Scholarly article on topic 'Pseudo-ductility and damage suppression in thin ply CFRP angle-ply laminates'

Pseudo-ductility and damage suppression in thin ply CFRP angle-ply laminates Academic research paper on "Economics and business"

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{"A. Laminates" / "B. Delamination" / "C. Laminate mechanics" / "D. Mechanical testing"}

Abstract of research paper on Economics and business, author of scientific article — J.D. Fuller, M.R. Wisnom

Abstract Composite materials usage is limited by linear elasticity and the sudden, brittle failure they often exhibit. It is possible to mitigate this inherent limitation and enlarge the design space by using thin plies. This paper presents an experimental study, using a spread tow thin ply carbon–epoxy prepreg material with a cured ply thickness of 0.03mm, which shows that highly non-linear stress–strain behaviour can be achieved with angle-ply laminates, whilst suppressing the damage mechanisms that normally cause their premature failure. Several angles between 15° and 45° are investigated in a [ ± θ 5 ] s layup. It is shown that for all angles delaminations are suppressed, allowing considerable pseudo-ductile strains to develop. Significant fibre rotations take place, permitted by matrix plasticity, leading to a post-yield stiffening of the laminate, as the fibres reorient towards the direction of loading.

Academic research paper on topic "Pseudo-ductility and damage suppression in thin ply CFRP angle-ply laminates"


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Composites: Part A

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Pseudo-ductility and damage suppression in thin ply CFRP angle-ply laminates

J.D. Fuller *, M.R. Wisnom

Advanced Composites Centre for Innovation and Science, University of Bristol, Queen's Building, Bristol BS8 1TR, United Kingdom


Composite materials usage is limited by linear elasticity and the sudden, brittle failure they often exhibit. It is possible to mitigate this inherent limitation and enlarge the design space by using thin plies. This paper presents an experimental study, using a spread tow thin ply carbon-epoxy prepreg material with a cured ply thickness of 0.03 mm, which shows that highly non-linear stress-strain behaviour can be achieved with angle-ply laminates, whilst suppressing the damage mechanisms that normally cause their premature failure. Several angles between 15° and 45° are investigated in a [±05]s layup. It is shown that for all angles delaminations are suppressed, allowing considerable pseudo-ductile strains to develop. Significant fibre rotations take place, permitted by matrix plasticity, leading to a post-yield stiffening of the laminate, as the fibres reorient towards the direction of loading.

© 2014 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY license (http://



Article history:

Received 30 July 2014

Received in revised form 28 October 2014

Accepted 1 November 2014

Available online 10 November 2014


A. Laminates

B. Delamination

C. Laminate mechanics

D. Mechanical testing

1. Introduction

Carbon fibre reinforced polymer (CFRP) composites are well known to possess high stiffness and strength. They are, however, limited by brittle failure, which often occurs without warning and is catastrophic. This linear-elastic to failure stress-strain behaviour reduces design allowables and precludes the realisation of the materials' full potential. Achieving non-linear stress-strain behaviour, with the ability to yield, as metallics do, is therefore highly desirable. Non-linear behaviour and high strains to failure have been demonstrated previously with ±45° angle-ply laminates, often used to determine the shear properties of materials. In these cases, however, the large fibre angle leads to relatively low values of modulus and failure stress. Reducing the fibre angle (towards the loading direction) leads to a higher initial modulus, but despite promise of high strains to failure, laminates of ±0 (where 0 is in the range 15-30°) often fail prematurely - before the development of non-linearity. Failure of angle-ply laminates is primarily due to matrix cracking and delaminations due to high free-edge interlaminar stresses [1]. These failure mechanisms have been widely studied [2-6].

O'Brien [2,3] characterised the onset and development of delam-inations in [+Qn/-Qn/90n]s laminates. Delaminations at Q/-Q and -0/ 90 interfaces were seen to initiate from the edges of the specimens following matrix cracking in 90° plies. A non-linear stress-strain

* Corresponding author. E-mail address: (I.D. Fuller).

behaviour was associated with a 'stiffness loss' brought about by the accumulation of damage. This stiffness loss was coupled with a strain energy release rate (G) approach employed to predict the initiation of delamination. The value of G was found to depend only on the laminate stacking sequence and location of delaminations.

