COMPARISON OF STATIC AND DYNAMIC TEST
METHODS FOR DETERMINING THE STIFFNESS
PROPERTIES OF GRAPHITE/EPOXY LAMINATES
by
MICHAEL DERRYCK TURNER
B.S., University of Washington(1977)
SUBMITTED IN PARTIAL FULFILLMENT
OF THE REQUIREMENTS FOR THE
DEGREE OF MASTER OF SCIENCE
at the
MASSACHUSETTS INSTITUTE OF TECHNOLOGY
June, 1979
( Massachusetts Institute of Technology
Signature of AuthorDepartment of Aeronautics and Astronautics
--April 19, 1979
Certified byA'Thesis Supervisor
Accepted by /Cha i rman, De pa mentaI -Graduete-Emi-ttee-
ARCHIVESMSSACHU HTS INSTITUTE
OTECHNOLOGY
JUL 3 1979
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4. TITLE AND SUBTITLE Comparison of Static and Dynamic Test Methods for Determining theStiffness Properties of Graphite/Epoxy Laminates
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14. ABSTRACT Various test methods for determining the in-plane stiffness properties of composite materials areconsidered. These tests include tensile coupon, sandwich beam, cantilever beam, and static and dynamicflexure tests. A series of tests are carried out to determine these properties for ASI/3501-6 Graphite/Epoxylaminates. The results indicate a significantly lower longitudinal modulus is found by static and dynamicflexure tests than by the other tests used. Computer software is developed to analyze the data quickly andaccurately and also produce graphical output that can be easily interpreted to evaluate each test. Resultsfrom several test methods are compared and differences analyzed.
15. SUBJECT TERMS
16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT Same as
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121
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Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std Z39-18
COMPARISON OF STATIC AND DYNAMIC TEST
METHODS FOR DETERMINING THE STIFFNESS
PROPERTIES OF GRAPHITE/EPOXY LAMINATES
by
Michael Derryck Turner
Submitted to the Department of
Aeronautics and Astronautics on April 19, 1979
in partial fulfillment of the requirements for
the degree of Master of Science.
ABSTRACT
Various test methods for determining the in-plane stiffnessproperties of composite materials are considered. These tests includetensile coupon, sandwich beam, cantilever beam, and static and dynamicflexure tests. A series of tests are carried out to determine theseproperties for ASI/3501-6 Graphite/Epoxy laminates. The results indicatea significantly lower longitudinal modulus is found by static and dynamicflexure tests than by the other tests used. Computer software is developedto analyze the data quickly and accurately and also produce graphical out-put that can be easily interpreted to evaluate each test. Results fromseveral test methods are compared and differences analyzed.
Thesis Supervisor: James W. MarTitle: Professor of Aeronautics
and Astronautics
ACKNOWLEDGEMENT
To all those who cooperated on this project the author would like
to express deep appreciation. First to Professor J. W. Mar for his
support and understanding in the area of composite materials. Also, to
Professor J. Dugundji for his advice and encouragement which was so
useful in carrying out the work. Thanks should go to Fred Merlis for
his assistance: particularly, for all the photographs that are in this
report. The author is grateful to Al Supple for his help in setting up
and running the test equipment. Acknowledgement should also be given
to Jerome Fanucci and Maria Bozzuto for their cooperation with the
author while writing and debugging computer programs. Thanks also, to
Ms. N. Ivey for typing the manuscript. Finally, the author wishes to
acknowledge all the other people who cooperated in this project.
This work was supported by the U.S. Air Force Materials Laboratory
under Contract No. F33615-77-C-5155: Dr. Stephen W. Tsai, technical
monitor.
TABLE OF CONTENTS
Page
Introduction
Test Specimens
Construction and Measurement ofTest Specimens
Test Equipment and Procedure
Theory and Data Analysis
Comparison of Test Results
Conclusions and Recommendations
References 39
Appendices
A Ritz Analysis of Effect of Beam Thickness 40
Taper on ist Bending Frequency
Summary and Tabulation of Test Results 44
Section
2
3
4
5
6
7
5
LIST OF TABLES
Cure Cycle for Graphite/Epoxy
Effect of Beam Taper on Ist Bending Frequency
Summary of In-Plane Stiffness Properties ofASI/3501-6 Graphite/Epoxy
Effect of Per Ply Thickness on the StiffnessProperties of 2, 4, and 8 Ply Laminates
Summary of
Summary
Summary
Summary
Summary
Summary
Summary
Summary
Summary
Summary
Summary
Summary
of
of
of
of
of
of
of
of
of
of
of
(00)2 Sandwich Beam Data
(00)4 Sandwich Beam Data
(0)4 Tensile Coupon Data
(00)8 Tensile Coupon Data
(04)8 Cantilever Beam Data
(900)4 Sandwich Beam Data
(900)8 Tensile Coupon Data
(904)8 Cantilever Beam Data
(+/- 450)s Sandwich Beam
(+1- 450)s Tensile Coupon Data
(+/- 450)s Tensile Coupon Data
(+/- 45*)s Cantilever Beam Data
Table No.
1
2
3
Page
13
42
44
LIST OF ILLUSTRATIONS
Figure No.
2
3
4
5
6
7
8
9
10
1I1
12
13
14-39
40-55
56-74
Sandwich Beam Construction
Beam and Laminate Measurement Location
Strain Gage Locations
Tensile Coupon Construction
Coupon Measurement Locations
Cantilever Beam Construction
Cantilever Beam Measurement Locations
Sandwich Beam 3 in Test Jig After Failure
Sandwich Beam Test Setup
Tensile Coupon Test Setup
Beam 5 Being Tested at a Load of 740 Pounds
Effect of Beam Taper on Ist Bending Frequency
Graph of EL vs. Fiber Volume, Vf
Graphical Program Output:
(0)2, (0)4, (0)8 Laminates
(900)4, (900)8 Laminates
(+/- 450)s, (+/- 454)2s Laminates
Page
1 3
I5
15
17
17
18
18
120
I 20
20
121
43
58
59-84
85-100
101 -1 19
SECTION I
INTRODUCTION
Graphite/epoxy and other advanced composite materials are seeing
increasing use in aerospace and some non-aerospace structures. An advan-
tage of these materials is that their elastic properties can be tailored
to give improved buckling strength, stiffness and aeroelastic properties
as well as reduced weight when compared to structures made with conven-
tional materials. In order to use this advantage effectively it is
necessary to accurately determine the basic stiffness properties of the
material.
This study considers the problem of testing the stiffness properties
of graphite/epoxy. A series of tests were carried out to determine these
properties. Some existing testing and analysis methods were
employed and new ones developed. This test program al lows the comparison
of different test methods and may help determine if certain test methods
are applicable to certain design problems.
In the process of making the test specimens, testing them, and
analyzing the data there were several additional objectives. First,
extra effort was put into automating the production process and producing
test specimens that were precisely made and accurately measured. Second,
a test method was developed that allowed speedy, accurate, and consis-
tent collection of data. Third and last, computer programs were develop-
ed to speed up the analysis of data and present it in a useful form.
8
Consequently, the results of these tests should be accurate , easily
reproducible, and provide a means of comparing stiffness properties from
different types of tests.
SECTION 2
TEST SPECIMENS
2.1 Types of Specimens Tested
There were three general types of specimens tested. First, four
point bending sandwich beam specimens for testing laminates in tension
and compression. Second, tensile coupons that are relatively easy to
build, test and provide additional data for comparison. Third, cantilever
beams that were tested with static tip loads and also dynamically tested
to determine natural frequencies and modulus.
2.2 Types of Laminates Tested
Three basic types of laminates were tested:
1. (00)N where N = 2, 4, or 8
2. (904)N where N =4 or 8
3. (+/- 450)NS where N = I or 2: (+/- 450, +/- 45)
The (00)N laminates were used to determine longitudinal modulus, EL on
all three types of specimens as well as major Poisson's Ratio, V LT from
sandwich beam and coupon tests. Similarly, (904)N laminates determined
transverse modulus ET and minor Poisson's Ratio, VTL for sandwich beam
and coupon tests. Lastly, (+/- 450)NS laminates were tested to deter-
mine shear stress-strain behavior and shear modulus, G.
A total of 70 laminates were tested. Twenty-six laminates were
tested in 13 sandwich beams. Thirty-two laminates were tested as
tensile coupons. Twelve laminates were tested as cantilever beams.
2.3 Sandwich Beams
At first, sandwich beam specimens were tested rather than tensile
coupons for several reasons. The relatively thin laminates (2 to 4 plies)
tested were easier to handle and less susceptible to damage when bonded
onto a core material. Also, beam specimens could be tested easily at
low stress levels. When testing was began this was not true for tensile
coupons because hydraulic grips were not readily available for holding
and testing tensile coupons. The grips that were then available tended
to slip with only small loads at low stress levels. Most importantly
thought, sandwich beams allowed the testing of each laminate in tension
and compression without elaborate testing jigs.
2.4 Tensile Coupons
Sandwich beam tests with 2 and 4 ply laminates indicated very little
difference between tensile and compressive stiffness properties. Also,
a new testing machine with hydraulic grips suitable for testing tensile
coupons was purchased and installed. Consequently, a series of tests
were performed using tensile coupons made from 8 ply and some 4 ply
laminates. The 8 ply laminates had a lower per ply thickness and thus
they made it possible to test the stiffness properties of material with
a lower fraction of epoxy matrix, and a higher fiber volume. The 4 ply
laminates allowed the comparison of data with earlier beam tests.
Tensile coupon tests have some significant advantages. The test
specimens are easy to construct accurately. Also, unlike sandwich beams
their is no core material which may affect laminate properties particu-
larly Poisson's Ratio.
2.5 Cantilever Beam Specimens
Cantilever beam specimens are used to determine stiffness properties
under static tip loads, and dynamically from the determination of
natural frequencies.
