Supplement to
MECHANICAL PROPERTIES CIF
CROSS4AMINATED AND COMPOSITE
GLASS-FACIRIC-13ASE PLASTIC LAMINATES
March 1953
INFORMATION REVIEWED
AND REAFFIRMED
1959
LO COPY
masse
•
rot.ota to:
Wood Engineering Ker*earcit
!Crest Product* Laboratory
53705
giscorAl ilk
This Report is One of a Series
issued In Cooperation !with the
AIR FORCE-NAVY-CIVIL SUCCOMMIITEE
on
AIRCRAFT DESIGN CRITERIA
Under the Supervision of the
AIRCRAFT COMMITTEE
of the
MUNITIONS BOARD
No. 1821-A
uNrrED STATES DEPARTMENT OF AGRICULTURE
FOREST SERVICE
FOREST PRODUCTS LABORATORY
Madison 5, Wisconsin
In Cooperation with the University of Wisconsin
SUPPLEMENT TO
MECHANICAL PROPERTIES OF CROSS-LAMINATED AND
COMPOSITE GLASS-FABRIC-BASE PLASTIC LAMINATES!
By
ALAN D. FREAS, Engineer
and
FRED WERREN, Engineer
2 Forest Service
Forest Products Laboratory, —
U. S. Department of Agriculture
Abstract
Methods of predicting the properties of cross laminates and composite parallel
laminates were presented in Forest Products Laboratory Report No. 1821.
Reported herein are the results of tests on additional composite parallel laminates, together with the results of application of the computation methods previously presented. The conclusions of the original report are reasonably confirmed by the results of the present study.
Introduction
It was the purpose of this study to supplement the data of the original report h.y
tests on composite parallel laminates of a different type than that originally
studied. In this study, facings of from 1 to 11 plies of a fine-weave fabric were
applied to a core of a coarser-weave fabric. This type of construction might
be employed where it was desired to impart a relatively smooth surface to a.
laminate made primarily of a coarse-weave fabric.
-This
—This progress report is one of a series prepared and distributed by the Forest
Products Laboratory under U. S. Navy Bureau of Aeronautics Order No.
NAer 01338 and U. S. Air Force No. USAF-(33-038)51-4 2 26-E, Amend. 2
(53-131), Results here reported are preliminary and may be revised as
additional data become available.
?Maintained at Madison, Wis. , in cooperation with the University of Wisconsin.
Report No. 1821-A
Agriculture -Madison
-1-
Material
One 1/4-inch-thick panel of each of three constructions was tested in this
study. All panels were made with a core of 162-114 fabric with 1, 6, or 11
plies of 112-114 fabric on each face. The panels were laminated with resin
2, a laminating resin of the polyester (styrene-alkyd) type conforming with
the requirements of types I, II, and III of U. S. Air Force Specification
12049. For al] panels, the resin was catalyzed with 0. 8 percent of benzoyl
peroxide by weight.
Fabrication procedures were as described in the original report. Information on the make-up and physical properties of the various panels is given
in table 11.
Before the composite panels were fabricated, an attempt was made to obtain
162-114 fabric from manufacturer A, who furnished the 162414 fabric for the
earlier panels.! The manufacturer did not have the fabric available at that
time, so the material was purchased from manufacturer B. Both manufacturers
stated that the 114 finishes were identical and that the fabrics, as obtained from
either source, were considered to be the same. Subsequent tests, however
showed that this was not so. Therefore, a 162-114 control laminate was fabricated by using fabric from the same roll as that used in the composite laminates.
Fabrication procedures were the same as those used for the composite laminates.
Testing
All test procedures were the same as those described in the original report
Presentation and Discussion of Results
Average miscellaneous data and test results are given in tables 11, 12, 13, 14,
and 15. Control data for the 112-114 laminate and for the shear properties of the
162-114 laminate are from earlier tests of another study. 3 Tension, compression,
and bending properties of the 162-114 laminate are from the control laminate made
of the same fabric used in the composite laminate. Comparisons of test results
with computed results are shown in tables 16 through 20.
3
--Werren, Fred. Mechanical Properties of Plastic Laminates. Forest Products
Laboratory Report No. 1820. February 1951.
Report No. 1821-A
-2-
Methods of computing the various properties of the composite laminates from
the properties of the components are described in the original report and will
not be repeated herein except for the equations.
