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FAILURE ANALYSIS OF COMPOSITE LAMINATES WITH
A HOLE BY USING FINITE ELEMENT METHOD
by
MOUMITA ROY
Presented to the Faculty of the Graduate School of
The University of Texas at Arlington in Partial Fulfillment
of the Requirements
for the Degree of
MASTER OF SCIENCE IN MECHANICAL ENGINEERING
THE UNIVERSITY OF TEXAS AT ARLINGTON
December 2005
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To my parents, Mrs.Mamata Roy and Mr.Sadhan Chandra Roy, who have
made me what I am today.
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ACKNOWLEDGMENTS
This thesis could not have been written without Dr.Wen S.Chan, my supervising
professor. All his teachings and his open door policy have been of great benefit to
me throughout my term of study at UTA. My sincerest gratitude goes to him for
suggesting the subject of research, and for his valuable suggestions and
encouragement throughout the work and his guidelines for publishing this thesis. In
addition, the author would also like to thank Dr.Bo P. Wang and Dr. Seiichi Nomura,
for serving on her committee. I want to express my loving appreciation to my parents
and my brother for their support and encouragement throughout my life. I would also
like to thank my friend(s) here at UTA and elsewhere who were always a constant
source of inspiration and encouragement.
October 26, 2005
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ABSTRACT
FAILURE ANALYSIS OF COMPOSITE LAMINATES WITH
A HOLE BY USING FINITE ELEMENT METHOD
Publication No . ______.
Moumita Roy, M.S.
The University of Texas at Arlington, 2005
Supervising Professor: Wen S.Chan
Predicting the strength of a given laminate has been extensively studied.
There have been several fracture models proposed but still the immense tests of
composite laminates are still needed. Different material systems or lay-ups exhibit
different failure modes and failure mechanisms. This thesis focuses in investigation
of the failure mechanism of the [/02]s and [02/]s laminates with a hole. ANSYS
finite element models with and without a crack in vicinity of hole were developed to
investigate the effect of the stress distribution due to the presence of the angle ply
crack. The stress concentrations were obtained. It is found that
The stress concentrations increases as d/w increases for a given in [/02]s
and [02/]s laminates
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For a given d/w ratio, the stress concentrations increases as ply orientation,
increases
The high stress concentrations on 0o ply occurs at the location where the -
ply crack intercepts at the line perpendicular to the 0o
ply
Based upon the understanding the failure mechanism, a simple expression to
estimate the strength of laminates with a hole is established. A strength model
based upon the 0o
ply load carrying capability is proposed. Prediction is a good
agreement with the experimental results
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TABLE OF CONTENTS
ACKNOWLEDGEMENTSiii
ABSTRACTiv
LIST OF ILLUSTRATIONSviii
LIST OF TABLES....x
Chapter
1. INTRODUCTION....1
1.1 Applications of Composites in Aerospace Structures .1
1.2 Hole in Composite Structures ..2
1.3 The Objective of this Thesis.3
1.4 Outline of Thesis.4
2.FAILURE OF LAMINATE WITH A HOLE...5
2.1Introduction.....5
2.2Experimental Study.6
2.2.1 Failure Mode.6
2.2.2 Notched and Un-notched Strengths...8
2.3Predictive Model of Notched Strength9
2.3.1 Waddoups-Eisenmann-Kaminski (WEK) Model [6].........10
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2.3.2 Whitney-Nuismer (WN) Model [7]....12
2.3.3 Mar-Lin (ML) Criterion..13
2.4 Stress Analysis of Laminates with a hole..14
3.FINITE ELEMENT MODELING...19
3.1 Definition of Problem.19
3.2 Element Type and Size...20
3.3 Modeling and Meshing...22
4.RESULTS AND DISCUSSIONS...28
4.1 Model Validation28
4.2 [02/ ]s composite laminate....29
4.2.1 Composite plate with crack at + layer..30
4.2.2 Composite plate with crack at - layer...32
4.3 [
/ 02]s Composite laminate.36
4.3.1 Composite plate with crack at + layer..36
4.3.2 Composite plate with crack at - layer...38
4.4 Failure Strength Prediction Using Classical Lamination Theory.38
4.5 Conclusion.48
REFERENCES...50
BIOGRAPHICAL INFORMATION..52
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LIST OF ILLUSTRATIONS
Figure Page
2.1 Failure of Hole in laminate with 62% of 0o
ply.[Ref.4].6
2.2 Failure of Hole in laminate with 25% of 0o
ply.[Ref.4].....6
2.3 X-radiographs showing damages around hole of laminates with various d/w ..7
2.4 X-radiographs of fiber breakage of 0o
ply in the [02/]S laminates.....8
2.5 Effect of width on the strength.......9
2.6 Strength variations due to fiber orientation........9
2.7 WEK fracture model....11
2.8 (a) Schematic representation of the point-stress criterion, (b)Schematic
representation of the average-stress criterion, for a laminate containing a circular
hole................................................................................................................13
2.9 Axial stress distributions in a laminate with d/w = 0.85 subjected to uniform tensile
load....15
2.10 Axial stress distributions of notched [02/]sb laminates with various fiberorientations.16
2.11 The location of the peak SCFs along the laminate straight edge....16
2.12 In-plane shear stress, the inter-laminar normal and shear stresses distribution at the
0/+45 interface along the y-axis...17
2.13 The interlaminar stresses along the hole edge at the interfaces of 0/+45 (z = 2h)plies and +45/-45 (z = 1h) plies.18
3.1 SOLID 191 Geometry...21
3.2 The Element output definition notation.21
3.3 TARGE170 Geometry...22
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3.4 CONTA174 Geometry...22
3.5 Full and Sub Models with a crack..23
3.6 3-d modeling in which + crack is present in +layer for [,02]s laminate....24
3.7 Meshed volume & with applied boundary conditions...26
3.8 Meshed volume with applied boundary conditions for [02,90]s laminate....27
4-1 Aluminum plate with a hole under uniaxial loading.....28
4.2 Composite laminate under axial loading....29
4.3 Quarter model of composite plate with hole for d/w=0.5...30
4.4 Normalized stress distribution along the direction in the 0o layer when crack ispresent at + layer31
4.5 Normalized stress distribution with varying d/w ratio for [02/60]s ..33
4.6 Normalized stress distribution for [ 02/45]s for varying d/w ratio...33
4.7 Stress distribution along the y-axis in the zero degree layer when crack is present at -
layer....34
4.8 Stress distribution in 0 degree layer when crack is present at + layer for[ / 02]s laminate...36
4.9 Notched Strength in [02/ ]s laminate when crack is present in layer in therespective direction....42
4.10 Notched Strength in [ /02 ]s laminate when crack is present in layer in therespective direction.....45
4.11 Comparison of Notched Strength between [02/ ]s & [ /02 ]s when crack ispresent in layer next to 0 degree ....45
4.12 Comparison of Notched Strength between [02/ ]s & [ /02 ]s when crack is presentaway from 0 degree.48
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LIST OF TABLES
Table Page
3.1 Material Constants Used..19
4.1 Comparison of Calculated values to ANSYS values...30
4.2 Max. Stress in 0o layer with varying d/w ratio for [02/ ]slaminate for + crack ..35
4.3 Max. Stress in 0o layer with varying d/w ratio for [02/ ]slaminate for - crack....37
4.4 Max. Stress in 0o layer with varying d/w ratio for [ / 02]slaminate for + crack...40
4.5 Max. Stress in 0o layer with varying d/w ratio for [ / 02]slaminate for - crack....41
4.6 Notched Strength of [02/]s when crack is present at + directionin + layer....43
4.7 Notched Strength of [02/]s when crack is present at - direction
in - layer.44
4.8 Notched Strength in [ /02]s laminate when crack is present in layer..46
4.9 Notched Strength in [ /02 ]s laminate when crack is present in layer..................47
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CHAPTER 1
INTRODUCTION
A material containing two or more distinct constituents in a macroscopic scale is
called a composite material. These distinct constituents have significantly different properties
and composite properties are noticeably different from constituent properties. One of the
constituents is used to reinforce the other constituent(s). In case of unidirectional
continuously fiber-reinforced composites single fiber orientation in a layer but different
layers in a laminate may have different single fiber orientation in a given layer. This kind of
composite laminates reinforces the stiffness or strength along the fiber orientation. In this
thesis this type of composite laminate has been studied.