Crossman et al. [5] also used a strain energy concept to determine the failure mechanisms of [±25/90n]s laminates (n = 1, 2, 3), highlighting the importance of the ply thickness in calculating the value of G. It is shown that increasing the number of 90° plies, not only decreases the stress levels required for matrix cracking and delaminations to occur, but also alters the order in which they take place - showing a direct interaction between the damage modes. Treating the adjustment in the number of 90° layers as effectively changing ply thickness, it is postulated that reducing ply thickness could suppress microcracking and delaminations.

Investigating angle-ply laminates, Leguillon et al. [7] examined edge delamination initiation in [Qn/-Qn]s laminates (n = 1-8), comparing tensile test data with predictions. Of the two methods implemented, both showed decreases in delamination initiation stress with increased layer thickness. Herakovich [8] investigated these edge effects using [(+Q/-Q)2]s and [+Q2/-Q2]s laminates, where Q = 10°, 30°, 45°. For all angles tested, the tensile strength, tensile strain and toughness (in this case defined as the area under the stress-strain curve) were each increased for the laminates with dispersed plies. For the 30° and 45° laminates containing dispersed plies, the increased failure stress and strain allowed more non-linearity to develop, highlighting the potential this type of laminate possesses for ductility. 1359-835X/© 2014 The Authors. Published by Elsevier Ltd.

This is an open access article under the CC BY license (

For both the theoretical approaches taken, Leguillon et al. [7] present sharp increases in delamination initiation stress for ply thicknesses less than 0.125 mm. Expense and damage to fibres during the manufacturing process, however, have limited work on reducing the ply thickness below the standard. Recent advances in tow-spreading technology have allowed so called thin prepregs to be produced. Sasayama et al. [9] and the Industrial Technology Centre of Fukui Prefecture in Japan developed a pneumatic technique, which is described in detail by Sihn et al. [10].

Several studies have been undertaken [10-13] to experimentally investigate the general behaviour of thin ply laminates. Sihn et al. [10] performed both static and fatigue tension tests of un-notched and notched quasi-isotropic (QI) specimens, impact and compression-after-impact tests on thin ply laminates (ply thickness, tp = 0.04 mm). In all cases these laminates showed less cracking, delaminations and splitting than specimens with thicker plies (tp = 0.2 mm) of the same material. Results from fatigue testing of the un-notched specimens show the potential of thin ply material. It was demonstrated that, after 50,000 cycles at 60% of the strength, the thin ply laminates maintained stiffness and strength. The thick ply laminates lost in the region of 17% from both the original stiffness and strength. X-ray images taken prior to failure show very little development of damage within the thin ply laminates, indicating their superior damage suppression capabilities. Yokozeki et al. [11,14] conducted investigations covering the com-pressive strength and damage resistance of thin ply QI laminates under both in-plane [11]and out-of-plane [14] loadings. In all cases, the thin ply laminates were shown to be more resistant to damage accumulation. This is particularly noticeable in transverse indentation tests. Thick ply laminates (tp = 0.14 mm) exhibited considerable delaminations on the back face, whereas the thin ply specimens (tp = 0.07 mm) showed only internal delaminations at the same applied load. As presented by Sihn et al. [10], this suppression of damage led to sudden brittle failure. Ogihara and Nakatani [13] presented work on carbon/epoxy angle-ply laminates, also concentrating on the effect of ply thickness. Specimens of ±45° and ±67.5° both showed increases in tensile strength with ply thicknesses of 0.05 mm ([±012]s) rather than 0.15 mm ([04/ -h4]s). A mesoscale continuum damage mechanics model, devised by Ladeveze and LeDantec [15], was employed to show also that the thin-ply laminates were significantly more damage resistant. Highly non-linear strains, in excess of 15%, were recorded for the [±4512]s laminates tested under quasi-static tension. At these large strains, the effect of the geometric rearrangement of fibres towards the loading direction (known as fibre scissoring) becomes important, as stated by Wisnom [16] and Herakovich et al. [17]. Wisnom

[16] showed how taking account of fibre rotations for in-plane shear testing leads to a more accurate representation of both the shear stresses and strains in a [±45]s laminate. Herakovich et al.

