Scholars Journal of Engineering and Technology (SJET) ISSN 2321-435X

Scholars Journal of Engineering and Technology (SJET) ISSN 2321-435X
Scholars Journal of Engineering and Technology (SJET)
ISSN 2321-435X
Sch. J. Eng. Tech., 2013; 1(1):13-26
©Scholars Academic and Scientific Publisher
(An International Publisher for Academic and Scientific Resources)
www.saspublisher.com
Research Article
Fracture Properties of Glass Fiber Composite Laminates and Size Effect
Y. Mohammed1, Mohamed K. Hassan1, Abu El-Ainin H2, A. M. Hashem1
1
South Valley University, Qena, Egypt, 83521
2
Minia University, Minia, Egypt, 61111
*Corresponding author
Y. Mohammed
Email: [email protected]
Abstract: The fracture properties like fracture toughness and nominal strength of glass fiber reinforced epoxy laminates
are very important especially when using cohesive zone model. Compact tension specimen test for [0, 90] 2sand center
cracked specimen tension test for Quasi-isotropic laminates [0/45/90]2s and [0/45/90/-45]sare carried out. The open hole
tension test is performed on a matrix of specimen of various diameters (2, 4, 6, 8 and 10 mm) keeping the hole diameter
to width (d/w) equal 1/6. The fracture toughness of cross ply laminates is measured as 51.98 Kj/m 2whereas, for Quasiisotropic laminates [0/45/90] 2s and [0/45/90/-45]s are 32.98 and 31.5 KJ/m2 respectively. A strength reduction of 32 % is
observed with increasing the hole diameter from 2 mm to 10 mm, while this percentage was decreasing by inserting an
angle ply as 26 % for [0/45/90]2sand 14 % for [0/45/90/-45]s. Delamination are observed with thickness increasing for
un-notched specimens. Fiber orientation affects deeply the laminates carrying capacity.
Keywords: Nominal Strength, Fracture toughness, Quasi-isotropic laminates, Glass fiber reinforced epoxy
INTRODUCTION
Composite material has been widely applied in
industry, military structure and Marian. Also analytical
and numerical model such as; cohesive zone model
which is basely depended on two main parameters
which are un-notch nominal strength and fracture
toughness of the material [1-5]. Therefore, the precisely
experimental evaluation of the mechanical properties of
this material is very important for used in design,
modeling and simulation [5, 6] Pinho et al. [7]
investigated the fracture toughness of carbon fiber
reinforced laminates using compact tension test and
compact compression test specimens. It is concluded
that the initiation and propagation fracture toughness of
the cross ply laminates [0, 90] 8s are determined as 91:6
kJ/m2 and 133 kJ/m2 respectively and for fiber
compressive kinking, an initiation value of 79:9 kJ/m2.
It is used especially costly equipments. Donadon et al.
[8] studied
the tensile fiber fracture toughness
characterization of hybrid plain weave composite
laminates using non-standardized Over height Compact
Tension (OCT) specimen. Initiation and propagation
values around 100 kJ/m2 and 165 kJ/m2, respectively,
were obtained for the fiber toughness using the
compliance method. It was found that the application of
the ASTM E399-90 is fully questionable for composites
in general and it can overestimate the toughness values
if used in its original form. Three-point bend specimens
with a (0)40 layup to measure fracture toughness of
carbon PEEK composite, and surmised a mode I critical
energy release rate of 26 kJ/m2. The technique used to
introduce a pre-crack in the specimen was not discussed
by the authors [9]. A center notched compression
specimen was carried out [10, 11, 12]. Many length of
notch were used to study its effect on the fracture
energy of T800/924C Carbon fiber reinforced epoxy
laminate with [0, 902, 0]3S layup the critical energy
release rate for the laminate was reported as 38:8 kJ/m2
and no effect for crack length on the fracture toughness.