1.1 Applications of Composites in Aerospace Structures
Composite materials are one such class of materials that play a significant role in
current and future aerospace components. Composite materials are particularly attractive to
aviation and aerospace applications because of their exceptional strength and stiffness-to-
density ratios and superior physical properties. Among the first uses of modern composite
materials was about 30 years ago when boron-reinforced epoxy composite was used for the
skins of the empennages of the U.S. F14 and F15 fighters. Initially, composite materials were
used only in secondary structures, but as knowledge and development of the materials has
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improved, their use in primary structures such as wings and fuselages has increased.
Composite materials are used extensively in the Euro fighter: the wing skins, forward
fuselage, flaperons and rudder all make use of composites. The use of composite materials in
commercial transport aircraft is attractive because reduced airframe weight enables better
fuel economy and therefore lowers operating costs. The first significant use of composite
material in a commercial aircraft was by Airbus in 1983 in the rudder of the A300 and A310,
and then in 1985 in the vertical tail fin. In the latter case, the 2,000 parts (excluding fasteners)
of the metal fin was reduced to fewer than 100 for the composite fin, lowering its weight and
production cost. Composites also has deep-water applications like those in hydraulic cylinder
for underwater operations, small diameter low pressure pipes in composite materials, etc.
Due to the superior specific strength and stiffness characteristics and their adaptability to
given structural needs composite laminate materials are not only widespread in the fields of
aircraft and spacecraft constructions but also gain more spread in the classical fields of
mechanical engineering.
1.2 Hole in Composite Structures
Most structures are assembled by a number of individual structural elements connected to
form a load path. These connections or joints, in general, can be classified as adhesively
bonded, mechanical fastened (bolted or riveted) or combination of both. Mechanical joints
usually have some type of bolt holding the different pieces of the structure, which needs a
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hole to be drilled into the composite laminate. Besides this, hole in structural laminate can be
served as access area of any purpose. Predicting the notched strength of a given laminate has
been extensively studied in the last three decades. Most of both analytical solutions and exper-
mental investigations were concentrated for wide laminates with a hole. A thorough review
of strength prediction models has been conducted by Awerbuch and Madhukar [1]. A
systematic study of hole problems in laminate was given by Tan [2]. The damage
characteristics around hole for laminates with different percentage of 0 ply was
experimentally investigated by Chan [4]. The notched strength of the [02/]S laminates was
studied by Harn [5]. Numerous failure criteria and fracture models have been proposed by
Waddoups et al. [6], Whitney and Nuismer [7] . However, little work on laminates with a
large hole size was conducted.
1.3 The Objective of this Thesis
The objective of this research is to investigate the stress concentration around the hole
with various ratios of the hole size to the laminate width. The finite element analyses are then
conducted to aid understanding of the failure process of the laminate. The test observations
of failure process of laminate with a hole reveal the final failure of laminates occurring in the
00ply. With this in mind, a fracture model based upon the 0 ply load carrying capability is
proposed for predicting the notched strength.
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1.4 Outline of Thesis
A brief discussion on various works done in the field of composite laminate with a hole
is given in chapter two. In chapter three, modeling and meshing using ANSYS has been
described. In chapter four the results have been discussed and the conclusions are drawn
based on the results.
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CHAPTER 2
FAILURE OF LAMINATE WITH A HOLE
2.1 Introduction
Failure of structures with a hole is caused by excess stress/strain in the neighborhood
of the hole. The stress or strain concentration tangential to the hole edge has been
investigated extensively for isotropic materials for a century. The equations of the state of
stress have been developed for both infinite- and finite-width isotropic plates. The solution
for the finite width is often used the solution for infinite width of the plate by multiplying a
factor. This factor is termed as finite width correction factor. The finite width correction
(FWC) factors are function of d/w but not the material properties.
The analytical solutions to the stress or strain concentration for isotropic materials
with infinite plate size have been well-established couples of decade ago. The FWC factors
of isotropic material are usually applied to the infinite-geometry anisotropic or orthotropic
materials to obtain the finite-geometry data reduction. Recently, the FWC factors for
anisotropic plates have been reviewed in Ref[2]. It is a function of d/w and material
properties. However, the exact solution to the axial stress distribution in a finite-width
composite laminate with a hole has not been found yet.
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2.2 Experimental Study
2.2.1 Failure Mode
Chan[4] studied the damage characteristics around hole of laminates with various
percentages of 0o-ply. He found that the damage process of the hole is confined in areas
shown in figure 2.1 for laminates with 62% of 0o
ply.
Figure 2.1 Failure of Hole in laminate with 62% of 0o
ply.[Ref.4]
For laminates with 25% of 0o
ply the hole damage grows from the edge of the hole to the
edge of laminates as shown in Fig 2.2.
