strength and failure mechanisms of composite laminates subject to localised transverse loading

Composite Structures 69 (2005) 301–314

www.elsevier.com/locate/compstruct

Strength and failure mechanisms of composite laminatessubject to localised transverse loading

Gordon Kelly *, Stefan Hallstrom

Division of Lightweight Structures, Department of Aeronautical and Vehicle Engineering, Kungl Tekniska Hogskolan, S-100 44 Stockholm, Sweden

Available online 21 August 2004

Abstract

The behaviour of composite laminates subject to transverse load introduction has been investigated experimentally and numer-

ically. The effect of the specimen size, stacking sequence and material system on the failure load was determined experimentally and

the failure modes examined through fractographic analysis. Damage was found to initiate at low load levels, typically 20–30% of the

failure load. The dominant initial failure mode was matrix intralaminar shear failure which occurred in sub-surface plies. The dam-

age developed into a network of intralaminar and interlaminar shear cracks. Two different macromechanical failure modes were

identified, fastener pull-through failure and global collapse of the laminate. The internal damage and ultimate failure mode were

found to depend upon the laminate stacking sequence and resin system.

A three-dimensional finite element model was developed to analyse the stress distribution within the laminate and predict first-ply

failure. The results from the finite element model were found to be in general agreement with the experimental observations.

� 2004 Elsevier Ltd. All rights reserved.

Keywords: Composite laminate; Localised transverse loading; Failure; Strength; Experimental investigation; Finite element analysis

1. Introduction

Concentrated transverse loading in composite lami-

nates is commonly avoided due to the poor interlaminar

and through-thickness strength of the materials. How-

ever, with the increasing use of such materials in load

bearing structures has come a need to develop design

principles for both in-plane and out-of-plane loading.

In-plane load introduction has been investigated inten-sively in the form of mechanically fastened joints. A

wealth of literature exists concerning experimental, ana-

lytical and numerical analysis of parameters which influ-

ence the behaviour of composite laminates loaded

in-plane. The behaviour of composite laminates subject

to out-of-plane loading has not been investigated or

0263-8223/$ - see front matter � 2004 Elsevier Ltd. All rights reserved.

doi:10.1016/j.compstruct.2004.07.008

* Corresponding author.

E-mail address: [email protected] (G. Kelly).

understood to the same degree. Limited published liter-ature regarding the transverse load bearing capacity of

composite laminates exists. The few studies published

concern the selection of countersunk fasteners [1], fas-

tener pull-through failure associated with post-buckled

structures [2–4] and quasi-static modelling of transverse

impact [5,6].

Chen and Lee [1] investigated the behaviour of car-

bon fibre/epoxy (AS4/3501-6) cross-ply laminates sub-ject to bending loads. Rectangular laminates were

loaded in three-point bending by countersunk fasteners

with different head geometries. The authors conducted

a progressive failure analysis of the laminates with fail-

ure being predicted based on the maximum stress theory

and the Ye delamination criterion [7]. The effect of fric-

tion between the fastener and the hole was found to be

significant together with the bending stiffness of the lam-inate. Reasonable agreement was found between the

experimental and predicted failure loads.

mailto:[email protected]

composite plate

Fastenerload attachment

restraining plate

washer

Fig. 1. Schematic of the experimental set-up.

302 G. Kelly, S. Hallstr€om / Composite Structures 69 (2005) 301–314

Waters and Williams [2] conducted an experimental

investigation of bolt push through failure which was a

possible failure mode in skin-stiffener attachments in

aerospace structures at the time. The study considered

the response and failure modes of composite laminates

loaded by a countersunk fastener. Circular specimensof diameter 25.4 and 50.8 mm were clamped at the outer

boundary and a centrally located fastener was used to

apply the transverse load. A series of material systems

were investigated including combinations of different

carbon and kevlar fibres with thermoset and thermo-

plastic matrices. A number of conclusions were drawn

from the study. The effect of specimen size, or boundary

diameter, was not found to significantly affect the loadcarrying capability of the laminates. The behaviour of

laminates made with brittle matrices were improved

through the use of transverse stitching with the initial

failure load found to be dependent on the stitching pat-

tern. It was also apparent from the results presented in

the report that the load capability was not strongly

dependent upon the laminate stacking sequence. How-

ever, the effect of the laminate stacking sequence wasnot addressed explicitly.

Banbury and Kelly [3] conducted an experimental

investigation of fastener pull-through failure in lami-

nates made from woven fabric and unidirectional tape

pre-preg material. The experimental set-up was consist-

ent with [2] with clamped circular specimens of diameter

31.8 mm being centrally loaded via a 3.97 mm diameter

fastener. The effects of the specimen and fastener geom-etry on the failure load were determined and the failure

mechanisms analysed. The observed failure modes were

matrix shear fracture and tensile failure which were

found to be dependent upon the bending stiffness of

the laminate. Matrix shear cracking was found to occur

in almost all tests while tensile failure occurred in lami-

nates with reduced bending stiffness. The internal dam-

age pattern in the laminates was likened to thatobserved in laminates subject to low velocity impact.

Failure was attributed to a critical level of matrix strain

which was independent of the laminate and fastener

geometry and material type for a given resin system.

Banbury et al. [4] conducted a two-dimensional axi-sym-

metric finite element analysis to simulate the fastener

pull-through tests in [3]. A progressive damage model

based on the maximum strain criterion was used to pre-dict failure in the laminates although no consideration

was given to interlaminar failure. The results from the

model were consistent with the experimental observa-

tions and the predicted failure loads were shown to be

in good agreement with the experimentally determined

values.

There are currently no widely accepted design rules or

established methods to accurately predict the structuralperformance of composite laminates subject to concen-

trated transverse load introduction. The aim of this

investigation is to provide a greater understanding of

the mechanisms which govern the structural behaviour

and failure of composite laminates under such loading

conditions.

2. Experimental investigation

The main objectives of the experimental work were to

study the effect of selected parameters on the structural

behaviour of laminates subject to transverse loading and

to investigate the micromechanical failure modes and

development of damage leading to ultimate laminate

failure. An experimental investigation was undertakenwhere the effects of the laminate stacking sequence, lam-

inate thickness, specimen size and material system were

considered.