In the cantilever beam test, the strain is linearly distributed
through the thickness such that the strains on the top and bottom are
approximately equal in magnitude and opposite in direction. This is
considerably different from sandwich beam or coupon tests where there is
little or no variation in strain through the laminate. The strain
distribution found in cantilever beams may be similiar to that found in
many aerospace structures including, vibration of fan blades or the
buckling of shel I structures. Therefore, it will be worthwhile to
compare results from cantilever beam tests to other test methods.
SECTION 3
CONSTRUCTION AND MEASUREMENT OF TEST SPECIMENS
3.1 Construction of Laminates
Al I laminates were made from 12 inch wide 3501/ASI-6 pre-preg tape.
Layups for each type of laminate were made by using sheet aluminum
templates to cut out pieces to the correct size, shape, and fiber orien-
tation. These pieces were stacked up to produce the desired sequence
and orientation of plies. Each layup was then placed between aluminum
plates with peel ply, porous teflon, correct number of fiberglass
bleeders, and non-porous teflon on each side of the layup. The laminate
was then cured in a hot press according to the cure cycle shown in
Table I.
After curing, laminates were cut from each layup using a table saw
with a diamond coated, water cooled saw blade.
3.2 Sandwich Beam Construction
The beam cores are constructed of styrofoam and mahogany as shown
in Figure I. The mahogany was cut roughly to size (2 x 7.5 x 13 cm) in
a table saw. The styrofoam was cut roughly to size (2.5 x 7.5 x 13.5 cm)
with a hot wire. The mahogany and styrofoam were glued together with
Titebond glue and allowed to set overnight. The beams were sanded down
until they were flat in a milling machine with the milling head replaced
by a sanding disk.
13
TABLE I: Cure Cycle
rMAHOGANY
STYROFOAM
TOP LAMINATE
o oo o C 0
LOWER LAMINATE
130 ----- +- - 135 S.130
DIMENSIONS: mm
FIG. I: SANDWICH BEAM CONSTRUCTION
TEMP PRESSURE TIME(*F) (PSI) (MINUTES)
275 15 18
RAISE TO 15 5300
300 100 30
RAISE TO 100 7350
350 100 35
SMAHOGANY
/ o
Before the beams were bonded together thickness measurements were
taken on the cores and the laminates at the 18 locations shown in Fig. 2.
The laminates were bonded to the cores using Smooth-on EA-40 Epoxy
adhesive. This bonding process was carried out on a jig constructed from
aluminum and placed inside a vacuum bag during the bonding process.
The jig and vacuum bag assured that the laminates were kept flat and
correctly aligned; also importantly, the adhesive was squeezed out so
that only a thin layer remained.
After the beams were removed from the vacuum bag the edges were
sanded down to remove excess dried epoxy. The beam thicknesses were
then measured at the same 18 locations as before and widths were measured
at the 6 locations shown in Fig. 2.
Four strain gages were glued onto each beam. The strain gages used
were Micro-Measurements type EA-09-125AD-120 or type EA-06-125AD-120.
Each laminate had two strain gages glued on to give longitudinal strain
and transverse strain as shown in Fig. 3.
After the strain gages were glued on and wires soldered on, the
beams were ready to be tested.
3.3 Construction of Tensile Coupons
Tensile coupons consisted of a test laminate and loading tabs as
indicated in Fig. 4. The gage length was 275 mm for the (+/- 45 *)NS
laminates and 200 mm for other laminates.
I I
Width Measurements Top andBottom
I I
T--+-+--F ±H---I---t----i-
Dimensions: centimeters
FIG. 2: SANDWICH BEAM AND LAMINATE MEASUREMENT LOCATIONS
Longitudinal Gage
/-Transverse Gage
FIG. 3: STRAIN GAGE LOCATIONS
I
Test laminates were cut as described previously. Then sanded to a
constant width. After which width measurements were taken at five
locations and thickness measurements at ten as indicated in Fig. 5.
The loading tabs were cut from (0*, 904)2S sheets of 3M Scotchply,
fiberglass/epoxy. These were cured in the same way as graphite/epoxy
except that the cure cycle consisted of 40 minutes at 50 psi and 330*F
fol lowed by a gradual cool down to room temperature.
The loading tabs were bonded onto the test laminate with Cyanamid
FM123 film adhesive cured at 2404F, 40 psi for 90 minutes.
Longitudinal and transverse strain gages of the same types used on
beam specimens were then attached to the test coupons. They were centered
at the equivalent locations to those shown for beam specimens in Fig. 3.
3.4 Fabrication of Cantilever Beam Specimens
To make the cantilever beam specimens the cured graphite/epoxy was
cut as before and sanded carefully to be straight and square. After
which the mass of each laminate was measured. Measurements of thickness
were taken at 12 locations and width was measured at 6 locations as
indicated in Fig. 7. The next step was to bond onto the base a 25 mm x
25 mm loading tab machined from 1/8" aluminum as indicated in Fig. 6.
Finally, a strain gage was bonded on each laminate 5 mm from the loading
tab.
Fiberglass Loading Tabs
Film Adhesive
/ G/E Laminate
1 7 7 Z=
7.5- - -- Gage Length - 7.5
Dimensions: centimeters
FIG. 4: TENSILE COUPON CONSTRUCTION
Width Measurements
2
S4
-2 -
3
Dimensions: centimeters
FIG. 5: COUPON MEASUREMENT LOCATIONS
L7IIZZZI
'7±F7++
+ ± + + +
r 7 7 7
4 --- =
Aluminum Loading Tabs
Film Adhesive
G/E Laminate
25
- -1- 7 5 - - - - - - - - -----
Dimensions:
FIG. 6: CANTILEVER BEAM CONSTRUCTION
2 3 4 5 6
25
S- 3+ -1-
> 2. 5 K 30 30 30 WE 30- 30 12.5<
Dimensions: mm
FIG. 7: CANTILEVER BEAM MEASUREMENT LOCATIONS
SECTION 4
TEST EQUIPMENT AND PROCEDURE
4.1 Sandwich Beam Tests
The sandwich beams were tested in a four point bending test jig
made from aluminum I-beams. The test jig transfered the load from the
Baldwin-Emery SR-4 test machine to the test specimen through four
cylindrical rollers. The strain gages were attached to four BLH-1200
strain indicators. Figure 8 shows the aluminum test jig in the test
machine with beam 3 after failure.
One person ran the test machine and called out the load every 10 or
every 20 pounds and four volunteers wrote down the strain readings:
Fig. 9.
Each beam was first tested upside down in the test jig up to a load
between 100 and 200 pounds depending on the type of laminate. Then the
beam was removed and tested right side up until it reached failure load.
This prodedure was followed so that data could be collected for each
laminate both in tension and compression.
4.2 Tensile Coupon Tests
The tensile coupons were tested in a 100,000 pound MTS testing
machine using hydraulic grips: Fig. 10. Strain indicators were used
as before.
The test procedure was similar to that used for sandwich beams:
one person ran the test machine and called out the load every few
hundred pounds and two others wrote down the strain readings. However,
the coupons were only tested in tension. As before, each specimen was
tested to failure.
FIG. 10: TENSILE/COUPON TEST SETUP
4.3 Cantilever Beam Experiments
The cantilever beams were first tested with static tip loads and
then tested dynamically with a shaker to find natural frequencies.
The static test was performed by first, clamping the test specimen
onto a 12" x 12" x 3" base of aluminum. Then a Kevlar thread was taped
on and draped over the center of the end of the specimen. Three
different weights were hung from the thread and the tip displacment was
measured for no load and then'the three weights individually. An Edmund
direct measuring microscope, NO. 70,266, was used to measure these dis-
placements accurately.
After each beam was tested statically it was tested to find the
frequencies of the first 3 natural modes of vibration. This was done
by clamping the specimen in an aluminum block attached to a Ling Model
420 shaker.5 An Endevco 7701-50 "Isoshear" accelerometer was mounted
to the aluminum block. This accelerometer and the specimen strain gage
were used to produce a signal that was amplified and displayed on an
osci l loscope.
By monitoring these signals resonances could be determined by
maximum signal amplitude and most clearly from a 90* phase shift.
SECTION 5
THEORY AND DATA ANALYSIS
5.1 Sandwich Beams With (0) 2(00) 4 and (900)4 Laminates
The sandwich beam data is analyzed by a computer program on an
IBM 370. The beam, laminate dimensions, and load vs. strain data are
input into the program. The program converts the load into metric units
and then calculates moments. The program finds the two best straight
lines through the moment vs. longitudinal strain data by linear regres-
sion.
SLOPE I
SLOPE 2
IL' S2L
M = Moment
e IL = Upper Laminate Longitudinal
62L = Lower Laminate Longitudinal
Strain
Strain
The location of the neutral axis is calculated:
ZNA t + SLOPE ISLOPE 2
The moment of inertia for each laminate about the neutral axis is calcu-
lated:
I A [t 2/l2 + (Z - Z 21 ~ 1 I 1 ZNA
I = A [t /12 + (Z - Z ) 22 1 2 2 NA
Where
II = Moment of Inertia
I2 = Moment of
for the Upper Laminate
Inertia for the Lower Laminate
t
t2
Beam Cross Section
A, =W It
A2 W2 2
W = Width of Upper Laminate
W 2 Width of Lower Laminate
Moment and force equilibrium yield the formulas used to determine
Young's Moduli, E1 , E2
A (SLOPE )Y
2 Z NA Z2A2 2NZz2 + I2A
2 Z 1 NA,
E = [(SLOPE )Y - E2 Y2 ]/I
The stresses are
E IY IM
I E I + E2 2
E2 22TE2 I 2+ E 2
Poisson's Ratio's are
IT
IL
2TV2 c2L
The program plots stress vs. strain, Poisson's Ratio vs. strain,
and the best straight line through the stress-strain data for each
laminate in tension and compression.
The graphical results are Figs. 14 to 23 and Figs. 40 to 47.
5.2 Sandwich Beams with (+/- 45) Laminates
The sandwich beams with (+/- 45) laminates are used to determine
the shear stress-strain behavior and shear modulus, G. Using a (+/- 45)
laminate in a uniaxial stress state to determine shear properties was
proposed by Petit and others. 2,34 Testing the laminates on beams made
it possible to further check the validity of the test by seeing if it
worked equally well for a laminate in tension and compression.
