Liam O Sullvian FYP Mar 2013

The Validity of Compact Tension
Testing in Fracture Toughness
Calculations for Thick Laminated
Composites
Liam O Sullivan – 09006876
I hereby declare that the work contained in this report is my own and
that any contributions from other persons have been appropriately
identified and acknowledged.
A final year project submitted in partial fulfilment of the
requirements for the degree of Bachelor of Engineering in
Mechanical Engineering.
Submitted: 22nd
March 2013

Final Year Project 2013 ID. No: 09006876
i
Abstract
Compact tension (CT) specimens are validated for use in the determination of mode 1
fracture toughness, KIc, in homogeneous metallic materials using American Society
for Testing & Materials (ASTM) standard E399-90. This study primarily investigated
the validity of CT testing of composite laminates. The fracture characteristics of
HTA-6376 composite laminates under similar test conditions was analysed when
subjected to both plane stress and plane strain loading.
Three laminates of [0/90]8s lay-up and three of [0/90]16s lay-up were tested to failure
and the fracture toughness of each specimen was calculated through both ASTM
methods and through the use of clip-on displacement (COD) gauges. The strains
induced in the specimen were analysed using digital image correlation (DIC) software
and large compressive stresses were noted across the rear walls of the specimens. A
finite element study was carried out using conventional shell elements to evaluate the
properties of concept designs. This was done in an attempt to overcome the
compressive stresses which caused failure of the specimen by shearing across the rear
wall. Thicker 64 ply specimens were made to the original specimen geometry and
tested to establish whether plane strain loading would overcome the shearing at the
rear wall. The first of these specimens failed in compression, and so modifications
were made to the geometry of the next two, and mode 1 fracture was achieved.
The KIc values obtained from the COD gauges were found to be highly inaccurate due
to errors in the calculation of crack-tip opening displacement (CTOD) using empirical
data. The compressive failure that occurred in four of the six specimens caused errors
in the ASTM method for KIc calculation, but the two successfully modified specimens
produced accurate values for the fracture toughness of the material. It was concluded
that CT specimens are not suitable for the measurement of plane stress fracture
toughness, but with alterations to the geometry, they can be used to accurately predict
plane strain fracture toughness.

ii
Acknowledgements
I would like to thank my project supervisor Dr. Ronan O’Higgins for his help and
guidance over the course of this project. Help was always available when needed in
progressing the project, and his contribution was greatly appreciated.
Senior technical officer Adrian McEvoy also deserves huge thanks for devoting his
time to help in the running of this experiment. Both Mr McEvoy and chief technical
officer John Cunningham showed great enthusiasm in progressing the project and
their help was invaluable in completing the project.
I would like to thank my peers, who have helped me throughout the four years of my
college degree, but especially during this project. They were always available to lend
a hand, and an accumulated knowledge of the subject area helped no end to progress
this study.
Finally, I would like to thank my family, my parents in particular, for supporting me
in my studies over the past four years. They have strived to help me with any
problems encountered and their contribution to my degree cannot be overstated.

iii
Table of Contents
1. Introduction............................................................................................................1
2. Objectives ..............................................................................................................4
3. Literature Review ..................................................................................................5
4. Theoretical Analysis ............................................................................................ 14
5. Finite Element Analysis....................................................................................... 17
5.1. Specimen Modelling..................................................................................... 17
5.2. Design Modifications ................................................................................... 19
6. Experimental Work.............................................................................................. 23
6.1. Specimen Manufacture.................................................................................23
6.2. Experimental Apparatus...............................................................................27
6.3. Experimental Procedure ...............................................................................29
7. Results and Discussion ........................................................................................ 30
7.1. 32 Ply Square Specimen...............................................................................30
7.2. 32 Ply Bevelled Specimen............................................................................35
7.3. 32 Ply Chamfered Specimen. .......................................................................39
7.4. 64 Ply Specimen A....................................................................................... 43
7.5. 64 Ply Specimen B. ...................................................................................... 46
7.6. 64 Ply Specimen C. ...................................................................................... 49
7.7. Discussion of Failure Mode and Test Validity.............................................53
8. Conclusions..........................................................................................................56
9. Recommendations for Future Work ....................................................................58
References................................................................................................................... 60
Appendices................................................................................................................A-1

iv
List of Figures
Figure Title Page
3.1 Specimen geometry used by Pinho (2005). 5
3.2 Failure sites in compact tension specimens (Blanco et. al. 2011). 10
3.3 Examples of multidirectional laminated composite failure in
tension (Kaman 2011).
12
4.1 Critical dimensions used in KIc. calculations according to ASTM
standard E399 (ASTM 1997).
14
5.1.1 Image of lay-up and partitions used in FE study of 64 ply laminate. 18
5.1.2 Image of mesh used on original geometry in FE study. 18
5.2.1 (a) Standard square specimen. 20
5.2.1 (b) Bevelled specimen. 20
5.2.1 (c) Chamfered specimen. 21
5.2.1 (d) Double chamfered specimen. 21
5.2.1 (e) Tapered specimen. 22
6.1.1 Vacuum bag used in lay-up process. 24
6.1.2 Autoclave used to cure panel. 25
6.1.3 Composite cutter used to machine specimens. 25
6.1.4 Wafering saw used to put pre-crack in specimens. 26
6.1.5 Specimen after speckling process. 27
6.1.6 Knife edges applied to specimen. 27
6.2.1 Dartec 100 kN tensile tester. 28
6.2.2 COD gauges attached to specimen. 28
7.1.1 Load vs. extension for 32 ply square specimen. 31
7.1.2 Failure of 32 ply square specimen. 32
7.1.3 Fracture surfaces of 32 ply square specimen 32
7.1.4 Load vs. crack opening displacement for 32 ply square specimen. 33
7.1.5 Strain fields in 32 ply square specimen. 35
7.2.1 Load vs. extension for 32 ply bevelled specimen. 36
7.2.2 Failure of 32 ply bevelled specimen. 37
7.2.3 Fracture surfaces of 32 ply bevelled specimen. 37
7.2.4 Load vs. crack opening displacement for 32 ply bevelled
specimen.
38
7.2.5 Strain fields in 32 ply bevelled specimen. 38
7.3.1 Load vs. extension for 32 ply chamfered specimen. 39
7.3.2 Failure of 32 ply chamfered specimen. 40
7.3.3 Fracture surfaces of 32 ply chamfered specimen. 41
7.3.4 Out of plane deformation of the three 32 ply specimens. 41
7.3.5 Load vs. crack opening displacement for 32 ply chamfered
specimen.
42
7.3.6 Strain fields in 32 ply chamfered specimen. 43
7.4.1 Load vs. extension for 64 ply specimen A. 44
7.4.2 Failure of 64 ply specimen A. 45
7.5.1 Load vs. extension for 64 ply specimen B. 46
7.5.2 Failure of 64 ply specimen B. 47
7.5.3 Load vs. crack opening displacement for 64 ply specimen B. 48
7.5.4 Strain fields in 64 ply specimen B. 49
7.6.1 Load vs. extension for 64 ply specimen C. 50
7.6.2 Fracture zones on rear wall of 64 ply specimen C. 51

v
7.6.3 Fracture surface of 64 ply specimen C. 51
7.6.4 Load vs. crack opening displacement for 64 ply specimen C. 52
7.6.5 Strain fields in 64 ply specimen C. 53
7.7.1 Strain fields for 64 ply specimen B plotted with DIC and FEA. 55
List of Tables
Table Title Page
5.2.1
Comparison of compressive and tensile strains for alternative
geometries.
19
5.6.1 Material propertied for HTA-6376. 23
7.1 Critical specimen dimensions for ASTM procedure. 30

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Nomenclature
Subscripts:
Symbol Description
11 In the direction of loading
22 Transverse to the direction of loading
cr Critical value
Q Fracture value
o Initial value
Symbol Description Unit
a Crack length m
ao Initial crack length m
aQ Crack length at fracture m
do characteristic damage parameter m
e11 Strain in the direction of loading (%) ---
B Specimen thickness m
E11 Young's modulus in direction of loading GPa
E22 Young's modulus in transverse direction GPa
G12 Shear modulus GPa
GIc Mode 1 Critical strain energy release rate kJ/m2
J J integral kJ/m2
KIc Mode 1 Fracture Toughness MN/m1.5
Kt Stress intensity factor ---
P Load N
Qc General toughness parameter ---
W Specimen width m
εut Ultimate tensile strain --
σ Stress GPa
σo Un-notched laminate strength GPa
υ12 Poisons ratio ---

1
1. Introduction
As technology advances, and science and engineering are facing more complex
challenges, the demand for materials that are stronger, lighter and cheaper has
increased greatly over the past 30 years. Composite materials have become a popular
solution as they offer an excellent strength to weight ratio over some of the more
common construction materials of the past, such as steel and aluminium. Composite
materials are currently used in the automotive industry as the main materials in the
bodies of high performance cars, as they are lighter than any metal previously used
while maintaining the required rigidity for application at high speeds. Other common
uses of composites, such as carbon fibre, include speed boats hulls, airplane fuselages
and sports equipment such as tennis rackets, baseball bats, and lacrosse sticks.
A composite material is said to have two main constituents; a matrix material such as
a plastic or resin, and a reinforcement material which is suspended within the matrix.
There is a distinct boundary between both these materials, where both maintain their
own physical and chemical properties, but produce a combination of properties that
cannot be achieved by either constituent acting alone (Mallick 2008). The
reinforcement material, usually in fibrous form, carries most of the load applied to the
material, while the matrix material serves to keep the fibres in the desired orientation.
The matrix material also serves to provide desirable properties such as heat
resistance, corrosion resistance and the transfer of the load between fibres. Most
composite materials are structured into laminates, which are panels of the material
which have a specified number of layers, or plies. In each individual ply, all fibres are
running parallel with the other fibres in that ply across the length of the panel. The
individual plies can then be layered on top of each other at specific orientations to a
fixed reference axis. This results in a laminate which is equally strong under all types
of loading in the plane of the fibres. As all fibres in a given ply are running parallel to
each other, and have identical properties, a laminated composite panel can be
described as an orthotropic material. This means that it has differing, but constant
properties under longitudinal and transverse loading.
This study investigates the physical property of materials known as fracture
toughness. This is a property which describes the amount of energy required to cause
a crack to propagate through a material. It is a critical criterion to be examined before