The experimental apparatus used in the current study

consisted of a loading fixture and a restraining fixture.

The loading fixture comprised of a steel cylinder which

was fastened to the load cell of the tensile testing ma-

chine and to the test specimen via a protruding headedtitanium fastener (Hi-Torque / 6 mm). The fastener

was finger tightened to exclude any possible effects of

the clamping load on the failure load and mode. Stain-

less steel washers of thickness 1 mm and outer diameter

12 mm were used together with the fasteners. The test

specimens were simply supported at the outer boundary

by a steel restraining plate with a circular opening.

Restraining plates with circular openings of diameter40, 80 and 120 mm were manufactured to allow investi-

gation of specimen size effects. The restraining plate was

attached to the base of the fixture by four bolts located

at the corners of the plate. The experimental set-up is

illustrated in Fig. 1. The experiments were performed

using an Instron universal test machine (Instron 4505)

under displacement control at a rate of 1.0 mm/min.

The load and transverse displacement of the fastenerwere recorded continuously during testing using a PC

data acquisition system. A series of specimens were

instrumented with strain gauges in order to measure

the strain field in the laminate at different locations dur-

ing the test.

Table 1

Test specimen configurations

Notation Stacking sequence Resin Thickness (mm)

½0=45=90=�45�es [0/45/90/�45]s Epoxy 1.68

½0=45=90=�45�es2 [0/45/90/�45]s2 Epoxy 3.42

½0=90=45=�45�es [0/90/45/�45]s Epoxy 1.48

½0=90=45=�45�vs [0/90/45/�45]s Vinylester 2.42

½0=90�vs2 [0/90]s2 Vinylester 2.42

½0=90=45=�45�vs2 [0/90/45/�45]s2 Vinylester 4.68

G. Kelly, S. Hallstr€om / Composite Structures 69 (2005) 301–314 303

The outer boundary diameter of the test specimens

was simply supported in the current study while the pre-

vious studies [2–4] of transverse load introduction con-

sidered specimens with a clamped outer boundary. Asimply supported boundary was chosen to be represent-

ative of a concentrated load introduction in a larger

structure. In contrast to previous studies [2,3], a range

of specimen sizes were investigated.

A fractographic investigation was undertaken in

order to determine micromechanical failure modes

occurring within the specimens. Several tests were

stopped at sub-critical load levels and the specimens un-loaded to allow for microscopic investigation. The spec-

imens were sectioned along planes parallel to the fibre

directions, polished and examined using an optical

microscope.

2.1. Materials and specimen preparation

Two different material systems were considered forthe investigation. The first system consisted of carbon

fibre non-crimp fabric and epoxy matrix (T700/Shell

Epicote LV828) with laminates being manufactured

using the resin transfer moulding (RTM) technique.

The second system was a carbon fibre non-crimp fabric

and vinylester matrix (T700/Jotun 9100) with the lami-

nates being manufactured using the vacuum infusion

technique. The laminate configurations used in theinvestigation are listed in Table 1.

Square specimens were machined from the laminates,

10 mm wider than the diameter of the restraining plate,

using a diamond tipped saw. A centrally located hole of

diameter 6 mm was machined in the specimens using a

dagger drill which was used in combination with a back-

ing plate to limit delamination at the exit side of the

laminate.

3. Results

3.1. Load–displacement behaviour

The load–displacement behaviour of the test speci-

mens subject to transverse load introduction was foundto be different, depending on the effective flexural stiff-

ness of the laminates. The effective flexural stiffness in-

creases with increase in the laminate thickness and

reducing boundary diameter. The load–displacement

relationship for specimens with high flexural stiffness

(low /) was slightly non-linear in the beginning followedby an almost linear load–displacement relation until fail-

ure. Specimens with lower flexural stiffness illustrated

non-linear deformation behaviour with a significant in-

crease in stiffness evident at larger displacements. Typi-

cal load–displacement graphs for the ½0=45=90=�45�esand ½0=90=45=�45�vs specimens are illustrated in Fig.

2(a) and (b) respectively.

The ½0=45=90=�45�es laminate exhibited a number ofminor load drops as the load increased. The minor load

drops were typically between 0.05 and 0.1 kN and the

resulting damage did not appear to have a significant ef-

fect on the global stiffness of the specimen. The minor

load drops occurred more frequently as the laminate ap-

proached final failure. The load levels at which the

minor load drops occurred were found to be independ-

ent of the specimen diameter. A similar independenceof the specimen size on the load when first cracking oc-

curred was noted by Chen et al. [8] in a study of biaxial

flexure of rectangular cross-ply laminates and by Cap-

rino et al. [6] who considered transverse loading of

CFRP plates via an indenter. The first audible fracture

within the specimens occurred at load levels significantly

lower than the first minor load drops visible on the

load–displacement curve. The first audible fractureoccurred at approximately 20–25% of the ultimate load.

Audible fractures were heard periodically as the load

was increased. At higher load levels, the audible frac-

tures coincided with minor load drops visible on the

load–displacement curve. At final failure, a number of

smaller load drops preceded a sudden loss of all load

bearing capacity. In the current study, failure of the

specimen was deemed to occur at the first major loaddrop illustrated by point (F) in Fig. 2(a) and (b). The

load drop is attributed to a major loss of stiffness or

pull-through failure of the laminate. The ultimate failure

mode of each of the ½0=45=90=�45�es laminates was fas-

tener pull-through failure where the fastener penetrated

through the laminate.