The test data was again analyzed with a computer program on an IBM
370. This program calculates the longitudinal stresses alL' a2L in the
same way as the previous program. Rotating the axis 450 gives the shear
stresses:
T 2 IL
2 2 2L
and the shear strains
= 'IL IT
Y2 =2L - 2T
The program performs a linear regression analysis on the shear
stress-strain data to calculate the shear moduli:
dT
I dy
dT2
G2 dy2
The shear stress-strain behavior becomes nonlinear above a strain
of 3000 microstrain. Therefore, only data points with a shear strain of
less than 3000 microstrain are used in the linear regression analysis.
The program plots shear stress vs. strain and the best straight
line through the data for each (+/- 45) laminate in tension and
compression.
The graphical results are Figs. 56 to 63.
5.3 Analysis of Tensile Coupons
Computer programs are also used to analyze data from tensile coupons.
Longitudinal stresses are just calculated on the basis of load divided
by cross sectional area. Longitudinal and transverse strain data having
been read during the test.
The modulus of the (0*)N and (900)N laminates is calculated directly
by doing a linear regression analysis on the stress-strain data.
The program plots stress vs. strain, Poisson's Ratio vs. strain, and the
best straight line through the stress-strain data for each laminate.
The graphical results are Figs. 23 to 36 and Figs. 45 to 52.
For the (+/- 45)NS laminates the shear stress is half the longitu-
dinal stress and the shear strain is the difference between the longitu-
dinal and transverse strain, The shear modulus, G is determined by per-
forming a linear regression on the shear stress-strain data for shear
strain of less than 3000 microstrain. Shear stress vs. strain is plotted
along with the best straight line through the data for each of the
(+/- 45)NS laminate. The graphical output is given in Figs. 64 to 74.
5.4 Analysis of Cantilever Beam Experiments
In the static tip load
simple beam theory:
test the modulus can be determined from
L3q = Q
differentiating
dq L3
dQ 3EI
Then
E = d 3
where
= length of
= moment of
beam: from tab to tip
inertia: assumed constant along the beam
= the slope of tip load vs. displacement
1 being determined from a linear regression analysis done ondq
tip load vs. displacment data.
For the (0*)8 laminates E is the longitudinal modulus, EL and for
the (900)8 laminates E is the transverse modulus, ET. For the (+/- 45*)2S
laminates E is some effective longitudinal modulus, the significance of
which will be discussed later.
The cantilever beams are also tested dynamically to determine the
lowest 3 resonances. From beam theory the first 3 natural frequencies
of a clamped-free beam are
I = (1.8751041)2 E
mL
2 = (4.6940911)2 EImL
3 = (7.8547574)2 /mL
Employing these formulas the modulus can be determined from the
frequencies. As with the static tests E and ET are found from the (00)8
and (90*)8 laminates.
Now to consider how to effectively analyze cantilever beams made
with (+/- 450)2S laminates. One difficulty with analyzing these
laminates is that they exhibit bending-twisting coupling. That is to
say, if one considers one of these laminates as plate, the bending stiff-
ness terms D 112, D2212 / 0. If simple beam theory is to be applied it
is necessary to neglect these terms.
Therefore, with this approximation a straightforward analysis can
be performed assuming the only stress acting is a stress along the axis
of the beam:
_ Mz(YI :- I
G22
For each +45* ply
[45]
[45]
[45]
2E1 1 12[ 45]
[45]
[45]
With the approximation of no twist, E 12 = 0, this becomes
E [45]
E221 1[45]
E 2[45]] r1 122 1Il
E 2222[~45] 22
This last relationship also holds for the -45* plies. Inverting the
above relationship and using &22 = 0:
[45]- CFI
Therefore, the effective modulus determined from tip loads and beam
natural frequencies is
E[ ].E45]-1
[45] [45]
E 2222 E45]
In terms of the orthotropic properties for a 04 ply
E [E45] = IE * +i-E * + E *42222 21122 11
a,
'12
E 1111 45 E 1122
E2211 E2222
L 1211 E1222
La2
22
E 2 [ 45]- E [~12 45]
+ 1212
E [45] _ E * + E * + E * -E *11224 4 2222 2 1122 1212
Also,
E t45] = E2222 45]
E 1122 45] = E221 [45]
For convenience define
A = El* + E22 + 2E*24 111 4 2222 2 1122
and note that
E * G1212
Then
E =4 GG4G
or
G=
Putting in previously determined properties to calculate A and G shows
that changing the value of A by 20% only changes G by 2.5%: showing the
results for G determined by this method are not overly sensitive to
values assumed for other stiffness properties. Therefore, the values of
G are calculated from cantilever beam tests with (+/- 450)2S laminates.
SECTION 6
COMPARISON OF TEST RESULTS
6.1 Difference in Test Methods
Of the three different test methods employed: sandwich beams,
tensile coupons, and cantilever beams, the first two methods are
basically similar and give comparable results. Therefore, results from
these two methods will be considered together. Then the results from
the cantilever beam tests will be analyzed and compared to the other
test methods.
Al I test results are summarized and tabulated in Appendix B for
easy comparison.
6.2 Stiffness Properties Determined from Sandwich Beams &Tensile Coupons
When looking at the test results it is worthwhile to try and deter-
mine what parameters seem to affect the stiffness properties. One would
expect the fraction of material that is graphite fibers, the fiber
volume, VF should be one of the parameters.
Test results clearly show the effect of fiber volume, V F on material
properties. Laminates of only a few plies have a greater per ply thick-
ness and consequently a lower fiber volume. The effect of this on
material properties is apparent in Table 4 where the average stiffness
properties are summarized for 2, 4, and 8 ply laminates.
The dependence of E on fiber volume is shown by Fig. 13. ThisL
includes data from (ON I, 2 and 4 ply, laminates tested on sandwich
beams. This graph shows similar results for E in tension and com-
pression. Also, a linear regression through this data has a near zero
intercept indicating that it is possible to approximate the affect of
VF on EL by neglecting the stiffness of the epoxy matrix.
Table 4 indicates that the transverse modulus, ET goes down very
slightly with fiber volume. This i
just be an anomaly from a small amo
results for the (90*)8 laminates on
cut from one sheet have a 15% lower
However, the slope of the load vs.
indicate there is
(900)8 coupons. T
indicate some loca
significant proble
likely that ET is
On the other
direction expected
fiber volume becau
than the epoxy mat
fiber volume as ex
behavior is shown
s certainly not expected and could
test data.
II it appea
n those cut
Looking at the
rs that laminates
from another sheet.
stroke graphs made during each test
less than 3% variation in the overall stiffness of all
herefore, it is felt that the variation in ET must
I soft or hard spots in the material. This may be a
m in testing (90*)N laminates. Consequently, it is
not reduced for high fiber volume G/E.
hand the properties v LT and G vary with VF in the
. The major Poisson's Ratio goes down with increasing
se the graphite fibers have a lower Poisson's Ratio
rix. Similarly, G increases slightly with increasing
pected. The nonlinear nature of the shear stress-strain
in Figs. 56 to 74. Figure II is a photograph of beam 5
being tested. It shows the large deflection and distinct anticlastic bend-
ing caused by the large longitudinal and transverse strains of (+/- 450)s
specimens. For the laminates tested on sandwich beams, the measured value
of 'VTL is about half that needed to satisfy the relation ELVTL = ETVLT
which should hold for ideally orthotropic laminates. However, for the
8 ply laminates tested on coupons the average properties agree with this
relationship within 10%. Perhaps testing the laminates on sandwich beams
restricts the transverse strains and affects the measured Poisson's
Ratios. Also, there is a certain amount of inaccuracy in the measurement
of VTL for small stresses. This is the result of a small amount of drift
in the strain readings due to temperature variation. This has its
greatest effect on the smal I strain readings of the transversely mounted
gage (parallel to the fibers). Looking at the plots of vTL, Figs. 37 to
52 and comparing the results from sandwich beams to those from tensile
coupons, it is noticeable that, temperature drift was not a problem for
the tensile coupons. Consequently, the coupon data for vTL is probably
more reliable than that from sandwich beams and the coupon data agrees
closely with the previous relatinship indicating that vTL can be deter-
mined from the other stiffness properties.
Taking into account the variation of stiffness properties with fiber
volume, values for these properties are calculated for the manufacture's
specified per ply thickness. These values are included in Table 3.
6.3 Stiffness Properties from Cantilever Beam Tests
In looking at the stiffness properties determined from cantilever
beam tests there are several important considerations. First, are the
test results consistent and are there exp-lanations for any variation.
Second, how do stiffness properties determined statically and from the
first three bending frequencies compare. Third, how do the cantilever
34
beam results compare with results from the other test methods and
particularly with tensile coupons cut from the same sheets of cured G/E.
To address the first consideration, look at Tables 9, 12, and 16.
It is clear that, there is little variation in the ET determined for the
4 (900 )8 laminates. However, for the (0*)8 and (+/- 45*)2S laminates
both have a test specimen that appears to have signific
ness properties than the other laminates. In the first
where this difference is most pronounced the (0) 8- 2
modulus 12% lower than the other 3 (00)8 cantilever bea
(+/- 454)2S - 2 - D specimen has a modulus 16% lower th
(+/- 450)2S laminates. The measurements of these 2 spe
they are thicker at the tip than the root and the 6 rem
(+/- 45 0)2S specimens are thicker at the root than the
possible to approximate this thickness variation as str
root to tip. A taper involving a difference between ro
ness of about 4% for the (04)8 specimens and as much as
(+/- 454)2S specimens. A Ritz analysis is performed in
effect of beam taper on first bending frequency. This
that to get the 12% and 16% difference in moduli found
(+/- 450)2S laminates would require a thickness taper of
pectively. This compares fairly well with the 4% and 8%
ness variation. Consequently, the variation in measured
easily explained.
antly lower stiff-
bending frequency
- B specimen has a
ns and the
an the other
cimens indicate
aining (0)8 and
tip. It is
aight taper from
ot and tip thick-
8% for the
Appendix A on the
analysis indicates
in the (0)8 and
5% and 6% res-
measured thick-
stiffness is
35
The second consideration is how the properties determined from the
first 3 bending frequencies and from static tests compare for the canti-
lever beam specimens. The difference between static and dynamic modulus
is only 2 to 3% for the (00)8, (900)8 specimens, and as much as 6% for
the (+/- 450)2S specimens. A 2 to 3% difference is insignificant and the
small 6% difference for the (+/- 450)2S could easily be caused by the
bending-twisting-coupling they exhibit, or the variation in thickness.