2
a material is chosen for a particular application as fast crack propagation is likely to
lead to catastrophic failure of a component, which can lead to financial losses,
environmental disasters, and loss of life. Fracture mechanics assumes that all
materials contain pre-existing flaws and cracks which can be caused during
production or during the use of the material. Heat can be a major factor in this as
repeated heating and cooling of an object can lead to thermal fatigue, where cracks
begin to form in the substructures of materials due to expansion and contraction. The
sharp point at the site of a crack, known as the root, is the site of large stress
concentrations which cause cracks to propagate as these concentrations get larger. An
object will fracture when unstable crack growth occurs. This happens when an
increment of crack growth causes more stored strain energy to be released than is
absorbed by the formation of the new crack surfaces (IME 1978). It is critical to study
the stress intensity factor when analysing any fracture mechanics problem as it
describes the loading of the specimen. The stress intensity factor, Kt, in a fracture
mechanics problem refers to the ratio of local stress around the crack tip and the
normal stress throughout the specimen. Factors which affect the normalised stress
level at the root of a crack include specimen geometry, the magnitude of the load
applied and the direction of the load. If the specimen fractures, the value for Kt is
known as the fracture toughness of the specimen, denoted by KIc.
These values are calculated for plane-strain homogeneous materials using American
Society for Testing & Materials (ASTM) standard E399-90. This standard sets out
guidelines to be followed during testing in order to obtain consistent, valid results.
Compact tension (CT) tests, as outlined in ASTM standard E399, are frequently
carried out on metals and isotropic plastics to evaluate their performance in a given
situation, as well as for inspection of a manufacturing process to ensure the material
made has the correct physical properties.
The aim of this study is to analyse the failure of CT specimens and to conclude
whether or not composite CT specimens of a marginally different geometry to that set
out in ASTM standard E399 will yield accurate results for the material fracture
toughness. As composite materials are inherently heterogeneous, it cannot be said that
the test methods laid out in ASTM standard E399 yield accurate results for these
materials. While the geometry of the specimen may be altered, the method used in
this standard will still be followed and should still yield similar results. The effect of

3
laminate thickness will be examined and alternative specimen geometries will be
tested to discover whether compressive stresses on the rear wall, which cause failure
in 32 ply specimens, can be overcome to give valid test results at that thickness.
These test specimens will be laid up by hand in a laboratory and cured before being
machined to the desired geometry. A 100kN tensile tester will be used to load the
specimens and measurement equipment, such as Digital Image Correlation (DIC) and
Clip-On Displacement (COD) gauges will be used to monitor the tests and record
data.
A finite element (FE) study will be carried out to consider the effect of specimen
geometry on the compression across the back wall of the specimen experienced by
Hannon (2012). 32 ply specimens will then be tested using these geometries to
evaluate the effectiveness of the alterations. 64 ply laminates will be tested with the
standard geometry and, based on the outcome of initial testing, may be tested with
alternative geometries similar to those studied using FE analysis. The results for
fracture toughness and critical strain energy release rate will be compared between
specimens and conclusions will be made regarding the validity of the test, based on
experimental results.

4
2. Objectives
This study covers many areas from design and manufacture, to testing and analysis.
During the study, the following goals are to be achieved:
 To manufacture carbon fibre specimens to a high degree of accuracy for use in
mechanical testing. A panel shall first be created, and the final test samples
will then be machined to specification from this panel.
 To carry out numerical analysis of the problem using ABAQUS 2.0 FE
software to investigate possible alterations to the design of CT specimens
which could reduce compressive stresses on the rear wall.
 To carry out testing on CT specimens to failure using a Dartec 100kN tensile
test machine.
 To obtain data regarding fracture load and specimen extension using DIC and
COD gauges as well as visual inspection under a microscope.
 To calculate the fracture toughness of the specimens using data gathered from
the mechanical testing carried out, such as the load at which fracture occurred.
 To compare these results with those obtained through theoretical calculations
and to then conclude on the validity of CT testing of composite materials as a
means of calculating material fracture toughness.

5
3. Literature Review
While the widespread use of composite materials is a reasonably new trend, the study
of the behaviour and characteristics of composites has been taking place for much
longer. As it was outlined in the introduction to this report, much of the information
gained about the fracture toughness of composites has come from testing, such as
single edge bend tests and open-hole tensile testing. The CT testing that will be
carried out in this report, with reference to ASTM standard E399, is only validated
for homogeneous metallic specimens. The interlaminar fracture toughness, the
resistance of the laminate to fracture between plies, has been researched to a greater
extent than intralaminar fracture toughness, the resistance to fracture across the plies.
For this reason, experimental procedures have been developed to accurately calculate
the interlaminar fracture toughness of a specimen, but further study is required to
assess the reliably of CT specimens as a means of intralaminar, or translaminar
fracture toughness.
Pinho (2005) carried out a test on carbon/epoxy T300/913 in order to calculate the
fracture toughness of the material. The lay-up used in this experiment was a [0/90]8S
lay-up and the specimens were machined to the geometry shown in figure 3.1. In the
lay-up of a composite material, [0/90] refers to the orientation of the fibres in the
panel with respect to the direction of loading. The subscript 8s states that the pattern
mentioned in brackets is repeated eight times, and is symmetrical, mirrored about the
centre plies. As a result, the specimen tested by Pinho (2005) was 32 plies thick. The
crack was created using a razor saw with a 0.2 mm thick blade, and the crack tip was
sharpened using a razor blade to give a final root radius of 0.005 mm.
Figure 3.1: Specimen geometry used by Pinho (2005)

6
The specimen was loaded at a rate of 5 mm/min and both DIC and a digital video
recorder were used to record the crack extension through to fracture. The load vs.
displacement graph obtained from this study is linear to approximately 2 mm, but
then becomes very staggered as the load begins to decrease close to failure.
Microscopic analysis of the samples carried out by Pinho (2005) suggests that fibre
bridging may have occurred, which would explain the oscillating results. Fibre
bridging involves only a portion of the fibres fracturing during crack growth. While
the crack continues propagating, the remaining fibres are put under greater stress until
they fracture, causing a sudden release of energy which results in incremental crack
extension and a drop in load. The value obtained for the mode 1 critical strain energy
release rate, GIc, of the carbon epoxy T300/913 during this test was 91.6kJ/m2
. From
this study, it was concluded that FE software provided a more accurate measurement
of the stress intensity factor at the crack tip than the method laid out in ASTM
standard E399 when analysing composite laminates and that the value of toughness
may be lay-up dependant due to the variations in GIc at the beginning of the test. As
valid results were obtained from this experiment, it was decided that test methods for
the investigation in this report should match as accurately as possible to those used by
Pinho (2005). Variations in test method in both laminate thickness and composite
material used shall be investigated to confirm the results found by Pinho (2005) and
the validity of using ASTM standard E399 in these calculations.
Jose et al. (2001) carried out a similar study to Pinho (2005) in which the relationship
between the fracture toughness of a cross ply laminate and that of its sub laminates
was examined. Three different specimens were examined; a cross ply laminate with a
[0/90]15 lay-up, and two representative sub-laminates with a [0]30 and [90]30 lay-up
respectively. The composite material tested was M55J/M18 carbon/epoxy and the test
was carried out in accordance ASTM standard E399. A razor blade was again used to
create the initial crack in the specimen, however the thickness of the razor blade used
to sharpen the crack was not specified. An FE study was carried out and fracture
toughness was calculated using a modified crack closure integral (MCCI) method.
This involves using information from the FE study about the forces and
displacements around the crack to calculate the strain release rate, GI, at the crack tip.
Clip-on displacement (COD) gauges were used to monitor the opening of the crack

7
during testing. A similar technique will be used in this report with the aim of
calculating KIc from the data.
The average value for the fracture toughness of the cross ply laminates, similar to
those under investigation in this report, was 808.6 N/mm1.5
. Jose et al (2001) noted
that the load required to fracture the [0/90]15 laminate and the load required to
fracture the other two sub-laminates show a good conformance to the relationship
below:
Ps[0/90] = Ps[0] + Ps[90]
This relationship can be used to compare the fracture toughness of the cross ply
laminate and its sub-laminates. This formula can be seen below, where n represents
the numbers of pairs of alternate 0o
and 90o
plies and B is the laminate thickness:
KIc[0/90]n = n/B{ KIc[0] + KIc[90]}
It was concluded in this study that the fracture toughness of a cross ply laminate can
be predicted with reasonable accuracy using a MCCI method coupled with an FE
analysis of its sub-laminates. A good agreement was seen between values of crack
opening displacement for both the COD gauges and the FE study, with a maximum
error of 8% for the unidirectional plies. Jose et al. (2001) found that the fracture
toughness of the laminate decreased as the laminate thickness increased, which is
consistent with linear elastic fracture mechanics of homogenous materials. These
results suggest that the testing of composite CT specimens in accordance to ASTM
standard E399 would yield valid results.
Harris and Morris (1984) carried out extensive research into the relationship between
the fracture toughness of a composite laminate and the thickness of the laminate. The
investigation focussed a graphite/epoxy T300/5208 composite material which was
tested under CT, centre-cracked tension and three-point bend loading conditions. This
study was carried out to confirm the findings of other investigations which found no
correlation between the fracture toughness of centre-cracked specimens (Hahn and
Morris 1977) and of three-point bend specimens (Cruse and Osias 1974).
Three different lay-ups were used in this investigation, [0/±45/90]ns, [0/±45]ns, and
[0/90]ns. These lay-ups were used to create test samples for three-point bend and

8
centre-cracked tension tests at thicknesses of 8, 32, 64, 96 and 120 plies. For the CT
tests, only the 64, 96 and 120 ply laminates were tested as it was noted that out of
plane bending could occur in thinner specimens which would lead to mixed mode
fracture. The studies carried out by Hannon (2012) and Pinho (2005) both use thinner
32 ply specimens, and it is documented well in Hannon (2012) that out of plane
buckling occurs in these specimens. Due to the greater strength in the out of plane
direction, thicker specimens are less likely to deform enough to cause significant
damage to the laminate. Harris and Morris (1984) also mention the issue of shear
forces on the loading holes in CT specimens as another reason to avoid testing the
thinner specimens. This is noted by Pinho (2005) when altering the specimen
geometry to use smaller loading holes and to move them further from the edge of the
specimen to reduce the possibility of failure at this point.
Crack opening displacement data is used by Harris and Morris (1984) to obtain values
for the fracture toughness of the graphite epoxy composite. It was noted in this
section of the investigation that the tensile forces which lead to crack propagation
also cause compression across the rear wall of the specimen. This compressive force
acts to prevent further crack growth. This compression was also be seen in Hannon
(2012) where the forces on the rear wall led to delamination and failure of the
specimen before mode 1 fracture could occur. While thinner specimens were used in
the studies by Pinho (2005) and Jose et al. (2001), none of the authors reported the
damage observed by Hannon (2012).
It was noted during this study that the fracture toughness of the [0/90]ns laminates
varied quite significantly with changes to both the thickness and crack length of the
specimens. Harris and Morris (1984) noted that as the thickness of the laminates
increased, the fracture toughness decreased. This confirmed the findings of Jose et al.
(2001). This was due to extensive damage at the site of the crack tip. As the load
applied to the specimen increased, the stress at the crack tip intensified. This caused
the crack to propagate, but it also caused damage to occur in fibres and matrix
running perpendicular to the crack. This damage released some of the strain energy in
the region, as well as blunting the crack tip, and causes retardation of the crack
growth. This could be another source of the unsteady crack growth seen by Pinho
(2005).