Camanho et al. [13] performed series of center crack
specimen tension tests on different lay up of carbon
fiber reinforced epoxy to validate the proposed
analytical model without illustration the damage
mechanics or failure mechanics induced in these
techniques. The size effect or scaling effect which is
the reduction of nominal strength with increasing of
specimen size of carbon fiber reinforced polymer is
investigated with a lot of authors [14-16]. Camanho et
al [16] investigated the size effect of IM7-8552 carbon
epoxy Quasi-isotropic laminates of [90, 0, 45,-45]3s
stacking sequence. It is reported that there is a clear size
effect based on strength as large specimen decreases
strength, but they don‘t investigates how to overcome
this phenomena. The Glass fiber reinforced epoxy has
an importance like carbon fiber. It has application in
automobile industry, aerospace [6-16] and in Seawater
Pipe System Offshore[17]. Therefore, the fracture
properties of glass fiber reinforced epoxy laminate
should be given a considerable investigation with more
accuracy as there are very a little study which deal with
fracture energy and size effect.
The main goals of the present study is to measure
the very important fracture properties which is known
as the fracture toughness for both cross ply of [0, 90]2s
and Quasi-isotropic of [0/45/90]2s and [0/45/90/-45]s
glass fiber reinforced epoxy laminates. The size effect
is investigated for these types of materials. Also a
simple in plane shear test method will be illustrated to
measure the in-plane shear modulus. A solution for the
size effect defects has been suggested.
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
MATERIAL
AND
EXPERIMENTAL
PROCEDURES
Glass fiber reinforced laminate of [0/90]2s,
[0/45/90]2s and [0/45/90/-45]s stacking sequence using
hand layup technique [18] are used where fiber bundles
warping over a molding frames of equally step bolts
using hand layup techniques, the curing process were
on the room temperature. The material constitute
properties are shown in tables 1 and 2. The fiber
volume fracture is calculated using the ignition
technique according to BS 3691. It is found 34%. The
elastic properties and strengths of the unidirectional
lamina are measured using ASTM D3039 test standers
[19]. Five specimens are used for each test performed.
The mean measured values of the ply elastic properties
are listed in Table 3. While the in-plane shear modulus
was obtained using ±45 tensile in-plane shear test
method which will illustrated in the next paragraphs.
Table 1 the constituent materials of the composite
laminates (CMB international Co.)
Material
Type
Matrix
ResinKemapoxy(150RGL)
Reinforcement fiber
E-glass
(Alkialian)roving-pl=2200 gm/km
Table 2 Mechanical and physical properties of Eglass fiber and epoxy resin, [20,21,22]
Properties
E-glass
Kemapoxy(150RGL)
Density(kg/m2)
2540
1.07 ±0.02 kg/litres
Tensile
strength 2000
50-100
(MPa)
Tensile
modulus 76
1.2-4.5
(GPa)
Passion ratio
0.25
0.37-0.39
Table 3 Ply elastic properties
property
Longitudinal young‘s modulus,
E1(GPa)
Transverse young‘s modulus,
E2(GPa)
In-plane shear modulus, G12(GPa),
(±45 shear test)
Major passion ratio, υ12
Longitudinal strength (Xt), MPa
Transverse (Yt), MPa
Mean value
27
stress [24]. Thus, the calculated shear stress and strain
values at failure should only be used with caution.
There are several test standards/guides based on this test
method, i.e., ASTM D3518 [25].
The ±45 ° tensile specimen has the following
merits: good reproducibility, simple to make, is a
conventional tensile test, economical in material
requires, simple data reduction and is easy to test at
high or low temperatures. The cross–ply [0, 90]2s
laminate was cut at 45o to gives the ±45 tensile in-plane
shear test of stacking sequence [45,-45]2s
The quasi-static tensile tests were done in a
displacement-controlled manner with a displacement
speed of 2 mm/min, during which the force F, the
longitudinal and transverse strains, εxx and εyy were
recorded. With these values, the shear stress τ12 and
shear strain γ12 can be calculated as:
1
  F / 2wt
12
 12   xx   yy
Where w is the width of the specimen and t is the
thickness. The longitudinal and transverse strains are
measured using two perpendicular-element strain
gauges (Model, FLA-6-11 of gage factor 2.1). Fig. 2
show the digital strain meter attach with the specimen,
in the machine grippes.