Tension and Compression
Modulus of elasticity. --Modulus of elasticity of composite laminates may be
computed from the equation:
=n
>
E a =
(It)
E1A1
i=1
where E a = modulus of elasticity of the composite laminate
A = cross-sectional area of the composite laminate
Ei
= modulus of elasticity of the ith ply in the direction of stress
A i = cross-sectional area of the ith ply.
Values of E i were taken from the data for parallel laminates, tables 12 and 13.
A compariMn of computed and test results is shown in tables 16 and 17. Computed and test values for modulus of elasticity in tension and in compression
generally agree within 10 percent.
Strength properties. --Three possible procedures for computing the tensile and
compressive strength properties of composite laminates from those of the CQMponent materials are represented by the equations:
Method 1
Sa - A
[A l 4
Ez
A2
(2)
Method 2
_
Report No. 1821-A
52 [ E
A
Ez
A1
-3-
+
Az]
(3)
Method 3
Sa=
1
i =n
S•A.
(4)
i =1
where Sa = strength property of composite laminate
S / = strength property of type 1 layers parallel to direction of stress
S 2 = strength property of type 2 layers parallel to direction of stress
Si = strength property of the ith ply parallel to stress
E l = modulus of elasticity of type 1 layers parallel to direction of stress
modulus of elasticity of type 2 layers parallel to direction of stress
E
2 =
A / = total cross-sectional area of type 1 layers
A2 = total cross-sectional area of type 2 layers
Ai = cross-sectional area of the ith ply
A = total cross-sectional area.
In method 1, the computations were based on the properties of the 112 laminate;
that is, Sl equals the strength of the 112 laminate. In method 2, the computations were based on the strength of the 162 laminate.
Ratios of strength values computed by all three methods to strength values obtained by test are shown in tables 16 and ]7.
As was to be expected from the results given in the original report, neither
method 1 nor method 2 gives good correlation between computed and test values.
By using method 3, there is reasonably good correIation between the tensile
values. Compression values are also in reasonable agreement except for the
composite laminate with facings of one ply of 112-114 fabric.
Properties in tension at a 45° angle. --Two methods of predicting the tensile
properties of composite laminates at an angle of 45° to the warp direction were
discussed in the original report. The first involved the use of the formulas of
Report No. 1821-A
-4-
Forest Products Laboratory Reports Nos. 1803 and 1803-A, with the
tensile and shear properties of the composite laminates at angles of 0° and
90°. The second was based on using the weighted average of the 45° properties of the component parallel laminates. Parallel-laminate properties
were weighted according to the proportion of the area occupied in the composite laminate. Results of both methods of computation are shown in table
20.
Both methods, except for stress at proportional limit, give reasonable
correlation between computed and test results in tension, with the second
method being somewhat better. The lack of correlation for stress at proportional limit may be attributed largely to difficulties in accurately locating the
proportional limit.
Discussion of tensile properties. --Previous discussion has mentioned that a
new panel of 162-114 fabric, parallel-laminated, was made to furnish control
values for tension, compression, and bending tests. This was necessary
because the laminates made from the 162-114 fabric from manufacturer B were
appreciably lower in tensile properties than comparable laminates made with
similar fabric from manufacturer A. Other mechanical properties were
reasonably comparable.
In two other instances in tests of plastic laminates, laminates made with
fabric from manufacturer B were appreciably lower in tensile properties than
those made with fabric from manufacturer A. This may be seen by a comparison of items 1-3 and 1-13 (128-114 laminates) and items 1-8 and 1-20
(181-114 laminates) in Forest Products Laboratory Reports Nos. 18202 and
1820-A. it Examination of the data will also show that laminates made with
fabric from manufacturer A were thinner, had a higher specific gravity, and
had a lower resin content.
Analysis of these data indicates that there is apparently some difference
between 114 finish fabrics furnished by the two manufacturers, even though
the materials are supposed to be identical. This difference or variation,
which might be in :he finish, substantially affects the tensile properties of
the laminate. Some differences might also be noted in tensile tests of the
fabric, as are shown in the results of tests of 162-114 fabric, table 21.
Static Bending
Modulus of elasticity. --The modulus of elasticity in static bending may be
computed from
4
—Werren, Fred. Supplement to Mechanical Properties of Plastic Laminates_
Forest Products Laboratory Report No. 1820-A. February 1953.
Report No. 1821-A
-5-
EiIi
(5)
where E a = modulus of elasticity of the composite laminate
Ei
= modulus of elasticity of the ith ply parallel to span
Ti
= moment of inertia of the ith ply about the neutral plane of the
laminate
= moment of inertia of the laminate about its neutral plane.