9
A similar explanation is given in this literature to explain the variation in fracture
toughness with crack length. In the [0/90]2s specimens tested by Harris and Morris
(1984), a deply investigation took place, which involved analysing each ply of the
laminate separately after testing. It was noted that fibre breakage occurred in the two
inner 0o
plies while on the outer plies of the specimen a matrix failure occurred that
resulted in the splitting of the ply in the 0o
direction. This highlighted the importance
of accurate crack length measurement at the beginning of each test. The damage
described above can also cause discontinuities in crack opening displacement
information during the testing. This is due to the sudden onset of damage which can
cause crack retardation followed by sudden increases in crack opening displacement.
Poe (1983) investigated the validity of using a term known as the general toughness
parameter, Qc/εut, as a means of fracture toughness calculation. Qc, known as the
toughness parameter, is the strain failure criterion for fibres in the principal load-
carrying plies (Poe 1983). The existence of such a parameter would mean that the
fracture toughness of any material could be calculated using only the elastic
properties and ultimate tensile strength of the fibres. Testing was carried out on 44
different combinations of matrix material, reinforcement material and lay-up.
Poe (1983) noted that the mean value for Qc/εut for laminates that did not delaminate
during testing was 1.5 mm0.5
. However, there are a number of limitations to the
application of this method. It was observed that specimens that showed low to
medium levels of damage at the crack tip did not agree with this value. Also, this test
was validated for specimens no greater than 16 plies in thickness. This could result in
values for plane strain scenarios varying from the value calculated in that study.
Harris and Morris (1985) used the theory mentioned by Poe (1983) to calculate the
general toughness parameters for various specimens used in their testing which
included a [0/90]24s laminate. The results provided by Harris and Morris (1985)
suggested that this was a valid means of fracture toughness calculation for plane
stress specimens. As a result, the value for Qc/εut from Poe (1983) will be used during
the COD gauge analysis later in this report.
A critical factor in analysing any fracture mechanics problems is the stress state of the
material. The methods for calculating KIc and GIc values for an object under loading
are dependent on whether the object is in a state of plane stress or plane strain.

10
Jayatilaka (1979) describes plane strain as a loading condition where stress
components exist in the x, y and z directions. In a notched specimen, such as the one
shown in figure 3.1, the specimen is loaded in the y direction, perpendicular to the
crack plane. Due to longitudinal extension of the part, the material contracts in the z
direction by Poisson’s effect. This movement is inhibited by the material around it
which causes a stress in the z direction. This effect is reduced up to the very edge of
the specimen, where no stress exists in the z direction due to the material only being
constrained on one side. Plane stress conditions arise when the dimensions of the
specimen in the z direction are negligible in comparison to the other two dimensions.
In this scenario, the stress in the z direction is reduced due to the small amount of
material that prevents contraction, and the reduced Poisson’s effect.
Blanco et al. (2011) conducted a study into the different modes of failure that occur in
CT specimens. These different failure modes can be seen in figure 3.2. FM1 in this
figure illustrates the compressive stress discussed by Hannon (2012) and also
highlights the potential for out of plane bending to occur, FM6 as described by Harris
and Morris (1984). This study was carried out in an attempt to create specimens that
would always fail due to crack propagation before any other failure mode occurred. A
number of alternative designs were tested, and extensive numerical modelling was
carried out to accurately replicate the composite structures. These design
modifications will be investigated further in section 5.2 of this report.
Figure 3.2: Failure sites in compact tension specimens (Blanco et al. 2011)
Out of plane bending, shown by FM6 above, is a common issue with thin laminates.
This bending has been seen in testing by Hannon (2012) and Slepetz and Carlson

11
(1976). As a load is applied to the specimen in the y direction, the fibres at the rear
wall of the specimen are compressed. This may lead to out of plane bending as the
fibres move so as to relieve the stress which can lead to buckling of the fibres,
shearing across the rear wall, or delamination. Slepetz and Carlson (1976) carried out
CT tests on strengthened glass/epoxy and graphite/epoxy specimens similar to those
used by Blanco et al. (2011). The specimens had a smaller crack/width ratio, but a
key difference was the use of 1/8 inch lubricated steel plates inside the clevises to
prevent the out of plane bending seen by Hannon (2012). These plates constrained
movement in the specimen to x and y directions, without affecting the forces acting in
these directions. However, these plates exert a force on the specimen in the out of
plane direction, meaning that the loading conditions experienced by the piece are not
strictly mode 1 fracture, but a mixed mode condition. It is for this reason that plates
were not machined to prevent this motion during the experimentation carried out in
this report. Instead, it is hoped that design modifications discussed in section 5.2 will
aid in preventing out of plane bending.
Kaman (2011) carried out an investigation into the effect of fibre orientation on the
fracture toughness of single edge bend specimens. Tensile testing was carried out on
a number of laminates that contained four plies, two 0o
plies on the outside and two θo
plies in the middle, where θ= 15, 30, 45, 60, 75 and 90 . Numerical modelling of the
setup was used in an attempt to predict the fracture toughness of the laminated
composite to a reasonable degree of accuracy. The calculations used to obtain fracture
toughness experimentally are similar to those used by Pinho (2005) and Jose et al.
(2001), however, a different characteristic equation was used when calculating the
fracture toughness value. This was done to account for the difference in the geometry
of the specimens tested.
It was found that the fracture toughness of the specimens decreased as θ increased.
This would be expected as a larger stress component was being exerted on the matrix
material, which was significantly weaker than the fibres. It is also interesting to see
that the failure that occurred was always parallel with the θo
fibres, as shown in figure
3.3. This is due again to the weak nature of the matrix material with respect to the
fibres. This confirmed the findings of Jose et al. (2001) regarding the fracture of
unidirectional laminates. From this, it would be expected that the crack growth should
follow a path perpendicular to the line of loading in the specimens to be tested in this

12
study, as the lay-up is of a [0/90]ns nature. Fracture toughness values obtained by
Kaman (2011) range from 4405MN/m1.5
to 2488MN/m1.5
. As Harris and Morris
(1984) concluded that fracture toughness decreases with increasing laminate
thickness, it is expected that the specimens in this study would have a considerably
lower fracture toughness as they are eight and sixteen times thicker respectively than
the specimens tested by Kaman (2011).
Figure 3.3: Examples of multidirectional laminated composite failure in tension (Kaman 2011).
The results obtained by Kaman (2011) also suggest that FE analysis can be performed
on composite single edge bend specimens to yield accurate results, with an average
error for a conventional shell element analysis of just 0.43%.
As noted by Harris and Morris (1984), damage at the crack tip of a notched specimen
had a large impact on the fracture toughness of that specimen. O’Higgins et al. (2008)
carried out experimentation regarding the strength of open hole tension specimens of
both glass fibre reinforced plastic and HTA 6376, the same carbon fibre reinforced
plastic to be used in this investigation. As well as comparing the two composites and
their respective properties, this study also highlighted the damage that occurred at the
notch in such specimens. Penetrant radiography was used to visualise the flaws
created in the material at different damage percentages. Iodomethane 99% was
applied to the damaged areas where it filled the voids created during the tensile
testing.
X-rays of the specimens then revealed the extent of the notch damage. It is noted that
most of the damage was a result of cracking in the 0o
and 45o
plies causing the release
of strain energy. Delamination can also be seen around the hole and it is noted that
the damage causes a blunting of the crack tip, which causes a reduction in the stress
concentrations around it. It is also noted that the lay-up has a large bearing on the

13
damage that occurs in the specimen. The [0/90]4s specimens that were tested showed
up far less damage than the [02/902]2s specimens. The authors concluded that the
presence of 0o
fibres between the 90o
plies helps to arrest crack development in the
region due to the reinforcement material being more evenly distributed throughout the
specimen.

14
4. Theoretical Analysis
ASTM Standard E399 Method:
A value for KIc can be obtained from ASTM standard E399 testing using the
following formula:
√
{1}
Where PQ is the load required to fracture the specimen, KIc is the critical fracture
toughness of the specimen, aQ is crack length at fracture, and Y(a/W) is a function of
specimen geometry such that:
( )
( ) ( ) ( ) ( )
{2}
The formula for Y (a/W) listed above is only validated for specimens of the geometry
shown in figure 4.1. For other geometries, different formulae may be required.
Figure 4.1: Critical dimensions used in KIc calculation according to ASTM standard E399. (ASTM, 1997)
The strain energy release rate of a body is the rate at which strain energy built up in
the body is converted into the surface energy on the faces of a new or extended crack
(IME 1978). This value for KIc obtained from Eq. 1 above can then be used to
calculate GIc, the critical strain energy release rate, using Eq. 3 outlined overleaf.