Fig. 1 ±45 tensile in-plane shear test standard
specimen
5.3
1.75
0.31
645
15
±45 tensile in-plane shear test method
In this shear test method, a [±45°]2s laminate is
loaded in axial tension to determine the in-plane shear
properties. This test method is frequently used because
the specimens are easy to be fabricated and no special
test fixture is required, the specimen is shown in Fig. 1.
It is a simple test method for predicting in-plane shear
modulus with an acceptable precision [23]. However,
the laminate is not in a state of pure in-plane shear
Fig. 2 tension Specimen between the machines
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Compact specimen tension test
1. This test methods is carried out according to
ASTM E399 [26]. The specimen for fracture
toughness testing is Compact Tension. It was
machined from the laminates in accordance with the
dimension given in ASTM E399 as shown in Fig.3.
of five specimens are used to determine KIC values
(.
) as a measure of fracture toughness by
using the following data reduction scheme. According
to ASTM standard E399 [26], valid for an isotropic
material, the critical stress intensity factor for a fracture
load PQ, is given by
2
p
K 1c 
Q
h w
f (a / w )
Where h = specimen thickness, mm, W = specimen
width, mm, a = crack length, mm and f
 aw  is
shape correction factor.
 w   1  a / w 
f a
Fig. 3 Typical CT specimen with dimension
(dim. in mm)
The initial portion a V notch has to be machined
with a milling cutter or with a diamond saw and a
starter crack has to be introduced at the root of the
notch by tapping or sawing a fine razor blade [7]. The
pre-cracked fracture specimen is loaded with suitable
loading devices. For Compact tension specimen a
loading clevis is required as shown in Fig. 4, special
care is taken to create the loading holes to prevent
delimitation and damage, therefore, they were cut using
carbide tungsten drill while clamping the specimen
between two sacrificial glass/epoxy plates of the same
material. The fracture loads PQ, obtained from the tests
2  a /w
1.5
0.886  4.64(a / w3)  13.32(a

3
4
 14.72(a / w )  5.6(a / w )
where h is the thickness of the specimen, w is the
dimension from the load line to the right hand edge of
the specimen, as indicated in Fig. 3 and a is the crack
length, whose initial value ao is also indicated in Fig. 3.
The critical energy release rate of the laminate can be
calculated from KIc as [7]:
4
2
G1c  J 1c 
K 1c
2E x E y
Ey
Ex

Ey
2G xy
 xy
Where Ex, Ey, Gxy and  xy are the Young‘s moduli in
the x and y directions (see Fig. 3), the shear modulus
and the Poisson‘s ratio of the laminate, respectively,
these properties are determined experimentally in the
previous steps.
Fig. 4 one part of the clevis used in compact tension test specimen
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
The layup used is [90, 0]2S of glass fiber reinforced
epoxy laminates with the 00direction is parallel to the
loading direction as shown in Fig. 3. Five sample are
used. Figure 5 shows the test set up. Fig. 6 shows the
photograph of the compact tension test.
 a 
K 1C    a sec 

w 
5
Where (KIC) is the fracture toughness of the laminates,
(a) is semi center crack length and (W) is laminates
width. After substituting the specimen dimension in the
equation the fracture toughness is calculated, then
implemented in Eqn. 2.
Fig. 5 Compact tension set up
Fig. 6 Sketch of the compact tension test [7]
Center crack plate specimen tension test
The tension test of center crack plate specimen is
carried out according to Soutis- Flick model [27] to
measure the surface release energy of multidirectional
composite laminates. The test is performed using the
Quasi-isotropic laminates [0, -45, 90, 45]s and
[0/45/90]2s. The manufacturing technique used in CT
specimen is used. The specimen dimensions are shown
in Fig. 6. Five specimens are used. The test is simple to
be performed and can be summarized as follows:
1-Five specimen are used for tension test of the
following nominal dimension; Width W=45 mm, gauge
section length- L=90 mm, thickness- t=4 for [0, -45, 90,
45]s and t=7 for [0/45/90]2s, finally the center crack
length -2a=15 mm.(see Fig. 6).
2-After manufacturing the five specimens for each lay
out, they loaded until failure and the specimen‘s failure
load were obtained.