The difference in modulus determined from each of the 3 bending modes is
insignificant when the moduli are averaged for the 4 specimens of each
type. Something that is noticeable is that the moduli data is less
scattered for the higher natural frequencies of the (0*)8 and (+/- 45*)2S
which could be because the thickness variation of these laminates has
less effect on the frequencies of the higher modes. Consequently, there
are no significant differences between the moduli determined from the 3
lowest natural frequencies and the static tests of the cantilever beam
specimens.
The third and most interesting consideration is how the cantilever
beam test results compare to moduli from the other test methods. Table
3 provides a summary of moduli from the cantilever beam tests compared
to the results from the other test methods and design stiffness proper-
ties used by Grumman. The shear modulus, G is not significantly differ-
ent from that determined from the other tests. The transverse modulus,
ET is somewhat lower. However, the longitudinal modulus, E is some 30%
lower than that found in other test methods.
36
The difference in measured E is a direct result of the test method
rather than a difference in material properties between the cantilever
beam specimens and other test specimens. Consult Table 8, the (0)8 - 2
laminates cut from the same sheet of cured G/E as the (04)8 cantilever
beams and tested as tensile coupons have a much higher E than found in
the cantilever beam tests. As a final confirmation the (0*)8 cantilever
beam specimens were made into tensile coupons by cutting off the aluminum
tabs and bonding on 25 cm x 25 cm fiberglass loading tabs. The specimens
were strain gaged and tested like other tensile coupons. The results
are included in Table 9 and the test data is Figs. 37 to 39. These re-
sults agree with the other tensile coupon data. This indicates that the
same material tested with different methods exhibits a different modulus
E L.
In summary, it appears that cantilever beam specimens give consis-
tent results from the beam natural frequencies and static tip loads but
E is significantly lower than that found by other test
methods. The cantilever beams are sufficiently long and thin that
transverse shear will have little effect on test results. Therefore,
it would appear that the stiffness may vary through the thickness perhaps
due to the distribution of fibers.
SECTION 7
CONCLUSIONS AND RECOMMENDATIONS
The test results and analysis in this report make it possible to
draw some significant conclusions about the stiffness properties of
Graphite/Epoxy and the test and analysis methods used to determine those
properties for composite materials. First, useable stiffness properties
and some variables that may affect those properties have been determined.
Second, the effectiveness of the test and analysis methods has been con-
firmed but some important differences have been found in stiffness pro-
perties from tests that involve laminate bending or flexure.
Considering the stiffness properties first, Tables 3 and 4 give a
good summary of stiffness properties that can be expected from ASI/3501-6
G/E used at M.I.T. One important conclusion is that these properties are
the same in tension and compression. Also, the per ply thickness or
fiber volume has some effect on all the stiffness properties. The longi-
tudinal modulus is most sensitive: the quantity of fiber being the most
important item in determining this property.
From comparison of the test and analysis methods several conclusions
can be drawn. First, tests using coupon specimens are easier to perform
than those using sandwich beams but they both give similar results.
Also, cantilever beam tests indicate that laminates exhibit different
material properties in bending.
The results in this report indicate some techniques that may be
useful in the future and some areas that warrant further investigation.
The use of several types of laminates such as (0*) N (90*)N, and
(+/- 450)NS laminates to determine basic material properties can be use-
ful in finding other characteristics of composite materials. This
approach could be applicable to finding strength characteristics, damping
properties, and fatigue damage. One area that warrants further investi-
gation is the determination of stiffness properties in flexure: particu-
larly, E . Making (0*)N laminates with different numbers of plies and
then testing them as 4 point bending flexure specimens could provide in-
sight into why EL is apparently lower in bending.
In conclusion, this work has determined stiffness properties,
compared test methods, and also in indicated where more research could
be worthwhile.
REFERENCES
1. Petit, P. H., "A Simplified Method of Determining the In Plane ShearStress-Strain Response of Unidirectional Composites," CompositeMaterials: Testing and Design, ASTM STP460, American Society forTesting and Materials, 1969.
2. Rosen, B. W., "A Simple Procedure for Experimental Determination ofthe Longitudinal Shear Modulus of Unidirectional Composites," J.Composite Materials, Vol. 6 (1972), pp. 552-554.
3. Sims, D. F., "In-Plane Shear-Strain Response of Unidirectional Com-posite Materials," J. Composite Materials, Vol. 7 (1972), pp. 124-126.
4. Hahn, H. T., "A Note on Determination of the Shear Stress-StrainResponse of Unidirectional Composites," J. Composite Materials,Vol. 7 (1973), pp. 383-386.
5. Crawley, E. F., "The Natural Mode Shapes and Frequencies of Graphite/Epoxy Cantilevered Plates and Shells," S.M. Thesis, June 1978, M.I.T.
6. Young, D., and Felgar, R. P., "Table of Characteristic Functions Re-
presenting the Normal Modes of Vibration of a Beam," EngineeringResearch Series, No. 44, July 1, 1949, University of Texas, Austin,
Texas.
APPENDIX A
RITZ ANALYSIS OF EFFECT OF BEAM THICKNESS
TAPER ON FIRST BENDING FREQUENCY
A Ritz analysis is performed using the first mode shape for a uni-
form cantilever beam. This analysis will yield a good approximation of
the frequency of the first mode for a slightly tapered beam.
For the harmonic transverse vibration of a beam the displacement is
of the form
w(xt) = $(x)e
The maximum potential and kinetic energy are
V = fL EI(x)(") 2dxf0
T = I 2 fL m(x) 2dx0
For convenience a new variable is introduced:
x = - I
The beam thickness of a uniformly tapered beam is
h( = + (h - hR T2 TIP ROOT
Where h is the average thickness. If m and EI are the mass distribution
and stiffness for a uniform beam of thickness h, then for the tapered beam
41
m h
M h
EI (h 3
f+1(h 3 (L 2 ,,)2 d2 ET -l h
m~L- +l I(h 42 d-- h
The function used for $ is the first bending mode for a uniform beam:
= os - - a (sin x - sin )L LL L
Values of a and 8 are in Ref. 6 along with tables of $(x) and $"(x).
However, # and $" can easily be calculated on a programmable calculator.
The expression for w2 is evaluated for 5 cases: a uniform beam and
tapered beams with the tip thickness 4% less, 8% less, 4% greater, and
8% greater than the root thickness. The necessary integrals were eval-
uated numerically using Gauss quadrature on 6 points. In the case of the
uniform beam the integrals are equal to the exact result up to the sixth
decimal place.
The results of this analysis are presented in Table 2 and plotted
in Fig. 12.
TABLE 2: EFFECT OF BEAM TAPER ON FIRST BENDING FREQUENCY
2 2 -2h -ROO h TIPROOT TIP -2
L4 1
.08 13.63453 .10291
.04 12.9833 .05023
0 12.362364 0
-.04 11.7703 -.04790
-.08 11.2058 -.09356
h = Average beam thickness
m = Mass distribution for uniform beam of thickness h
EI = Bending stiffness for uniform beam of thickness h
w = First bending frequency for uniform beam of thickness h
2
1,12
I .O8-
1 , 06
t 04
1.02
-10 -8 -6
0.9
0.90
2 4 6 8 10% ROOT TH ICKERTHAN T I P
(ROOT TI P
a h
FIG. 12: EFFECT OF BEAM TAPER ON FIRST BENDING FREQUENCY
APPENDIX B
TABLE 3: SUMMARY OF IN-PLANE STIFFNESS PROPERTIES OF
AS1/3501-6 GRAPHITE/EPOXY
VALUE USED FROM SANDWICH 8 PLY LAMINATESPROPERTY BY GRUMMAN BEAM AND* IN FLEXUREt
COUPON DATA
E 128 GPa 134 GPa 98 GPaL (18.5 msi) (19.4 msi) (14.2 msi)
11.0 GPa 10.0 GPa 7.9 GPaET (1.60 msi) (1.45 msi) (1.15 msi)
VLT .25 .28 ---
4.5 GPa 5.7 GPa 5.6 GPaG (.65 msi) (.83 msi) (.81 msi)
msi = 106 psi
*
Values estimated for manufacture's per ply thickness = .13335 mm.
tBased on cantilever beam tests with per ply thickness = .130 mm.
