the author through private communications. Visible damage-and
 
failure thresholds have been developed. Preload and impact energy
 
necessary to initiate catastrophic failure of specimens were determined.
 
* Numbers refer to bibliography. 
I­
2 
The impact velocity was in the range of 16-67 meters/second (52-220
 
ft/sec.). The residual strengths of specimens that survived the impact
 
at a particular energy level were also measured.
 
The effect of moderately high-and low-temperatures on the strength
 
carrying ability of composites under impact is reported here. Since 
the number of specimens tested was small, only few conclusions with re­
gards to the existance of failure threshold and residual strength may
 
be drawn. However, in conjunction with room temperature test results,
 
some trends could be seen in the behavior of selected composites under
 
impact damage.
 
SPECINENS 
The specimens were of sandwich bean type with a honeycomb core. 
The nominal dimensions of the specimen were 56 cm (22 in.)Jlong by 
8 cm (3 in.) wide with a thickness of 2.5 cm (1 in.). The test face 
(front) was a laminate and the back of the honeycomb was supported by
 
a steel plate. A view of the specimen is shown in Figure 1. A typical 
core, 8 cm (3 in.) by 8 a (3 in.) with a thickness of 2.5 cm (1 in.), 
was located in the test area where uniform flexural stress was induced 
through a four-point beam bending apparatus. Since the two outer sections 
of the specimen were subjected to high shear stress, a dense aluminum 
core was used in these two sections. 
Two types of composite sandwich beam specimens were tested. The details of these
 
specimens are given below;
 
Face Sheet Material No. of Plies Lamina Orientation Honeycomb Core in Impact Area 
Narmco 5208 - T300 8 (90, ±45, 0)S Alum. 130 kg/m3 '(8.1 Ibm/ft3) 
G-5208 - T300, 
K-5208 - Kevlar 6 (OK, 90, OG)S Nomex 48 kg/m3 (3.0 ibm/ft3) 
The 8 - ply specimen is designated as Graphite/Epoxy.(G/£) and the 6 - ply specimen as
 
Graphite-Kevlar/Epoxy (Hybrid). All the specimens were supplied by the Langley Research Center,
 
NASA.
 
4 
EOUIPMENT AND TESTS 
The schematic of the equipment set-up is shown in Figure 1. The 
description and operation of the air gun to propel the projectile, an
 
aluminum sphere having a diameter of 1.27 cm (0.5 in. with a mass-of 
2.9 grams (0.0065 ibm), was given'in [3]. The velocity of the projectile 
was measured with an electronic counter using start/stop diodes. The 
four-point beam bending.apparatus was enclosed in a heating/cooling 
chamber. The description and schematics of chamber construction were 
reported in an earlier retort [6]. The testing techniques such as the 
measurement of load, strain and velocity were described in a paper by 
Rhodes[3]. 
Testing at high or low temperatures was perfoned after attaining 
a steady temperature in the test chamber. Six temperature sensing 
probes - one at load cell, one on the surface of the specimen in the 
vicinity of impact zone, and four others at various points in the chamber 
volume - were used. A temperature control probe was located in the ­
vicinity of the specimen. Initiation of actual testing. (loadJg/_i,nrTLg:­
took place after the predetermined surface temperature was reached. 
In general, two specimens at low temperatures were tested back to-back 
whereas only one specimen was tested'at a time at high temperatures. 
The average total heating/cooling times and the corresponding temper­
atures are given below: 
Heating Cooling 
Surface Temp: 366 ± 10 K (1990F) 223 ± IK (-590F) 
Ambient Temp: 369 ± 1 K (2040F) 216 ± IK (-710F) 
Temp. at load cell: 371 ± 10 K (208°F) 248 ± 1°K (-140F) 
Time to attain
 
surface temp. 135 minutes 150'minutes
-
Trial runs were conducted prior to actual testing of snecimens to
 
assess the free expansion/contraction characteristics of the G/E beam
 
as well as the four-point loading frame. The'load cell was built into
 
the loading frame. The absolute value of thermal strains (loading screw 
was not contacting the specimen) on the surface of a 12-ply G/E laminate 
was found to be less than 10-3 . Since thermal environment also creates 
deformations of the loading frame structure which in turn may induce 
thermal load on the specimen, it was decided to measure the temperature 
induced load on the load cell. It was found that this load was less 
than 7 kg (15 lb.)' In actual testing of specimens these initial values 
were rebalanced to zero.
 