After measuring the failure load for each material the
fracture toughness is measures as:
Fig. 6 Centered crack specimen
un-notch tension test
The size effect needs to be compared with the un
notch nominal strength of composite laminates; (laminates without holes)-.Therefore the specimen
which is shown in Fig. 7 is prepared. The nominal
dimension is as follows: w=40, Lg= 150 and L= 250
mm, the thickness t=4 mm for 8 layers laminates or 7
mm for 12 layer laminates. Two-end tabs are attached
in order to distribute the stress along the cross section of
the specimen. Tension test is performed according to
ASTM D638M-93 [28] standard to obtain tensile
properties of plastic using five specimens and the
nominal stress strain curve is drawn. The nominal unnotch strength is average of the five specimens. This
test is performed on [0, 90]2s, [0/45/90]2s and [0, 45,
90,-45]s stacking sequence. The longitudinal
displacement measured using a strain gage (2.1 gage
factor), which is bonded on the surface of two of the
specimens.
Fig. 7 Un notch composites coupon
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Open hole tension test for similar shape specimen
The tension test of notched composite laminates is
carried out to quantify the size effect and to obtain
experimental data. The quasi-isotropic, cross ply and
unidirectional laminates are of glass fiber/epoxy of
stacking sequence [0, 90]2s, [0, -45, 90, 45]s and
[0/45/90]2s. These are also to study the effect of
stacking sequence on the size effect phenomenon. The
unidirectional specimen is not tested for size effect, as it
has no industrial application.
The presence of a stress raiser, in this case a
circular hole, leads to enhanced complex damage and
failure mechanisms, causing a wide range of effects not
present in un-notched components. A matrix of
specimens of different hole diameter and width are
shown in Table 4, but all specimens keep that hole
diameter to width ratio is constant ( d/w =1/6) [29].
Diameter
d1=2
d2=4
d3=6
d4=8
d5=10
Table 4 Experimental matrix program
Width Ratio
Number of specimen
w/d
used
12
6
5
24
6
5
36
6
5
48
6
5
60
6
5
A typical specimen of end tab used in open-hole
tensile tests is presented in Fig. 8. The displacement
will be recorded with the machine load cell, as the
fracture behavior during the test is important from
fracture
mechanics
point
of
view.
Fig. 8 Typical OHT specimen with dimension
RESULTS AND DISCUSSION
Shear test
Figure 9 shows stress-longitudinal strain curve
for [45, -45]2s tension test specimen. This test is curried
out to determine the shear strength and modulus of the
material under consideration. Defining shear strength is
still debatable as to which load value should be used
[18, 30]. Bhatanager et. al. [30] considers the first load
drop to be the shear load responsible for material
failure. Khashaba [18] defined shear strength as the
ratio of the load just prior to the nonlinear behavior, to
the cross-sectional area. Some investigators [30]
defined the in-pane shear strength as the stress value
corresponding to the ultimate load. The latter definition
for shear strength is more suitable for nominal strength
failure criteria [30]. Fig. 10 shows the relation between
the shear stress and strains measured in both
longitudinal and transverse direction. From this figure,
the relationship between the shear stress τxy and shear
strain γxy is constructed as illustrated in Fig. 11. The
values of the in-plane shear stress and strain are
calculated from Eqn.1:
The in plane shear modulus has a reasonably
acceptable value compared with the value obtained
from constituent material and volume fracture based on
lamination theory [32] as 2 GPa. With 12.5 % error
which is accepted from scientific and industrial point of
view [31]. Shear stress-strain curve has proportional
behaviors in the beginning. Just beyond proportional
limit, it becomes nonlinear due to the accumulation of
matrix cracks. The specimens tend to deform nearly
into ‗dog bone‘ like shape Fig.12. Furthermore, it must
be remarked that the narrowing of the specimen to the
‗dog bone‘ like shape does not happen in a uniform
manner over the entire specimen, but starts near the
clamped ends and then gradually grows along the entire
specimen length. A more severe delamination is
observed extending away from the fiber breaks.