Figure 2.2 Failure of Hole in laminate with 25% of 0o
ply.[Ref.4]
(62/31/7)%S 90/0/45/0/45 22
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In study of [02/]s laminates, Harn[5] observed the fiber splitting of 0o
plies and the
delamination at the interfaces of 0o/+45
oand +45
o/-45
o. Figure 2.3 shows intermediate and
final failures of the [02/45]S laminates with different d/w ratios. As shown, the fiber splitting
within 00plies is due to the steep strain and stress gradients. Delamination was first observed
within the regions where the material is not continuous.
Figure 2.3 X-radiographs showing damages around hole of laminates with various d/w
The extensive delamination starts from the edge of the hole and propagates rapidly toward
the edge of the specimen. The laminate with a smaller hole size exhibits more extensive
delamination compared to laminate with a large size of the hole. At the final failure, the
breakage of 00ply is clearly shown along the direction of 45
0, which is the fiber orientation
of the neighboring ply. The phenomenon of the fiber breakage of the 00ply was also
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observed along the direction of the neighboring ply in other laminates, such as [02/15]S,
[02/30]S and [02/60]S as shown in Figure 2.4
Figure 2.4 X-radiographs of fiber breakage of 0o
ply in the [02/]S laminates
2.2.2 Notched and Un-notched Strengths
Harn [5] studied the effect of notched strength due to d/w ratio and ply fiber orientation.
Figure 2.5 shows the ratio of ultimate failure and onset of delamination strengths to the
unnotched strength of the [02/450]S laminate with various d/w ratios. Figure 2.6 shows It is
observed that both notched strength and delamination strength decrease as the hole size
increases. It is also shown that the difference between two strength ratios is smaller as the d/w
ratio increases. The strength variation due to fiber orientation of the laminate is shown in Fig2.6.
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Figure2.5 Effect of width on the strength
Figure 2.6 Strength variations due to fiber orientation
As shown, except laminate with=300, both notched and unnotched strengths decrease as
increases. The higher value is, the lower the delamination strength is.2.3 Predictive Model of Notched Strength
The notched strength of a composite laminate depends on laminate configuration, laminate
stacking sequence, hole size, and laminate width. Different lay-up or material system may
exhibit different failure mechanism. A brief review of fracture prediction models is described
as following:
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2.3.1 Waddoups-Eisenmann-Kaminski (WEK) Model [6]
This model is based on Mode I fracture analysis for a laminate containing either a
hole or a line crack. This model assumes that there exists a small but finite region of intense
energy at the edges of the hole or crack in the direction transverse the to the loading
direction. Figure 2.7 shows the WEK fracture model. These intense energy regions were
intuitively seen as the damage region ahead of the hole or crack in the laminate. Then the
model further assumes that failure strength of the laminate will occur at the vicinity of the
crack.
The strength of a laminate with no hole can be obtained from Eq. (2-1) by setting R
equal to zero,
)/( Raaf
KICN
= (2-1)
The ratio of notched strength for infinite-width laminate, N and unnotched, o is
given as
)/(
1
Raf eo
N =
(2-2)
where the functionf(ae/R) is tabled by Paris and Sih[10] ae is a parameter that is determined
from a set of coupon tests. The assumptions of the existence of an intense energy region of
length ac results in the following equation at failure.
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Figure 2.7 WEK fracture model
[ ] 2/1)( cNIC acK += (2-3)
For the case of unnotched laminates, the strength of the laminate can be determined
by letting c equal to zero. Combining these two equations yields2/1
+=
c
c
o
N
ac
a
(2-4)
This model involves two parameters, the unnotched strength o and the characteristics
length ac to be determined. It should be noted that ac was assumed to be independent of the
original crack length and thus considered a material parameter. The concept of the intense
energy zone region is a fractious region treated the laminate as a homogeneous material.
IntenseEnergy
Region
a
R
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2.3.2 Whitney-Nuismer (WN) Model [7]
The model hypothesized that the strength of the laminate with a hole can be evaluated
at a characteristics distance based on the point or average stress across the region from the
edge of the hole. Both criterions assume that fracture occurs when the stress at some
characteristics distance from the edges of the hole or the crack tip reaches the unnotched
strength.
Point-Stress Criterion.
In point stress criterion, it is assumed that failure occurs when the stress y at some
distance do away from the edges of the discontinuity is equal to or greater than the strength of
unnotched laminates. Figure 2.8(a) shows this criterion schematically. The ratio of the
notched to unnotched strength is given as
[ ])75)(3(322
8
1
6
1
3
1
2
1
++=
To
N
K(2-5)
whereodR
R
+=1 (2-6)
The appropriate value of do can be determined from the test data.
Average Stress Criterion.
The average stress failure criterion states that failure will occur if the average value of
the stress y over a distance ao is equal to the unnotched strength. Fig 2.8(b) shows this case.
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x x
Figure 2.8 (a) Schematic representation of the point-stress criterion, (b)Schematic
representation of the average-stress criterion, for a laminate containing a circular hole
dxxa
oaR
R
y
o
o )0,(1
+
= (2-7)
))(3(2
)1(28
2
6
2
4
2
2
2
2
+=
To
N
K(2-8)
where
oaR
R
+=2 (2-9)
2.3.3 Mar-Lin (ML) Criterion
This model is based on the classic LEFM. Mar and Lin proposed that the fracture of
laminates is governed by
(a)
Y
y
o
do
Y
y
o
ao
R R
(b)
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n
cN cH = )2( (2-10)
where c can be either the hole radius or half of the crack length and H c is the laminate
fracture toughness similar to KC or KIC. The exponent n is the order of the singularity of a
crack with its tip at the interface of two different materials. The value ofn is a function of the
ratio of the shear modulus and Poissons ratio of the matrix and fiber.
2.4 Stress Analysis of Laminates with a hole
This paper investigates the failure mechanism and notched strength of the [02/]S
laminates with hole size to width ratio greater than 0.5. It is found that the stress
concentration increases as the hole size increases for a given in the laminate. For a given
d/w ratio, increasing the stress concentration as the ply orientation decreases. The
experimental investigation reveals that delamination damage occurs around the hole is more
extensive for a small hole than for a large hole in the laminates, resulting in a higher notched
strength. For d/w = 0.5, damage around the hole for laminate with = 15o is isolated in the
area where the material is continuous. How-ever, for greater than 15 extensive
delamination occurs around the hole region. A fracture model based upon the 0 ply load
carrying capability is proposed for predicting the notched strength.
The typical axial stress distributions along boundaries and across the laminate width
in the two dimensional analyses are plotted in Fig. 2-9.
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Figure 2.9 Axial stress distributions in laminate with d/w = 0.85 subjected to uniform tensileload
In case of two-dimensional stress distributions the stress reaches the peak value along
the plate straight edge about one radius away from the hole center. The stress drops below the
applied tensile load at point A, the intersection of the y-axis and plate straight edge. The
highest stress concentration is located at point B, the junction of the y-axis and the hole edge.