A comparison of computed with test values is shown in table 18.
--Three possible methods of computing the static-bendingstrength properties of the composite laminates from the strength_properties of
the component materials are represented by the equations:
Strength properties.
Method 1
1
E
[11
(6)
+1 2
a
Method 2
f
1
E
2 12
Er
2tf
[1
a
(7)
Method 3
=n
f
a
=
1
I
fiIi
i =
Report No. 1821-A
1
-6-
(8)
where fa = strength property of the composite laminate
f 1 = strength property of the type 1 layers parallel to span
f2 = strength property of the type 2 layers parallel to span
fi = strength property of the ith ply parallel to span
= sum of moments of inertia of the layers of type 1 about the
neutral plane of the laminate
1 = sum of moments of inertia of the layers of type 2 about the
2
neutral plane of the laminate
Ii = moment of inertia of the ith ply about the neutral plane of the
laminate
I = moment of inertia of the composite laminate about its neutral plane
E 1 = modulus of elasticity of the layers of type 1 parallel to span
E 2 modulus of elasticity of the layers of type 2 parallel to span
tf
d
= ratio of thickness of surface layer to total thickness of laminate.
In method 1, computations were based on the strength of the 112 laminate; that
is, f 2 equals the strength of the 112 laminate. In method 2, computations were
bas7F on the strength of the 162 laminate. The value of tf was taken as the total
thickness of each 112 facing. Results of computations birThese methods are
shown in table 18.
As was found in the original report, methods 1 and 2 do not give consistently
good correlation between computed and test results. Method 3, however, gives
generally good correlation except for stress at proportional limit.
Shear
The original report showed that reasonably good correlation between computed
and test results in shear could be expected by use of the following equation:
=n
FiAi
Fa = 1
=I
Report No. 1821-A
-7-
(9)
where F a = property of composite laminate
A = total cross-sectional area of composite laminate
F 1 = property of ith ply
A 1 = cross-sectional area of ith ply.
A comparison of computed and test results is given in table 19. In view of the
small number of tests (two per laminate), the correlation is considered satisfactory.
Conclusions
The results of tests of three composite laminates made of 112-114 and 162-114
glass fabrics confirm reasonably well the conclusions of the original report..
The conclusions of the original report were:
(a) The properties of composite or cross laminates stressed in tension or
compression parallel or perpendicular to the warp direction may be taken a s the
average of the properties of the component layers, parallel to stress, weighted
according to the proportion of the area they occupy.
(b) The properties of cross laminates or composite laminates stressed in bending parallel or perpendicular to the warp direction may be taken as the average
of the properties of the component layers, parallel to span, weighted according
to the proportion they contribute to the moment of inertia of the laminate.
(c) The properties of cross laminates or composite laminates stressed in shear
with the shear loads applied parallel and perpendicular to the warp direction may
be taken as the average of the properties of the component plies when loaded in
the same way, weighted according to the proportion of the area they occupy. For
cross laminates, the properties are the same as for parallel laminates.
(d) The properties of cross laminates or composite laminates stressed in tension
at 45° to the warp direction may be taken as the average of the properties of the
component layers, parallel to stress, weighted according to the proportion of the
area they occupy. For cross laminates, the properties are the same as for the
parallel laminates.
(e) Since the laminates tested involve fabrics having a range of relative strength
parallel and perpendicular to warp typical of the range used in aircraft applications,
the relations developed should be applicable to any other glass fabrics now being
used for aircraft laminates.
Report No. 1821-A
-8-
Table 11.--Miscellaneous data relative to composite parallel laminates
made of 112-114 and 162-114 fabrics and resin 2
Panel : Number of plies1 :Specific: Resin : Barcol : gravity: content: hardness
No. : :Total:112-114:162-114:
: tabri2:-fObr12: :
:
:Percent :
: Thickness
:
:
:
:
In.
69
:
69
:
0.250
.250
112-114 Control Laminates
5
6
:
:
84 : 44 •
84 : 284
....