TABLE 4: EFFECT OF PER PLY THICKNESS ON THE STIFFNESS
PROPERTIES OF 2, 4, AND 8 PLY LAMINATES
BASED ON SANDWICH BEAM AND TENSILE COUPON TESTS
2 PLY LAMINATE 4 PLY LAMINATE 8 PLY LAMINATEMEASURED PER MEASURED PER MEASURED PER
PROPERTY PLY THICKNESS PLY THICKNESS PLY THICKNESS= .169 mm = .146 mm = .130 mm
E L(GPa) 104 125 142
E T(GPa) --- 10.6 9.4
V LT .33 .29 .27
G(GPa) --- 5.5 6.0
TABLE 5: SUMMARY OF (0*)2 SANDWICH BEAM DATA
AVERAGE E L(GPa)RUN BEAM LAMINATE LAMINATE VLT
THICKNESS TENSION COMPRESSION(mm)
1 3 (0)2-2-3 .341 103.255 100.477 .325
1 3 (0)2-2-2 .338 102.105 104.564 .338
8 7 (0) 2-2-4 .332 103.469 104.835 .325
8 7 (0)2-2-1 .341 100.702 101.340 .325
7 10 (0)2- 1-2 .348 106.754 111.468 .350
7 10 (0) 2-1-3 .341 104.476 106.999 .338
5 II (0) 2-1-4 .327 99.876 102.853 .338
5 II (0)2-1-1 .332 99.411 104.329 .325
Average EL Tension = 102.505 GPa (14.876 msi)
Standard Deviation = 2.489 GPa (2.4%)
Average E Compression = 104.608 GPa (15.172 msi)
Standard Deviation = 3.459 GPa (3.3%)
Average of E Tension & E Compression = 103.557 GPa (15.020 msi)
Standard Deviation = 3.107 GPa (3.0%)
Average v LT = .333
Standard Deviation = .009 (2.7%)
Average Thickness = .338 mm
Standard Deviation = .007 mm (2.1%)
TABLE 6: SUMMARY OF (00) SANDWICH BEAM DATA
AVERAGE E (GPa)RUN BEAM LAMINATE LAMINATE L V LT
THICKNESS TENSION COMPRESSION(mm)
12 16 (0) -1-4 .575 127.213 120.205 .313
12 16 (0) -I-1 .571 118.174 114.231 .318
Average E Tension = 122.694 GPa (17.795 msi)
Standard Deviation = 6.392 GPa (5.2%)
Average EL Compression = 117.218 GPa (17.001 msi)
Standard Deviation = 4.224 GPa (3.6%)
Average of EL Tension & E Compression = 119.956 GPa (17.398 msi)
Standard Deviation = 5.437 GPa (4.5%)
Average V LT = .315
Average Thickness = .573 mm
TABLE 7: SUMMARY OF (0*) TENSILE COUPON DATA
Average EL = 126.830 GPa (18.395 msi)
Standard Deviation = 4.263 GPa (3.4%)
Average VLT = 0.284
Standard Deviation = .024 (8.5%)
Average Thickness = 0.567 mm
Standard Deviation = .015 mm (2.6%)
AVERAGE
RUN LAMINATE LAMINATE E (GPa) VLTTHICKNESS L L
(mm)
II (0) 4-2-I .572 129.224 .287
12 (0) -2-2 .584 130.768 .264
13 (0) -2-3 .549 121.093 .317
I4 (0)4-2-4 .564 126.234 .268
TABLE 8: SUMMARY OF (0*)8 TENSILE COUPON DATA
RUN LAMINATE AVG. THICKNESS E (GPa) v(mm) L LT
I (0) 8- 1-1 1.053 134.183 .272
4 (0) 8-1-2 1.080 141.784 .281
5 (0)8-1-3 1.055 142.091 .259
3 (0)8-I-4 1.034 140.708 .280
6 (0) 8- 1-5 1.000 142.165 .257
7 (0) 8-2-I 1.020 144.325 .292
8 (0) 8 2-2 1.069 142.679 .297
9 (0) 8-2-3 1.070 144.430 .270
10 (0) -2-4 1.038 145.265 .254
Average EL = 141.959
Standard Deviation =
Average x)LT = 0.274
Standard Deviation =
GPa (20.589 GPa)
3.265 GPa (2.3%)
.015 (5.6%)
Average Thickness = 1.047 mm
Standard Deviation = .026 mm (2.5%)
TABLE 9: SUMMARY OF (00)8 CANTILEVER BEAM DATA
E (EGPa) v LT
BEAM THICKNESS STATIC Ist MODE 2nd MODE 3rd MODE STATIC STATIC(mm) (TIP LOAD) (COUPON) (COUPON)
(0) 8-2-A 1.029 100.599 101.433 98.636 98. 129 142.236 .304
(0) 8-2-B 1.055 91.696 89.252 92.283 93.176 138.330 .310
(0) 8 2-C 1.044 99.716 101.060 97.669 96.627 142.253 .296
(0)8-2-D 1.031 101.996 102.447 99.900 98.941 ------- ----
AVERAGE 1.040 98.502 98.548 97.122 96.718 140.940 .303
STD.DEV. .012 4.633 6.225 3.353 2.549 2.260 .007(1.2%) (4.7%) (6.3%) (3.5%) (2.6%) (1.6%) (2.3%)
TABLE 10: SUMMARY OF (904)4 SANDWICH BEAM DATA
Average ET Tension = 10.4
Standard Deviation = .47
Average ET Compression =
Standard Deviation = .441
Average of ET Tension and
Standard Deviation = .465
59 GPa (1.517 msi)
7 GPa (4.3%)
10.807 GPa (1.567 msi)
GPa (4.1%)
Compression = 10.633 GPa (1.542 msi)
GPa (4.4%)
Poissons Ratio .016 for all Laminates Tension and Compression
Average Thickness = .584 mm
Standard Deviation = .006 mm (1.0%)
AVERAGE E (GPa)RUN BEAM LAMINATE LAMINATE T COMPRESSION
THICKNESS TENSION COMPRESSION(mm)
3 13 (90) -3-2 .591 10.470 10.154
3 13 (90) 4-3-3 .595 9.807 11.275
9 1 (90) -4-3 .583 10.263 10.472
9 I (90) -4-2 .581 10.477 10.532
Il 4 (90) -3-4 .581 11.342 11.231
Il 4 (90) 4-3-I .587 10.380 11.380
10 15 (90) -4-4 .577 10.740 10.702
10 15 (90) -4-l .578 10.190 10.7024
52
TABLE II: SUMMARY OF (904)8 TENSILE COUPON DATA
Average
Standard
ET = 9.351 GPa (1.356 msi)
Deviation = .809 GPa (8.7%)
Average vTL = .019
Standard Deviation = .002 (13%)
Average Thickness = 1.041 mm
Standard Deviation = .010 mm (1.0%)
RUN LAMINATE AVG. THICKNESS E (GPa) v(mm) T TL
14 (90) 8-II 1.037 10.187 .020
15 (90) 8-1-2 1.029 10.083 .020
16 (90) 8- 1-3 1.041 9.877 .023
17 (90) 8- 1-4 1.056 10.223 .020
21 (90) 8-2-I 1.054 8.423 .016
20 (90) 8-2-2 1.045 8.701 .017
18 (90) 8-2-3 1.029 8.485 .016
19 (90)8 2-4 1.040 8.825 .018
TABLE 12: SUMMARY OF (900)8 CANTILEVER BEAM DATA
BEAM THICKNESS ET (GPa)
(mm) STAT IC(TIP LOAD) Ist MODE 2nd MODE 3rd MODE
(90) 8-2-A 1.071 7.724 7.961 7.998 7.878
(90) 8-2-B 1.083 7.777 7.960 7.809 8.001
(90) 8-2-C 1.073 7.822 7.971 8.362 8.168
(90) 8-2-D 1.064 7.841 7.854 7.921 8.025
AVERAGE 1.073 7.791 7.937 8.023 8.018
STD.DEV. .008 .052 .055 .239 .119(.7%) (0.7%) (0.7%) (3.0%) (1.5%)
TABLE 13: SUMMARY OF (+/- 450) SANDWICH BEAM DATA
AVERAGE G(GPa)RUN BEAM LAMINATE LAMINATE
THICKNESS TENSION COMPRESSION(mm)
6 5 (+/- 450) -3-4 .596 5.184 5.238
6 5 (+/ 45*) -3-1 .611 5.974 5.964
13 8 (+/- 45*)S-4-3 .589 5.623 5.673
13 8 (+/- 450)S-4-2 .590 5.710 5.740
14 9 (+/ 450)-4-4 .595 5.619 5.542
14 9 (+/- 45*) -4-1 .601 6.057 5.936
4 12 (+- 450)-4-3 .594 5.008 4.789
4 12 (/ 45)-4-2 .594 5.048 5.109
Average G Tension = 5.528 GPa
Standard Deviation = .405 GPA
(.802 msi)
(7.3%)
Average G Compression = 5.499 GPa (.798 msi)
Standard Deviation = .418 GPa (7.6%)
Average of G Tension and G Compression = 5.513 GPa (.800 msi)
Standard Deviation = .398 GPa (7.2%)
Average Thickness = .596 mm
Standard Deviation = .007 mm (1.2%)
TABLE 14: SUMMARY OF (+/- 450) TENSILE COUPON DATA
Average G = 5.258 GPa
Standard Deviation = .305 GPa (5.8%)
Average Thickness = .586 mm
Standard Deviation = 0
TABLE 15: SUMMARY OF (+/-
RUN LAMINATE AVG. THICKNESS G(GPa)(mm)
22 (+- 45*)2S-- 1.042 6.509
23 (+- 45*)2S- I-2 1.005 6.142
24 (/ 450) 2S- 1-3 1.021 5.583
25 ( 4 5 ) 2S- 1-4 1.032 6.145
26 (+/ 450) 2S-1-5 1.063 5.765
27 (±/- 450)2S-2-1 1.089 5.867
28 ( 45 )2S-2-2 1.059 6.117
29 (+/- 454) 2S-2-3 1.032 6.162
30 (/ 450)2S-2-4 1.040 5.423
Average G = 5.971 GPa (.866 msi)
Standard Deviation = .335 GPa (5.6%)
Average Thickness = 1.043 mm
Standard Deviation = .025 mm (2.4%)
45*) 2S TENSILE COUPON DATA
TABLE 16: SUMMARY OF (+/- 450)2S CANTILEVER BEAM DATA
BEAM THICKNESS G(GPa)
(mm) LTATD) Ist MODE 2nd MODE 3rd MODE
(+1- 450)2S -2-A 1.095 5.172 5.685 5.466 5.641
(+/- 450) 2S-2-B 1.074 5.453 6.170 5.692 5.856
(+1- 450 ) 2S -2-C 1.076 5.483 6.193 5.843 5.984
(+/- 45) 2S-2-D 1.100 5.082 4.952 5.320 5.692
AVERAGE 1.086 5.298 5.750 5.580 5.793
STD. DEV. .013(1.2%) .201(3.8%) .581(10.1%) .233(4.2%) .157(2.7%)
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+ TENSION - YOUNG5 MODULUS = 1Q6.75GIGRFRSCAL[C] CMPRES5ItN - TOUNGS MGOULU5 = 111.47GIGRPR5CALSX FISSONS RRTIO FOR TENSIONt PGQI5SGNS RATIO FOR COMPRE55IN
1000 30002000
MICROSTRRIN
400 -
LRMINRTEBERM 10RUN 7
cn
a:-)
CE
f)Lu
c-ncn
LU
300
200
100
1.00
0.75 D
H-
0.50
0.25
0.00
a e
STRESSFR
pOM
NDFPG
OURISSONS RRTIOPOINT BENDIN
VS.G TE
STRRINST
(0) 2-1-3
x
X
XXX XX XXXXXX
XXX
+ TEN51N - YOUNGS MODULUS = 104.48GIGRSCRLSED COMPRES5ION - YOUNGS MODULUS = 107..00GIGRPRSCRL5X P15SUNS RATIO FOR TENS1N6Z> PI55SUN5 RATIO FOR COMPRESSION
1000 30002000
MICROSTRRIN
-m
A t
400
LAMINATEBERM 10RUN 7 1.00
300
200
100
0.75 E)
Fr-
cnCI)
0.50 zED
0.25
0.00
S RND POISSONROM FOUR POIN
RRTI VS.BENDING T
STRFRINEST
x
x
+ TENSIN - YOUNGS MODULU5 = 99.D7 GIGAPASC[D COMPRES5ION - YOUNGS MODULUS = 102.DSGIGAX POI53ON3 RATIO FOR TENSIONZ POISONS RATIO FOR COMPRE55ION
-m I
2000
MICROSTRRIN
A A 6
STRESF
400
LAMINATEBERM 11RUN s
(0)2-1-4
1 .00
300
200
100
0
0.75 D
F-
crcc
0.50 z
cncn
0.25
0.001000
ALSPA3CALS
3000.