The longitudinal and transverse strains are
measured using two perpendicular-element strain
gauges (Model, FLA-6-11 of gage factor 2.1). Hence,
the in-plane shear modulus has been determined from
the slope of the shear stress-strain diagram at 0.5% as:
G12 
shear stress
shear strain

8.75
100  1750MPa  1.75GPa
0.5
Fig. 9 Tensile stress longitudinal strain
curve for [45,-45]2s specimen
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
there is a tendency for the crack depth to vary through
the thickness. Substituent these results in Eqn.2, the
average fracture toughness K1c is measured as;
with
stander
deviation
24.098 MPa. m ,
1.82 MPa. m . Substituent this value in Eqn. 4,
fracture energy release rate G1c is measured as 51.915
kj/m2 with the stander deviation is 7.523kj/m2.
Fig. 15 shows the post failure picture of the CT
specimen, it is observed that the fiber bridging the two
face of cracks and the crack advances straightly through
the pre-crack direction, due to the highly stress intensity
factor induced at the crack tip.
Fig. 10 Shear stress longitudinal and
transverse strain curve for [45,-45]2s specimen
Fig.13 Typical load verse displacement curve
Fig. 11 Equivalent shear stress verses shear
strain for [45,-45]2s specimen
Fig. 12 Failure mode of [45, -45]2s Shear test
Compact tension
For the Compact Tension specimen (CT) test,
crack growth is neither smooth nor continuous: instead,
several crack jumps of a few millimeters each time
were observed, Fig.13. The fracture loads PQ, obtained
from the tests of five specimens and according eqns.2, 3
and 4 respects to the laminates elastic properties which
are listed in Table 3 the fracture toughness KIC values
(.
) can be calculated. The average load value
at 5% secant from the Fig. 13 equal 1900N, with the
specimen dimension and total crack length (a0+afpz),
where aFPZ is average fracture processing zone length
and is shown in Fig. 14 and measured experimentally as
approximately 3.5 mm. The average value as since
Fig. 14 the identification of the crack tip
just before Maximum load
Fig. 15 Failure Mode of CT specimen
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Center crack specimen
Soutis and Flek [27] showed that the fracture
toughness of Quasi-isotropic laminates is independent
of the center-crack size. Therefore, number of
specimens that need to be tested is decreasing (only one
length of the center crack was used). Figs. 16 and 17
show the load displacement curve for the test of center
crack specimen, it is observed that the curve is smooth,
a little jump seen, the extension records for laminate of
stacking sequence [0/45/90/-45]s is less than that of
[0/45/90]2s this return to the -45o plies which increase
the shear stress action. The curve obtained from a
compact specimen has a longer length compared to one
obtained from a center-cracked specimen, because the
gradient of K (stress intensity factor) in a compact
specimen is decreasing whereas the gradient of K in a
center-cracked specimen is increasing [27]. Fig. 18
shows the post-failure picture of one specimen for each
stacking sequence of specimen. The post–failure load is
determined from the load displacement curve for each
laminates and with the help of Eqn.6, the failure stress
is calculated and fracture toughness which are listed in
tables5 and 6. After measuring the failure stress for
each specimen, the fracture toughness is determined by
using the real dimension of the specimen:
6
p
 X  max
started at the ends of the slits, growing parallel to the 0o
plies, through the specimen thickness and towards the
gripping, 2) crack in the 45o fibers originating at slits
45o in addition, some cracking were observed between
0o plies and adjacent 45o plies. There is a more severe
delamination was observed extending away from the
fiber breaks. It is observed that the fiber mode of failure
is tension mode with inclination angle about 45o, for
angle ply and 90 for 0o ply. Focus look for Figs. 16, 17
it is shown that there is knee occurred at about 5kN this
is because first 90o ply failure occurs.
The stacking sequence has a visible effect on
notched failure strength and failure modes as an
increase of angle plies -450 for laminates results in
strength reduction approximately 7% as this angle
layers introduce shear stress in both hole sides lift and
right as shown in failure post image Fig.18
A
Where: σ = tensile strength, MPa,
Pmax = maximum load prior to failure, N,
A = cross-sectional area, m2.