[17] coupled fibre rotations with the Ladeveze and LeDantec model to emphasise the importance of their inclusion when predicting the stress-strain response of [±453]s laminates.

In this paper, experimental studies of thin ply angle-ply CFRP laminates loaded under quasi-static tension are presented. The effect of fibre rotation on the laminate stress-strain behaviour and the possibilities for a pseudo-ductile response are investigated. Analyses of the fractured laminates, including X-ray computed tomography (XCT) scans are presented to examine the damage resistance of spread tow thin ply prepreg material.

2. Experimental methods

All testing has been performed using Skyflex USN020A, a commercially available spread tow carbon fibre/epoxy prepreg produced by SK Chemicals. This material consists of Mitsubishi Rayon TR30 carbon fibres (E11 = 234 GPa, strain to failure,

e^ = 1.9%) and SK Chemicals K50 resin, a semi-toughened epoxy. Prior to testing of angle-ply laminates, full characterisation of the material was necessary. Quasi-static tensile tests were performed on [016], [9016] and [±455]s laminates. To ensure sufficient data was collected, batches of 10 specimens were fabricated for each layup. Unidirectional (UD) samples had a gauge length of 100 mm and width of 10 mm, with glass fibre/epoxy prepreg end tabs of length 40 mm. In all tests end tabs were a [(90/0)2]s cross-ply laminate of 2 mm thickness. The [±455]s samples had a gauge length of 150 mm, width of 15 mm and end tabs of 40 mm. Three-point micrometer measurements of laminate thickness, performed prior to testing, yielded a cured ply thickness (CPT) of 0.03 mm (CV = 0.34%). All tests were conducted, using an Instron hydraulic-actuated test machine, under displacement control, using cross-head rates of 1 mm/min for UD samples and 2 mm/min for the [±455]s. The results are shown in Table 1. Determination of £n allowed a fibre volume fraction (Vf) of 42% to be calculated using the rules of mixtures.

The angles chosen for further tensile testing were ±15°, ±20°, ±25°, ±30°. All layups were of the same stacking sequence: [±Q5]s, as used for the ±45° laminates. The dimensions and rate of displacement for these samples were also the same as for the ±45°. Batches of five specimens were prepared for each layup.

All strain data was captured using an Imetrum Video Extensom-eter and associated software. A rectangular grid of video gauge targets was set up, in order to record both longitudinal, ex and transverse, ey, strains. Calculation of fibre rotations and shear stress and strain requires knowledge of both of these. In all cases, the true stress and strain have been computed from the captured engineering strains to account for the change in cross-sectional area at high strains.

2.1. Calculation of fibre rotations

Fibre rotations have been considered in a similar fashion to the approaches taken by other studies [17,16,18]. The fibres are taken as inextensible and idealised to act in a scissoring motion, realigning towards the direction of applied stress. This gives rise to the concept of 'excess length', whereby the reorientation of the fibres allows further strain to be taken by the laminate. The updated fibre angle, Q', is related to the strains, ex and ey, in Eq. (1), where Q is the original fibre angle of the laminate.

h = arctan

Jtan(h) -

2.2. Definition of yield and pseudo-ductility

'Pseudo-ductility', in this case, refers to the geometric effect of fibre reorientation as well as yielding of the matrix. For clarity, yield stress, rY, and pseudo-ductile strain, ed are shown graphically in Fig. 1. The yield stress is defined as the point of intersection between the laminate stress-strain curve and a straight line of the initial modulus offset by 0.1% strain (shown as position 'A' on Fig. 1). The pseudo-ductile strain is the failure strain minus the strain at the same stress level on a straight line of the initial modulus.

2.3. Determination of shear stress and strain

As a change in the orientation of the fibres is accounted for in this study, it is therefore important to apply this to the calculation

Table 1

Elastic properties of Skyflex USN020A.

E11 G12

101.7 GPa 2.4 GPa

E22 V12

6.0 GPa 0.3

/ £d J*

// N.B. Position 'A' on

°Y / £„ axis indicates

0.1% modulus offset

Fig. 1. A graphical explanation of the method used to determine the yield stress and pseudo-ductile strain. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

of in-plane shear stress, s12 and shear strain, c12. Using classical lamination analysis, both are calculated from knowledge of the applied stress, rx, and longitudinal and transverse strains, ex and ey respectively. The value of s12 is found using a formulation presented by Herakovich et al. [17], which is summarised below.

