epoxy composites under combined stress

Composites

and Technolqv 53 (IYYS) 253-2.58 0 199.5 Elsevier Science Limited

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THROUGH-THICKNESS FAILURE OF LAMINATED CARBON/EPOXY COMPOSITES UNDER COMBINED STRESS G. Y. Department of Engineering

Cui & C. Ruiz

Science, University of Oxford, Parks Road, Oxford,

UK, OX1 3PJ

(Received 29 March 1994: accepted 27 September 1994) A simple modification to the plain C-specimen consists of reducing the width at one side of the apex. In this way, interlaminar shear stresses are added to the radial through-thickness tension and the tangential tension or compression. In this paper the use of the waisted C-specimen for studying delamination under combined stresses is described. Both experimental data and SEM fractographic pictures reveal very different failure mechanisms depending on the state of stress.

Abstract The through-thickness failure of laminated composites has been investigated experimentally by the use of waisted C-specimens. In this configuration, the test section is under in-plane tension or compression combined with interlaminar shear and tension. The results show that the interlaminar shear stress has an important effect when combined with in-plane tension, resulting in delamination. The effect is less pronounced under in-plane compression. Fractographic SEM pictures also reveal very different features for each of the stress states. Keywords: failure, through-thickness, laminate, composite

THE WAISTED

combined stress,

C-SPECIMEN

TEST

In the configuration shown in Fig. l(c), by reducing the width of the specimen over a region at one side of the apex, it is possible to localise the failure away from the apex and thus combine shear stress, rrtr with the radial and tangential components or and cr,. According to the theory of elasticity solution for isotropic materials4 the state of stress, ignoring stress concentrations, is given by

INTRODUCTION Interest in the through-thickness strength of laminated composites has prompteti substantial research effort. With more complicated composites structures being used, there is an increasing awareness of the through-thickness strength of the laminate, which is very low compared with the in-plane strength. It is obvious that interlaminar normal and shear stresses, acting either by themselves or together with in-plane stress, can lead to the occurrence of delaminationthe main through-thickness failure mode. Two kinds of curved beam specimen configurations (semicircular and elliptical) have been proposed’ for determining the interlaminar tensile strength of a laminate composite subjected to tensile loads. Wu et al2 calculated the through-thickness stresses generated in curved laminates by bending moments. Hognestad3 tested C-specimens made of woven carbon-fibre-reinforced composite (FRC, O/90) Subjected to tensile loads. The plain C-specimen is shown in Fig. l(a). The maximum stresses occur at the apex, which is the region where failure takes place. The stress system at that point consists of in-plane bending and through-thickness tension, as shown in Fig. l(b). In this configuration, there are no interlaminar shear stress in the failure region.

ffr= -4-

Pe

a2b2

-lnb+b21n~+a21n1

N, ( r2

r

a

(a” + b’) r

cr,=-4-

Pe

N,

a2b2

b

r2

a

--ln-+b21n~+a21na+b2-a’

r

(1)

(a’ + b2) r

(a’+

b’) r

where

N2=a2253

b2+(az+b2)lnb

a

G. Y. Cui, C. Ruin

254

Distance Along Width (mm) (a)

Distance Along Width (mm) Fig. 1. Plain C-specimen

C-specimen.

and waisted C-specimen. (a) Plain (b) The stress state at the apex of a plain C-specimen. (c) Waisted C-specimen.

In the tests reported here, the specimens consisted of 12 layers of O/90 woven-carbon-fibre prepreg with a total thickness of 2-90 mm. The anisotropy is therefore mainly in the radial (through-thickness) direction. The dimension and waist configuration were the same for all specimens except for the position of the waist, which is defined by the angle 8 in Fig. l(c). The width of the specimen was 6 mm, the minimum width at the waist was 3 mm, the inner radius was 15.7 mm and the waist radius was 20 mm. The waist was carefully machined by grinding with much attention being paid to minimising any mechanical damage. A detailed finite element stress analysis by ABAQUS’ for the waisted C-specimen used here was found to be in very good agreement with the analytical formulation for plain C-specimen@ and the results also showed that the additional stresses induced by the waist were very low and could be neglected. Figure 2(a) shows the