Then Substituting in Eqn. 5 the surface release
energy can be calculated. The average value of the
fracture toughness K1C for [0/45/90]2s and [0/45/90/45]s is
19.88 MPa. m with Stander deviation
1.078 MPa. m and
19.066 MPa. m
Fig. 16 load displacement curve for (0, 45,
90)2s
with
stander deviation 1.2310 MPa. m
repetitively.
Whereas, the fracture energies for these laminates are
35.48 Kj/m2 with Stander deviation 3.763 Kj/m2 for
[0/45/90]2s, while for [0/45/90/-45]s is 32.57 kj/m2 with
stander deviation 2.417 Kj/m2. Fig. 18 (a) shows image
of post failure surface of the center crack specimen of
[0/45/90]2s. It can be reported that it is observed that the
crack propagates through the notch corner and
advanced approximately direct to the loading direction
as the eight layer of 0o plies and the 90o plies are of
highest stiffness more 45o plies which is the main
reason for the shear band appearing in the fracture
surface of Fig. 18 (b). It is appeared that matrix
damage to take two forms. 1) crack in the 0o plies which
Fig. 17 load displacement curve for (0, 45,
90,-45)s
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Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Fig.18 Failure modes for the centered crack composite laminates a)[0/45/90/-45]s, b)[0/45/90]2s
Specimen number
1
2
3
4
5
Table 5 Fracture toughness test results
(0,45,90)2s
(0,45,90,-45)s
Stress (MPa)
Stress (MPa)
95.23
91.77
89.77
86.33
83.14
83.77
92.88
83.3
93.65
82.6
Table 6 Fracture Toughness and energy release rate
Specimen types Fracture Toughness, K ( MPa. m ) Fracture energy , GIC (KJ/m)
[0/45/90]2s
19.88
35.48
[0/45/90/-45]s
19.066
32.57
Un-notch Tension Test
Figure (20 a) shows the stress-strain diagram, of
cross-ply glass fiber reinforced epoxy (GFRE)
specimen, printed from the PC of testing machine in
tension test as the machine draw the relation between
load and displacement in time. The main characteristic
of this curve is the knee at about 35 MPa. This knee
was due to the failure of the transverse layers (90 o) in
the cross-ply laminates. Redistribution of stress
between longitudinal fibers and matrix was occurred
leading to another increasing in the tensile stresses with
apparent tensile modulus lower than that for the initial
linear portion. The final failure is catastrophic without
any yielding.
Figure (20 b) show the stress-strain diagram
obtained from the strain gage attached to the testing
specimen. The actual modulus of elasticity was
calculated as 20.3 GPa at 0.5 % strain. This value is
agreed with the theoretical results 21 GPa obtained
using Micromechanical theories of lamina [32]. The unnotch average strength of five the cross ply laminates
was calculated as 167 MPa with stander deviation equal
23.9 MPa. The failure mode is net tension with a little
warping due to moment occurs, results from not perfect
axially of fiber in 0o layer, also the failure mode tends
to be near the gripping region in some specimen due to
increase of stress concentration factor just above this
region, matrix cracking with delamination near the
fracture region through thickness.
The stacking sequences plays important role in the
anisotropic material strength like this composite
laminates. as the average strength for laminates of
[0/45/90]2s is 94.75 MPa with stander deviation 6.5
MPa and it is decreases than [0/90]2s by about 43.3%
20
Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
as shown in Fig. 21 and are listed in Table 7, this results
is agree with the published results for other material
laminated structure [33]. This can be attributed to that,
the angle plies insert in the laminates produce shear
stress across the fiber direction, which make a stat of
complex stresses. Whereas the strength reduction not
affected a lot with the addition of the -45 angle plies as
the average strength of [0/45/90/-45]s is 98.5 MPa with
stander deviation 4.7 MPa. The shear stress in the
direction of fiber is the main reason for strength
reduction as clearly appear in the failure mode of
[0/90]2s which is due to pure tension and cracks are
perpendicular to the loading direction (Matrix
cracking), while for angle ply it is propagated along the
fiber direction. It is clear from Table 7 and Figs. 21 and
22 that the brittleness decreases with insertion of angle
ply for the laminates stacking sequence structure; this
can be attributed to that, the laminates get some
isotropy by inserting these angle to the material. Fig. 23
shows A more severe delamination was observed
extending away from the fiber breaks, but no damage
was visible during the tests. These results are by
looking carefully to the failure region and noticing that
the Surface 45◦ plies were completely delaminated. The
edges of the specimens were also examined away from
the fracture location and clear evidence of matrix
cracking and delamination of the surface ply was found,
as shown in Figure 20 b(b) for the smallest specimen,
which is occurred before fiber failure. The specimens
presented different behavior: delamination occurred
sooner and for more plies as the thickness increased.