Similar load–displacement behaviour was noted for

the ½0=45=90=�45�es2 laminates. Slightly different behav-iour was noted for the specimens with /boundary = 40

mm which illustrated a marked reduction in stiffness at

0 2 4 6 8 10 12 140

1

2

3

4

5

6

7

8

Displacment (mm)

Load

(kN

)

φ=40 mm φ=80 mm φ=120 mm

F

First audible cracking

0 2 4 6 8 10 12 140

1

2

3

4

5

6

7

8

Displacment (mm)

Load

(kN

)

φ=40 mm φ=80 mm φ=120 mm

First audible cracking

Minor load drop

F

(b)(a)

Fig. 2. Typical load–displacement curve for the quasi-isotropic 8-ply laminates: (a) ½0=45=90=�45�es laminates, (b) ½0=90=45=�45�vs laminates.


approximately 45% of the failure load. However, the

specimens continued to carry load beyond this point

but exhibited a reduced stiffness. The reduction in

stiffness was attributed to internal damage in thelaminate.

The ½0=90=45=�45�es laminate illustrated similar char-

acteristics to the ½0=45=90=�45�es laminate. However, the

load drops were slightly larger (0.1–0.2 kN) and were

more visible on the load–displacement curve. The first

audible fractures occurred at 25% of the failure load

and the ultimate failure mode of the ½0=90=45=�45�eslaminates was fastener pull-through failure.

The ½0=90=45=�45�vs laminates exhibited similar

load–displacement behaviour to the ½0=45=90=�45�eslaminates. Audible fracture could be heard from the

specimens at load levels around 25% of the failure load

although no detectable loss of stiffness was observed as

illustrated in the load–displacement curves (see Fig.

2(b)). The ½0=90=45=�45�vs specimens continued to frac-

ture as the load was increased and exhibited a series ofminor load drops prior to the first major load drop.

The slope of the curve following the load drops re-

mained unaltered indicating that no appreciable loss of

global stiffness occurred. Final failure of the laminate

was characterised by a sudden major drop in load. Fail-

ure occurred through global collapse of the laminate

with interlaminar fractures extending to the boundary

Epoxy matrix laminates

Fig. 3. Schematic of obse

of the test specimens. Compressive failure of the lami-

nate at a plane orthogonal to the fibre orientation of

the surface plies. The failure patterns of the epoxy and

vinylester matrix laminates are illustrated schematicallyin Fig. 3.

Visual inspection at failure of the vinylester speci-

mens showed that the inter-ply damage extended to

the edge of the specimen. The laminates were shown

to be capable of carrying additional load past this failure

point, however, the post-failure behaviour was consid-

ered to be beyond the scope of this study. The magni-

tude of the load drop at failure was found to bedependent upon the specimen size with specimens of lar-

ger diameter experiencing a larger drop in load. The lar-

ger drop occurred as a result of the greater about of

elastic energy accumulated in the larger diameter speci-

mens before failure.

The behaviour of the ½0=90=45=�45�vs2 laminates were

found to be similar to the behaviour of the

½0=45=90=�45�es2 laminates with a distinct reduction instiffness occurring at approximately 45% of the failure

load for the /boundary = 40 mm specimen. The marked

reduction in stiffness was not evident in the larger diam-

eter specimens. A more sparse distribution of load drops

were found on the load–displacement curves of the

½0=90=45=�45�vs2 laminates in comparison to the

½0=45=90=�45�es2 laminates.

Vinylester matrix laminates

rved failure modes.


Comparison of the ½0=90=45=�45�vs and ½0=90�vs2 spec-imens illustrated no significant difference in the load–

displacement behaviour. The behaviour was found to

be similar with the exception that the damage propa-

gated along different paths as a result of the different

stacking sequence.

3.2. Strain distribution

A series of specimens were instrumented with strain

gauges as shown in Fig. 4. Uniaxial strain gauges of

gauge length 2 mm were placed in orthogonal directions

on both sides of the laminate in order to investigate

bending behaviour and the radial strain field aroundthe hole. Strain measurements from the ½0=45=90=�45�es(/boundary = 80 mm) case are illustrated in Fig. 5(a)–

(d). The deformation behaviour of the ½0=45=90=�45�eslaminate was shown to alter as the load increased and

damage developed. The radial strain on the top surface

of the ½0=45=90=�45�es laminate is illustrated in Fig. 5(a)

and (b). The radial strain was initially tensile as a result

of the laminate bending. However as the load was in-creased, local deformation occurred around the load

introduction and the radial strains became non-symmet-

ric. At 40% of the failure load, the radial strain close to

the hole in the 90� direction became compressive. This

corresponded to the plane of the laminate with the lowest

bending stiffness. The same behaviour did not occur in

the 0� direction until 60% of the failure load. The change

in sign of the radial strain indicated a shift from bendingto membrane deformation behaviour in the specimen.

5 mm

7 mm

8 mm

T0_R1

T0_R2

T0_R3

T90

_R1

T90

_R2

T90

_R3

T

Orientation of surface ply

Loading direction

B

Fig. 4. Strain gauge locations on instrumented specimens.

The non-linear deformation was apparent when consid-

ering the underside of the laminate. The initially com-

pressive radial strains changed sign at load levels of

25% and 40% of the failure load in the 90� and 0� direc-tions respectively. At 60% of the failure load, the tensile

radial strain close to the hole was significantly larger inthe 90� direction (�rr = 0.45%) in comparison to the 0�direction (�rr = 0.25%). The difference between the radial

strains in the 90� and 0� directions continued to increase

as the load was increased. The unequal strain distribu-

tion around the hole indicated that damage initiation

may occur at a preferred location, or weakest plane.

3.3. Effect of the specimen size

The effect of the specimen size on the specimen failure

load is illustrated in Fig. 6. The failure load of the

½0=45=90=�45�es laminates was found to be independent

of the specimen diameter. The ultimate failure mode

was similar for each of the specimens and the result

was deemed reasonable given the local nature of the fail-

ure mode. The experiments produced repeatable resultswith the variation in strength being relatively constant

for each of the specimen types. The ½0=45=90=�45�es2laminates illustrated a very similar behaviour with regard

to specimen size. However, the laminates with a diameter

of 40 mm were found to withstand slightly higher loads

with the load decreasing by 15% as the boundary diam-

eter was increased to 80 mm. The failure load remained

unchanged for the specimens with a boundary diameterof 120 mm. The failure loads of the ½0=90=45=�45�es lam-

inates were also found to be independent of specimen

size. The failure loads showed a similar trend to the

½0=45=90=�45�es laminates, with the load levels being

lower as a result of the reduced laminate thickness.