15
√
√√ {3}
Where E11 and E22 represent the Young’s Modulus of the laminate in the directions
parallel and perpendicular to the direction of loading respectively; υ12 is the Poisson’s
ratio of the laminate and G12 is the shear modulus of the laminate.
J-Integral Method:
The J-integral is commonly found in practice using FE software. An example of this
software, ABAQUS 2.0 is used in section 5 of this report. The FE solver calculates
the J-integral using information about the loading conditions and the plastic
behaviour of the material in front of the crack tip. Material properties and geometry
are compared and a final value for the J-integral is given. In plane stress conditions,
the J-integral is equal to the value for GIc which can then be filled back into Eq. 3
above to obtain a value for the fracture toughness, KIc, of the material. This is a
complex procedure and requires very accurate modelling of a composite laminate
which is beyond the scope of the FE study carried out during this investigation.
(CTOD)cr Method:
Fracture toughness can be calculated using the data gathered from clip on
displacement (COD) gauges. The following formula is obtained from Harris and
Morris (1985):
√ {4}
Where (CTOD)cr is the crack tip opening displacement at fracture, CODcr is crack
opening displacement measured during experiment, ao is the initial crack length and
do is a distance parameter used to define the damage that forms around a crack under
mode 1 loading. As this value can be very difficult to measure experimentally, the
following expression empirically proven by Poe (1983) can be used to find do:
√ ⁄ {5}

16
Where Qc is the general toughness parameter of the material and εut is the ultimate
tensile strain of the fibres. The general toughness parameter, which is a property
independent of lay-up, defines the critical strains in load carrying plies. According to
Poe (1983), the relationship between Qc and εut can be assumed to be a material
constant for all composite materials at 1.5 mm0.5
. This value can be used to then work
out do and subsequently calculate (CTOD)cr.
Once (CTOD)cr values are known, they can be filled into Eq. 6 to obtain the fracture
toughness of the material.
{6}
Where E represents the Young’s modulus in the direction of loading, Kt is the fracture
toughness of the composite, and σo is representative of the strength of an un-notched
laminate. The un-notched strength of a composite laminate can be found by carrying
out a tensile test on a specimen containing no pre-machined flaws. However, this was
not an experimental objective at the time of specimen manufacture and so a scaled
value obtained from O’Higgins et al. (2007) was used. Alternatively, un-notched
laminate strength can be estimated by finding the product of the ultimate tensile strain
of the fibres in the direction of loading and the Young’s modulus of the material.

17
5. Finite Element Analysis
As part of this study, a qualitative analysis of the strains induced in the specimens
during testing was carried out. As well as aiding in visualising the stress fields across
the laminate, this analysis also provided information which helped in choosing the
new 32 ply specimen geometries to be tested. It was hoped that these alterations
would reduce the out of plane bending and compressive stresses on the rear wall of
the specimens which caused invalidation of the results of Hannon (2012). Finite
element (FE) analysis is a powerful tool for assessing the behaviour of composite
materials. As mentioned in the theory section of this report, FE analyses can be used
to calculate the J-integral value for composites during a fracture scenario which can
in turn be used to calculate the critical strain energy release rate, GIc, and fracture
toughness, KIc.
5.1. Specimen Modelling
The first step in this process was to model the original specimens used by Hannon
(2012) and Pinho (2005). ABAQUS 2.0 FE software was used for this analysis and
the sketch of the original geometry was created. In order to accurately model
composite materials, the specimen had to be modelled using conventional shell
elements similar to those tested by Kaman (2011). While this is not fully
representative of the 32 ply specimens as it does not take thickness into account, it
was deemed acceptable due to the thin nature of the specimen. This investigation
aimed to find suitable alterations to specimen geometry, so the absolute values for
stress and strain obtained from the study were not the main focus. Rather, it was the
proportion of these stresses and strains throughout the specimen that was of most
interest.
The material was then assigned a lay-up of 32 plies alternating between 0o
and 90o
and symmetrical about the centre ply. A full list of the material properties can be seen
in appendix B. An image of the laminate and lay-up used can be seen in figure 5.1.1.
By using the composite lay-up feature in the software, it is possible to assign
orthotropic properties to a laminate. Each ply was given properties for Young’s
modulus in the longitudinal and transverse directions, as well as shear modulus and
Poisson’s ratio values. The plies were then assigned an orientation with respect to
fixed reference axes. The orientation of the fibres in the model created on ABAQUS
can be seen in figure 5.1.1 by the red lines running perpendicular to each other.

18
Figure 5.1.1: Image of lay-up and partitions used in FE study of 64 ply laminate.
The pre-machined crack was present in the material at this stage and partitioning of
the specimen was completed to provide an even and dense mesh, which provides the
most accurate results. Once the specimen was created and all relevant material
properties assigned, the loading conditions and mesh had to be applied to the part. An
arbitrary load of 30 kN was applied to the specimens. This was applied as two 15 kN
loads acting around the faces of the loading holes so as to cause mode 1 failure. Once
this was done, a mesh of the part was created. This mesh can be seen in figure 5.1.2.
The element type used was an S4R element, a 4-node doubly curved thin shell with
reduced integration, hourglass control and finite membrane strains.
Figure 5.1.2: Image of mesh used on original geometry in FE study

19
5.2. Design Modifications
The four design modifications that were tested were chosen based on investigation
carried out by Blanco et al. (2011) where the six main causes of failure in the
specimen other than mode 1 fracture are discussed. The new geometries aimed to
tackle the compressive forces on the sides of the specimen as well as on the rear wall.
As out of plane bending is caused by the compression and buckling across the rear
wall of the specimens, it was also hoped that the reduction in compressive stress
would result in the out of plane motion during testing being reduced. A rounded rear
wall is investigated due to the good load distribution qualities of circular structures.
As well as this, three other variations of chamfered specimens were tested which can
be seen in figures 5.2.1 (c)-(e).
The results of the FE analysis are given in figure 5.2.1 (a)-(e) and in table 5.2.1.
ABAQUS applies a colour scale to the strain fields based on the maximum and
minimum strains through the model. Areas of bright red indicate high tensile strain
and areas of dark blue indicate high compressive strain. As it is the compression in
the direction of the 0o
plies that is of interest in this section, the strain plotted is ε11.
From the analysis of these new designs, it was noted that the specimen with the
bevelled rear wall caused significantly less compression in the rear wall than the
standard test specimen (see figure 5.2.1 (b)). The two designs with the chamfered rear
wall both gave very similar results, and it was decided that specimen (c) would be
manufactured. These geometries can be seen in figures 5.2.1 (c) and (d). The final
design shown in figure 5.2.1 (e) with the tapered rear wall showed a reduction in
stress over the standard specimen, however the change in ε11 obtained was smaller
than the other alternatives. There was also the increased risk of the loading pins
shearing out of the load holes as detailed in Blanco et al. (2011).
Table 5.2.1: Comparison of tensile and compressive strains for alternative geometries.
Specimen Strain at Crack Tip (μe) Strain at rear wall (μe)
Standard Square 4.598x108
-2.612 x108
Bevelled 7.645 x108
-2.064 x108
Chamfered 9.218 x108
-2.149 x108
Double Chamfered 9.228 x108
-2.146 x108
Tapered 9.316 x108
-2.235 x108
Dimensioned drawings of the three specimens chosen can be found in appendix A.

20
Figure 5.2.1 (a): Standard square specimen.
Figure 5.2.1 (b): Bevelled specimen.

21
Figure 5.2.1 (c): Chamfered specimen.
Figure 5.2.1 (d): Double chamfered specimen.

22
Figure 5.2.1 (e): Tapered specimen.

23
6. Experimental Work
6.1. Specimen Manufacture
The first step in carrying out the experimental testing in this study was to
manufacture the test specimens. The geometry of these specimens is shown in figure
3.1 in section 3. This geometry was chosen to match the specimens used by Pinho
(2005) and Hannon (2012). The only alteration to this design was the thickness of the
laminate. In both studies, 32 ply laminates were used to make the test specimens. This
test shall investigate one 32 ply specimen to verify that testing was carried out under
similar conditions to these studies, and three thicker 64 ply specimens. In an attempt
to reduce the compressive stresses on the rear wall of the specimen experienced by
Hannon (2012), two alternative 32 ply specimen geometries are also tested. Details of
these new designs are outlined in Section 5.2. The 32 ply specimens were already
available from a previous study which meant only the 64 ply specimens needed to be
manufactured from the lay-up stage. Both types of specimen required machining to
the desired geometries.
The material used in this study was a HTA-6736 composite material. This material
comes in the form of a unidirectional roll of material which is pre-impregnated with
resin which cures at high temperatures. A list of the elastic properties of the material
properties are given in table 6.1, with a detailed list of properties listed in Appendix
B. The rolls of material are stored in cold rooms adjacent to the composites laboratory
in the University to avoid the material becoming tacky, which would make the lay-up
process more difficult. The lay-up used in this panel was of a [0/90]16s symmetric and
balanced design. This resulted in an 8 mm thick specimen, as each ply measures
0.125 mm in thickness. To get laminates from which test specimens can be machined,
a large panel of the composite material had to first be manufactured. It was decided
that 3 specimens were to be obtained from the panel, so a panel of 210 mm x 80 mm
was constructed through a hand lay-up process.
Table 6.1.1: Material properties for HTA 6376.
Material HTA 6376
E11 (GPa) 139
E22 (GPa) 10
G12 (GPa) 5.2
v12 0.32

24
This involved cutting 64 pieces from the sheet of material, 32 of which had the fibres
running parallel with the 80 mm side, and 32 running parallel with the 210 mm side.
Care was taken to ensure that all pieces were cut accurately and also that the amount
of scrap material produced was minimised. The roll holds the pre-preg fibres between
a sheet of thin card and plastic to prevent the fibres from adhering to adjacent layers.
The pieces of material were placed into stacks according to fibre orientation and a
chart was drawn to ensure the correct ply was placed on the panel at the correct time.
This was an important step, as removing a ply from the panel once it had been placed
down can cause damage to other plies and render the panel unusable.
Each ply was placed alternately on the panel and then flattened using a small roller to
remove any air pockets that would cause delamination in the finished panel. After
every 8 plies, the panel was placed into a vacuum bag which was used to consolidate
the panel and make it more compact. This bag can be seen in figure 6.1.1. This
process was repeated until a 64 ply panel was produced, which was then stored in a
sealed plastic bag to wait for curing. This bag was kept in the cold room with the roll
of material to prevent the resin from setting.
Figure 6.1.1: Vacuum bag used in lay-up process.
To cure the panel, the autoclave shown in figure 6.1.2 was used. The panel was
placed inside a vacuum bag which applied a pressure of 7 bar to the composite while
it was cured inside the autoclave at 180o
C for 2 hours. A thin sheet of Teflon was
placed over the panel to prevent it from adhering to anything else in the vacuum bag,
while a breather fabric was placed on top to allow an even vacuum to be achieved
across the bag.