Failure was located near the tab ends with part of the
material ejected, which suggests that stress
concentration may be partly responsible for the failure.
Failure combined also longitudinal splitting and fiber
breakage.
Stiffness of the composite laminates is affected by
stacking sequence as clearly shown in Figs. (20-b, 21
and 22) and values of Young‘s modulus which are
listed in Table 7. Reduction of 46 % is occurred in
stiffness of same number of layer from [0, 90]2s to
[0/45/90/-45]2s. This reduction in modulus is attributed
to yielding of the matrix in the 90º plies. This data
indicates that laminate stiffness depends on the stiffness
of the reinforcing fibers as well as the percentage of
fibers aligned in the direction of loading, i.e. the stiffer
laminates have a greater percentage of plies aligned
with the direction of loading, as expected. Although, the
laminates of 12 layer of stacking sequence [0/45/90]2s
have an equally aligned plies to [0, 90]2s but decreases
by about 30 % stiffness. This can be attributed to the
increasing thickness in this laminates (7-9) mm for
5mm for cross ply laminates this increase in thickness
induce stresses in this direction and the specimen can be
in sate of tri-axial stress which reduced both stress and
strength [34].
Fig.20 Stress strain relation for [0, 90]2s a) Apparent strain b) Actual strain
21
Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Fig. 21 Effect of stacking sequence on un notch nominal strength
Fig. 22 Stress actual strains relation for a) [0/45/90]2s b) [0, 45, 90,-45]s stacking sequence
Stacking sequence
[0/90]2s
[0/45/90]2s
[0/45/90/-45]s
Table 7 Un notch strength of composite laminates
Mean nominal strength (MPa) STDV (MPa) Young‘s modulus (GPa) at 0.5 % strain
167.5
23.9
20.3
94.75
6.5
15
98.5
4.7
12
Open hole Tension Test
The experimental results presented in Table 8
and Fig. 24 clearly identify a specimen size effect: an
increase in the hole diameter from 2 mm to 10 mm
results in 29.5% a average reduction in the strength.
The observed size effect is caused by the development
of the fracture process zone, which redistributes the
stresses and dissipates energy. In small specimens, the
fracture process zone extends towards the edges of the
specimen and the average stress at the fracture plane
tends to the un-notched strength of the laminate. Size
effect test results are listed in Table 8, the increase of
size lead to increase of brittleness. The stress
concentration at the hole is completely blunted. For this
material it is, observed that the specimen failed
suddenly without visible damage due to the low
interfacial toughness.
Figures 25 and 26 show the stress strain
diagram for [0/45/90]2s and [0/45/90/-45]s composite
laminates. It is show the same results that there is a
specimen size effect appears for Quasi-isotropic
laminates, but the size effect can be reduced by
increasing numbers of 45o ply in the laminates stacking
sequence with reducing the plate thickness. The 45 o ply
introduces shear stress around the hole that in role
reduced the stress concentration factor and stress
intensity factor. These angle ply make like stress
releaser factor [7, 18]. Tables 9 and 10 show these
results.
22
Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Table 8 Test results for [0, 90]2s laminates open hole
:
diameter Mean nominal strength STDV
(MPa)
(MPa)
2
135
4.12
4
121
3
6
116.25
5.8
8
105
3
10
95.5
1.11
Table 9 Test results for [0/45/90]2s laminates open
hole
diameter Mean nominal strength STDV
(MPa)
(MPa)
2
86.25
3.3448
4
82.5
3.27
6
79.25
1.92
8
71
1.58
10
65.2
0.8672
Delamination occurs in combination with splitting at
the notch, relieving the stress concentration [33].