The stress along the centerline starts from 0 at the hole edge, then drops, and increases to the
far field stress. The stress recovers to the applied stress about 4 times of the radius away from
the hole edge. The recovery of the far field stress along the plate centerline is closer to the
hole edge as the hole size decreases. The existence of a hole affects the stress distributions
tremendously. The axial stress distribution along the y-axis and along the plate straight edge
are plotted in Fig. 2.10
It is indicated that the peak stress concentration along the straight boundary shifts
toward the plate net cross-section as is increased. The location of the peak SCFs is
observed at the place where the continuous fibers tangential to the hole edge reach the edge of
the plate as shown in Fig. 2.11.
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Figure 2.10 Axial stress distributions of notched [02/]sb
laminates with various fiber
orientations
Figure 2.11 The location of the peak SCFs along the laminate straight edge
In case of three-dimensional stress distribution The distributions of the in-plane shear
stress, the inter-laminar normal and shear stresses at the 0/ +45 (z = 2h) interface along the
x-axis are shown in Fig. 2.12.The interlaminar stresses near the curved edge are prominent
than those near the straight edge. The distributions of interlaminar normal and shear stresses
are affected by the d/w. The in-plane shear stress,xy is zero on the plate straight edge, but
finite elsewhere except on the hole edge. For every d/w ratios, the value of the stresses
reaches the maximum near the hole boundary.
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Figure 2.12 In-plane shear stress, the inter-laminar normal and shear stresses distribution at
the 0/+45 interface along the y-axis
The sign of the interlaminar stresses also change the sign from the hole edge to the laminate
straight edge. As the hole size increases, the interlaminar normal and shear stresses become
more significant. The interlaminar stresses exist everywhere along the y-axis in the 0/+45
interface as d/w increases. The interlaminar stresses along the hole edge at the interfaces of
0/+45 (z = 2h) plies and +45/-45 (z = 1h) plies are plotted in Fig. 2.13. Among these
interlaminar stresses, is the dominant stress at z z=1h, 2h along the hole edge. The figure
also indicates that the peak stress ofz occurs at = 80 and 100 at the 0/+45 interface.
This suggests that the initiation of the delamination may occur at these locations.
The effect of angle ply orientation is that the unnotched strength decreases as is
greater than 30. However, the notched strength reaches a fairly constant as is equal or
greater than 30. The delamination occurs at the interface of 0/+45 plies and interface of
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+45/-45 plies. In Fig.2.13, the interlaminar normal and shear stresses near the hole
boundary are much higher than those near the plate edge. The high interlaminar stresses near
the hole edge may contribute to the initiation of delamination starting from the hole edge.
The delamination is found near the hole edge and bounded by 0 ply splitting.
Figure 2.13 The interlaminar stresses along the hole edge at the interfaces of 0/+45 (z = 2h)
plies and +45/-45 (z = 1h) plies
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CHAPTER 3
FINITE ELEMENT MODELING
3.1 Definition of Problem
A composite laminate plate with hole, with and without crack is analyzed .The test
cases are [02/ ]s and [/02]s plate where ranges from 15o, 30
o, 45
o, 60
o, 75
o. Different
diameter to width ratio d/w has been studied. The d/w ratios studied are 0.5, 0.6, 0.7, 0.8,
0.85. The crack is modeled in + and - direction for + and - layer, respectively. Full
model has been used for analysis. Sub-modeling of full model has been done for more
accurate results. The material constants used are listed in table 3-1[11]
Table 3.1 Material Constants Used
Material
System
E1 Msi E2= E3 Msi 23 12= 13 G23 Msi G12= G13 Msi
IM6/3501-6,Gr/Ep *
23.3 1.395 0.342 0.2965 0.5198 0.9161
* ply thickness = 0.005
Three-dimensional modeling using ANSYS has been done. There is only one
symmetric plane in the model, namely, x-y plane (z=0). Therefore, half of the laminate is
needed to model. The dimension of the model that has been used is 1.0 inch long, 0.1 inch
breadth and 0.2 inch thick where the thickness of each layer is 0.005 inch.
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3.2 Element Type and Size
The element type and size in the mesh affects the results in the model. For the same
number of elements in a model, the choice of tetrahedral and quadrilateral element can
usually lead to different results. The difference may or may not be significant for solving
some problems. Therefore, the comparison of finite element results and experimental result,
are important for determining the element type and size.
Twenty-node quadrilateral element SOLID191 is used in this study. The quadrilateral
element can be degenerated to the tetrahedral to be best fitted into the boundary of the hole.
Fig.3-1 shows SOLID191 geometry and Fig.3-2 shows SOLID191 element output definition
notations. The number of elements and nodes in the model varies for different cases under
consideration.
Gap elements have been used for crack at zero degree layers for 90o
half model where
two cracks were present. Surface-to-Surface Contact elements, TARGE 170 and
CONTA174, have been used. Figure 3-3 shows TARGE170 geometry and Figure 3.4 shows
CONTA174 geometry. Each contact pair is identified via same real constant number.
CONTA174 is a 8-node higher-order quadrilateral element that can be located on the
surfaces of 3-D solid or shell elements with mid-side nodes.
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Figure 3.1 SOLID 191 Geometry
Figure 3.2 The Element output definition notation
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modeling. The local model is a sub model of the global model which models either the entire
of the laminate or portion of the laminate. A typical model of this study contains 2500
elements and 4500 nodes. Both full and sub models have been analyzed for all the cases,
except 15o crack. For15o crack, the sub model requires a large size in order to include
both the hole and the crack.
Full model Sub modelFigure 3.5 Full and Sub Models with a crack
For this care, the benefits of using the sub model are not significant. A sub model is
selected for refining the meshes to capture the high stress gradient around the vicinity of the
hole and the crack. Hence the size of the sub model should be large enough to contain a crack
which is emanated from the hole.
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The cases under study are [02 ,]s and [, 02]s laminates with hole and crack
where different values of are 15o, 30o, 45o, 60o, 75o, 90o. Different d/w ratio is analyzed.
The different d/w ratios analyzed are 0.5, 0.6, 0.7, 0.8 and 0.85.
Crack is present at + or - layer along the fiber orientation. For ,15o, 30o, 45o, 60o
and 75o
half model of the complete model is modeled as it is been assumed that a crack is
present at both the angle ply, respectively.
For composite modeling the fiber orientation is assigned through real constant
commands R and RMODIF commands assign real constant values. Because of laminate
symmetric to its mid-plane of the thickness only half of the laminate is considered.