• 1.71 : 43.5 :
1.70 : 43.4 :
162-114 Laminates--for Shear Controls
52
53
17 •
•
•
17 2 17 : 1.76 : 37.7 :
217 : 1.76 : 38.2 :
62
:
.252
62
:
.254
63
:
.257
69
:
.262
69
69
:
:
.264
162-114 Laminate--for Tension,
Compression, and Bending Controls
223
:
16 •
°
316 : 1.72 : 41.0 :
Composite Laminates
217
218
212
219
:
18 :
:
26:
:
:
34 :
E 2 : 216 : 1.73 : 41.1 :
214 : 1.71 : 42.0 :
222 : 212 : 1.71 : 42.0 :
.260
1
–Layers of 112-114 fabric formed a facing on a core of 162-114 fabric.
One-half the number of plies of 112-114 fabric shown above at each face.
Fabric from manufacturer A.
Fabric from manufacturer B.
Report No. 1821-A
Table 12.--Results of tension tests of parallel laminates made of 112-114
and 162-114 fabrics and resin 2
Stress at -°.
Fabric :Layers of: Direc- : Modulus of
:----- --------------- ------ ----: 112-114 :tion of: elasticity
:fabric at: test :------ ----------- :Proportional limit: Ultimate
:
:Initial:Secondary: :each face:
:Initial :Secondary:
:
:
2
6
--------g --------- :66-66-..---: ------- g...... ...... : :
:
:Degrees : 1,000: 1,000 : P.s.i. : P.s.i. : P.s.i.
: No.
:
: 2:ILtit : 2.i. :
.
Controls - Parallel-laminated
112-114 g.........:
.
:
0
45
90
: 2,690 : 2,390 : 11,800 : 29,500 : 42,700
:.......: 1,540 : ----- ...: 3,760 : 20,550
: 2,640 : 2 ) 240 : 9,800 : 27,050 : 38,650
0
45
90
: 2,760 : 2,440 : 7,940 : 19,750 : 37,600
:e... .e: 1,360 :........: 3,460 : 16,350
: 1,960 : 1,520 : 4,560 : 10,200 : 20,600
.
°.
162-114 :.........:
.
°.
:
°.
Composite Laminates - Parallel-laminated
112-114 8 :
162-114:
:
:
112-114;:
162-114:
:
:
112-11482
162-114:
:
1
:
:
:
0
45
90
0
6
11
Report No. 1821-A
:
:
:
:
:
:
:
0
45
90
0
45
90
: 2,760 : 2,370 : 6,400 : 18,500 : 35,600
0
. 1,450 :.. ------ : 2,580 : 16,000
: 2,040 : 1,380 : 4,330 : 11,750 : 19,900
0
.
:
:
..
: 2,670 : 2,250 : 8,120 : 20,950 : 36,050
° 2,580 : 16)400
° 1,350 :
°
: 1,960: 1,400 : 4,660 : 12,850 : 22)400
.
:
:
:
.
: 2,690 : 2,240 : 8,490 : 24,500 : 37,900
° 2,380 : 16,200
:.......: 1,530 °
: 2,000 : 1,490 : 5,520 : 14,750 : 24,300
Table 13.--Results of compression tests of parallel laminates made of
112-114 and 162-114 fabrics and resin 2
Fabric :Layers of: Direc- : Modulus of
: 112-114 :tion of: elasticity
:fabric at: test :
:.
:each face:
:
No.
:Degrees :
1,000
Stress at -:
:
:Proportional: 0.2 percent: Ultimate
:
offset
:
limit
:
:
P.s.i.
:
P.s.i.
: P.s.i.
'
•
Controls - Parallel-laminated
0
90
:
:
162 .. 114 :MO000090:
0
:
:
90
:
112-114
:70000000.:
.
2,820
2,630
23,250
21,600
:
:
36,850
32,900
: 36,850
: 32,900
2,850
2,100
11,250
10,200
:
:
19,100
18,300
: 19,100
: 18,300
Composite Laminates - Parallel-laminated
112-114;:
162-114:
1
112-114j:
162-114:
6
112-114;:
162-114:
11
Report No. 1821-A
:
0
90
:
:
2,810
1,900
:
:
0
90
:
:
2,660
2,020
:
0
90
:
:
2,810
2,050
e
0
:
0
16,950
15,700
:
:
23,900
21,750
16,450
17,300
:
:
22,100
21,450
16,750
17,800
23)400
23,250
: 23,900
: 21,750
22,100
: 21,450
t
23,400
23,250
Table 14.--Results of static-bending tests of parallel laminates made of
112-114 and 162-114 fabrics and resin 2
:Modulus
Stress at --
Fabric :Layers of: Direc- Modulus of :
: of
112-114 :tion of: elasticity : :Proportional: 0.2 percent :rupture
:fabric at:
test
:
:
offset
:
limit
:
:each face:
:
:
:
:
P.s.i.