Damage in the GFRP [0/90]2s laminate is
characterized by axial splits at each side on the hole in
the 0° plies. As the load increases, the damage zone
increases in size, but the matrix cracks in the 90° plies
only occur outside (i.e. towards the laminate edge) of
the 0° ply axial splits indicating that the splits
effectively blunt the stress concentration from the hole.
The apparently lower level of matrix cracks in the 90°
plies in the GFRP OHT specimens may be due to the
less brittle nature of the matrix material. The failure
mode changed from fiber-dominated to matrixdominated with decreasing hole diameter, and that
change was accompanied by an increase in
delamination and much less change in strength with
hole size than expected on the basis of the WhitneyNuismer or Mar-Lin hole size models. It is attributed to
change of the failure mode to the interlaminar stresses
in the region around the hole boundary, which decrease
with increasing ratio of hole radius to laminate
thickness.
Table 10 Test results for [0/45/90/-45]s laminates
open hole
diameter Mean nominal strength STDV
(MPa)
(MPa)
2
68.84
1.29
4
67.35
3.96
6
67.27
4.755
8
62.3
2.87
10
57.75
3.25
The failure mode observed in all specimens is
net-section tension as shown in Figs. 27,28 and 29.
Damage mostly consists of matrix cracking in the ±45°
and 90° plies in the vicinity of the hole, accompanied
by some delamination around the hole; cracks in the 90°
plies extend to the laminate edge. However, large
triangular delamination zones emanate from the hole,
decoupling plies and leading to rapid failure of the
laminate. The shear band appears on surface inclined at
45o for GFRP [0/45/90/-45] at both side of hole, anther
words in direction of 450 fibers. Carefully look for
crack propagation it is found that the crack path change
direction a little before reach the specimen end. this
because this location near the ends of 45o fiber,
therefore fiber matrix debonding failure occurs and the
crack propagates directly perpendicular to the load
direction (Matrix cracking) as the inclined fiber cannot
at this location resist the crack.
As the thickness increase a more severe
delamination appeared in the structure which has
dangerous damage for small size (small width respect to
thickness) Fig.30 shows the delamination through
thickness of [0/45/90]2s specimen thicker laminates.
Fig. 24Stress strain diagram for [0, 90]2s size effect
specimen
Fig. 25 Stress strain diagram for [0/45/90]2s size
effect specimen
23
Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Fig. 26 Stress strain diagram for [0/45/90/-45]s size effect specimen
Fig. 27 Post failure image of [0/90]2s
Fig. 28 Post failure image of [0/45/90]2s
24
Mohammed Y et al., Sch. J. Eng. Tech., 2013; 1(1):13-26
Fig. 29 Post failure image of [0, 45, 90,-45]s
Fig. 30 through thickness delamination
Conclusion
The fracture properties of glass fiber composites
laminates are measured and following conclusion are
summarized:
 Compact tension test specimen is simple and give
acceptable results for fiber tension fracture
toughness G1c of laminates of stacking sequence
[0/90]2s as 51.915 kj/m2 which agreement with the
allowable value in text book. However Center
cracked plate specimen is suitable for measuring
fracture toughness for Quasi-isotropic laminates
[0/45/90] 2s and [0/45/90/-45] switch have a value
of 32.98 and 31.5 KJ/m2 respectively.
 A strength reduction of 32 % is observed with
increasing the hole diameter from 2 mm to 10 mm,
while this percentage was decreasing by inserting
an angle ply as 26 % for [0/45/90]2s and 14 % for
[0/45/90/-45]s.
 Delaminations are observed with thickness
increasing for un-notched specimens.
 Fiber orientation affects deeply the laminates
carrying capacity.
 Three failure modes are summarized for glass fiber
composite laminates, fiber pull out, fiber tension
and delamination.
 ±45 in plane shear test give acceptable results for
shear modulus while underestimates the shear
strength value by about 50% when compared with
the value measured with the lamination theories.
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