Figure 3.6 3-d modeling in which + crack is present in +layer for [,02]s laminate
In creating meshes for ANSYS model, the KEYPOINTS are first assigned as nodal
points. The area and volume commands were executed. Two volumes are created in each
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layer and are glued together in the same plane except the layer with a crack. Then the volume
of each layer is stacked according to the stacking sequence of the laminate. For the layer with
a crack, double points with the same co-ordinates remain to simulate the crack. A cylinder is
built, which is then subtracted from the entire volume, which creates a hole in the model. The
hole is located at the center of the laminate and is tangential to the crack if the crack is
required in the model.
The model is constrained on the surface of one end in x-direction. The middle line in
z-direction on this surface is fixed in y-direction to simulate the symmetric surface of the
laminate width. The bottom surface of the model is constrained along z-direction. After
solving the full model sub-modeling was developed. In sub-model, a cut boundary command
of ANSYS, CBDOF is utilized to translate the displacement of each node in the full model
along the cut boundaries into corresponding displacement in the sub-model. The
interpolations of displacements are applied to the new created nodes of the sub model. Fig3-7
shows meshed volume with boundary condition . Unlike the other laminate, only a quarter of
the model is needed for [02/902]s laminate because of symmetry even if the crack in 90o
layer
is present. For this case, symmetry condition is enforced for the quarter model. Fig 3.8
shows meshed and with boundary condition of this case.
Another 90-degree case has been studied in which another crack in 0 degree layer is
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Figure 3.7 Meshed volume & with applied boundary conditions
modeled with the 90degree crack. This crack is of length 0.02 in and is tangential to the hole
at 0 degree. For this case two types of elements are used namely, SOLID191 and Gap
Elements. Two type of Gap Elements have been used, namely TARGE170 and CONTA174.
The combination of these two types of elements creates a surface-to-surface contact. To
create a contact pair same real constant number is assigned to both TARGE170 and
CONTA174.
This gap element pair is used for the crack in 0 degree layer. The reason for
choosing this element is that it prevents the nodes to penetrate into each other. Atypical
inputs for creating ANSYS models are attached to Appendix A. 90 degree ply with the only
difference that their will a crack at 0 degree layer which is not glued to any of the volumes.
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Figure 3.8 Meshed volume with applied boundary conditions for [02,90]s laminate
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CHAPTER 4
RESULTS AND DISCUSSIONS
4.1 Model Validation
Two steps were adopted before we use to investigate the stress distribution around the
hole. First, the material constants of aluminum were used in the model to check the
connectivity of node in the model, the boundary conditions and the load application.
The aluminum plate has 0.1 inch wide and 0.05-inch diameter. Since accuracy of stress
concentration factor is not the primary concern in this study a coarse mesh of ANSYS was
used. A quarter model of aluminum block has been studied and an axial load of 10 psi is
applied to it. Fig 4.1 shows Aluminum plate with a hole.
Figure 4.1 Aluminum plate with a hole under uniaxial loading
The result got from ANSYS plots is that the maximum value of stress is 46.09psi
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and the minimum value of stress is 3.895psi. The calculated stress concentration factor, Kt
is Kt = 4.25. However, ANSYS results gives Kt = 4.61. Negative sign indicates compression.
Then, a composite laminate of [02/45]s without a hole is modeled. The purpose of this study
is to check the applicability of the anisotropic material input. The ANSYS results are used to
compare the results obtained from the MATLAB results calculated by lamination theory. A
composite laminate of [02/45o] s is being studied. Fig 4.2 shows the ANSYS results for this
case.
.
Figure 4.2 Composite laminate under axial loading
A comparison of the calculated MATLAB results and ANSYS values are listed in After
this a composite plate with a hole under uni-axial loading is considered. Fig4.3 shows the
plot of composite plate with hole under uni-axial loading. The results were in agreement with
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Table 4.1 Comparison of Calculated values to ANSYS values
4.2 [02/ ]s composite laminate
There are two cases studied, one in which a crack is present at + layer and the other
crack is present at - layer.
Figure 4.3 Quarter model of composite plate with hole for d/w=0.5
4.2.1 Composite plate with crack at + layer
There are five d/w ratios studied, d/w = 0.5, 0.6, 0.7, 0.8, 0.85.There are six angles that
are being considered, = 15o, 30o, 45o, 60o, 75o. Fig, 4-4 shows the stress distribution along
Layer x MATLAB Results x , ANSYS results
-45o
2.6969psi 2.5898psi
+45o
2.6969psi 2.5898psi
0o
17.303psi 17.553psi
0o
17.303psi 17.553psi
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normalized increases as the ratio of d/w increases. For d/w ratio of 0.8 and 0.85 the point of
intersection lies outside the laminate plate and consequently the maximum stress is reached
at the free edge
From Fig.4-6 we can see that with increase in d/w ratio the point of maximum
stress shifts towards the free edge and beyond a particular d/w ratio the maximum value is
reached at the free edge. The orientation of the breakage of the 0o
plies tends to be tangent to
the hole edge. Referring to Fig 2.3 the final failure of the laminate is not along the -
direction for +y region .For=45o, as shown in Fig 4.5 the trend of the normalized x is
similar to the case of=60o. The peak value of the normalized stress, x is listed in Table 4.2
for all of d/w ratios and various s. For any given , the peak value ofx increases as the
d/w ratio increases
4.2.2 Composite plate with crack at - layer
There are five d/w ratios studied, d/w = 0.5, 0.6, 0.7, 0.8, 0.85.There are six angles that are
being considered, = 15o, 30o, 45o, 60o, 75o. Fig, 4-7 shows the stress distribution along the
-direction in the zero degree layer when crack is present at - layer for d/w=0.5.From this
plot we can see that with increase in the maximum value of stress increases but if we
compare the results when crack was present at + layer the stress value is less in case of -
layer crack compared to + layer. The table shows that the peak value ofx increases as
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-2
0
2
4
6
8
10
12
14
16
-0.06 -0.04 -0.02 0 0.02 0.04 0.06
-y +y
NormalizedStress
d/w=0.5
d/w=0.6
d/w=0.7
d/w=0.8
d/w=0.85
Normalized stress=60 for different d/w ratis
Figure 4.5 Normalized stress distribution with varying d/w ratio for [02/60]s
-1
0
1
2
3
4
5
6
7
-0.06 -0.04 -0.02 0 0.02 0.04 0.06
normalizedstress
d/w=0.5
d/w=0.6
d/w=0.7
d/w=0.8
d/w=0.85
+y -y
0.
x
y
Figure 4.6 Normalized stress distribution for [ 02/45]s for varying d/w ratio
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increases for a given d/w ratio. For any given , the peak value ofx increases as the d/w
ratio increases.