04 Fig. 2. Tangential

stress distribution along width of the specimen at the waist: (a) on the outer surface; (b) on the inner surface.

variation of tangential stress along the width of the specimen at the waist on the outer surface. The stress is compressive: the maximum value is at the edges and exceeds the average stress by less than 10%. The same conclusion is reached for all of the other stress components through the specimen waist. Figure 2(b) shows the tangential stress at the inner surface, where the stress concentration factor is no more than 10%. The specimens were loaded in an Instron testing machine and the tests were monitored with a video tape recorder which permitted the identification of the instant of failure and the visualisation of the process of crack growth. Six groups of specimens were tested, with the angular position of the waisted region between 30” and 70”, as detailed in Table 1.

255

Failure of laminated carbon /epoxy composites Table 1. Waisted C-specimen test results Calculated

stresses

Failure load (kN)

Failure position (PRYno.)

r, @@a)

0, (MPa)

r,, (MPa)

145.6 210.2

3 11

225.58 -501.5

16.23 7.66

17.21 8.0

94.4 81.1 153.3

3 3 10

175.93 151.15 -355.78

12.62 10.84 12.48

10.08 8.66 9.85

50

132,3 90.6

3 5

276.79 59.03

19.82 16.59

11.77 9.82

55

166.7 182.8

7 7

31.22 34.24

15.68 17.2

60

89.8 45.1 01.2

4 6 3

135.12 -1.63 234.0

17.09 9.10 16.74

7.21 3.84 7.08

70

14.5 65.1 22.8 114.9

3 5 3 9

280.5 49.92 300.8 - 265.24

20.05 13.92 21.5 16.94

5.46 3.78 5.86 4.58

Waisted position (0”)

Specimen no.

30 : 40

a

d

RESULTS

Table 1 summarises the results of the tests. In each specimen the plies are numbered from 1 to 12 starting with the inside ply. In any given specimen geometry, delamination did not always occur between the same plies. For example, for the first two tests, with 8 = 30”, delamination occurred in one case between plies 3 and 4, on the tension side of the section (towards the inside face) while in the other case it was between plies 11 and 12, i.e. near the outside face and in compression. This was expected from the results of plain C-specimen tests and is due partly to the unavoidable variability of the mechanical properties of the material. But the main reason is the presence of at least two competing modes of failure, one in which the fibres are in tension and the other in which fibre buckling under compressive stress is possible. The results are plotted in Fig. 3 together with those obtained from plain C-specimens (no shear).3 Note that the three axes have different scales reflecting the relative magnitudes of the three stress components. The composite is much weaker in the throughthickness direction (in tension or in shear) than along the fibres. The experimental variability is indicated by the scatter between the results and the best fit curve for the no-shear case, i.e. about i10 MPa. In absolute terms this is a fairly small magnitude but it is, of course, of large relative importance when compared to the value of the radial stress, which is below 40MPa, and of the shear stress, which is below 20 MPa. This means that it is not possible to draw any firm quantitative conclusions from a single experiment and

_

- 126.6 - 138.8

0,

Fig. 3. Waisted C-specimen test results, together with the data from plain C-specimen test (no-shear).’

that all the experimental data must be considered together. For simplicity, all the results have been projected onto the ((T,-uJ plane, against the no-shear data, as shown in Fig. 4. The fracture surfaces of the specimens were carefully examined using scanning electron micros: copy (SEM). In order to prevent charging, the surfaces were coated with gold prior to SEM examination. The SEM pictures taken are shown in Figs 5 and 6.

G.

256

4

Y.Cui. C. Ruiz

WW

Fig. 4. Effect of shear stress on failure by comparing the results possessing shear with no-shear data in CT,- CT,plane.