P.s.i.
: P.s.i.
:Degrees :
1,000
No.
..
: p.s.i.
:
Controls - Parallel-laminated
112-114
.....
0
90
162-114
.....
o
90
:
:
2,590
2,400
:
:
31,000
26,450
:
58,250
48,350
:
:
58,250
48,350
17,200
11,300
:
:
32,550
22,850
:
:
35,000
26,900
•
2,470
1,720
:
:
Composite Laminates - Parallel-laminated
112-1148:
162-114:
1
:
0
90
112-114,1:
162-114:
6
:
:
0
90
112-114::
162-114:
11
:
:
0
90
Report No. 1821-A
2,690
2,000
24,550
12,200
:
:
2,680
2,110
24,050
15,250
:
:
2,730
2,270
28,200
14,950
•
:
38,600
26,250
: 38,800
28,800
43,750
33,750
:
54,65o
36,800
:
43,750
33,750
55,050
37,800
Table 15.--Results of panel shear tests of parallel laminates made of
112-114 and 162-114 fabrics and resin 2
Stress at -Fabric :Plies of : Modulus
: 112-114 :
of
:fabric at: rigidity : Proportional : 0.2 percent : Ultimate
offset
:
:
limit
:each face:
:
12---
000
No.
°
P.s.i.
P.s.i.
P.s.i.
Controls - Parallel-laminated
112-114 :.........:
660
1,920
:
:
162-114 :.........: 580
2,220
:
:
:
4,070
:
:
13)600
4,810
:
11,700
:
10,300
Composite Laminates - Parallel-laminated
112-114)':
162-114:
.
112-114;:
:
2,240
570
:
:
:
6
:
2,160
:
:
4,580
•.
10,850
2)540
:
:
4,850
:
10,900
560
:
162-114:
112-114;:
4,930
1
11
162-114:
Report No. 1821-A
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Table 19.--Ratios of computed shear properties to properties obtained
from test
Ratios
:
Construction
Stress at -:Modulus :
:
of
:rigidity:Proportional:0.2 percent: Ultimate
:
: offset
limit
:
:
:
:
: 1.02 :
1 ply of 112-114 fabric
on each face
0.99
•.
:
0.97
:
1.14
1.01
:
:
1.03
:
:
1.10
.84
:
.95
:
1.12
6 plies of 112-114 fabric
:
:
: 1.05 :
:
11 plies of 112-114 fabric :
:.
on each face
: 1.18 :
on each face
Table 20.--Ratios of
pro pe rties,
properties obtained in test
Construction
1 ply of 112-114
fabric on each face
Stress at
:
Proportional : Ultimate
lim
limit
:
:Method 1:Method 2: :Method 1:Method 2:Method 1:Method 2
Modulus of
elasticity
•.
:
0.99
plies of 112-114
fabric on each face
1.06
11 plies of 112-114
fabric on each face
.89
6
Report No. 1821-A
:
.
•
:
.
•
:
•.
:
:
1.74
1.03
:
1.67
.92
:
2.13
0.94
.•
•.
:
1.34
°.
1.11
:
:
1.03
•.
•.
:
:
:
:
.•
:.
:
1.36
:
1.15
:
1.03
1.49
:
:
1.19
.°
:
1.08
Table 21.--Physical properties of 162-114 glass fabric from two different manufacturers)
Properties
Samples from : Manufacturer A : Manufacturer B
Breaking strength (pounds per inch)
(Grab method)
Warp
Fill
:
:
•.
459
277
398
Elongation (percent)
Warp
Fill
•
:
:
8.0
4.0
7.3
4.5
Breaking strength of thread (pounds) Warp
Fill
:
:
:
17.0
18.1
17.5
15.9
Elongation of thread (percent)
Warp
Fill
:
:
1.7
1.7
Thread count per inch
Warp
Fill
•
:
:
28.5
16.5
Yarn ply
Warp
Fill
:
:
5
5
:
5
(2)
:
(2)
:
253.4
1.87
1.49
•
Weave
Twist per inch (yarn)
Warp
Fill
28.5
16.5
4.5
4.6
4.1
4.2
-'Properties based on tests at the Wright Air Development Center,
Wright-Patterson Air Force Base, Ohio.
Plain, 1 up and 1 down.
Report No. 1821-A
5