[B(1 - 2m2) + m2]rx

where m = cosh, n = sinh and B is expressed as follows:

m2(2m2 - 1) + 4m2n2 (f22 v12 + l)

4m2n212 (f22 + 2 f22 V12 + 1) +(2m2 - 1 )(m2 - n2)

It has been assumed that the elastic properties £n, £22 and G12 remain constant over the course of the loading. In this work and differently to [17], m and n are updated as the fibres reorient, leading to a change in the value of B.

Shear strain, c12, takes account of fibre rotation via the strain transformation equations for the material directions:

en" m2 n2 mn " ex"

e22 = n2 m2 - mn ey

712. -2mn 2mn m2 - n2 .Cxy.

Isolating c12 and accounting for yxy = 0 leads to:

C12 = -2mn(ex - £y)

where as above, the values of m and n are updated to reflect the change in fibre angle.

3. Experimental results

Large non-linearity in stress-strain behaviour was observed for all laminates, except in the case of [±155]s, which were largely linear. Borne out by the low values for coefficient of variation in Table 2, the response of samples was consistent across each batch. As such, Fig. 2 shows only a representative stress-strain curve for each angle tested.

Also consistent across all samples was the lack of delaminations before final failure of the laminates. Laminates with fibre angles of Q = ±25°, ±30°, ±45° went through three distinct regions on the stress-strain curves: an initial linear part, followed by a yielding and finally a stiffening behaviour as the fibres rotate towards the direction of applied load.

3.1. [±155]s and [±205]s laminates

The [±155]s laminates exhibited an almost completely linear behaviour (Fig. 2). As can be seen in Table 2 yield was not reached. The loading in this case was dominated by the fibre direction ply stress, r11, with very small contributions from the transverse, r22 and shear, s12 stresses, as shown in Fig. 3. Fibre rotations were minimal, reducing the original fibre angle by only 1.5° over the course of the loading.

Non-linearity is increased for the [±205]s layup (Fig. 2), though only in the final stages of loading. The response is still governed by the fibre direction, with a low level of shear stress developed, as shown in Fig. 3. Table 2 shows that the strength remains high, but the non-linearity in the response allows some pseudo-ductility to develop. The fibre rotation in these laminates is more pronounced, with an overall reduction in fibre angle of between 3° and 4°. This is predominantly controlled by the larger, non-linear transverse strain, ey, that develops, as shown in Fig. 2. The effect of the non-linear ey is clear in Fig. 4, as the fibre rotation increases with ex.

The minimal amount of non-linearity demonstrated by these layups can be attributed to the relatively low initial value of G12 (2.4 GPa) for the Skyflex prepreg, meaning that the shear yield point of the material was not reached.

The stress-strain behaviour of these laminates (Fig. 2) was highly non-linear, whilst retaining an initial modulus, £°, of 39 GPa. Laminate failure strains in excess of 3.5% are exhibited, with a small degree of stiffening in the response before complete failure of the sample, which allows strengths of over 950 MPa to be reached. This stiffening of the laminate is shown also by the increase in fibre rotation at high strains. Unlike the previous two layups, there is a promising level of pseudo-ductility with Table 2 showing a ed of over 1%. Consultation of the in-plane shear stressstrain behaviour (Fig. 3) shows that both s12 and c12 are increased compared with the [±155]s and [±205]s laminates. There is, therefore, more influence from the shear relative to the transverse stress on the failure of the sample. The effect of the increased level of fibre rotations is also clear, as the level of s12 decreases following a short plateau region beyond the matrix yield point. This decrease indicates that the matrix has undergone significant plastic flow, allowing the reorientation of fibres that transfers stress away from the matrix on to the fibres.

It is noted from Fig. 3 that there are small differences in the post-yield shear response of each laminate. As the shear stressstrain curves for each laminate should overlay due to being expressed in the principal material directions and so independent of fibre angle, it is thought that this is a consequence of the assumption that the material properties (£22, G12 and V12) remain constant when determining the shear stress from Eqs. (2) and (3).