25
Figure 6.1.2: Autoclave used to cure panel.
Once the finished panel had been created, the three test samples had to be cut from it.
This was done using the composite cutter located in the University workshop, shown
in figure 6.1.3. This machine uses a circular cutting blade on a crosshead to cut
through the material. Composite materials are difficult to cut, which can generate a
lot of heat. To counteract this, coolant is constantly delivered to the cutting area. Due
to the forces exerted on the cutting blade, the measured distance on the digital display
and the actual distance of the blade from the guide can vary. For this reason, a piece
of scrap composite material was used to calibrate the machine. The first step in
cutting the 65 mm x 60 mm specimens was to create two square edges from which
measurements can be taken. Once this was completed, the three specimens were cut
to size. PPE (earplugs, a dust mask and goggles) was worn to prevent any injury
during the process, and the guide for positioning the panel was checked before each
cut to ensure nothing was caught between it and the flat surface of the panel.
Figure 6.1.3: Composite cutter used to machine specimens.

26
Carbide tooling bits were required to machine the specimens to the desired geometry.
For this reason, this part of the manufacture was dealt with by technicians at the
University workshop. Drawings were generated for the three 64 ply specimens and
the two 32 ply specimens that were to be altered. These drawings were submitted to
the workshop and completed in early February. The two loading holes in the
specimens were created by drilling through two sacrificial pieces of plastic which
prevent delamination from occurring on the outer plies of the specimen. The 4 mm
slot in the side of the specimens was created using a carbide coated milling bit which
is designed especially to cut through composite materials. These steps were only
required for the 64 ply specimens as both 32 ply pieces had these features already
machined. The changes in geometry to the 32 ply specimens were carried out using a
computerised machine designed specifically for the exact cutting of complex shapes
into composite materials. A thin starter crack had to be inserted in the specimens.
This was done using the wafering saw shown in figure 6.1.4. A thin circular blade of
0.2 mm thickness was used at a speed of 530 rpm to slice through the material to a
depth of 25 mm. Finally, a small razor blade was used in a sawing motion to give the
crack a final root radius of 0.05 mm.
Figure 6.1.4: Wafering saw used to insert pre-crack in specimens.
The specimens were then prepared for testing by applying the speckled surface that
would be required for the DIC investigation. This finish can be seen in figure 6.1.5.
Inside a fume hood, a base coat of white primer was applied to the specimens and
allowed to dry. This process was repeated a second time and black spray paint was
used to produce the black spots on the surface. The can was held almost 40 cm from
the specimens and lightly sprayed to produce a fine mist that speckled the pieces.

27
These were left to dry to prevent smudging of the paint which would cause the DIC
readings to be less accurate.
Figure 6.1.5: Specimen after speckling process.
Knife edges were cut to size and adhered to the specimen using an epoxy mixture.
The blades were scratched to give the epoxy a better surface to adhere to and clamped
to ensure a good bond. These knife edges can be seen in figure 6.1.6.
Figure 6.1.6: Knife edges applied to specimens.
6.2. Experimental Apparatus
The mechanical testing in this project was carried out on a Dartec 100kN tensile
testing machine located in the composites laboratory. An image of this tester can be
seen in figure 6.2.1 in which all major components are shown. The machine was then
hooked up to software which was used to operate the equipment while also recording
the data obtained during a test. During a tensile test, the machine’s crosshead moves
at a constant speed while a load cell on the crosshead measures precisely the amount

28
of force applied to the test specimen. The load cell had a resolution of 0.001N and
measured the load ten times per second. Two clevises were used to hold the specimen
during testing and were held in position by load blocks.
Figure 6.2.1: Dartec 100 kN tensile tester.
Clip-on displacement (COD) gauges were attached to the specimen to track the
opening of the crack during testing. Two knife edges had to be fixed to the front of
the specimen into which the COD gauges were placed. Figure 6.2.2 shows the COD
gauge when attached to the specimen before loading. The COD gauges have two
cantilever metal panels in them which are connected to an LVDT which measures the
displacement of each panel and converts this to an electrical signal which was picked
up by System6000 software and recorded. This software was also hooked up to the
load cell on the Dartec machine which allowed the crack opening displacement to be
measured against the load applied to the specimen.
Figure 6.2.2: COD gauge attached to specimen.
Load cellLoad blocks
Clevises
Crosshead

29
Two dimensional DIC was used to monitor the strains induced in the specimens
throughout loading. This system involved a high resolution camera tracking the
movements of numerous reference points on the surface of the specimen and then
compared these to a datum image. From these images, the displacement of a point
from its original position can be measured and the strain at that point can then be
calculated.
6.3. Experimental Procedure
The experiment took place in the Composites Research Laboratory in the University.
The DIC rig was setup beside the Dartec machine and System6000 data acquisition
software and the camera was adjusted to view the specimen perpendicular to the front
face of the specimen. This ensured that the images picked up by the camera were
proportionate. The DIC system was connected to the Dartec tester which allowed the
images recorded by the camera to be associated with a particular load reading. The
cameras were focused to give the clearest image, and the exposure was adjusted to
give the optimum brightness for the testing.
The COD gauge was then calibrated. This involved pinching the two arms of the
gauge together to give a zero reference point, it was then allowed to extend to a
known value and this was also set. The gauge was then placed between the knife
edges on the specimen and the software was prepared for recording. The System6000
software was also connected to the Dartec machine to allow the force applied to the
specimen to be compared with the crack opening displacement.
The Dartec machine was set so that all initial values for applied load and extension
were zero and the crosshead speed was set to 5 mm/min as outlined in Pinho (2005).
The DIC system and System6000 then began recording and the tensile test was then
started by the Dartec. Each test lasted between 2-4 minutes and once failure of the
specimen could be seen occurring on the Dartec load cell read out, the test was
stopped and all data recording equipment was paused. The raw data extracted from
the test equipment was then compiled and prepared for analysis.

30
7. Results and Discussion
The results obtained from the six specimens tested are outlined in the sections below.
Values for the fracture load and (CTOD)cr were recorded and analysed and the
calculated fracture toughness and critical strain energy release rates are discussed as
well. Images of the fracture zones of the specimens taken through a microscope and
DIC results are analysed to provide insight into the failure mechanisms within the
laminate. Table 7.1 lists the critical measurements of the specimens taken prior to
testing. All specimen drawings can be found in appendix A and a full list of the data
acquired from all of the specimens can be found in Microsoft Excel file format on the
CD appendix D.
Table 7.1: Critical specimen dimensions for ASTM procedure.
Specimen a (mm) W (mm) (a/W) Y(a/W) B (mm)
Square specimen 24.9 50.2 0.496 9.541 4.17
Bevelled Specimen 25.4 50.3 0.505 9.810 4.16
Chamfered Specimen 25 49.8 0.502 9.719 4.17
Specimen A 25.7 50.6 0.508 9.902 8.14
Specimen B 25.6 51.1 0.501 9.689 8.14
Specimen C 26 51.3 0.507 9.871 8.14
7.1. 32 Ply Square Specimen.
The first specimen tested was a 32 ply specimen of the geometry laid-out in ASTM
standard E399. The only difference between the two geometries is the decreased
loading hole diameter to reduce the likelihood of failure by shearing as mentioned by
Pinho (2005). This specimen is identical to the specimens tested by Hannon (2012)
during his study and was tested primarily to provide a benchmark against which the
other specimens could be compared. However, this specimen also serves to prove that
the testing carried out closely matches that completed by Hannon (2012). Figure 7.1.1
shows the relationship between the load applied to the specimen and its extension.
The sudden decrease in load which signifies the point of failure can be clearly seen
and occured at 5456.43 N.

31
Figure 7.1.1: Load vs. extension for 32 ply square specimen.
This load is quite similar to the maximum load sustained by identical specimens
tested by Hannon (2012) with a deviation of just 1.02%. This proves that the testing
carried out in this report matches the work carried out during that study. Calculating
fracture toughness KIc using the formulae outlined in ASTM standard E399, a value
of 55.7 MN/m1.5
is obtained. The critical strain energy release rate for this specimen
was 240.991 kJ/m2
. Both these values are similar to those seen by Hannon (2012).
The graph maintains good linearity up to fracture, which is similar to the findings of
Jose et al. (2001). However, neither of these values are truly accurate measures of
fracture toughness due to the specimen failing by compression of the rear wall rather
than crack propagation. For this reason, the actual fracture toughness of this specimen
would be expected to be higher than the experimentally obtained value as the
laminate would have sustained a higher load before fracturing.
Microscopic Analysis:
The failure of this specimen can be seen in figure 7.1.2 overleaf. It can be seen in this
image that the specimen failed at an angle close to 45o
across the rear wall. This is
due to the plane stress nature of loading across the specimen. This type of cracking is
expected and is due to maximum shear stress occurring at 45o
to the principle axes.
However, there is also evidence of compressive failure in this specimen, including
delamination [1], and shearing of the fibres [2]. This replicates the results found by
Hannon (2012). The test is therefore invalidated due to failure not occurring due to
propagation of the pre-machined crack under mode 1 loading. Small amounts of
0
1000
2000
3000
4000
5000
6000
0.004
0.219
0.443
0.667
0.890
1.114
1.336
1.559
1.782
2.006
2.228
2.453
2.675
2.899
3.122
3.347
3.572
3.797
4.023
4.247
4.472
4.697
4.921
Load(N)
Extension (mm)
Load Vs. Extension Square Specimen
Load Vs. Extension
PQ=5456.43 N