Both the ½0=90=45=�45�vs and the ½0=90�vs2 laminates

were found to behave in a similar manner with regard

to the specimen size. This supported the theory thatthe same failure mode was responsible for damage prop-

agation and ultimate failure of the laminates. The failure

load was found to increase by approximately 30% as the

diameter increased from 40 to 80 mm with no additional

increase in failure load being noted for the /boundary =

120 mm laminates. The failure load of the ½0=90=45=�45�vs2 laminates showed a similar trend to the

½0=90=45=�45�vs laminates although the failure loadwas shown to increase with specimen size for the range

of diameters selected in the current study.

The differing dependence of the failure load on the

specimen size was assumed to be a result of the different

failure modes occurring in the laminates.

3.4. Fractographic investigation of failure modes

The load–displacement behaviour of the laminates

did not provide any clear indication of damage initiation

-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 10

1

2

3

4

5

6Lo

ad (

kN)

Strain (%)

T-0-R1T-0-R2T-0-R3

-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 10

1

2

3

4

5

6

(Loa

d kN

)

Strain (%)

-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 10

1

2

3

4

5

6

Load

(kN

)

Strain (%)-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1

0

1

2

3

4

5

6

Load

(kN

)

Strain (%)

(a) (b)

(c) (d)

B-0-R1B-0-R2B-0-R3

T-90-R1T-90-R2T-90-R3

B-90-R1B-90-R2B-90-R3

Fig. 5. Radial strain distribution in [0/45/90/�45]s epoxy laminate: (a) top surface, 0� orientation; (b) top surface, 90� orientation; (c) bottom surface,

0� orientation; (d) bottom surface, 90� orientation.

20 40 60 80 100 120 1400

5

10

15

20

Fai

lure

Loa

d (k

N)

φ

[0/90/45/–45]

[0/45/90/–45]

[0/90/45/–45]

[0/90/45/–45] s2v

[0/90] s2v

[0/45/90/–45] se

se

s2e

sv

boundary (mm)

Fig. 6. Effect of specimen size on the ultimate failure load.


and the development of damage within the laminates

prior to failure. It was also evident from visual inspec-

tion of the specimens that two different final failure

modes occurred, namely fastener pull-through failure

and global laminate collapse. It was considered neces-

sary to identify the failure mechanism responsible for

damage initiation, the location and the manner in whichthe damage propagated.

The planes which were examined are shown in Fig. 7.

The observations made for each laminate type are dis-

cussed in the following sections. The individual plies in

the laminates are numbered from top to bottom with

ply 1 being the outermost ply on the top side of the lam-

inate. The term off-axis plies refers to plies where the

fibre orientation does not coincide with the directionof the sectioned plane. This convention is adopted

throughout the following sections.

A B C

D

O

+45o

-45o

0o

90o

Fibre directions

Fig. 7. Sectioned planes for the fractographic investigation.


3.4.1. [0=45=90=�45]esn laminates

The first audible fractures occurred in the ½0=45=90=�45�es specimens at load levels approximately 20–

25% of the failure load. There was no indication of dam-

age on the load–displacement curve at this load level

and inspection of optical micrographs taken of speci-

mens exposed to this load level revealed only limited

damage. A minor intralaminar shear failure in the sub-

surface ply (ply 7) beneath the edge of the washer onplane O–D was noted together with limited tensile frac-

tures in the resin rich layer on the top surface of the lam-

inate. After 50% of the failure load, more significant

structural damage was apparent in the form of trans-

verse tensile failures in the top surface of the laminate.

On plane O–D, the transverse tensile cracks propagated

through the upper 0� (ply 1) and 45� (ply 2) plies and

were arrested at the interface of the fibre bundles. Typ-ical transverse tensile failures are shown in Fig. 8. To-

gether with the intralaminar matrix shear failure, this

failure mode was assumed to be responsible for damage

initiation in the ½0=45=90=�45�es laminates. No indica-

tion of the transverse tensile failures occurring at 50%

of the failure load were visible on the load–displacement

curves. The fracture location correlated with the radial

Fig. 8. Micrograph of ½0=45=90=�45�es epoxy laminate, /boundary = 40

mm loaded to 50% of the failure load (plane O–D).

strain in the 90� direction which was tensile at this load

level. There was no apparent damage at this load level in

the other planes examined.

At 90% of the failure load, a number of different

micromechanical failure modes were found to be present

in all of the examined planes. Fibre kinking and matrixintralaminar shear cracking was found directly under

the edge of the washer. The most significant intralami-

nar shear cracking was found along planes O–A and

O–C. Along plane O–A, inclined shear cracking had

propagated through plies 3–6 which were orientated

off-axis to the sectioned plane. Matrix shear cracking

was also evident in the upper plies along the plane O–

C. Inclined intralaminar shear cracks and in-plane inter-laminar shear cracks were found to exist between plies 1

and 3 with the damage extending approximately 10 mm

from the edge of the hole. Limited damage was noted in

plane O–B in the specimens inspected. The dominant

failure mode in the O–D plane was transverse tensile

failure of the matrix in the top ply (ply 1). The trans-

verse tensile cracks were most apparent in the region

adjacent to the edge of the washer.Final failure of the ½0=45=90=�45�es specimens in-

volved the fastener penetrating through the laminate.

The through-thickness shear cracks which propagated

conically upward from the edge of the washer resulted

in local collapse with the material unable to withstand

the transverse load. No significant in-plane damage

was noted in the failed specimens except in the upper

plies where splitting occurred causing a �volcano� likefailure topology. The behaviour was found to be consist-

ent for each of the specimen diameters.

The first audible fractures in the ½0=45=90=�45�es2laminates occurred at 25–30% of the failure load. The

initial damage occurred in the form of matrix intralami-

nar shear cracking in sub-surface plies at the edge of the

washer. Damage was most evident along plane O–D. At

50% of the failure load, internal damage within thelaminate was more pronounced. The specimen with a

boundary diameter of 40 mm illustrated a distinct reduc-

tion in stiffness at this load level. Inspection of the spec-

imen revealed damage in all of the examined planes.