POISSONS RRTIOOUR POINT BENDIN
v 5G
STRRINEST
(0)2-1-1
xxxx xxxxxxxxxxxxXXXXXXXXXXXXXXXXXX X
TENSION - YOUNGS MO0ULU5 = 99.41 GIGAPASCALSCOMPRES5ION - YOUNGS MODULUS = 104.33GIGAP'A5CAL5PummISNS RATIO FOR TENSIONPG155N3 RATIO FOR COMPRESSION
1000 30002000MICROSTRRIN
RM m
STRESF
NDF
400
LAMINATEBERM 11RUN 5
(ncJ
(n
c:CL
c:aouLJ
cncnu-ir-
Cm
300
200
100
1. 00
0.75 O
H-
0 n0 . 5o z
0.00
STRESS RND POISSONFROM FOUR POIN
RTIO VS.BENDING T
STRRINEST
LAMINATE (0)4-1-4BEAM 16HUN 12
+ TE5NN - YOUNG3 MUDULU5 = 127.21GIGAPASCALS[D COMPMES5IN - YUUNG3 MGDULU5 = 120.20GIGAPASCAL3X PGI5GON3 RATIG FGR TENSIN' PI55ONS RATIO FOR COMPRES3IGN
0 1000 2000MICROSTRAIN
400
300
200
100
1 00
0.75 D
cEHr-
(n
0.50 zC
CD
CL
0.25
0.003000
STRESSFR
RND POISSONS RRTIO VSOM FOUR POINT BENDING
STRPINEST
(0)4-1-3
:xx)omoxooooxxxx
TEN5IGN - YOUNG5 MODULU5 = 113-17GIGAPA5CALSCOMPRE5IGN - YOUNG5 MODULU5 = 114.23CAPASCAL3FI55N3 RATIO FOR TENS1GNP0I53ON3 RATIO FOR COMPMES5IGN
0 1000 300020L10
MICROSTRRIN
0.75 eD
H-a o,
cn0.50 z
cncn
0.25
0.00
400
LAMINATEBEAM 16RUN 12
cn
I-'Jcn
Cn
r-fl
cn
cnf
1 .00
300
200
100
4 6
SSONS RRTLE COUPON
10T
VS.EST
STRRIN
x X X x xx x XX X X XXX Xx x
1.00
0.75 eD
Cn
0.50 z
(ncno_
C.
0.25
+ LGNGITUDINAL STRESS VS.YGUNGS MGOULUS = 129.22
X POISSGNS RATIO
STRAINGIGAPASCALS
I I I8000
MICROSTRRIN
STRESFS RROM
NDT
LAMINATERUN I1
2000
PDINSI
(0) 4-2-1
C-)a:
CE
Lii
crzH-wO
1500
1000
500
4000 120000.00
4 4 4
RPTIO VS.PON TEST
STRRIN
x X X X x X xx x X X X X x
+ LGNGITUDINAL STRESS VS.TGUNG5 MODULUS = 130.77
X PGISSGNS RATIG
4000
X X
STRAINGIGAPASCALS
120008000MICROSTRAIN
STRESF
S RROM
NDTE
LAMINATERUN 12
2oooL
PDINSI
(0) 4-2-2
SsLE
ONScou
CE
(-)
CE
CD
LUJ
coc-
cn
1.00
15001
1000
500
0.75 ED
Fr-
a:
0.50 zEDcD
0.25
0.00
L #
& 4
STRESSFIR
AND POISSOM TENSILE
ONSCO
RFTUPON
IO VS.TEST
STRRIN
LAMINATE (0)4-2-3RUN 13
x X X X XX x
x x x x x x x x x X
LGNGITUDINAL STRESS VS.YOUNGS MODULUS = 121.09POISSONS RATIO
STRAINGIGAPASCALS
2000
cn-J
CL
C-CE
Co
co
1500
1000
500
1.00
0.75 ED
CCH-
0. 50
0. 25
0. 000 4000 8000 12000
MICROSTRRIN
POISSONS RRTIONSILE COUPON TE
VS.ST
STRRIN
x x X X X X X X X XXX
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 126.23 GIGAPASCALS
X POISSONS RATIO
120008000MICROSTRRIN
1.00
0. 75 ED
ff-
0.50
0.25
0.00
6
STRESF
RGM
NDTE
LAMINATERUN 13
(0)4-2-4
2000
1500
1000
500
4000
4 1
STRESSFR
RND POOM TENS
ISSILE
ONScou
RTIOON TE
S.T
STRAIN
(0) 8-1-1
1500.
1000
5ooL x x xxx xxxxx XXXXXXXXXXXXXXXX xxxx
0.75 D
H-
(r
Co
0.50 zco(nCD
0.25
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 134.18
X POISSONS RATIO
LAMINATERUN I
2000
cn
L)J
ci:
CEC)w
cnCo
1.00
STRAIN
0 4000 8000 12000
MICROSTRRIN
0.00
4 4
STBESSFR
RND POOM TENS
SSONSLE CO
RATIUPON
0 vTES
STRAIN
(0) 8-1-2
X X X X XAI x x x x x X X X X X X X
+ LONGITUDINAL STYOUNGS MODULUS
X POISSONS RATIO
RESS VS. STRAIN= 141.75 GIGAPASCALS
120004000 8000MICROSTRRIN
LAMINATERUN 4
2000
cr)
-J
co
(n~U-)
Fy-
cn~
1.00
1500
1000
500
0.75 OD
Hr-
a:
0.50 z(nCD
0-
_0.25
I I 0.00
STRESS RND POISSONS RRTIO VS. STRPINFROM TENSILE COUPON TEST
LAMINATE (0)8-1-3RUN 5
2000- - 1.00
(n-J
E 1500- - 0.75 C)
c(n
Ln cnLLJW5i 1000- 0.50 =
wc CL
Hr-n
S 500_ x x x x x x x x x x x x x x x x x x x x - 0.25
+ LGNGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 142.09 GIGRPASCALS
X POISSONS RATIO
0 I I I I I 0.000 ~4000 8000 12000
MICROSTRRIN
4 4
STRESSFR
AOM
ND POISSONSTENSILE CO
RUP
RTIO VS.ON TEST
STRAIN
(0) 8-1-4
xX X xX XX XX XX x xx X
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 140.71
X POISSONS RATIO
STRAINGIGAPASCALS
I I I I I I
8000MICROSTRRIN
LAMINATERUN 3
2000 1.00
1500
1000
500
0.75 ED
CE
ci
0.50 zED
(nU)
ED0.