32
delamination can be seen but shearing of the fibres is the most prominent failure
mode seen in this specimen. Fibrous composites are much weaker in compression
than they are under a tensile load. This causes the laminate to fail at the rear wall
before the stress intensity factor at the crack tip reaches the critical value required for
fast crack propagation.
Figure 7.1.2: Failure of 32 ply square specimen.
Figure 7.1.3 was taken after the specimen had been separated into two individual
limbs. It shows the fracture surfaces of the specimen and also highlights the area of
subcritical crack damage that occurred during testing prior to failure.
Figure 7.1.3: Fracture surfaces of 32 ply square specimen.
In the image on the left, the three distinct areas of crack growth can be seen, with [1]
fast fracture zone, [3] subcritical notch-tip damage and [3] pre-machined crack zones
all visible. The image on the right is a greater magnification of the transition zone
between subcritical notch-tip damage and fast fracture. It was observed that the area
of subcritical notch-tip damage was characterised by a rough fracture surface, with 0o
fibre fracture occurring at different lengths in the fibres across the crack area. The fast
fracture zone by comparison was a much smoother surface with delamination and
fibre buckling evident across the rear wall of the laminate.
1
2
1 2 3 1
2

33
Crack Opening Displacement:
The crack opening displacement data obtained during the test is shown in figure 7.1.4
below. Similar to figure 7.1.1, the point of failure can be seen quite clearly in this
graph. The displacement measured by the COD gauge at the point of failure was 2.78
mm. However, shortly after failure occurred, the bond between the top knife edge and
the specimen was broken, and so the gauge opened fully and all results beyond that
point were invalid. This was not a major problem however, as the drop off in load
after failure was recorded accurately.
Figure 7.1.4: Load vs. crack opening displacement for 32 ply square specimen.
The value for crack opening displacement was used in conjunction with formulae
from Harris and Morris (1985) and Poe (1983) to obtain values for the (CTOD)cr at
failure. These formulae can be seen in the theory section of this report. These values
were also compared with images taken using the DIC equipment to confirm their
accuracy. A (CTOD)cr of 0.455 mm was calculated for this specimen when using the
COD gauge information. However, when analysed using the DIC images, it was
found that the actual (CTOD)cr was 0.25 mm. There are a number of reasons that are
likely to have caused this error.
As outlined in the theoretical section of this report, the value for the general
toughness parameter, Qc/εut, is taken to be 1.5 mm0.5
from Poe (1983). This value was
obtained empirically from laboratory experiments and is assumed to be reasonably
accurate for most composite materials, but outlying results are possible. This value
for general toughness parameter was calculated using several different material types,
0
1000
2000
3000
4000
5000
6000
0.016
0.019
0.098
0.25
0.407
0.582
0.777
0.99
1.214
1.445
1.678
1.914
2.151
2.381
2.66
2.963
3.196
3.432
3.773
4.028
Load(N)
COD Extension (mm)
Load Vs. COD Square Specimen
Load Vs. COD

34
and lay up patterns to cover a broad range of possible panels. However, no panel
greater than 16 plies in thickness was tested, which could lead to errors in the
calculated value as the mechanics of fracture change significantly once plane strain
conditions are met. It is also mentioned by Poe (1983) that the general toughness
parameter of 1.5 mm0.5
may not be accurate for specimens experiencing low-medium
levels of notch damage.
The un-notched laminate strength for a HTA 6376 [0/90]8s laminate is based on scaled
results from O’Higgins et al. (2007) where a [0/90]4s laminate was tested. A more
accurate result could have been obtained for this value if an un-notched specimen had
been manufactured and tested along with the test pieces. Another possible means of
calculating this value is to use the product of the ultimate tensile strain of the fibres
and the Young’s modulus of the fibres to find the stress which would cause tensile
failure. However this method is less accurate as it does not account for any stress in
the matrix material, and assumes that all fibres are in ideal condition, which would
yield the highest possible value for un-notched laminate strength.
As well as errors in the theory, experimental errors could have occurred while using
the COD gauges. As mentioned earlier, the bond between one of the knife edges used
and the specimen broke shortly after failure occurred. This suggests that some
deformation may have occurred in the adhesive prior to breaking, which would affect
the values of crack opening displacement as they are very precise measurements.
Also, the out of plane bending seen in this specimen could have led to inaccuracies in
the COD gauge readings.
The value for fracture toughness using values for (CTOD)cr calculated from COD
gauges for this specimen is 223.8 MN/m1.5
. When the value of 0.25 mm for (CTOD)cr
is used, KIc is found to be 165.204 MN/m1.5
. Both of these values are significantly
higher than the KIc given by ASTM standard E399 of 55.7 MN/m1.5
and this error can
be attributed to any number of causes listed above.
Digital Image Correlation:
DIC results for the 32 ply square specimen are shown in figure 7.1.5. The DIC
software used was LaVis® 7.4, created by LaVision®. The two images show the
strain fields around the area of interest at the point of fracture (a) and after significant
fracture has occurred (b).

35
Figure 7.1.5: Strain fields in 32 ply square specimen.
Figure 7.1.5a shows the large amount of strain that occurs around the region of the
crack tip up to and during failure of the specimen. This is the region which
experienced propagation of the pre-machined crack prior to failure of the specimen. It
is this propagation which causes the damage at the tip of the crack. Figure 7.1.5b
shows the specimen shortly after major compressive failure along the rear wall. This
can be seen clearly in the image by the dark blue colour concentrated around the
centre of the rear wall. Each image taken by the DIC system calculates the maximum
and minimum strains experienced by the specimen in a given frame, and then scales a
colour coded legend to fit. For this reason, images of the specimen prior to significant
deformation can be misleading as a very minute strain gradient across the specimen
will still have a full colour scale applied to it, which can give a user the impression of
much greater deformation. From this point onwards, only images of the specimens at
initial crack growth and after significant propagation shall be investigated.
7.2. 32 Ply Bevelled Specimen.
This was the first of the modified specimens to be tested. It was hoped that the
circular rear wall would distribute the compressive forces throughout the specimen
more evenly than the original square specimen geometry. The load vs. extension data
can be seen in figure 7.2.1. Again, the point of initial fracture can be seen in the graph
and the test was continued until the catastrophic failure occurred.

36
Figure 7.2.1: Load vs. extension for 32 ply bevelled specimen.
The failure load for this specimen was lower than that of the square specimen at just
4493.37 N, but this decrease is marginal. This is likely due to the reduction in
compression at the rear wall which would otherwise act to arrest crack development
(Harris and Morris 1984). Reduction in compression at the rear wall was an objective
of this test, and the associated decrease in the fracture load of the specimen is
accounted for in the ASTM method for KIc determination. When this method is used,
a KIc value of 52.013 MN/m1.5
and a GIc value of 210.143 kJ/m2
are obtained. Again,
due to failure in compression across the rear wall of the specimen, the point at which
the load begins to drop in this graph is not an accurate representation of the load
which would cause fibre rupture. As a result, the experimental value for KIc is found
to be conservative.
A magnified image of the damage that occurred in the bevelled specimen can be seen
in figure 7.2.2. Upon initial visual inspection, this specimen appeared to have failed
in the x-z plane, which would be indicative of plane strain failure and also of mode 1
fracture. However, when viewed closely under a microscope, it can be seen that the
fibres sheared due to compression, this time in intermittent sections of ± 45o
. This is
consistent with the plane stress nature of the specimen.
0
500
1000
1500
2000
2500
3000
3500
4000
4500
5000
-0.007
0.495
1.006
1.517
2.027
2.538
3.050
3.565
4.079
4.591
5.104
5.615
6.128
6.640
7.153
7.667
8.178
8.690
Load(N)
Extension (mm)
Load Vs. Extension Bevelled Specimen
Load Vs. Extension
PQ=4493.37 N

37
Figure 7.2.2: Failure of 32 ply bevelled specimen.
Fibre fracture can be seen in the area around the crack tip of the specimen with fibre
pull out being a significant feature of the damage at the crack tip. Extensive widening
of the pre-machined crack is also evident along with delamination across the rear wall
[1]. The fracture surfaces of the specimen are shown in figure 7.2.3. The image on the
right highlights the areas of delamination [1] and shearing [2] which were caused by
the compression across the rear wall. The image on the left shows the transition from
subcritical crack-tip damage to fast fracture which occurred 17 mm from the rear wall
of the specimen.
Figure 7.2.3: Fracture surfaces of 32 ply bevelled specimen.
Figure 7.2.4 overleaf shows how crack opening displacement varied with applied
load. The crack opening displacement measured 3.062 mm at the point of initial
fracture for this specimen. A similar problem arose with this specimen whereby the
knife edges broke away from the specimen during loading. Again, this occurred after
failure of the piece and as a consequence no critical results were affected. This
resulted in a (CTOD)cr of 0.513 mm which was again compared with values obtained
from DIC images to confirm whether it was an accurate measurement. From the DIC
1
1
2

38
images, a (CTOD)cr of 0.25 mm is measured. The disagreement between the two
values can be attributed to the same factors outlined in section 7.1 of this report.
Using the visually acquired value for (CTOD)cr, fracture toughness of the material is
found to be 165.2 MN/m1.5
which is much larger than the value of approximately 110
MN/m1.5
found by Pinho (2005) for a specimen with similar material properties.
Figure 7.2.4: Load vs. crack opening displacement for 32 ply bevelled specimen.
Images of the strain fields plotted for this specimen can be seen in figure 7.2.5 below.
Due to errors that were encountered when processing the strain fields with a circular
border to match the contours of the specimen, it was decided that a rectangular area
would be processed. Although this does affect the accuracy of the results marginally,
the stress concentrations around the crack tip, and compression near the rear wall are
still quite clear.
Figure 7.2.5: Strain fields in 32 ply bevelled specimen.
0
500
1000
1500
2000
2500
3000
3500
4000
4500
5000
0.044
0.31
0.595
0.894
1.207
1.528
1.86
2.204
2.53
2.892
3.368
3.771
4.146
4.513
4.86
5.137
5.425
5.782
Load(N)
COD Extension (mm)
Load Vs. COD Bevelled Specimen
Load Vs. COD
(a) (b)