Single shear cracks through off-axis plies were evident

in planes O–B and O–D. The most significant damage

was however noted in planes O–A and O–C. A network

of inclined intralaminar shear cracks was visible to-gether with interlaminar shear failures which occurred

at different interfaces through the thickness. A micro-

graph of plane O–A is shown in Fig. 9. The matrix

intralaminar shear cracking which initiated in sub-sur-

face plies beneath the edge of the washer, was shown

to have propagated at an angle of 45� to the applied

load. The cracks propagated through fibre bundles be-

fore reaching the interface between the fibre bundleand the matrix. Upon reaching the fibre bundle inter-

faces, the cracks changed direction and continued to

Fig. 9. Micrograph of ½0=45=90=�45�es2 laminate, /boundary = 40 mm

loaded to 50% of the failure load (plane O–A).


propagate along the interfaces before being arrested.

Similar fracture patterns were also found above themid-plane of the laminate where inclined shear cracks

propagated through off-axis plies and formed interlami-

nar cracks at fibre bundle interfaces. Very limited crack-

ing was found in the resin rich areas which were present

between fibre bundles, with the cracks tending to prop-

agate through the off-axis fibre bundles. The extent of

the damage in the larger diameter specimens was not

as pronounced in comparison to the 40 mm diameterspecimens. Intralaminar shear failure was shown to be

present in most planes with the most significant damage

occurring in the O–D plane. Final failure of the

½0=45=90=�45�es2 laminates was found to occur in a sim-

ilar manner as in the ½0=45=90=�45�es laminates with the

fastener penetrating through the laminate.

Fig. 10. Intralaminar and interlaminar shear cracking in the

½0=90=45=�45�vs laminate, /boundary = 40 mm at 50% of failure load

(plane O–A).

3.4.2. [0=90=45=�45]es laminates

The first audible fractures in the ½0=90=45=�45�es lam-

inates were recorded at approximately 25% of the failure

load. There was no indication of damage on the load–

displacement curve and inspection of the laminates

loaded to this load level revealed only minor transverse

tensile cracking on the upper surface of the laminate (ply

1). Transverse tensile failure was regarded as the first ply

failure mode for these laminates.At 50% of the failure load, more significant damage

was observed in the laminates. The most notable dam-

age was on plane O–B where intralaminar shear fracture

was found in ply 6 close to the edge of the washer. Sim-

ilar intralaminar shear cracks were observed on plane

O–D where cracks had propagated through ply 6 and

through plies 4 and 5.

At 90% of the failure load, the most significant dam-age was found on planes O–B and O–D. The damage

was consistent with the damage pattern observed at

50% of the failure load. On plane O–B, the most signif-

icant damage was an interlaminar shear crack between

plies 5 and 6. Intralaminar cracks branched from the

main crack through ply 5 in the region close to the edge

of the washer. Transverse tensile fractures were also ob-

served in ply 7 which was orientated at 90� to the sec-

tioned plane. Damage was also observed on plane O–

C where transverse tensile fractures were observed on

the upper surface (ply 1). There were also interlaminar

shear cracks which had propagated between plies 1

and 2 and between plies 2 and 3.Final failure of the ½0=90=45=�45�es laminates oc-

curred with the fastener penetrating through the lami-

nate. The results were found to be consistent for all

specimen diameters.

3.4.3. [0=90=45=�45]vsn laminates

Damage initiation in the ½0=90=45=�45�vs laminates

occurred at approximately 20% of the failure load.The initial failure was found to be similar as in the

½0=45=90=�45�esn laminates with inclined intralaminar

shear cracking occurring in sub-surface plies beneath

the edge of the washer. At 50% of the failure load, dam-

age was found to be significant in all of the examined

planes. In plane O–A, interlaminar shear failures were

found to extend 10 mm from the edge of the hole. Inter-

laminar cracks were found to be most prominent be-tween the +45� and �45� plies (3 and 4 and 5 and 6).

The interlaminar shear cracks were connected by in-

clined intralaminar shear cracks through off-axis plies

(4 and 5) in the centre of the laminate as shown in

Fig. 10. Similar fractures were noted on each of the ob-

served planes. Additional damage was noted on plane

O–D with interlaminar failure occurring between plies

1 and 2. The interlaminar cracks appeared to have devel-oped from transverse tensile matrix cracks which oc-

curred in the top ply (ply 1) in this plane.

At 90% of the failure load, interlaminar cracks were

found to extend to the free edge of the specimen along

several planes. The failure modes were found to be con-

sistent with those observed at 50% of the failure load. An

additional failure mode, namely shear failure in the

upper plies (1–3) was observed. The failures were foundto occur toward the outer boundary of the specimen.

Fig. 11 illustrates the failures which were present in off-

axis plies on plane O–C. An interlaminar shear crack be-

tween the plies parallel (ply 4) and perpendicular (ply 3)

to plane O–C was also evident. The damage pattern was

found to be similar for each of the specimen diameters.

Fig. 11. Shear failure in the upper plies of the ½0=90=45=�45�vslaminate /boundary = 40 mm at 90% of the failure load (plane O–C).


At final failure, the interlaminar shear cracks were

found to have propagated to the boundary of the spec-

imen on all of the examined planes. Visual inspection re-

vealed compressive collapse of the laminate along the

plane O–D at one side of the hole. The collapse of the

laminate was most likely a secondary failure modewhich occurred as a result of instability due to the loss

of bending stiffness associated with the extensive interla-

minar shear failures. The failure modes were consistent

for each of the specimen diameters.