0. 25
4000 120000.00
STRESF
ND POTENS
SSONSLE CO
STRRIN
x X X X X x x x x x x x X X X
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 142.16 GIGAPASCALS
X POISSGNS RATIO
4000 8000MICROSTRRIN
12000
V
RoM
LAMINATERUN 6
(0)5-1-5
RUP
IoTE
S.T
2000
1500
1000
500
0
1.00
0.75 O
F-
(n0.50 z
ED(ncn
0.25
0.00
STRESS RND POISSONS RRTIG VS. STRAINFROM TENSILE COUPON TEST
LAMINATE (0)8-2-1RUN 7
2000 1.00
cn
E 1500_ 0.75 O
~((
Qcn
CE Co
(n n
LLJ
cr
Co 500 0.25
+ LGNG1TUDINAL STRESS VS. STRAINY1OUNGS M1ODULUS = 1441.33 GIGAPARSCALS
X PXISSGNS RATIn
0 I I I I I I 0.000 4000 8000 12000
MICROSTRRIN
4 & 4
STRESSFR
RNOM
POISSONSENSILE CO
RRTIO VS.UPON TEST
STRRIN
LAMINATE (0)8-2-2RUN 8
xx X X X XX X Xx x XX XX x X
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 142.68
X POISSONS RATIO
STRAINGICRPRSCRLS
2000
1500
1000
500
1.00
0.75 D
H-
02
(n
0.50 z
cn(n
0.25
0.000 4000 8000 12000
MICROSTRRIN
POISSONSENSILE CU
RUP
RTIOON TE
VS.ST
STRAIN
LAMINATE (0)8-2-3RUN 9
x x X X X x x x x x x x x x x x x x
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 144.43
X PGISSONS RATIO
4000 8000 12000MICRHSTRRIN
STRESFSRO
NDT
2000
1500
1000
500
1.00
0.75 E3
crcn
0.50 =cncn
0-~
0.25
0.00
STRESS RNDFROM T
POISSENS ILE
ONS RRTCOUPON
I0TE
S.T
ST RRIN
(0) B-2-4
x x x x x Ar x x x x x x x x x x x x x x x
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 145.27
X POISSONS RATIO
8000
MICROSTRRIN
STRAINGIGAPASCALS
12000
LAMINATERUN 10
2000
1500L
1.00
1000
500
0.75 ED
H-co
0.50 z(ncn
0.25
0.004000
STRESSFRO
AND POM TENS
SSONSLE CO
HRTUPON
STRRINIoTE
LAMINATE (0)8-2-RRUN 3
xx xX ~xx x x x x x Xx x x xx x x
LONGITUDINAL STRESS VS.YOUNGS MODULUS = 142.24POISSONS RATIO
20
(n
0I
CD
u
Cn
Cr)
w
CO
1.00
0.75 ED
Hr-
cn0.50 z
ED(n
0.25
0 4000 8000 12000
MICROSTRRIN
STRAINGIGAPA:
0.00
RND POGM TENS
SsLE
ONS BRTIOCOUPON TE
STRRIN
XXX X X X X X X X X X XX X x X x x X X X
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 135.33
X POISSONS RATIO
STRAINGIGAPRSCALS
120004000 8000MICROSTRRIN
G)
OD
2000
STRESF
LAMINATERUN 2
(D) 5-2-B
(n
-E
(-)
cii
CECDLUJ
cn
co-
U)
1.00
1500
1000
500
0. 75 ED
F-
cn0.50 z
U)0 n
0.25
0.00
ONS RRTIOCOUPON TE
STBRIN
x x x x x x x x x x x x x x x x x x x x x x x X
1 .00
0. 75 ED
Hr-
0.50 =ED)
0.25
+ LGNGITUDINAL STRESS VS.TGUNGS MGDULUS = 142.25
X PGISSGNS RATIG
STRAINGIGAPASCALS
G') STRESF
MNDTE
LAMINATERUN 1
200
P01SSNSILE
(0) --2-C
-)CE
-
cn
LLJ
(f)
1500
1000
500L x
0 i4000 8000 12000
MICROSTARIN
0.00
STRESS PND POISSONFROM FOUR POIN
RRTIO VSBENDING
. STRRINTEST
LAMINATE (90) 4-3-2BEAM 13RUN 3
x x
X
x x xx x Zxo
0.04
0.03 O
clco
r>
0.02 z
. n
0.01
TE1SIGN - YaUNC MGOULU5 = 10.47 GIAFPA5CALSCUMFME5IGN - YOUNG5 MODULU5 = 10.15 GIGAPASCAL5PGI53ON5 RATIO FOR TENSINPI55GN5 RATIO FOR CUMPRES5IGN
2000 60004000MICRMSTBRIN
80
60
40
0. 00
ND POISSONFOUR POIN
RRTBEN
VS. STRRINNG TEST
X
xXXX XX
xX X X X X X X X X X X X X XxX
+ TENSION - YOUNGS MIOULU5 = 9.1 GIGAPA5CALSE] CGMFES5IGN - YGUNGS MODULU5 = 11.27 GIGAPASCALSX PISSGNS RATIO FOR TEN51Nb PGISSGNS RATIG FGR CMPrMES5IGN
2000 50004000MICROSTRRIN
0.04
0.03 OD
cn
0.02 zED
0.01
0.00
STRESF
(90) 4-3-3
ROM
80
L AMINATEBEPM 139UN 3
60
40
cn--J
(I
(n
LUJ
7)
r )
S RNO POISSONROM FOUR POIN
RRTIO VS. STRRINBENDING TEST
+ TEI51N - YOUNG3 MGOULU5 = 10.26 GIGAPA5CALSE] COMrMESIG1N - YGUNGS MIOULUS = 10.47 CIGAPA5CALSX PI55GNS RATIO FOR TENSIGNZ PGI55N5 RATIG FOR COMPRESSIGN
m II I I I
2000 4000MICROSTHRIN
STRESF
(90)4-4-3LAMINATEBEAM 1RUN 9
- x
80
20
0.04
0.03 D
cr:
(n
c-nc-n
0.01
0.006000
STRESS AND PFROM FOU
ISSONS RRTIO VS.POINT BENDING TE
(90) 4-4-2
TEN5IGN - YOUNG5 MOULU5 10.4B GIGRA5CAL5CGMMESSIGN - YOUNGS MUDULU5 = 10.53 GIGAPASCAL5PUI55GN5 RATIG FGR TENI51NPGISSGNS RATIG FGM CUMPRESS1IGN
4000MICROSTRRIN
LAMINATEBERM 1HUN 3
TRRINT
60
40
0.04
x- +
0.03 D
mCC(n
0.02 2f
(n
a-0.01
0.002000 6000
a a
STRESS RNFROM OUR
I SSONSPOINT
RRTIO VSBENDING
. STRRINTEST
xx
Sxxx
4 TEN5IGN - YOUNG5 MOOULU5 11.34 GIGAPASCALS[D COMPRESSIN - YOUNG3 MGOULU5 = 11.23 GIGAPFACALSX PGISSN5 RATIG FOR TEN5ION
G I53UNS RATIO FOR COMPRE55ION
I II I4000
MI CROSTRR IN
m
80
LAMINATEBERM 4RUN 1
-x( x
(90] 4-3-4
60
40
Cr(-
rlCLLJ
CCCn
0.03 D
GE
0.02 z
0.01
0.002000 6000
STRESS RND PFROM FOU
ISSONS RRTIO VSPOINT BENDING
. STRINTEST
(90) 4-3-1
X
XXX
XXXX Xx x x x x X X XX X X X X x x x x X
+ TEN5ION - YGUNG5 MODULU5 = 10-3b GIGAPASCALSCD CGMrMES5ION - TYUNG3 MGQULU5 = 11.35 GIGArASCAL5
X PG1I5SGN5 RATIG FOR TENSIGNl PU153N5 RATIO FOR COMPrESSIGN
2000 4000MICROSTRRIN
5000
80
LAMINATEBEAM 4RUN 11
60
t40
CI-)
al:57-
cncn~
(f)
0.04
0
0.03 ED
f-
an0.02 z-
E0c.cn
CO-
0.01
0.00
STRESSFR
AND PGOM FOUR
(90)4-4-4
x x X x
+ TENSIGN - YTUNGS MGOULUS = 10.74 GIGAPA5CAL3C] CUMPrESSIGN - TGUNG3 MOULU5 = 10.71 GIGAPASCAL5X PGISSN5 RATIO FOR TENSI1N6 FI553N3 RATIO FOR COMPRE5IGN
2000
MICROSTRRIN
LAMINATEBEAM 15RUN 10
I SSONPIN
RATIOBEND I
80
60
VSNG
. STES
TRAINT
0.04
0.03 D
l:
CD
0.02 z
CD-
0.01
0. 006000
STRESS RND POISSONFROM FOUR POIN
RRTIO VS. STRPINBENDING TEST
x
4- TEN51(114 - YIOUJ5 MU0ULU5 = 10-IS GIGPA'R5CAL5+D CT1RE55NIN - YOUNG MOULU5 = 10.70 GICPA5CAL5X PI5SONS RATIO FUM TENSIGN
PGISSGNS RATIG FGR CMrESSIGN
4000
MICROSTRRIN
.
LAMINATEBEAM 15RUN 1080
60
40
20
QQ4
0.03 D
a-
cn0.02 z
CD
0.01
0.00
2000 6000
(90) 4-4-1
STRESS AND POISFROM TENSIL
ONS RRTIO VCOUPON TES
S.T
STRRIN
x x
x x x x
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MGOULUS = 10.19 GIGAPRSCALS
X PGISSONS RATIO
4000
MICROSTRRIN
LAMINATERUN 14~
(90) 5-1-1
40
30
20
10
0.04
0.03 CD
CE
0.02 zc-nc-n
Q.