3.3. [±305]s laminates

As Fig. 2 shows, these laminates have a highly non-linear stress-strain response, exhibiting significant pseudo-ductility. Table 2 shows that the strain to failure exceeds 5%, with strengths of 700 MPa reached after a section of stiffening due to fibre rotations of 7°. Pseudo-ductile strains reach almost 2.9%. The ply-level shear strain, c12, reaches over 10%, showing the increased influence of the shear on the laminate behaviour. In this case, the shear stress does not decrease immediately after matrix yielding, instead showing a longer plateau region up to laminate failure. Not as pronounced as the response shown for the [±255]s layup, this plateau of shear stress still shows how stresses are transferred to the fibre

3.2. [±255]s laminates

Table 2

Experimental results for key parameters. Coefficient of variation is shown in brackets after each value. Batch sizes of 5 specimens were tested, except for the [±455]s, which had 10.

r'x (MPa) (%) ry (MPa) ey (%) ed (%)

[±155]s 1423 (5.54%) 1.75 (4.15%) - - -

[±205], 1220(2.64%) 2.35(4.24%) 1063(5.48%) 1.90(5.23%) 0.28 (24.25%)

[±255]s 952 (7.34%) 3.60 (6.29%) 439 (5.40%) 1.20 (5.03%) 1.22 (6.49%)

[±305]s 727 (1.60%) 5.40 (1.54%) 228 (10.41%) 0.90 (9.32%) 2.88 (3.51%)

[±455]s 390(9.35%) 17.94(7.91%) 64(8.88%) 0.76(7.88%) 13.90(7.50%)

-14.00 -12.00 -10.00 -8.00 -6.00 -4.00 -2.00 0.00 2.00 4.00 6.00 8.00 10.00

Ey [%] £» [%]

Fig. 2. Applied longitudinal stresses are shown against longitudinal and transverse strains for all tested layups. Note that, for clarity, the [±455]s strains have been truncated to half the value reached at failure. Also only one response from each batch of five (10 for [±455]s) is shown. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

-+-15 .......+-20 ----+-25 ---+-30 .......+-45

/ /' / / ,

> 7/ 7 M.........................

0 2 4 6 8 10 12 14 16 is 20 £„ [%]

Fig. 4. The fibre rotations for each layup are shown against applied strain, ex. The increase in reorientation for each layup is clear. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

3.4. [±455]s laminates

Fig. 3. In-plane shear stress-strain behaviour for each laminate tested. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

direction with the onset of matrix yield. The response of this layup shows the same initial modulus as the other orientations, but the shear stress reaches a higher value at the yield point. This increased absolute value of T]2 is related to the amount of fibre rotation that takes place. From Fig. 4, it is seen that the [±305]s laminates exhibit a similar amount of reorientation as the [±255]s. This is, relative to applied strain, a lower amount than expected. It is the fibre rotation that influences the material direction shear stress, so relatively low reorientation leads to higher level of T]2.

Initially tested to provide shear property data for the Skyflex material, [±455]s laminates show highly non-linear stress-strain behaviour, Fig. 2. Shown in Table 2, strains to failure in the region of 20% have been recorded. The stress-strain curve shows three distinct regions: an initial linear response, followed by a large reduction in modulus near 100 MPa and lastly a stiffening of the laminate once strains in excess of 8% are reached. The modest initial stiffness means that the ry is somewhat lower than the other layups tested. There is, however, considerable pseudo-ductile strain of 13.9% - resulting from the low yield stress. Fig. 4 shows how, once the yield point has been passed, the fibre rotation increases with ex. This coincides with what is seen in the latter stages of the stress-strain response, as fibre rotation brings about the large increase in stiffness before laminate failure.

3.5. Pseudo-ductility and yield

Whilst the general increase in non-linearity can be seen from Fig. 2, the interaction of the pseudo-ductile strain, ed, with the yield point is not so obvious. Fig. 5 presents characteristic stress-strain curves for [±205]s (Fig. 5A), [±255]s (Fig. 5B), and [±305]s (Fig. 5C) laminates that have been annotated to show the locations of the yield and the amount of pseudo-ductility achieved.