The ½0=90=45=�45�vs2 laminates illustrated a similar

development of damage to the ½0=90=45=�45�vs lami-

nates. The initial failure mode was inclined intralaminar

shear failure which occurred in the sub-surface plies atthe edge of the washer. As the load was increased, the

damage developed through intralaminar shear cracks

in off-axis plies as shown in Fig. 12. A staircase of shear

cracks extending from the edge of the washer through

the thickness of the laminate was evident in all of the in-

spected planes. The staircase form of the intralaminar

and interlaminar shear fractures was also reported by

Banbury and Kelly [3]. The most prominent failuremode was interlaminar shear failure in the plane of the

laminate where cracks propagated along the interface

between the +45� and �45� plies. Interlaminar failure

along the interface between ply 2 (90�) and ply 3

Fig. 12. Intralaminar shear cracking in the ½0=90=45=�45�vs2 laminate

at final failure /boundary = 40 mm (plane O–D).

(+45�) was also evident, especially on plane O–D. In

the larger diameter specimens, transverse tensile failures

occurred close to the hole in the 0� plies in the centre of

the laminate on plane O–D. More pronounced indenta-

tion and fibre kinking at the edge of the washer was evi-

dent in the ½0=90=45=�45�vs2 laminates in comparison tothe ½0=90=45=�45�vs laminates. Final failure of the

½0=90=45=�45�vs2 laminates was found to occur in a sim-

ilar manner to the ½0=90=45=�45�vs laminates with com-

pressive collapse occurring on plane O–D.

3.4.4. [0=90]vs2 laminates

The initiation of damage in the ½0=90�vs2 laminates was

consistent with the ½0=90=45=�45�vs laminates with ini-tial audible fractures occurring at 20–25% of the failure

load. The initial failure mode was also found to be sim-

ilar with matrix intralaminar shear failure occurring in

sub-surface plies. At 50% of the failure load, damage

was evident on all of the examined planes. On plane

O–B, which is parallel to the fibre direction of ply 1,

interlaminar shear cracks were found to have propa-

gated between plies 1 and 2 and between plies 6 and 7.The interlaminar cracks extended approximately 10

mm from the edge of the hole. Limited damage was

found on plane O–C with only minor intralaminar shear

cracks found in ply 6 close to the edge of the washer.

There are no fibres aligned parallel to plane O–C which

most probably influenced the damage pattern. On plane

O–D, very limited damage was evident with only minor

fractures in the bottom ply. The damage on plane O–Ashowed similar characteristics to that found on plane O–

C but in addition, interlaminar shear failure was evident

between plies 1 and 2 in the region above the edge of the

washer.

At increased load levels, the failure modes remained

unchanged but the extent of the in-plane and through-

thickness damage increased significantly. The ½0=90�vs2laminates exhibited more pronounced inclined intrala-minar shear failures in comparison to the ½0=90=45=�45�vs laminates. At final failure, the inclined intralami-

nar cracks extended through the thickness of the lami-

nate which resulted in numerous compressive shear

buckling failures along the planes parallel to the fibre

directions. Permanent out-of-plane deformation of the

specimens was found to occur as a result of the

through-thickness shear failures.

3.5. Effect of the laminate stacking sequence and resin

system

The laminate stacking sequence was shown to affect

the failure modes and the propagation of damage within

the laminates. Microscopic inspection of failed lami-

nates where ply angles between adjacent plies was 90�revealed widespread delamination. The delamination

was not as pronounced in laminates with 45� between

Table 3

Strength properties of the carbon fibre/epoxy lamina

Xt (MPa) Xc (MPa) Yt (MPa) Yc (MPa) S

1250 730 60 200 70


adjacent plies. The dominant failure mode in these lam-

inates was inclined interlaminar shear cracking. Similar

damage patterns were reported by Banbury and Kelly

[3] who investigated laminates with [0/90/45/�45]ns and

[0/90]8s stacking sequences. Again it can be noted that

the difference in ply orientation between adjacent plieswas 90� at several locations within the laminate.

The difference in ply orientation within the laminates

causes high interlaminar stresses which in turn promote

delamination. Tao and Sun [9] investigated the influence

of ply orientation on delamination in composites plates

made from carbon fibre epoxy pre-preg (AS4/3501-6)

subject to mode II fracture loading. The critical strain

energy release rates of 0�/h� interfaces were determinedexperimentally and the interlaminar fracture toughness

was found to decrease as the off-axis angle (h) increased.After initial failure, transverse loading imposes a mixed

mode fracture load on the laminates with the dominant

mode being mode II. Thus, the results of Tao and Sun,

while based on different materials, can be used to under-

stand the development of damage in the laminates after

initial failure. Similar correlation between stacking se-quence and delamination was reported by Liu [10]

who investigated impact induced delamination. The

delamination area in [0�/h�] laminates subject to trans-

verse impact was shown to increase with increasing h.A theoretical impact damage model by Clark [11] sup-

ported the hypothesis of delamination occurring prefer-

entially at certain interfaces.

The resin system appeared to have the most signifi-cant effect with regard to the final failure mode. Ulti-

mate failure of the epoxy matrix laminates resulted in

the fastener penetrating through the laminate. The

epoxy matrix has a lower failure strain (�f = 1.5%) in

comparison to the vinylester matrix (�f = 11%) and this

was reflected by the internal damage within the lami-

nates where a large network of inter- and intralaminar

cracks was present in the epoxy laminates. In contrast,the ultimate failure of the vinylester matrix laminates

was global collapse where in-plane damage propagated

to the specimen boundary. The damage in the vinylester

laminates was characterised by delamination between

0�/90� interfaces which propagated in-plane. The com-

plex network of smaller fractures under the fastener

head was not evident in the vinylester laminates. After

initial failure of the laminates, the strength appears tobe governed by the interlaminar strength properties.

The results from this study show that the stacking se-

quence influences the mode of failure within the lami-

nate and in particular the extent of in-plane damage.

Table 2

Elastic properties of the carbon fibre/epoxy lamina

E11 (GPa) E22 (GPa) E33 (GPa) G12 (GPa) G13 (GPa

98 7.8 7.8 4.7 4.7

The ultimate failure mode appeared to be governed by

the resin system with the failure mode being shown to

be independent of laminate stacking sequence for

the configurations tested. No conclusive relation be-

tween the stacking sequence/resin and the ultimate fail-

ure load could be deduced from the performedexperiments.