-2
0
2
4
6
8
10
-0.06 -0.04 -0.02 0 0.02 0.04 0.06
-y width +y
NormalizedStress
2
2
2
2
2
2
Variation of sx/so for d/w=0.5 for -x
-x
y
Figure 4.7 Stress distribution along the y-axis in the zero degree layer when crack is present
at - layer
The pattern of stress distribution for the entire d/w ratio remains same. With varying d/w
ratio only the values of maximum stress varies. Table 4-3 shows the value of maximum stress
in zero degree layers with varying d/w ratio. The patterns remain same as that in the case of
+ crack i.e. with increase in d/w ratio the maximum stress value increases and with increase
in also the maximum value of stress increases. Comparing the results from table 4.2 & 4.3
for crack in + layer and - layer, respectively one can see that the value of maximum stress
is lower in case of - layer crack compared to +layer crack. From this we can conclude that
the farther is the crack layer from the 0 degree layer the lower will be the maximum stress
value in zero degree layer.
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Table 4.2 Max. stress in 0o
layer with varying d/w ratio for [02/ ]s laminate for + crack
+ Crack layer Maximum x/in Location from ANSYS Location of intersection0o layer (Max. stress location) point(Max. stress location)
d/w =0.5
+15 2.88 0.0244 0.0241+30 3.95 0.0213 0.0211
+45 4.22 0.0201 0.015
+60 4.57 0.0092 0.0+75 5.17 0.0043 outside plate
d/w =0.6
+15 3.79 0.0213 0.0189+30 4.21 0.0147 0.0153
+45 4.78 0.0067 0.00757
+60 5.86 0.0023 outside plate+75 6.12 0.0 outside plate
d/w = 0.7+15 4.05 0.0133 0.0138
+30 5.32 0.0067 0.0094
+45 5.97 0.0 0.0005
+60 6.27 0.0 outside plate+75 6.89 0.0 outside plate
d/w = 0.8+15 6.75 0.0067 0.0086
+30 7.32 0.0067 0.0038
+45 7.15 0.0 outside plate+60 8.85 0.0 outside plate
+75 10.837 0.0 outside plate
d/w = 0.85
+15 7.44 0.0067 0.006+30 8.24 0.0 0.0
+45 9.96 0.0 outside plate+60 12.34 0.0 outside plate
+75 14.12 0.0 outside plate
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From the table we can see that the point of maximum stress is close to the calculated value
and this means that in the case of crack layer, respectively the maximum stress value
point remains same. The little difference in the results obtained from ANSYS is because of
different number of elements in various cases. For the meshing of 15o
and 30o
very fine
elements had to be used because of the geometry of the figure which was not the case for
+15o
and +30o.Even for the other angles the total number of elements were not same.
4.3 [/ 02]s Composite laminate
4.3.1 Composite plate with crack at + layerThere are five d/w ratios studied, d/w = 0.5, 0.6, 0.7, 0.8, 0.85.There are six angles
that are being considered, = 15o, 30o, 45o, 60o, 75o. Fig., 4-8 shows the stress distribution
along the y-axis in the 0o
layer when crack is present at + layer for d/w =0.5.
-0.5
0
0.5
1
1.5
2
2.5
3
-0.1 -0.05 0 0.05 0.1
-y width +y
NormalizedStress = 15
= 30
= 45
= 60
= 75
Variation of Normalized Stress with different ford/w=0.5
Figure 4.8 Stress distribution in 0 degree layer when crack is present at + layer for[ / 02]s laminate
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Table 4.3 Max. Stress in 0o
layer with varying d/w ratio for [02/ ]s laminate for - crack
- Crack layer Maximum x/in Location from Ansys Location of intersection0o layer (Max. stress location) point(Max.stress location)
d/w = 0.5-15 2.77 0.0256 0.0241
-30 3.68 0.0233 0.0211
-45 4.01 0.0133 0.0147-60 4.41 0.0008 0
-75 4.79 0.0005 outside plate
d/w = 0.6-15 3.16 0.0196 0.0189
-30 3.88 0.0164 0.0154
-45 4.50 0.0067 0.0076-60 4.97 0 outside plate
-75 5.27 0 outside plate
d/w = 0.7
-15 3.75 0.0164 0.0138
-30 4.21 0.0133 0.0096-45 4.85 0.0066 0.0005
-60 5.23 0 outside plate
-75 5.75 0 outside plate
d/w = 0.8
-15 4.22 0.0067 0.0086
-30 5.30 0.0048 0.0038-45 6.66 0 outside plate
-60 7.04 0 outside plate
-75 8.52 0 outside plate
d/w = 0.85-15 5.27 0.0048 0.0060
-30 6.46 0 0.00009
-45 8.17 0 outside plate-60 9.85 0 outside plate
-75 12.78 0 outside plate
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From this figure we can see that the pattern of stress distribution remains same as that in the
[02,]s but the value of maximum stress has decreased compared to that in the case of
[02,]s. Table 4-4 shows the maximum value of stress with varying d/w ratio.
4.3.2 Composite plate with crack at - layer
There are five d/w ratios studied, d/w = 0.5, 0.6, 0.7, 0.8, 0.85.There are six angles
that are being considered, = 15o, 30o, 45o, 60o, 75o.The pattern of stress distribution even in
this case remain same as in the other cases. Table 4-5 shows the maximum value of stress
with varying d/w ratio. The value of maximum stress in this case is larger than that in +
crack. This pattern is same as in the previous case of [02/]s in which the layer that was
farther from the 0 degree layer had lower value of stress compared to that which was next to
0 degree layer.
4.4 Failure Strength Prediction Using Classical Lamination Theory
A fracture model for prediction strength of laminate with a hole is proposed. This
model is based on the fact that final failure of the laminates is controlled by fiber breakage in
zero degree laminas. The model is described below:
T
ON
pred
N XCC 1= (4-1)
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where CN is called the notched correction factor,TX1 is the strength of the 0
oply and Co is the
axial stress ratio of the 0
o
ply stress to the laminate stress in the same direction. The x of 0o
plies in a laminate, which is calculated using CLT ,o
atelax
0
min, is related to the applied load,
which is x of the 0o
lamina,o
alax
0
min, through co-efficient C0. The C0 is defined as
o
o
atelax
alaxC
0
min,
0
min,
0
= (4-2)
CN, which is the notched correction factor, is different for different lay-ups in a specific
material system and it is determined by taking the inverse of stress concentration factor in 0o
ply. The calculated notched strength is shown in table 4.6. Comparisons of the experimental
and calculated results have been done for [02 /45]s laminate. From table we can see that for
all the d/w ratio except 0.5 the results obtained are close to that of the experimental results.