Increase in original fibre angle from 20° to 30° shows a ninefold increase in ed from 0.33% to 2.92%. Yield stress is shown to decrease by approximately four times from 965 MPa to 230 MPa over the same change in angle. This highlights that these angle-ply laminates allow the response to be tailored to suit the requirements. There is, obviously a compromise to be made, as a high ed is not possible whilst retaining a high yield stress. The response

of the [±255]s in this case looks most promising, as it represents

Fig. 5. Plots show the development of pseudo-ductility and change in yield point with increase in original fibre angle. (A) [±205]s; (B) [±255]s; and (C) [±305]s. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

the middle ground. Yield takes place at rY = 450 MPa and eY = 1.24%, with a ed of 1.23%, while retaining a strength in excess of 900 MPa. Values of ±0 in this region, for the material in question, clearly show the potential for the approach to achieve pseudo-ductility in a high-performance laminate.

4. Post-failure analysis

All samples were tested to failure, following which, a visual inspection of each was carried out, both by eye and optical microscopy. The global failures of each layup were seen to be progressively different from one another, but linked in the level of damage accumulated.

Figs. 6 and 7 present images taken of the typical failure points for [±05]s (0 = 15, 20,25, 30), where each shall be dealt with in turn.

• [±155]s (Fig. 6A) - Failure has taken place perpendicular to the direction of load, indicating that the failure in this layup was clearly dominated by fibre failure. There is an absence of splits running parallel to the fibres, highlighting the local nature of the failure.

• [±205]s (Fig. 6B) - Failure in these layups is via both fibre failure and ply splitting. The introduction of splits parallel to the fibre direction correlates with the non-linear nature of the stressstrain curve, indicating an increased level of s12 in the plies. Most notably, however, fibre failure occurs at the same point as the splits and, as the inset image (Fig. 6E) of the meeting of the fracture surface and free edge, shows, there is a complete absence of delaminations between the split and fractured plies.

• [±255]s (Fig. 6C) - A more interactive failure took place in these laminates. Fibre failure of every ply seems to be predominant, taking place away from the end tab region. Some splitting is also seen to occur that coincides with the fibre fractures. The image (Fig. 6D) showing the highlighted area in more detail at the free edge of the sample, again shows the lack of major del-aminations at the position of failure along the free edge.

• [±305]s (Fig. 7) - Similarly to the [±255]s layup, these samples exhibited both fibre fracture and split dominated failures. In these cases, however, the sample fragments were much larger. The pictured fibre fracture (Fig. 7A) and split (Fig. 7B) being the only points of failure in the gauge length for those particular specimens, with the rest of each intact and bonded to the end tabs. As can be seen on the left of Fig. 8, there was a small amount of metallic-like necking of the sample near the end tab. Both failures shown took place with no delaminations prior to the eventual failure of the samples.

• [±455]s - Pictured in a separate image (Fig. 9), the failure of the [±455]s laminates was a combination of fibre fracture and shear-driven splitting. Predominantly, however, the laminates did not fail within the gauge length, with 80% exhibiting extensive necking (pictured right in Fig. 8) and finally fracturing and pulling out from within the end tab region. Some samples reached axial strains in the region of 20% and did not actually exhibit complete failure, but remained permanently deformed. The fibre rotations in these laminates are clearly visible and one example, is highlighted in Fig. 9, with reorientations of 17°. This, allowing for error from the manual measuring, correlates well with the fibre rotations in the region of 15° calculated from the captured strains.

Common to all samples tested, are the absence of the free-edge delaminations that govern the failure of angle-ply laminates with standard thickness plies. Reduction of the ply thickness to a quarter of the standard 0.125 mm has resulted in a suppression of this damage and has allowed other mechanisms to take place. Fibre rotations are seen to increase with the degree of non-linearity in the stress-strain response and all samples show an increase in the rate of these rotations with applied strain, ex. This increase comes about once the shear yield point is surpassed and the plastic flow of the matrix allows continued deformation. The pseudoductile strains exhibited by the ±25, ±30 and ±45 laminates occur due to this combination of plasticity and fibre reorientation, which would not have possible using standard thickness plies.