4. Finite element modelling

Finite element analysis was used to investigate the

stress distribution within ½0=45=90=�45�es laminates

when subject to a concentrated transverse load. Thestress distribution around a hole in the laminate is mul-

ti-axial and thus a three-dimensional finite element

model was deemed necessary. Finite element models of

the laminates were developed using the ABAQUS [12] fi-

nite element software. Each ply of the laminate was

modelled using one layer of quadratic solid brick ele-

ments (C8D20). Each ply of the finite element model

was modelled as an orthotropic solid. The in-planeproperties of the lamina were experimentally determined

[13] and the through-thickness properties estimated val-

ues. The material properties which have been used in the

model are presented in Tables 2 and 3.

The load was applied to the model in the form of a

prescribed displacement on the lower surface of the lam-

inate representing the washer. The washer has been as-

sumed to be rigid due to the difference in stiffnessbetween the washer/bolt head and the thin laminate

plate. The circular boundary of the plate was simply

supported i.e. only being prevented from moving in

the transverse direction. A half section of the finite ele-

ment model is illustrated in Fig. 13. It was evident from

the experimental results that the laminates underwent

large deformation and thus the calculations were per-

formed using large deformation analysis.The results from the finite element model were com-

pared with the strain measurements recorded during

the tests. A comparison of the surface strains is illus-

trated in Fig. 14. The strains from the finite element

analysis are shown to compare reasonably well with

the measured strains.

) G23 (GPa) m12 (GPa) m13 (GPa) m23 (GPa)

3.2 0.34 0.34 0.44

Fig. 13. Finite element model of the laminate subject to transverse

load introduction.

0 5 10 15 20 25 30 35 40–0.8

–0.6

–0.4

–0.2

0

0.2

0.4

0.6

0.8

Str

ain

(%)

Radial position (mm)

FE Top 90FE Top 0FE Bot 90FE Bot 0Exp Top 90Exp Top 0Exp Bot 90Exp Bot 0

Fig. 14. Comparison of predicted and experimental strains (P = 0.7

kN).

0 2 4 6 8 10 120

1

2

3

4

5

6

7

8

9

10

Displacement (mm)

Load

(kN

)

Experiment: φ=40mmExperiment: φ=80mmExperiment: φ=120mm Predicted: φ=40mmPredicted: φ=80mmPredicted: φ=120mm

Fig. 15. Comparison of predicted and experimental load–displacement

behaviour.


Through knowledge of the stress field within the lam-

inate, the load at which first-ply failure occurred could

be predicted together with the location of the failure

within the laminate. The strength based failure crite-

rion proposed by Hashin [14] has been used to predict

first-ply failure within the laminate. The criterion distin-

guishes between fibre and matrix failure modes in ten-

sion and compression and requires measurableproperties of unidirectional laminae as input. The

three-dimensional form of Hashin�s criteria is given in

Eqs. (1)–(4),

Fibre tensile fracture, r11P 0:

r11

X t

� �2

þ r212 þ r2

13

S2

� �P 1 ð1Þ

Fibre compressive fracture, r11< 0:

� r11

X c

� �P 1 ð2Þ

Matrix tensile or shear cracking failure, r22 + r33 P0:

ðr22 þ r33Þ2

Y 2þ r2

12 þ r213 þ r2

23 � r22r33

S2P 1 ð3Þ

t

Matrix compressive or shear cracking failure, r22 +

r33 < 0:

1

Y c

Y c

2S

� �2

� 1

" #ðr22 þ r33Þ þ

ðr22 þ r33Þ2

4S2

þ r212 þ r2

13 þ r223 � r22r33

S2P 1 ð4Þ

where Xt is the longitudinal tensile strength, Xc is the

longitudinal compressive strength, Yt is the transverse

tensile strength, Yc is the transverse compressive

strength and S is the shear strength. The rij terms arethe components of the stress tensor. The experimentally

determined strengths of a unidirectional ply of the car-

bon fibre epoxy laminate which have been used in the

calculation of first-ply failure are given in Table 3. The

failure criterion was implemented in ABAQUS in

the form of a user-subroutine (USDFLD) where each fail-

ure equation is defined as a field variable. The load is ap-

plied incrementally with the field variables (failurecriteria) being evaluated at each increment.

4.1. Results from the finite element analysis

The results from the finite element analysis are illus-

trated in Figs. 15 and 16. The load versus displacement

behaviour of the laminates with different boundary

diameters are shown in Fig. 15. The predicted load–dis-placement behaviour is shown to be in good agreement

with the experimentally determined behaviour. The

inclusion of damage propagation in the finite element

model could possibly improve the predictions further.

The results from the simulations were analysed

through plotting the values of the maximum Hashin

failure index at each tangential location (h) within the

0 20 40 60 80 100 120 140 160 1800

0.02

0.04

0.06

0.08

0.1

0.12

0.14

Has

hin

failu

re In

dex

θ (degrees)

Ply 1Ply 2Ply 3Ply 4Ply 5Ply 6 Ply 7Ply 8MAX

0 20 40 60 80 100 120 140 160 1800

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

Has

hin

failu

re In

dex

θ (degrees)


0 20 40 60 80 100 120 140 160 1800

0.2

0.4

0.6

0.8

1

1.2

Has

hin

failu

re In

dex

θ (degrees)


0 20 40 60 80 100 120 140 160 1800

0.2

0.4

0.6

0.8

1

1.2

1.4

Has

hin

failu

re In

dex

θ (degrees)

L8FC

L6FC

L1MT

L7FC

L7MC

L8FCL1

MT

L7MC

L6MC

L8FC

L8MC

L6FC

MC - Matrix compressionMT - Matrix tensionFC - Fibre compressionFT - Fibre tensionL - Ply number n

P=0.5kN

P=1.5kN

P=1.0kN

(a) (b)

(c) (d)

Fig. 16. Failure mode map for the ½0=45=90=�45�es , / = 80 mm specimen at a radius R = 6.5 mm from the hole centre: (a) ply failure indices (P = 0.5

kN), (b) ply failure indices (P = 1 kN), (c) ply failure indices (P = 1.5 kN), (d) failure mode map.


laminate. An example of the analysis is illustrated in Fig.