The following types of tests were performed:
 
/Hbrid
 
Heating: Tension/Compression Tension/Compression 
Cooling: Tension/Compression Tension/Compress ipn 
RESULTS
 
In the testing of G/E and hybrid composite sandwich beans; approx­
imately 5-8 specimens were tested for each. case such as heati.g/tension, 
etc. Two to three specimens were used in the-evaluation of ultimate 
strength and the balance of specimens were used to study the fail7re/ 
non-failure mode under impact. Those specimens that survived the impact
 
were subjected to further loading to calculate the residual strength. 
Graphite/Epoxy Beams 
Heating/Tension: The variation of normalized stresses and strains 
with respect to kinetic energy of impact is shown in Figures 2 and 3. 
The impact energy level, by design, was held almost constant. At an 
energy level between 1.2 joules and 1.5 joules (11 in-lb and 14 in~lb), 
two of the specimens survived the impact at a pre-stress level of 43% 
of the ultimate strength; The loading was continued to assess the
 
residual strength of the above two specimens. It was found that they 
both failed at about 51% of the ou (ultimate strength). Another 
virgin (no preload) specimen failed at 50% of. u. It was noticed that 
a fourth specimen,has failed by debonding at a stress level of 91%
 
of au . The term debonding is used here when there is a separation of
 
either the laminate or steel back plate from the core. The strain values
 
corresponding to the above stresses were found to be about 2%higher,
 
i.e., at about 53% of Cu(ultimate strain). An imaginary line.repre­
senting the failure threshold has been drawn through these points based 
on the present data and the data by Rhodes, published [3] and unpublished.
 
Heating/Compression: Six specimens were tested in this series and 
the data for a. and Fu against impact energy is shown in Figures 4 
and S. At an energy level of 1.5 joules (13.7 in-lbs), three specimens 
were impacted. The average pre-stress level for two specimens was 
around 57% of ou . These two specimens survived the impact having ,a 
residual strength of 67% of cu". The preload and, tesidad, sten-th"fat. 
a third -specimen were just about the same. The last specimen in this 
series failed catastrophically at higher nreload and energy. The strains 
followed a similar pattern as in the case of heating/tension. The strains 
for the third specimen, refered above, were not as close as stresses 
between impact and residual strength evaluation stages. 
Cooling/Tension: Seven specimens were tested in ftis s&ries.' The 
stress/strain variations with impact energy are shown in Figures- 6 and 
7. Even though the impact energy was not exactly the same, two of the 
specimens were pre-loaded to a stress level of 44% of au . These did not 
fail and had an average residual strength of 50% of au One of the.
 
7 
other specimens failed at 58% of ou even though the energy level was 
less than the two previous specimens. The values of strain at the level 
of residual strength was about- 48% of eu" One of the specimens appear­
ed to have failed by debonding but the absolute values of strains and 
stresses were in the range of values for au and ou . 
Cooling/Compression: The impact energy levels in this series of
 
6 specimens tested were higher. The variation of stress/strain with
 
energy as shown in Figures 8 and 9 has the same pattern as in the case
 
of heating/compression. However, the absolute value of ou in this case
 
was about 69 MPa (10 ksi) higher than in the other case. The strains
 
were also higher but this difference was of the order of 10"..
 
Kevlar-Graphite/Epoxy Hybrid .Beams
 
Heating/Tension and Compression: Due to debonding-of the specimens
 
in this series, ou in both tension and compression tests were not mea-­
sured. However, the-hybrid specimens were able to withstand .astress
 
level of 807 Wa (117 ksi) in tension and 276 Ia (40 ksi) in compression " 
before they failed by debonding. The corresponding strain levels were 
1% and 0.48%, respectively. Some of the specimens in this-series-wee-sub-­
jected to impact in tension. Itwas found that they were able to sustain­
a pre-load of 393 MPa (57 ksi) at 1.7 joules (15 in-lbs) of impact
 
energy. The residual strength for these specimens was found to be
 
around 490 MPa (71 ksi).
 
Cooling/Tension: Six hybrid specimens were tested in this series.
 