4.1. X-ray CT scanning

To further aid understanding of the damage resistance of these thin ply laminates, X-ray computed tomography (XCT) has been conducted, using a Nikon Xtek XT225 with CT Pro reconstruction,

Fig. 6. Micrographs show the typical failure characteristics of (A) ±15°; (B) ±20°; (C) ±25°; (D) highlighted area in C; and (E) highlighted area in B. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

Fig. 7. Micrographs show the observed failure characteristics of [±305]s laminates. (A) Fibre fracture dominated failure and (B) splits and fibre fracture of every ply. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

on one sample from each of the [±305]s and [±455]s layups. Both were tested fully, as part of the main batches, and no special interruptions were made. The [±305]s sample had failed as shown in Fig. 7A, reaching a strain, ex of 5.45%. The [±455]s reached an ex of 19.8% but did not exhibit failure, so was intact. These layups were selected as they exhibit the highest strains to failure and considerable fibre rotation and so represent, in terms of possible damage accumulation, the worst case. While microscopy carried out of the sample edges can give insight into their general condition at failure, XCT yields far better information as to the state of the laminate throughout the cross-section. Thus it allows efficient determination of the locations and extent of any damage that may have occurred. Each scanned sample was first immersed in a dye penetrant of zinc iodide solution to highlight any damage. To maximise the amount of detail visible, 40 mm sections of each of

the laminates were scanned. The visualised section of the [±305]s laminate was located between the end tab and the position of laminate failure. Complete failure did not take place in the scanned [±455]s sample so the section chosen was at the middle of the gauge length.

Using the VG Studio 2.1 Max post-processing software, multiple slices through the thickness of each laminate have been taken to visualise the internal condition of the samples. It was found that there is a complete absence of free-edge delaminations at any point in either sample. Representative images, taken from halfway between the mid-plane and surface of each sample are presented in Fig. 10, [±305]s is on the left and [±455]s the right hand side of the figure. In both images the fibre directions are clearly visible. In an effort to get images of each ply, a scan resolution of 0.018 mm was used - less than the 0.03 mm ply thickness, and

Fig. 8. Image shows metallic-like necking behaviour of [±305]s (upper) and [±455]s (lower) laminates. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

Fig. 9. [±455]s laminates that fractured showed a mixture of fibre and shear driven failure. The permanent deformation of the sample is clear from the largely decreased fibre angle. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

so sufficient to show delaminations if present. The specimens, however, were not perfectly flat, so it is possible to see both fibre directions in one slice.

Free-edge delaminations are visualised by a distinctive triangular shape in angle-ply laminates, made up of matrix cracking parallel to the fibre direction and the delamination front perpendicular to the fibres. It can be seen that there are none of these characteristic triangular areas of dye penetrant in either sample. The ±30 sample shows no signs of matrix cracking, as the image is a monochrome grey showing no ingress of dye penetrant. The ±45 image shows similar traits, though in this case there is some evidence of matrix cracking, shown by the faint white lines that extend across the sample parallel to the fibres. Considering both the higher fibre orientation that leads to some small transverse tensile stresses developing at the very start of the loading and the large strains reached by the ±45 laminates, some matrix cracking is not unexpected. The key is that these cracks have not led to the failure of the sample by acting as initiators for delaminations.

5. Angle-plies and pseudo-ductility

The thin ply angle-ply laminates studied in this work have shown that it is possible to achieve large pseudo-ductile strains with this type of [±hn]s layup. While previous studies of these laminates have centred on establishing the benefits compared with standard thickness material [13], the current work has made use of the effect to bring about a controllable, predictable non-linear

Fig. 10. X-ray CT scans of [±305]s (left) and [±455]s (right) laminates indicate that delaminations do not develop during loading of these angle-ply laminates. Note that the arrows indicate artefacts generated by the tape on the laminate surface used to identify the position of the area to be scanned.

stress-strain response. The damage suppression, in particular, allows a more confident estimate of failure to be made and gives a level of control to the overall stress-strain behaviour. For example, as well as predetermining the initial modulus and amount of linear strain shown before yield by choice of the original fibre angle, the thin ply effect ensures that the post-yield condition is well-defined and damage free. In this way, the structural integrity of the laminate is preserved.