0.01
0.002000 6000
STRESS AND POISSONS RATIO VS. STRRINFROM TENSILE COUPON TEST
LAMINATE (90)8-1-2RUN 15
40 0.04
30 0.03 ED
cnnU)_ Cn
W 20 X 0.02 zX x x x x
C(
(n c
10 -0.01
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 10.08 GIGAPASCALS
X POISSONS RATIO
0 2000 4000 6000
MICROSTRAIN
STRESSF R
(90) 8-1-3
H0M
NDT
P01ENSI
SSONSLE COU
RRTIOPON TE
VS.ST
STRAIN
0.04
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 9.98 GIGAPASCALS
X POISSONS RATIO
I I
2000 4000
MICROSTRRIN
LAMINATERUN 16
40
cn-Ja:CE)
(n
CE
CLa-
ancnLUi
mo
cn
x X x
30
20
10
x/ x
0.03 CD
CEa:
cn
C-
0.02 ED
Co
a_0.01
60000.00
STRESSFR
AND POIOM TENSI
SSLE
ONS RRTIO VS.COUPON TEST
ST RRIN
190) 8-1-4
+ LONGITUDINAL STRESS VS.YOUNGS MODULUS = 10.22
X POISSONS RATIO
STRAINGIGAPASCALS
60004000
MICROSTRRIN
0.03 CD
C-
cn~
0.02 z
E_
cncn
0.01
0.00
L AMINATERUN 17
40 0.04
30
20
10
Xx x x x
2000
STRESS AND POISSONS RATIU VS. STRAINFRM TENSILE COUPON TEST
LAMINATE (90) 8-2-1RUN 21
40 0.04
c 30 0.03 ED
LLJ
w 20 0.02
P x xS x x x RTcn
S10 0.01
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MO~DULUS - 8.412 GIGAPASCALS
X POISSONS RATIO
0 0.000 2000 4000 6000
MICRBSTRAIN
STRESSFR
A0M
ND POITENS I
SSONSLE CO
RRTUPON
ioTE
VS.ST
STRAI N
(90)8-2-2
40
30
20 kx x x x
0.04
0.03 ED
CCE
(n)0.02 z
(L
0.01
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 8.70 GIGAPASCALS
X POISSONS RATIO
4000
MICROSTRRIN
LAMINATERUN 20
2000 60000.00
STRESSFR
RND POISSOM TENSILE
ONS RRTCOUPON
ID VS.TEST
STRRIN
(90) 8-2-3
x / x x X X
+ LONGITUDINAL STRESS VS. STRAINYOUNGS MODULUS = 8.48 GIGAPASCALS
X POISSONS RATIO
I I I I I4000
MICROSTRAIN
LAMINATERUN 18
40 0.04
30
20
10
x
0.03 E3
H-
cr
0.02 z(D
0-~
0.01
0.002000 6000
STRESSFR
FND POM TEN
ISSILE
ONS RRTIO VS.COUPON TEST
STRRIN
(90) 8-2-4
x x x x x
+ LONGITUDINAL STRESS VS, STRAINYOUNGS MODULUS = 8.82 GIGRPASCALS
X PCISSNS RATIO
4000
MICROSTRAIN
LAMINATERUN 19
40
30
20
10
x x
0.04
0.03 ED
F-CEry(n
0.02 z
. D
0.01
0.002000 6000
STRESS VS.FOUR POINT
SHERR STRRINBENDING TEST
LAMINATE (+/-45) 5-3-4BEAM 5RUN 6
+ TENSIN - HEAfEl CMPRE5ITN -
10000
C9
MaDULU5 = 5.13 GIGAPASCALS3MEAR MODULUS = 5.24 GIGAPASCAL5
P0000SHERR STBRIN(MICROSTBRIN)
SHERRFROM
100
75
so
ElD ID C
30000
SHEARFROM
STREFOUR
5 VS. SHEAR STRAINPOINT BENDING TEST
LAMINATE (+/-45) 5-3-1BERM 5RUN 6
+++
+TENSIUN - HEAR MODULUS 5.97 GCGRRCALSC9 CUMPRE33IGN - SHEAR MUDULUS = 5.96 GIGAPRSCALS
0 10000 20000 30000SHEAR STRRIN(MICROSTRRIN)
a 4
100
75
50
4 0
SHERR STRESS VS. SHERR STRRINFROM FOUR POINT BENDING TEST
LAMINATE (+/-45) 5-4-3BERM 5
100 HUN 13
(n
-I
cna- 75
LU E
Eil
(n 50 m l
LU 2 E
rr- 19
H-(F)
LLJ
+ TEN5IGN - 5HEAR MODULU5 5.62 GIGAPASCAL5(D 9 CEMl ME33IGN - SHEAR MODULUS = 5.67 GICAPASCAL3
0 10000 20000 30000
SHERR STRPIN(MICROSTRPIN)
SHERRFROM
STREFOUR
SS vs.POINT
SHERR STRRINBENDING TEST
LRMINRTE (+/-(j5) 3-Lj.-~BERM ~~iUN 13
++
-I-+
+±
+±
++
+4-
+±
++
±±
±+
±4-
+-I-
4- TENJ~1~J - 5l1ER~ M~QULU5 S71 GIcRrfl3CRL50 C~t1~E55I~N - ~t1EPM M~QULU3 = S7q ~IGPPR3C~L~
0 10000 20000 30000SHERR STRRIN(MICROSTRRIN)
100
50
4
SHERR STREFROM FOUR
SS VS.PINT
SHERRBENDIN
LAMINATE (+/-45)5-4-4BERM 9H9UN 114
- TEN51O - CER MUULUS =-62 5I.5 GASCAL SCD CMPRE5310W - 5HEAR MCIDULU5 =S-54 GICrAC1L5
0 10000 20000
SHERR STRAIN(MICROSTRAIN)
100
TRRINTEST
75
(n-J
ci:(J(1D
CE
CDLtJ
LLiC-
(F)
30000
STRESS VS.FOUR POINT
SHERR STRRINBENDING TEST
LAMINRTE (+/-45)3-4-1BERM 9HUN 14
++
+++
+ TEM51GN - SEAR MUDULUS 6.06 GIGRAACALS(1] COMPMESSIGN - SMEAR MODULUS = S.94 GIGAPASCALS
0 10000 20000
SHERR STPRIN(MICRPSTRRIN)
SHERRFROM
100
(nJ
Cn
cc
cL-
(O)c-nLU
cr
CE
Lu
c-n
75
50
25
030000
SHERR STRESS VS. SHERR STRRINFROM FOUR POINT BENDING TEST
LAMINATE (+/-45) S-3-3BEAM 12RUN 4
CG SG ME M 5 Gl RSCLS
EDP ED
+ TEN5ION - 5lIEAR MOtIULU5 =S.01 GIAACLCD COMFrRE3310N - 51ER MUOULU5 = 479 GICRAACAL5
0 10000 20000 30000
SHERR STRRIN(MICHOSTRRIN)
NJ
100
STRESSFOUR PC
vs.I NT
SHERR STRPINBENDING TEST
LAMINATE (+/-45)3S-3-2BEAM 12RUN 4
+~~~~~ ~ ~ ~ +EBO HA UUU5 S0 IAACL
- g ~ ~~~~~ + +OPE5O + +HA UUU S1 IAAC
0 10000 20000 30000SHEAR STRRIN(MICROSTRRIN)
SHERRFROM
100
-* SHERR STRESS VS. SHERR STRAINFROM TENSILE COUPON TEST
LAMINATE (+/-45)5-1-3RUN 31
100
(n
CE)
(n
a: 75
Cn 50 --
SHEAR MODULUS = 5.47 GIGAPASCALS(n
010000 20000 30000
SHEAR STRAIN (MICROSTRAIN)
STRESSTENSILE
VS.cDu
SHERR STRRINPON TEST
LAMINATE (+/-45)5-1-4RUN 32
+A +
++
- SEA MDUUS= .0 GGAASAL
10000 20000
SHERR STRRIN(MICROSTRAIN)30000
SHERRFROM
100
(n
-
LLJcLa:
H--
UJ(n
a:Co
U)j
75
50
25
STRESSTENSILE
VS.cou
SHERRPON TE
STRAINST
SHEAR MODULUS = 6.51 GIGRPASCRLS
10000 20000
SHERR STRRIN(MICROSTRRIN)
SHERRFROM
LAMINATE (+-45)23-1-1RUN 22
100
75
50
25
30000
SHERR STRESS VS. SHERR STRAINFROM TENSILE COUPON TEST
LAMINATE (+-45)2S-1-2RUN 23
100
c-
cn(n
CIE + +cL.
+
(n 50 +(nLU+wc +CC +
(n
m- 25 +CE +LU
SHEAR MODULUS = 5.14 GIGAPASCRLS(n
0 -- I |
0 10000 20000 30000
SHERR STRAIN(MICROSTRAIN)
SHERR STRESS VS. SHERR STRRINFROM TENSILE COUPON TEST
LAMINATE (+-45) 2S-1-3RUN 24
100
cn
L.)CcE- 75
Cowi +
+
Cr 50
cn +
Cr +
+
a- 25 +aE +LJ
SHEAR MODULUS 5.58 GIGAPRSCALS(n
0 i0 10000 20000 30000
SHEAR STRAIN(MICROSTRRIN)
STRESSTENSILE
VS.Cou
SHEPON
RRTE
STRRINST
LAMINATE (+-45)2S-1-4RUN 25
SHEAR MODULUS = 6.15 GIGAPASCALS
10000 20000
SHERR STRRIN (MICROSTRRIN)
SHERRFROM
100
75
25
030000
SHERR STRESS VS. SHERR STRAINFROM TENSILE COUPON TEST
LAMINATE (+-45)2S-2-1RUN 27
100
n.J
ci:L)
ci: 75
LLJCE +Cn 5 + +
III + U
+
cn ++LL)
Lii +
oc 25 +
SHEAR MODULUS 5.87 GIGAPASCALS
0 .0 10000 20000 30000
SHEAR STRAIN(MICROSTRAIN)
STRESSTENSILE
VS.Cou
SHERRPON TE
LAMINATE (+-45) 2S-2-2RUN 28
(nCE.-J
cnCo
a-cE:
Lu
co(nLLJH-cn
cii
co
10000 20000
SHERR STRAIN(MICROSTRAIN)
4
SHERRFROM
100
STST
BRIN
SHEAR MODULUS = 6.12 GIGAPASCALS
75
50
25
30000
STRESSTENSILE
VS. SHERRCOUPON TE
STRAINST
LAMINATE (+-45)25-2-3RUN 29
SHEAR MODULUS = 6.16 GIGAPRSCRLS
10000 20000SHERR STRAIN(MICROSTRAIN)
SHERRFROM
100
75
50
25
030000
SHERR STRESS VS. SHERR STRRINFROM TENSILE COUPON TEST
LAMINATE (+-45)25-2-4RUN 30
100
cnCE)
-nci: 75n +a: +CD + +
+
crn 50 +(n +w +
cc: +
SHEAR MODULUS =5.42 GIGAPASCALS(n
0010000 20000 30000
SHARR STRIN(MICRSTRR IN)
STRESSTENSILE
VS. SHERRCOUPON TE
STRRINST
LAMINATE (+-45)25-1-5RUN 26
SHEAR MODULUS = 5.77 GIGAPASCALS
10000 20000 30000SHERR STRRIN (MICROSTRRIN)
SHERRFROM
100
75
50
25
S
120
FIG. 8: SANDWICH BEAM 3 IN TEST JIG AFTER FAILURE
FIG. 9: SANDWICH BEAM TEST SETUP
121
FIG. 11: BEAM 5 BEING TESTED AT A LOAD OF 740 POUNDS