The ultimate stress (Yu) level was found to be 820 P-Ta (119 ksi) with a
 
corresponding strain (Eu) of 1%. The variation of stresses and strains
 
with respect to energy of impact is shown in Figures.10 and 11., The 
residual strength of two specimens that survived the pre-load (43% of au)
 
impact was about 56% of ou. However, the corresponding values of strains
 
8 
were not proportional but higher (Figure 11). The failure threshold in
 
this case was slightly less than 50% of au.
 
Cooling/Compression: Nine specimens were tested inthis series.
 
The average value of au was found to be 352 Mla (51 ksi) with a corre­
sponding value of eu at 0.5%. As shown in Figures 12 and 13, the impact
 
energy range is wide for this series. The width of failure threshold
 
is also wide.
 
-DISCUSSION& CONCLUSION
 
One of the objectives of these tests was to see the degradation of
 
properties, if any, of composite sandwich beams at moderately high and'
 
low temperatures as compared to room temperature data reported by Rhodes
 
([3] 	and unpublished data). Consequently, it has been observed that:
 
1. 	Failure threshold can be established as a function of 
impact energy level. 
2. The G/B specimens in heating/tension have a value of
 
au of 631 MWa (91.5 ksi) which is about 8% lower than
 
the room temperature value of 686 MPa (99.5 ksi).
 
3. 	The G/B specimens have shown similar percentage drop of
 
ou in heating/compression, 455 MWa v.s 563 MTa (66 ksi
 
v.s 81.6 ksi), tests.
 
4. Although the ultimate strength, au, in heating/tension,
 
631 MPa (91.5 Ksi) is higher than in heating/compression,
455 NPa (66 Ksi), the specimens in compression were able to
 
withstand a higher percentage (57% ,vs43% of au) value
 
of pre-inpact loads. The residual strength of these
 
specimens also was higher (67% vs 51%).
 
5. The ultimate strength of G/B specimens remained same in
 
both cooling/heating (tension) tests.
 
6. 	Low impact energy would cause catastronhic failures
 
at higher pre-load can be seen in Figure 6.­
7. The ultimate strength in cooling/compression is about
 
68.9 MPa (10 ksi) higher than the corresponding value
 
in heating/compression.
 
9 
8. 	 The ultimate strain values in heating/cooling were in the 
range of 0.9% to 1%. The strain values in cooling appear­
ed to be slightly higher than those in heating with a 
difference of 0.1%. 
9. The Kevlar-Graphite/Epoxy composites (called hybrids in
 
this report) have failed by debonding in tests at high
 
temperatures. This failure was attributed to improper 
fabrication procedures.
 
10: 	 Eventhough debonding occured in heating, the ultimate strength
 
values for hybrids in tension (heating/cooling) were found
 
to be close: The corresponding strain values were also
 
very close.
 
11. 	 The ultimate strength in cooling/tension was found to be
 
about 	10% lower than that of the corresponding room temoer­
ture 	values whereas in cooling/compressioA, the difference
 
was about 4%.
 
12. The failure threshold for hybrids in cooling/compression 
was found to be higher (60% of au) than that in cooling/
 
tension (50%).
 
Based on the limited amount of testing, it can be said that de­
gradation of the strength of composites exists at moderately high/low
 
temperatures over that of room temperature values. This degradation
 
was found to be less than 10%. 
RECOMMENDATIONS
 
As a 	result of the limited number of tests perfnrmed,_ somF' trend_ 
in the behavior of sandwich composites under thenro-mechanical loading
 
and impact can be seen. However, it is desirable to have positive
 
conclusions leading to design criteria. In view of this, it is recom­
mended that:
 
1. 	 Further testing be conducted to ascertain the strength deg­
radation of composites under thermal environment; 
2. 	Specimens be tested after they have been exposed to thermal
 
cycles with a 12 hour duration of heating at constant temper­
ature followed by cooling;
 
3. The variables such as impact velocity, preload levels be
 
limited so that positive conclusions as to strength deg­
radation may be drawn. 
BIBLIOGRAPHY
 