16(a)–(d) for the case of the / = 80 mm laminate.

The four Hashin failure criteria (Eqs. (1)–(4)) were

applied to the individual plies of the laminate at each

load increment and the maximum value of the failure

index plotted as shown in Fig. 16(a)–(c). A failure mode

map was generated by plotting the maximum failure

index at each tangential location for a given load asshown in Fig. 16(d). At low loads (P = 0.5 kN), the most

highly stressed plies are the top ply (ply 1) where the ma-

trix tensile failure mode is dominant, and the sub-sur-

face plies beneath the washer (plies 6 and 7) where

matrix compression or shear failure is dominant. As

the load increased to 1 kN, the dominant failure modes

remained unchanged although the matrix tensile mode

was active over a larger area. At the predicted failureload (P = 1.5 kN), a more complex mode map exists

with the dominant failure modes being matrix tensile

failure at the top surface (ply 1), fibre compression fail-

ure in the bottom two plies (7 and 8) and matrix com-

pression failure in the sub-surface plies beneath the

washer (plies 6 and 7). The failure modes illustrated in

the failure mode map correlate well with those observed

in the fractographic analysis of the laminates. The failure

mode map also provides an indication of how efficient

the laminate is designed for carrying a specific load.

The predicted first-ply failure loads for the laminates

are presented in Table 4. The predictions are in generalagreement with the experimental values, however the

predicted load levels for the laminates with /boundary =

40 and 80 mm are slightly higher than those

observedexperimentally. A possible reason for the over

predicted failure loads could be that the lamina failure

strengths are ultimate values and thus initial damage

may occur at lower load levels. The initial failure mode

for the larger diameter specimens (/boundary = 80, 120mm) is transverse tensile failure which precedes intrala-

minar shear failure in the sub-surface plies. This change

in initial failure mode was also evident from the fracto-

graphic analysis.

Table 4

Predicted first-ply failure loads of ½0=45=90=�45�es laminates

/boundary (mm) Predicted first-ply failure load (kN) Failure mode Experimental first-ply failure load (kN)

40 1.69 Matrix shear (ply 7, h = 135�) 1.25–1.58

80 1.50 Matrix tensile (ply 1, h = 90�) 1.2–1.4

120 1.02 Matrix tensile (ply 1, h = 90�) 1.1–1.6

h = 0 is the orientation of the surface plies.


5. Conclusions

The behaviour of composite laminates subject to

transverse load introduction was investigated experi-mentally and numerically. The load–displacement

behaviour was found to depend on the effective flexural

rigidity of the specimens with the response becoming

non-linear at large deformations. The load at which

the first audible cracking occurred was found to be

20–30% of the failure load. The damage which resulted

from the initial cracking was not visible on the load–dis-

placement curve and thus did not result in any noticea-ble loss of stiffness. At higher load levels, the load–

displacement curve illustrated minor load drops which

were the result of damage accumulation.

The initial failure mode of the laminates was identi-

fied through fractographic analysis of specimens loaded

to sub-critical load levels. The initial failure mode was

predominantly intralaminar matrix shear failure which

occurred in sub-surface plies beneath the edge of thewasher. Transverse tensile failures were found in the

outermost plies of the laminates with low flexural rigid-

ity but these failures were not pronounced and not

deemed to cause the loss of structural integrity. The ini-

tial damage propagated both through the thickness and

in the plane of the laminate in the form of intralaminar

and interlaminar shear failures. In-plane damage was

found to be more widespread in laminates where adja-cent plies were oriented at 90� to one another. Further

research is required in order to quantify the effects of

the stacking sequence.

Two different ultimate failure modes were identified,

namely fastener pull-through failure and global collapse

of the laminate. Fastener pull-through failure occurred

in the epoxy matrix laminates, where inclined intralami-

nar shear cracks propagated through the thickness ofthe laminate. Global collapse was found to occur in the

vinylester matrix laminates where ultimate failure oc-

curred after interlaminar cracks had propagated to the

boundary of the specimen. The different failure modes

was assumed to be related to a critical level of matrix

strain. The vinylester matrix had a significantly higher

failure strain than the epoxy matrix used in this study.

The effect of the specimen size on the failure load wasinvestigated for each test configuration. Laminates

which failed by fastener pull-through failure showed lit-

tle dependence on the specimen size, however laminates

which failed by global collapse showed a stronger

dependence with the load capacity increasing with the

increasing specimen diameter.

A three-dimensional finite element model was devel-

oped to study the behaviour of the ½0=45=90=�45�es lam-inates subject to transverse concentrated loading. The

predicted non-linear load–displacement behaviour of

the laminates was shown to be in good agreement with

the experimental results. The prediction of the failure ini-

tiation load and location was also found to show reason-

able agreement with experimentally determined values.

Based on the results obtained in this investigation, it

is concluded that the ultimate failure of transverse loadintroductions occurs by a process of damage accumula-

tion. The interaction of intralaminar and interlaminar

cracks plays an important role in the damage develop-

ment within the laminates. Interlaminar failure has been

shown to be a prominent failure mode in the laminates

and its effect should be included in the analysis of trans-

verse load introductions after first-ply failure. It is re-

commended to avoid transverse concentrated loadingwhere possible in composite laminates given that the

first-ply failure occurs at 20–25% of the ultimate failure

load. The fact that the laminates accumulate damage at

such low load levels will have an important bearing on

the fatigue strength of such load introductions.

The present investigation has highlighted that further

research is required in order to gain a more thorough

understanding of the behaviour of composite laminatessubject to such loading and to develop suitable guide-

lines for design.

Acknowledgements

This work has been financially supported by the

Commission of the European Union through GrowthProject TECABS (Technologies for Carbon Fibre Rein-

forced Modular Automotive Body Structures) and by

the Swedish Foundation for Strategic Research through

the national Swedish research program �Integral VehicleStructures�.

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strength and failure mechanisms of composite laminates subject to localised transverse loading

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