39
As can be seen in figure 7.2.5 (a), the area that is investigated, which lies outside of
the specimen, shows areas of severe tension and compression. This is due to the
inconsistency in the patterns picked up by the cameras between the speckled surface
and dark background. As the specimen is loaded however, this error is reduced as the
compression at the rear wall of the specimen grows larger. Also, the out of plane
bending that occurred in this specimen caused large errors in the strain fields shortly
after image (b) was taken.
7.3. 32 Ply Chamfered Specimen.
This specimen was the second to be modified and had the two corners of the rear wall
cut away, so as to leave a 20 mm long portion of the rear wall remaining, and 35 mm
long side walls. Similar to the bevelled specimen in the previous section, it was hoped
that the new geometry would distribute the compressive forces on the rear wall more
evenly through the body than the original square geometry. The chamfered specimen
failed at a load of 3827 N, which is significantly lower than the other two 32 ply
specimens. Strong linearity was seen up to fracture which indicates steady damage
initiation at the crack tip up to PQ. The load vs. extension graph for this specimen can
be seen in figure 7.3.1. This test was again run until significant damage occurred
which would make post-test analysis of the fracture surfaces easier to carry out.
Figure 7.3.1: Load vs. extension for 32 ply chamfered specimen.
The fracture toughness of this specimen according to ASTM standard E399 is 40.048
MN/m1.5
and GIc was found to be 124.581 kJ/m2
. These values are significantly lower
than the values recorded for the 32 ply square and bevelled specimens. This can be
0
500
1000
1500
2000
2500
3000
3500
4000
4500
0.002
0.945
1.896
2.844
3.797
4.750
5.705
6.655
7.610
8.560
9.515
10.465
11.420
12.370
13.324
14.274
15.227
Load(N)
Extension (mm)
Load Vs. Extension Chamfered Specimen
Load Vs. Extension
PQ=3827 N

40
explained by the change in geometry of the specimen. In the formula used to calculate
the KIc value for the laminate, the term Y(a/W) is dependent on the geometry of the
specimen. It relates the crack length to the overall width of the specimen but the ratio
with which these two measurements are compared differs from one geometry to the
next. It is possible that the relationship that is validated for the square geometry
outlined in ASTM standard E399 gives a larger Y(a/W) value for square laminates
than for chamfered laminates.
Similar to the two previous 32 ply specimens, the chamfered specimen also failed due
to shearing of the fibres across the rear wall. This shearing can be seen in figure 7.3.2.
Again, the shear took place at approximately 45o
across the wall. Very little
delamination can be seen in this specimen with the two delaminations occurring due
to the movement of the sheared fibres across each other. A much larger crack is seen
in this specimen when compared with the two previous tests. Fibre pull-outs are
visible and occurred across both limbs of the specimen.
Figure 7.3.2: Failure of 32 ply chamfered specimen.
Figure 7.3.3 shows the fracture surfaces of the chamfered specimen. A trend can be
seen throughout all 32 ply laminates, whereby a distinct change can be seen between
the slow progressing crack associated with the notch-tip damage and the unstable fast
crack propagation zone towards the rear of the specimen. It was observed that this
specimen has a much shorter fast crack propagation area with the transition line [1]
occurring only 10 mm in from the rear wall. Compression was the cause of failure
again, evidence of which can be seen by the fibre pull-outs [2] in one of the 90o
plies
which is caused by matrix shearing (Slepetz and Carlson 1976).

41
Figure 7.3.3: Fracture surfaces of 32 ply chamfered specimen.
Out of plane deformation can be seen in this specimen similar to the deformation of
the square laminate. This bending can be seen in figure 7.3.4 below which compares
the bending of the three 32 ply test specimens.
Figure 7.3.4: Out of plane deformation of the three 32 ply specimens.
It can be seen from the images above that all three 32 ply laminates tested in this
investigation experienced a large degree of out of plane bending. This is due to the
compression towards the rear of the specimen which causes buckling in the fibres.
This buckling tends to move the fibres into a position out of the vertical plane which
exposes them to stress components transversely across the fibres. This is the force
which causes the limbs of the specimens, [1] and [2], to move in a direction
perpendicular to the plane of bending.
The crack opening displacement curve for this specimen is quite similar to the other
32 ply specimens. The displacement measured at failure in this specimen was 3.29
mm, the largest of the three specimens. The bond between one of the knife edges and
1
2
1
2

42
the specimen was again broken, but in this instance, it appears to have been the
sudden increase in the crack size that put too much pressure on the bonds, rather than
a more gradual increase in pressure as seen with the other two specimens. The details
of the crack opening displacement are highlighted in figure 7.3.5 below.
Figure 7.3.5: Load vs. crack opening displacement for 32 ply chamfered specimen.
Visually inspected (CTOD)cr values obtained from DIC images are again significantly
smaller than the value of 0.552 mm calculated from COD gauges. At 0.1875 mm, the
(CTOD)cr from the DIC images gives a fracture toughness of 143.2 MN/m1.5
, the
lowest KIc value calculated through this method.
The images of the strain fields at fracture and just prior to catastrophic failure at the
rear wall of the chamfered specimen can be seen in figure 7.3.6. The images are
largely accurate in visualising strain, however movement of the test piece in the frame
during testing resulted in slight errors near the edges of the chamfers of the specimen
[1] and [2]. A large compression zone can be seen forming around the rear wall of the
specimen in a similar fashion to the earlier tests. It is noted however that the
compression appears to be better distributed throughout the specimen. This could
explain the smaller load required to break the sample, as less compression around the
crack tip area will result in cracks propagating more easily.
0
500
1000
1500
2000
2500
3000
3500
4000
4500
0.018
0.02
0.024
0.03
0.105
0.285
0.498
0.727
0.973
1.233
1.501
1.778
2.06
2.351
2.644
2.939
3.231
3.551
3.875
Load(N)
COD Extension (mm)
Load Vs. COD Chamfered Specimen
Load Vs. COD

43
Figure 7.3.6: Strain fields in 32 ply chamfered specimen.
The larger crack opening displacement measured by the COD gauges before failure
can be explained by the reduction in the percentage of the applied force that acts in
compression which means that more of the strain energy applied is used to propagate
the crack than before.
7.4. 64 Ply Specimen A.
This specimen was the first of the 64 ply samples to be tested. The geometry of this
specimen is identical to the 32 ply square specimen in section 7.1, but with a 64 ply
lay-up. The increase in thickness of the specimen should more closely represent a
plane strain loading scenario and be less inclined to fail due to compression and
shearing across the rear wall of the specimen. As the thickness of the specimen
increases, the fracture toughness should decrease, in accordance with linear elastic
fracture mechanics theory, and findings by Jose et al. (2001). Although the amount of
force required to cause failure of the specimen increases, the larger amount of
material at the rear wall of the specimen may reduce the effects of the compressive
stresses felt in this region. It can be seen in figure 7.4.1 that the specimen failed at
11.12 kN which is just over twice as large as the failure load sustained by the 32 ply
specimen of the same geometry.
1
2

44
Figure 7.4.1: Load vs. extension for 64 ply specimen A
The fracture toughness for specimen A was calculated to be 59.9 MN/m1.5
and GIc
was calculated as 278.705 kJ/m2
. These values are quite similar to the values obtained
for the 32 ply specimen of the same geometry. However, these values cannot be
taken as true values for fracture toughness and critical strain energy release rate as
large scale fracture of the 0o
fibres did not take place. The drop in load seen in figure
7.4.1 above was caused by shearing of the fibres along the rear wall of the specimen
in a similar fashion to the thinner specimens tested in this study. DIC was not used
during the testing of this specimen, and as a result no video images of the specimen
are available. An explanation for the unusually high load sustained by the specimen is
that the compression throughout the specimen prevented the crack from propagating
at all. At the crack tip, some fracture of the matrix in 90o
plies can be seen, but most
of the fibres remained intact or fractured in an area away from the crack zone. As a
result, none of the strain energy applied to the specimen was used to create new crack
surfaces, and was instead transferred to the load cell. Had less compressive stress
been present throughout the specimen, crack propagation would have occurred at a
lower load. As a result, a larger value for the KIc of the specimen was obtained than is
likely to be accurate.
This specimen experienced no noticeable out of plane deformation, which is
consistent with the work of Harris and Morris (1984). However, it still failed due to
compressive forces on the rear wall which caused shearing and delamination. The
0
2000
4000
6000
8000
10000
12000
0.001
0.417
0.842
1.263
1.688
2.109
2.534
2.957
3.381
3.803
4.228
4.649
5.075
5.496
5.924
6.348
6.778
6.784
Load(N)
Extension (mm)
Load Vs Extension Specimen 64 A
Load Vs Extension
PQ=11,120 N

45
failure is almost symmetrical about the central plies, with intermittent ± 45o
shearing
of the fibres taking place across the thickness of the specimen. These features can be
seen clearly in figure 7.4.2 which shows the rear wall of the specimen and the crack
tip area under magnification.
Figure 7.4.2: Failure of 64 ply specimen A.
From the images in figure 7.4.2, it can be seen that no large scale fibre fracture took
place at the crack tip area. Large scale failure refers to the presence of a visually
apparent crack through all plies. A line was drawn on the surface of the specimen to
aid in placement of the pre-crack during manufacture. As there are no broken fibres
beyond this point, it is known that no propagation of the crack took place through 0o
plies. Fracture of the matrix did occur through the 90o
plies but this is due to the
widening of the crack tip rather than the progression of subcritical crack damage. As
a result of this failure mode, it was decided that different specimen geometries shall
be tested in the next two parts of this investigation to attempt to avoid repeated
invalid results.
Due to an error in the procedure while recording the crack opening displacement
using COD gauges, the data obtained from this test was lost. However, it is noted that
the bond between the knife edges and the specimen did not break. This is likely due
to the increased surface area over which the epoxy acted and suggests marginally
more accurate crack opening displacement values for the thicker specimens. Similar
to the results from the 32 ply specimens, the result for KIc would have been flawed
due to no propagation of the crack taking place. It is interesting to note that the
CTOD of the specimen increased despite the crack not propagating through the