1. Slepetz, John Y., et al., "Impact Damage Tolerance of Graohite/ 
Epoxy Sandwich Panels", AMC TR 74-20, Sept. 1974.
 
2. Rhodes, M. D., "Low Velocity Impact on Composite Sandwich Structures," 
Presentation at the Second Air Force Cenfenerce on Fibrous Com­
posites in Flight Vehicle Design., Dayton, Ohio, May 1974. 
3. Rhodes, M. D., "Impact Fracture of Composite Sandwich Structures", 
presentation at the ASME/AIAA/ASE 16th Structural Dynamics, and 
Materials Conference, Denver, Colorado, May 1975. 
4. 	Rhodes, N. D., Williams, J. G., and Starnes, Jr. J. 13.. 'T ffect' 
of Low Velocity Inpact Damage on the Compressive Strenglh of 
Graphite/Epoxy Hat-Stiffened Panels', NASA TM X-73988, December 1976. 
S. Awerbuck, J., and Halm, H. T., "Hard Object Impact Damage of 
Metal.Matrix Composites", J. of Composite Materials, vol. 10,
 
July 1976.
 
6. Sharma, A. V., "Effect of Composite Sandwich Structures Subjected
 
to Low Velocity Impact", Status Report, N. C. A &T State Univ., 
Greensboro, N. C., December 1976. 
RESERVOIR PROJECTLE PROBE 
SA E PE 
FIG. .. 
OEQMCEONT STP 
EARL OUNEST_ 
SCEMATIOF GU!FSPNTCST-EU 
CHAMBER 
12 
(9oj45,0)5 GRAPHITE / Epoxy 
AVERAGE SURFACE TEMPERATURE 368- I K 
-7U(ULT. IN -LBS 
4 [ 1 16 20 
1.2L ® FAILURE 
A NON-FAILURE 
DEBONDING 
0 RESIDUAL 
0.6 
OA 
1 I. 1. I1
 
0.0 o.4 0.8 l.a 1-6 z.0 .4 
KINETIC -ENERGY, JOULES 
FIG 2., HIGH TEMPERATURE, :TENSION 
13 
(9o,0 t 45,0) GRAPHITE Epoxy 
AVERAGE SURFACE TEMPERATURE 366" 1° K 
6 / CULT. IN-LBS 
0 4 8 12 16 20 
1O. 	 '9 FAI LURE 
A NON-FALURE 
E0 DEBONDING 
0 RES.! DUAL0.8 
A 
0.4 	 A 
0.2­
0.0- 0.4 0.8' 1.2 1.6 	 ?A2.094 
KINETIC ENERGY, JOULES 
FIG 3. HIGH TEMPERATURE, TENSION 
14 
(9o, ±45, 0) 5 GRAPHITE / Epoxy 
AVERAGE SURFACE TEMPERATURE 366 ±10 K 
4/CULT. IN- LBS 
0 4 8 
A, 
12I 16K-­ 20 
to 
0.8-
A 
0.4­
0 RESIDUAL. 
S FAILURE 
02-, A NON- FAILURE 
El DEBONDING 
0.0" 0.4 0.8 1.2 1.6 2.0 2.4 
KINETIC ENERGY JOULES 
FIG4. HIGH -:TEMPERATURE,. COMPRESSION
 
15 
(9o,±45,0)5 GRAPHITE/EPoXY 
AVERAGE SURFACE TEMPERATURE 366±10K 
C/ CULT. IN-LBS 
0 4 8 12 16 20I .	 I I ­
t.. 
0.8- '0 
0.6 
A 
0.4­
0 RESIDUAL
 
o FAILURE 
0.2-	 A NON -FAILURE­
rn DEBONDING 
I I 	 I i ... 
0.0 0.4 0.8 1.2 - 1.6 -2.0 '..2.4 
KINETIC ENERGY .JOULES 
FIG 5. HIGH.. TEMPERATURE-, :COMPRESSION
 
16 
(9o3±45,0)5 GRAPHTE Epoxy
 
AVERAGE SURFACE TEMPERATURE 223 ± VK 
d/dULT. 	 IN-LBS
 
0 4 8 12 1 2-0
I I I I I 	 1­
1.0 	 ®-FAILURE
 
" NON -FAILURE
 
DEBOND!NG 
08 RESIDUAL 
_0-	 A A 
0.2-­
- t I I l I
 
0.0 0.4 0.8 1.2 	 1.6 2.0 2.4 
KINETIC ENERGY, JOULES 
FIG. Go, 'Low TEMPERA-TURE-, "TE.N.SION 
17 
223 
(9-, ±45 0)5 GRAPHITE/ Epoxy. 
16 KAVERAGE SURFACE TEMPERATURE 
/CULT: . IN -LBS 
0 4 8 12 16 20 
1. - . FAILURE 
A NON -FAILURE 
8 DEBONDING 
5-- 0 RESIDUAL 
.0.6-