The reason for the difference in results in case of d/w ratio of 0.5 can accounted to the fact
that in case of a smaller d/w ratio delamination propagation is larger and in this model
delamination propagation has not been considered which in case of experiment al results has
been taken into account. Table 4-6 shows the notched strength of [02/]s when crack is
present at + direction in + layer. Table 4-7 shows the notched strength of [02/]s when
crack is present at - direction in - layer. Comparing both the tables we can see that the
notched strength in case of - layer crack is more compared to + crack. From this we can
conclude that if a crack is placed right next to 0 degree layer then its notched strength will be
lower compared to the case when the crack is placed away from the 0 degree layer.
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Table 4.4 Max. Stress in 0o
layer with varying d/w ratio for [ / 02]s laminate for + crack
- Crack layer Maximum x/in Location from Ansys Location of intersection0o layer (Max. stress location) point(Max.stress location)
d/w = 0.5+15 1.90 0.0233 0.0241
+30 2.11 0.0200 0.0211
+45 2.37 0.0133 0.0147+60 2.59 0.0067 0
+75 2.79 0 outside plate
d/w = 0.6+15 2.75 0.0200 0.0189
+30 3.20 0.0133 0.0154
+45 3.45 0.0067 0.0076+60 3.61 0 outside plate
+75 5.41 0 outside plate
d/w = 0.7
+15 2.94 0.0133 0.0138
+30 3.44 0.0067 0.0096+45 3.87 0 0.0005
+60 4.04 0 outside plate
+75 6.37 0 outside plate
d/w = 0.8
+15 3.36 0.0133 0.0086
+30 3.95 0.0067 0.0038+45 4.73 0 outside plate
+60 7.02 0 outside plate
+75 8.39 0 outside plate
d/w = 0.85+15 4.65 0.0067 0.0060
+30 5.73 0.0004 0.0000
+45 6.50 0 outside plate+60 7.84 0 outside plate
+75 8.47 0 outside plate
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Table 4.5 Max. Stress in 0o
layer with varying d/w ratio for [ / 02]s laminate for - crack
- Crack layer Maximum x/in Location from Ansys Location of intersection0o layer (Max. stress location) point(Max.stress location)
d/w = 0.5
-15 2.26 0.0256 0.0242-30 2.48 0.0196 0.0241
-45 2.69 0.0133 0.0147
-60 2.89 0.0066 0.0146-75 3.15 0 outside plate
d/w = 0.6
-15 3.21 0.0196 0.0189-30 3.68 0.0164 0.0154
-45 3.84 0.0067 0.0153
-60 4.56 0 outside plate-75 5.79 0 outside plate
d/w = 0.7-15 3.50 0.0133 0.0138
-30 3.97 0.0067 0.0096
-45 4.42 0 0.0005
-60 5.02 0 outside plate-75 6.89 0 outside plate
d/w = 0.8-15 4.00 0.0133 0.0086
-30 4.50 0.0066 0.0038
-45 5.27 0 outside plate-60 8.12 0 outside plate
-75 9.59 0 outside plate
d/w = 0.85
-15 4.94 0.00667 0.0060-30 6.20 0.00483 0.0001
-45 6.97 0 outside plate-60 10.99 0 outside plate
-75 12.8 0 outside plate
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In case of [ /02]s it can again be seen that for the case of crack which is placed right next
to 0 degree layer, which in this case is -
layer, the notched strength is lower compared to the
case where the crack is present away from 0 degree layer, which in this case is + layer.
Table 4-8 & Table 4-9 shows notched strength [ /02 ]s laminate when crack is present in
layer in the respective direction. Comparing these two tables again we can see that the
notched strength of + layer crack is more compared to - crack. Figure 4.13 shows Notched
Strength in [ /02]s laminate when crack is present in layer in the respective direction.
0
20
40
60
80
100
120
140
0.4 0.5 0.6 0.7 0.8 0.9
d/w
N
otchedStrength(ksi)
= +15
= +30
= +45
=+ 60
= +75
= 15
= 30
= 45
= 60
= 75
Figure 4.9 Notched Strength in [02/ ]s laminate when crack is present in layer in therespective direction
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Table 4.6 Notched Strength of [02/]s when crack is present at + direction in + layer
Laminate Lay-up d/w Co CN
pred
N
test
N
[02/15]s 0.5 0.92 0.35 112.170.6 0.92 0.25 81.66
0.7 0.92 0.25 79.57
0.8 0.92 0.15 45.480.85 0.92 0.13 43.38
[02/30]s 0.5 0.71 0.25 63.120.6 0.71 0.24 59.21
0.7 0.71 0.19 46.84
0.8 0.71 0.14 34.050.85 0.71 0.12 30.27
[02/45]s 0.5 0.59 0.24 48.79 640.6 0.59 0.21 43.06 480.7 0.59 0.17 34.48 39
0.8 0.59 0.14 28.78 250.85 0.59 0.10 20.66 20
[02/60]s 0.5 0.54 0.22 41.320.6 0.54 0.17 32.19
0.7 0.54 0.16 30.090.8 0.54 0.11 21.32
0.85 0.54 0.08 16.75
[02/75]s 0.5 0.53 0.19 36.450.6 0.53 0.16 30.360.7 0.53 0.14 26.96
0.8 0.53 0.09 17.14
0.85 0.53 0.07 13.15
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Table 4.7 Notched Strength of [02/]s when crack is present at - direction in - layer
Laminate Lay-up d/w Co CN
pred
N
[02/15]s 0.5 0.92 0.36 116.660.6 0.92 0.32 101.95
0.7 0.92 0.27 86.08
0.8 0.92 0.24 76.340.85 0.92 0.15 49.93
[02/30]s 0.5 0.71 0.27 67.740.6 0.71 0.26 64.35
0.7 0.71 0.24 59.16
0.8 0.71 0.19 47.070.85 0.71 0.15 38.59
[02/45]s 0.5 0.59 0.25 51.340.6 0.59 0.22 45.740.7 0.59 0.21 42.43
0.8 0.59 0.15 30.930.85 0.59 0.12 25.20
[02/60]s 0.5 0.54 0.23 42.810.6 0.54 0.20 37.96
0.7 0.54 0.19 36.040.8 0.54 0.14 19.43
0.85 0.54 0.10 19.15
[02/75]s 0.5 0.53 0.21 38.790.6 0.53 0.19 35.480.7 0.53 0.17 24.15
0.8 0.53 0.12 21.81
0.85 0.53 0.09 14.53
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Table 4.8 Notched Strength in [ /02 ]s laminate when crack is present in layer
Laminate Lay-up d/w Co CNpred
N
[15/ 02]s 0.5 0.92 0.53 170.000.6 0.92 0.41 131.000.7 0.92 0.34 109.60
0.8 0.92 0.25 81.27
0.85 0.92 0.17 56.25
[30/ 02]s 0.5 0.71 0.47 118.310.6 0.71 0.31 78.00
0.7 0.71 0.29 59.160.8 0.71 0.25 63.02
0.85 0.71 0.17 43.48
[45/ 02]s 0.5 0.59 0.42 86.960.6 0.59 0.29 59.66
0.7 0.59 0.26 53.20
0.8 0.59 0.21 43.500.85 0.59 0.15 31.64
[60/ 02]s 0.5 0.54 0.39 72.89
0.6 0.54 0.28 52.360.7 0.54 0.25 46.72
0.8 0.54 0.14 26.870.85 0.54 0.13 24.06
[75/ 02]s 0.5 0.53 0.36 66.600.6 0.53 0.18 34.370.7 0.53 0.16 29.17
0.8 0.53 0.12 22.14