46
material. This is due to stretching of the fibres in the y direction in conjunction with
notch-tip damage blunting the crack root.
No DIC study was carried out on this specimen. As this exact geometry was
anticipated to be tested at the same thickness three times, DIC was only going to be
used to plot the strain fields on one specimen as setup of the system requires a lot of
time. As the specimen was then found to yield further invalid results, it was decided
to change the geometry of the next test specimen to gather accurate fracture
toughness data. However, expected results would be similar to the DIC study carried
out in section 7.1 on the 32 ply square specimen with a similar compression
concentration seen around the rear wall of the specimen.
7.5. 64 Ply Specimen B.
As a result of the failure due to compression on the rear wall of 64 ply specimen A, it
was decided to modify the remaining thicker specimens in an attempt to reduce this
effect and achieve a valid mode 1 fracture test. The geometry of the 32 ply chamfered
specimen was used for this test as it produced the most conservative results during the
32 ply tests. The results for 64 ply specimen B can be seen in figure 7.5.1 where the
maximum load sustained by the specimen was 6824.56 N. This is significantly lower
than the load sustained by the unmodified 64 ply specimen A in the previous test.
Figure 7.5.1: Load vs. extension for 64 ply specimen B.
0
1000
2000
3000
4000
5000
6000
7000
8000
0.112
2.241
4.374
6.547
8.721
10.897
13.057
15.210
17.370
19.514
21.710
23.823
25.983
28.156
30.286
32.445
34.619
35.901
35.897
35.887
Load(N)
Extension (mm)
Load Vs. Extension Specimen B
Load Vs. Extension
PQ=6824.56 N

47
A KIc value of 36.063 MN/m1.5
was determined for the new geometry according to
ASTM standard E399 which also gives a GIc value of 101.022 kJ/m2
. These values
are significantly lower than the reported values for the 64 ply specimen A, which is
explained by the failure of this specimen due to self-similar crack growth. Less
compression was present at the rear of the specimen which resulted in crack
propagation taking place at a lower load. This yields a more accurate value of fracture
toughness for the laminate. It is also noted that the fracture toughness for this
specimen is lower in comparison to the 32 ply specimens, which is consistent with the
findings of Jose et al. (2001) and the linear elastic fracture mechanics of
homogeneous materials.
Figure 7.5.2 shows a close up image of the rear wall of the specimen after testing.
Similar to the failure on the rear wall of the 32 ply chamfered specimen, this piece
failed at an angle across the rear wall. However, this angle [1] is much shallower than
any seen previously, which can be attributed to the specimen approaching plane strain
failure conditions. Also, small 45o
lips [2] at the side of the specimen highlight the
plain stress conditions closer to the edge.
Figure 7.5.2: Failure of 64 ply specimen B.
These images highlight the change in the damage which occurs at the rear wall and
crack tip of the specimen. While fibre pull-out [3] is still very evident on the fracture
surface, a straighter, better defined crack can be seen when compared with other
specimens tested in this study. On the rear wall, the smaller angle between the
fracture surface axis and the horizontal axis is indicative of plane strain fracture. It is
worth noting that even though both 64 ply laminates tested so far were of equal
thickness, specimen A failed in plane stress, where specimen B failed in plane strain.
3
2
25o
1

48
This is due to the significantly larger in-plane forces sustained by specimen A, which
make the effect of any out of plane stresses negligible. This loading scenario then
becomes analogous to plane stress loading.
COD gauge data for specimen B is shown in figure 7.5.3. It was noted in the first 64
ply test that a better bond was seen between the thick specimens and the knife edges
holding the COD gauges. This was seen again in this test where the gauges remained
in place until an extension of greater than 10 mm was seen. The crack opening
displacement measured at the point of failure was 3.39 mm.
Figure 7.5.3: Load vs. crack opening displacement for 64 ply specimen B.
The fracture toughness measured for the thicker chamfered specimen B was 249.02
MN/m1.5
when calculated using the COD gauge data. This value is far larger than the
value predicted using ASTM standard E399 and can again be attributed to errors in
the theory used. The value for general toughness parameter, Qc, used by Poe (1983)
was calculated for thin specimens in plane stress states, where evidence from this test
suggests that a plane strain analysis would be a better representation of the situation.
Using values from the DIC imaging, KIc is calculated to be 142.344 MN/m1.5
. This
value is still much larger than is anticipated by the ASTM method of KIc calculation,
but it is a more accurate value than would otherwise be obtained as there are less
assumptions made regarding physical parameters, such as Qc stated above.
Inaccuracies are possible in the DIC imaging as well due to the high magnification
required to measure crack width, which can cause the image to become pixelated.
0
1000
2000
3000
4000
5000
6000
7000
8000
0.003
0.008
0.219
0.654
1.137
1.624
2.127
2.622
3.168
3.771
4.394
5.05
5.644
6.256
6.858
7.471
8.059
8.662
9.228
9.791
10.368
Load(N)
COD Extension (mm)
Load Vs. COD Specimen B
Load Vs. COD

49
The strain fields for this chamfered specimen are shown in figure 7.5.4 below. These
images are again hampered by movement of the specimen, but still give a reasonably
accurate visualisation of ε11 in the specimen. The specimen at fracture point can be
seen in (a) and shows the stress concentrations in the region of the crack. As
significant crack growth occurred in this specimen, the tensile strain field can be seen
extending towards the back of the specimen as the crack propagates through the
specimen.
Figure 7.5.4: Strain fields in 64 ply specimen B.
The compression seen in these DIC images [1] is not as severe as the compression
seen in analysis of previous specimens. This is reflected in the microscopic analysis
which showed substantially smaller amounts of compressive shearing and
delamination. While there was still slight damage across the rear wall of the
specimen, the test was still deemed valid due to failure occurring by crack
propagation. An improvement in the failure mode of the specimen could be seen with
further modification in the next specimen.
7.6. 64 Ply Specimen C.
While specimen B provided a much better resistance to compressive failure at the rear
wall, it was felt that further modification to the design of the test specimen would
ultimately yield mode 1 fracture without any damage due to compression. The length
of the rear wall of the chamfered specimen was reduced to just 8 mm while keeping
the side walls of the specimen at 35 mm long. This specimen reached a maximum
load of 6643 N before failing, a marginally smaller value than that sustained by
specimen B. This data can be seen in figure 7.6.1.
(a) (b)
1

50
Figure 7.6.1: Load vs. extension for 64 ply specimen C.
The fracture toughness according to ASTM standard E399 for specimen C was 33.96
MN/m1.5
and the GIc measured was 89.583 kJ/m2
. This is the lowest fracture
toughness measured across the 6 samples tested in this study. This highlights an
overall reduction in the compressive stresses throughout the specimen which allows
crack propagation to take place under a lower load. This test was intended to run until
the specimen was separated into two distinct limbs. However, the specimen began to
rotate with the clevises once the fracture surfaces had opened up to an angle of almost
90o
. The test was stopped at this point and the specimen’s limbs were separated
manually.
The specimen limbs can be seen in two distinct pieces in figures 7.6.2 and 7.6.3. A
reasonably flat fracture surface can be seen, with fracture occurring across all 0o
fibres. The rear wall of the specimen has a mostly horizontal failure plane [1] with
two 45o
lips [2] at the sides of the specimen. Exceptionally long fibres can be seen at
the rear end of one of the two limbs which was caused by the unusual loading
conditions that occurred after the crack had opened to a wide angle.
0
1000
2000
3000
4000
5000
6000
7000
0.056
1.642
3.223
4.810
6.395
7.992
9.593
11.192
12.804
14.400
16.000
17.597
19.195
20.798
22.398
24.000
25.600
27.195
28.795
30.386
Load(N)
Extension (mm)
Load Vs. Extension Specimen C
Load Vs. Extension
PQ=6643 N

51
Figure 7.6.2: Fracture zones on rear wall of 64 ply specimen C.
Figure 7.6.3: Fracture surface of 64 ply specimen C.
Figure 7.6.2 shows the separated top and bottom limbs of specimen C after fracture
had occurred. The axis of failure is reasonably horizontal and shows the two 45o
lips
that are characteristic of plane stress fracture at the free surfaces of the specimen.
Figure 7.6.3 is a view perpendicular to the axis of fracture which shows the uniform
nature of the crack propagation through the specimen. Small amounts of damage can
be seen at the rear wall of the specimen, but these can be attributed to rotation of the
piece during testing as the crack begins to open up to a significant angle, rather than
compression. The fibres which are still intact at this point then undergo bending and
eventually rupture, leaving the longer fibres [3]. No out of plane movement was seen
which resulted in this being a valid test which should yield quite accurate results for
fracture toughness and KIc.
The crack opening displacement for this specimen is plotted in figure 7.6.4 below.
The value reached when fracture occurred was 3.336 mm. The graph shows that the
load on the specimen after fracture becomes quite discontinuous as fibre bridging
occurs.
1 2
3

52
Figure 7.6.4: Load vs. crack opening displacement for 64 ply specimen C.
Similar to the specimens tested earlier, specimen C showed a large difference in the
values obtained for (CTOD)cr from the COD gauges and the DIC images. Using the
information gathered from the COD gauges, fracture toughness was calculated to be
260.52 MN/m1.5
while the value obtained using (CTOD)cr from DIC imaging is
133.81 MN/m1.5
. The lower value of fracture toughness is again more accurate as the
errors mentioned in section 7.1 are avoided.
Both specimens B and C show the largest crack opening displacements at failure from
the six specimens investigated. As failure does not occur at the rear wall, the
specimens reached the maximum load that they can sustain before fracture. With the
exception of specimen A, no other specimens tested in this study reached this critical
load and as a result, the crack opening displacements measured were not recorded at
the actual fracture point.
The final test results for the DIC study are shown in figure 7.6.5. An issue with the
recording equipment resulted in a black screen being seen after the crack had
propagated almost halfway through the specimen. As a result, the image on the right
in figure 7.6.5 is the last image that can be seen before this error occurs.
0
1000
2000
3000
4000
5000
6000
7000
0.013
0.016
0.251
0.643
1.08
1.537
1.998
2.518
3.076
3.692
4.37
5.058
5.718
6.378
7.028
7.677
8.313
8.947
9.578
10.199
10.816
11.429
Load(N)
COD Extension (mm)
Load Vs. COD Specimen C
Load Vs. COD

Liam O Sullvian FYP Mar 2013

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Liam O Sullvian FYP Mar 2013