A 
U4_; A 
0.0 0.4 0.8- 1:2 1.6 -2.0 _-2.4 
KINETIC ENERGY , JOULES 
FIG, 7. i..Low -T-EMPERATURE 'TENSION. 
18 
(9o,b±45,0) GRAPHITE/ Epoxy 
AVERAGE SURFACE TEMPERATURE 223 t I K 
;/CULT. . IN- LBS 
0 8 16 24' 32 40I1I 	 1 1' 
0.6- 03 
FRESIDUAL 
S FAILURE 
012- A NON -FAJLURE 
-KINETIC-• ENR6 ,-. OL 
--	 a. DEBONDING 
0.0 	- 0.8 11.6 2.4 "3.2 "4.0 .4.S 
A(.INETIC- ENERGY - , J4OULES 
FIG. 6. Low TEMPERATURE, COMPRESSION
 
19
 
(9o, ± 45, 0) 5 GRAPHITE/ Epoxy 
AVERAGE SUREACE TEMPERATURE 223 ± OK 
C,! ULT. IN ' LBS
 
08 16 24 32 40
 
It:o 
0.4T­
0K6 	 A 
A 
0RESIDUAL 
0 FAILURE 
102- A NON-FAILURE 
in DEBONDING, 
1 I 1 I 1
 
0.0 	 O. 1.6 2.4 3.2 .... 4.0 .. 
KINETIC ENERGY: ,- JOULES. 
FIG 93. Low TEMPERATURE)* CoMPRESSION 
.20 
(Or, 9o, 06 ) KEVLAR-GRAPHITE/ HYBRID 
AVERAGE SURFACE TEMPERATURE - 223 I OR 
U/aULT IN- LBS
 
0 8 16 24 32 40
 
1t0 -Q 9 FAILURE 
.A NON-FAILURE 
0 RESIDUAL 
0.8-­
0.6­
'00 
0--"A 	 A. 
0.2
 
.000. 1 	 1 I-­
0.0 	 0.8 1.6 2.4 3.2 4.0 4.8 
KINETIC ENERGY , JOULES 
FIG ±0. Low TEMPERATURE, TENSION
 
21 
(OK, 9C, Qt KEVLAR-GRAPHITE HYBRID 
AVERAGE SURFACE TEMPERATURE -223 ± I K 
C/CULT 	 IN - LBS 
0 8 16 24 32 40
 
1.0 	 ® FAILURE 
A 	NON-FAILURE 
RESIDUAL 
0. 8
 
0.6--
A 
0.4­
0.2-
I I I I ,I .
 
0.0, 0.8 1.6" 2.4 3.2 4.0' 4.8 
KINETIC ENERGY, JOULES' 
FIG I.- Low TEMPERATURE, TENSION
 
22 
(O, 90 Oc,)s KEVLAR.- GRAPHITE HYB RID 
AVERAGE SURFACE TEMPERATURE 223 t 10 K 
/ULT IN - LBS
 
0 8 16 24 32 40
1 	 I I 1.
 
1.00 
0 ­
0.8­
0 0SIUA 
A NFALURE 
0.6 	 0.8 1.6 2 32 -4.0 4.8
 
KINETIC ENERGY. JOULES
 
FIG.. Low TEMPERATURE, COMPRESSION 
23 
( OR, 90G, I jKEVLAR -GRAPHITE/ HYBRID 
AVERAGE SURFACE TEMPERATURE 22-3 t I*K 
C/CULT -IN- LBS 
0 8 16 24 32- 40 
0.8­
0-0 
.0

~0 
0.6-
A A 
.7 
0 RESIDUAL 
a FAILURE 
A NON-FAILURE 
I 1. I 1< ,. , 
0.0" .0.8 1.6 2.4 3.2 4.0 4.8 
KINETIC ENERGY, JOULES " 
FiG. ±3. Low TEMPERATURE, COMPRESSION 


Effect of temperature on composite sandwich structures subjected to low velocity impact

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