A key aspect of exploiting the pseudo-ductile response of these laminates will be careful choice of the fibre angle. The results presented in Fig. 5 show that there is a large change in yield stress, strength and pseudo-ductility as the value of Q is increased from ±25° to ±30°. At ±30°, the yield stress and strength are considerably reduced from the values shown by the [±255]s, though the pseudo-ductile strain is tripled. These findings suggest that the region between these angles may give even better results in terms of optimising these parameters.

This study has focused on one material. If the results are compared to other studies of angle-ply laminates, they are seen to be very promising in terms of pseudo-ductility and strain to failure. Ogihara et al. [19] performed tensile tests on [±0]s angle-ply laminates with angles of ±15°, ±30° and ±45°, using a T700S-2500 carbon-epoxy prepreg material, with a E11 of approximately 100 GPa, which is similar to the value found for the Skyflex material (Table 2), and ply thickness of 0.13 mm. The ±15°, similarly to the results presented here, displays a predominantly linear stress-strain response. The recorded strength and strain to failure, however, were 780 MPa and 0.8% respectively compared with 1423 MPa and 1.75% demonstrated by the Skyflex. More non-linearity developed in the ±30° and ±45° laminates with strengths of 400 MPa and 155 MPa, and strains to failure of 1.59% and 4.50% respectively. Using the stress-strain plots in [19] and applying the methods presented in Fig. 1, estimations of the yield point and pseudo-ductile strains can be made for these laminates. In

each case, the experimental values presented in this study are Acknowledgements given in parentheses. The ±30° exhibited ry of 250 MPa

(228 MPa), ey of 0.70% (0.90%) and ed of 0.50% (2.88%). The ±45° showed a ry of 85 MPa (64 MPa), ey of 0.65% (0.76%) and ed of 3.5% (13.90%). For both studies the point of yield is seen to be similar, but the ed demonstrated in this work, using thin ply material, is considerably higher. The standard thickness plies used in [19] clearly suffer from the sort of premature failure discussed above. As such, the thin ply composites reach values of strengths that are approximately double those exhibited by the standard thickness ply composites.

The promising results exhibited by the Skyflex, opens the door to extend this concept of achieving pseudo-ductility through 'excess length' to other spread tow materials. A key aspect of this would be to use a material with increased Vf. The volume fraction of the Skyflex, at 42%, is considerably lower than the industry standard range of 55-60% and as such, significantly improved results could be obtained by increasing the Vf to this level. Ultimately, the combination of fibre and matrix can be designed to maximise the desired properties of high initial modulus and yield stress whilst also reaching high levels of pseudo-ductility.

6. Conclusions

Pseudo-ductility has been demonstrated by angle-ply rotation with suppression of delamination using thin carbon/epoxy plies of 0.03 mm thickness. Laminates of [±Q5]s, where Q = 15, 20, 25, 30 and 45, have been tested under monotonic tensile loading. Progressively higher and more non-linear strains, recorded using video extensometry, were obtained as the initial ±Q angle was increased. The shear stress and strain were shown to increase similarly, showing that as the shear yield of the matrix is reached, the shear stress plateaus, as the yielding allows an increase in fibre rotation. This process effectively increases the stress on the fibres, which is evidenced by the stiffening behaviour seen in the stressstrain curves and also by the fibre fractures exhibited by all laminates at failure.

It has been shown, via X-ray computed tomography and microscopy, that the thin ply material suppresses delaminations that are the predominant cause of failure in angle-ply laminates. This suppression of damage allows considerable strains to failure and pseudo-ductile strains to be realised via fibre rotations and matrix plasticity. The fibre reorientations exhibited by the tested laminates show that the concept of 'excess length' is able to give pseudo-ductility in a CFRP laminate whilst retaining initial stiffness and high strengths. [±255]s laminates in particular show strengths in excess of 950 MPa, failure strains of 3.6% and pseudo-ductile strains of 1.2%. For [±305]s laminates, failure strain is improved to 5.4% (an increase by a factor of 1.5), pseudo-ductile strain is more than doubled to 2.88%, whilst strength reduces by only 25%.

This work is part of the EPSRC Programme Grant EP/I02946X/1 on High Performance Ductile Composite Technology in collaboration with Imperial College, London and is financially supported by Grant No. EP/G036772/1 (as part of the ACCIS Centre for Doctoral Training).


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