DISCUSSION While it is possible to draw a continuous surface based on the experimental results, the value of such an exercise is questionable not only because of the paucity of data but because to do so assumes that

Fig. 5. SEM fractographs

of delamination

failure is governed by a single mechanism. That this is not the case can be deduced from Fig. 4 which shows that while the shear stress, z,~, has very little effect on failure in the negative g, quadrant, it has a pronounced effect in the positive quadrant. This is a first indication of the existence of at least two modes of failure: when (T, is compressive, fibre buckling dominates; when it is tensile, interlaminar shear is the dominate mode. The SEM fractographs offer further evidence to support this conclusion. A distinction has to be made between interlaminar crack opening in Mode 1, due to through-thickness tension or augmented by the fibre buckling, and in Mode II as a result of interlaminar shear. Smith and Groves7 have conducted extensive fractographic research for Mode I and Mode II fracture of graphite/epoxy laminates and found that for Mode I (pure tension) the fracture surface was relatively smooth and featureless while for Mode II (pure shear) the surface revealed numerous inclined hackles of epoxy oriented normal to the direction of resolved tension. This conclusion was based on the study of pure failure modes. In practice failure follows a combined

layer caused by u,( - )-a,( + )-z,~ stress state.

257

Failure of laminated carbon Jepoxy composites

Fig. 6. SEM fractographs

of delamination

mode and it is more difficult to draw a clear distinction from the fractographic features. In the present case, the situation is complicated by the interaction between interlaminar ((T,, r,J and in-plane (vJ stresses. The existence of a tensile, through-thickness stress introduces a Mode I component. The effect of a tensile tangential stress is to maintain the plies straight and aligned, inhibiting the Mode I opening. The fractographic pictures shown in Figs 5 and 6 are therefore more complicated than those obtained under pure Modes I or II. When ot is compressive (Fig. 5) the resin imprint on the fibres exhibits many inclined hackles with smooth or sharp faces, not entirely unlike those found in the brittle cleavage fracture of metallic materials. Mode I is regarded as governing. On the other hand, when (T, is tensile (Fig. 6) the fracture surface of the resin exhibits a large number of round dimples of the type associated with failure as result of microvoid growth and coalescence as found in the ductile fracture of metallic materials. These observations confirm that there is a clear difference between the two stress situations.

layer caused

by o,( + )-a,(

+ )-I,,

stress state.

CONCLUSION

The plain C-specimen data have shown a significant effect of the tangential stress, (T,, on the delamination strength under the through-thickness tension, IT,. As the stress changes from tension to compression, the delamination strength drops significantly. This is associated with the effect of fibre buckling under compression and is a well known phenomenon, although still not fully understood. The use of the waisted C-specimen permits the study of the effect of interlaminar shear, z,.~. From this investigation, it follows that in the carbon/epoxy woven laminates tested, rrt affects the delamination strength when U, is tensile and a shear (Mode II) failure is dominant but it has very little effect when (T, is compressive and a cleavage (Mode I) prevails. A full answer as to why shear functions so differently in these two stress states has not yet been found. The opacity and poor reflectivity of the material used makes it very difficult to identify the onset of damage and the process of crack growth. A transparent glass/epoxy system, currently being used in preference to the

G. Y. Cui, C. Ruiz

258

carbon/epoxy, is a better investigate this question.

model

material

to

ACKNOWLEDGEMENTS The authors would like to acknowledge the help of Rolls-Royce plc, who supplied the experimental materials and supported the investigation. The support of the British Council is also gratefully acknowledged. REFERENCES 1. Hiel, C. C. et al., A curved beam test specimen to determine the interlaminar tensile strength of a

laminated composite. J. Comp. Mater., 25 (1991). 2. Wu, Y. S. et al., Delamination of curved composite shells due to through-thickness tensile stresses. Plastics, Rubber and Composites Processing and Applications, 19 (1993) 39-46. 3. Hognestad, G., DPhil thesis, Oxford University, 1993. 4. Timoshenko, S. & Goodier, J. N., Theory of Elasticity. McGraw-Hill, 1970. 5. HKS Inc. ABAQUS User’s Manual, Version 5.2, 1992. 6. Cui, G. Y., DPhil first year report, Oxford University, 1992. 7. Smith, B. W. & Groves, R. A., Determination of crack propagation directions in graphite/epoxy structures. In Fracture of Modern Engineering Materials: Composites and Metals, ASTM STP 948, ed. J. E. Masters & J. J. Au.

American Society for Testing phia, PA, 1987, pp. 154-73.

and Materials,

Philadel-