0.85 0.53 0.11 21.96
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Table 4.9 Notched Strength in [ /02 ]s laminate when crack is present in layer
Laminate Lay-up d/w Co CN
pred
N
[15/ 02]s 0.5 0.92 0.44 142.680.6 0.92 0.41 131.00
0.7 0.92 0.34 109.600.8 0.92 0.25 81.27
0.85 0.92 0.17 56.25
[30/ 02]s 0.5 0.71 0.47 118.310.6 0.71 0.31 78.00
0.7 0.71 0.29 59.16
0.8 0.71 0.25 63.020.85 0.71 0.17 43.48
[45/ 02]s 0.5 0.59 0.42 86.960.6 0.59 0.29 59.660.7 0.59 0.26 53.20
0.8 0.59 0.21 43.50
0.85 0.59 0.15 31.64
[60/ 02]s 0.5 0.54 0.39 72.890.6 0.54 0.28 52.36
0.7 0.54 0.25 46.720.8 0.54 0.14 26.87
0.85 0.54 0.13 24.06
[75/ 02]s 0.5 0.53 0.36 66.600.6 0.53 0.18 34.37
0.7 0.53 0.16 29.170.8 0.53 0.12 22.14
0.85 0.53 0.11 21.96
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The notched strength of decreases with increase in for a constant d/w ratio. Similarly with
increase in d/w ratio for a constant the notched strength decreases. The pattern remains the
same as in previous case. Figure 4.11 and Figure 4.12 shows a comparison of the two lay-ups
.
Figure 4.12 Comparison of Notched Strength between [02/ ]s & [ /02 ]s when crack ispresent away from 0 degree
We can again see the same pattern as in the previous case. The reason for this behavior
can be accounted to the fact that if all the 0 degree layers are placed next t each other it can
withstand more load compared to the case where the 0 degree layers are scattered. Hence the
notched strength for the case where all 0 degree layers are placed together is higher.
4.5 Conclusion
With increase in in a given d/w ratio the stress concentration factor in 0 degree layer
for [02/]s and [/02]s laminates increases .Even with increase in d/w ratio for a given
0
20
40
6080
100
120
140
160
180
0.4 0.5 0.6 0.7 0.8 0.9
d/w
NotchedStrength
= 15
= 30
= 45
= 60
= 75
= 15
= 30
= 45
= 60
= 75
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value of the stress concentration factor in 0 degree layer for [02/]s and [/02]s laminates
increases. The location of the layer due the matrix crack also plays an important role to
determine the stress concentration factor. The closer the crack layer to the 0 degree layer the
higher the value of stress concentration factor in 0 degree layer for [02/]s and [/02]s
laminates. If all the 0 degree layers are attached to each other then the value of stress
concentration factor in 0 degree layer is lower compared to the case where it is on the outer
side of the laminates.
The notched strength of a laminate decreases with increase in d/w ratio for a constant
. Similarly, for a constant d/w ratio, with increase in the notched strength decreases. In
case of lay-ups, the notched strength is higher in the case where all 0 degree layers, which in
this case is the load carrying ply, are placed next to each other compared to the case where
they are scattered. Also, the notched strength is dependent on the location of the crack layer.
If the crack layer is placed next to 0 degree layer then the notched strength is lower compared
to the case where the crack layer is not placed to the 0 degree layer.
This study can provide a guideline to the engineers that for a given d/w ratio what
value of laminate can they select based on the stress permissible for that application.
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REFERENCES[1] Awerbuch, J. and Madhukar, M. , Notched Strength of Composite Laminates:
Predictions and Experiments---A Review, Journal of Reinforced Plastics and Composites,
Vol.4, January 1985, p.3.
[2] Tan, S. C., Stress Concentrations in Laminated Composites, Technomic Publishing Co., 1994.
[3] Whitney, J.M. and Nuismer, R.J. , Stress Fracture Criteria for Laminated Composites
Containing Stress Concentrations,Journal of Composite Material, Vol. 8,July 1974,
p.253.
[4] Chan, W.S. "Damage Characteristics of Laminates with a Hole," Proceedings of the American
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[5] Harn, F. E., Notched Strength of [02/]S Graphite/Epoxy Laminates with various Hole Size
Ph.D. dissertation, University of Texas at Arlington, Aug. 1997.
[6] Waddoups, M.E., Eisemann, J.R., and Kaminski, B.E., Macroscopic Fracture
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[7] Whitney, J.M. and Nuismer, R.J., Uniaxial Failure of Composite Laminates Containing
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[8] http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MA5086H116 last
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[9] http://casl.ucsd.edu/data_analysis/carpet_plots.htm last visited on July 31,2005.
[10] Paris, P.C. and Sih, G.C., Stress Analysis of Cracks, Fracture Toughness Testing and
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[11] Konish, H.J. and Whitney, J.M., Approximate Stresses in an Orthotropic Plate
Containing a Circular Hole,Journal of Composite Materials, Vol 9, pp157-166, 1975
[12] Tan, S.C., Laminated Composite Containing an Elliptical opening I. Approximate
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[13] Tan, S.C., Stress Concentration in Laminated Composites, Lancaster:
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BIOGRAPHICAL INFORMATION
Moumita Roy was born at Calcutta(now Kolkata),India, in 1981. She received
her Master of Science in Mechanical Engineering from University of Texas at Arlington in
December 2005. Author completed her Bachelors degree in Ceramic Engineering from
Regional Engineering College-Rourkela in May 2002. Author has primary interests in the
fields of stress analysis, design and analysis of composite materials, finite element analysis.
She plans to pursue a P.HD in Mechanical Engineering.