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Composites fractography without an SEM — the failure analysis of a CFRP I-beam
Composites fractography w i t h o u t an S E M - - the failure analysis of a CFRP I-beam D. PURSLOW This paper describes the fractographic analysis of a typical aerospace structural element made from CFRP. The methods of fracture characterization and failure sequencing are described in detail and the cause of ultimate failure successfully determined using low-power optical microscopy only.
Keywords: composite materials; [-beams, fractography; optical microscopy; three-point bending; failure mechanisms.
The fractographic analysis of broken coupons and components made from carbon fibre-reinforced plastics (CFRP) is becoming an increasingly frequent and important facet of the development of successful composite aerospace structures. In the understanding of basic fracture processes and in many component failure investigations, the use of a scanning electron microscope (SEM) is indispensable. 1-16. However, extracting relevant samples, sometimes many from a large component, mounting, and coating with a conducting surface, is a time consuming occupation. Also, there are certain to be circumstances in which access to an SEM is limited or unobtainable. This paper therefore describes the successful failure analysis of an aerospace structural element using lowpower optical microscopy alone. It must be added, of course, that the ability to use this technique is based
on considerable experience in composites fractography using an SEM. n°'~3"~5~6 The component chosen to illustrate the efficacy of low-power fractography is part of a research programme to investigate the relationship between defects and performance in structural components? 7 in this case a typical I-beam loaded in three-point bending.
SPECIMENDESIGNAND TESTING The I-beam is illustrated in Fig. 1. Two back-to-back channel sections of -t- 45° CFRP formed the web which was strengthened in the loading areas Z by tapered 90 ° + 45 ° reinforcement pads on both sides (pad layup: [(04 °. + 45 °. - 45°)3]0. The web and reinforcement pads were co-cured: pre-cured 0° + 45 ° tapered caps were then bonded to the web flanges. Details of materials and fibre lay-up of the caps are given in Fig. 2.
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O010-4361/84/010043-06 $03.00 © 1984 Butterworth 8- Co (Publishers) Ltd COMPOSITES. VOLUME 15. NO 1. JANUARY 1984
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A series of beams with various types and positions of artificial defects were tested in three-point bending both statically and in fatigueJ 7 The specimen described in this paper was manufactured as a 'non-defective' control and was loaded to failure statically.
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GENERAL DESCRIPTION OF FAILURE The broken specimen is shown in Figs 1, 3 and 4. (The lettering is consistent between all of the figures). The compression cap delaminated and split in several places between A and B and broke at positions C, D and E. The undelaminated cap plies and web flanges failed at E and one central reinforcement pad broke away from the web. The tensile flanges and cap remained intact and the web was therefore cut along xx to allow the fractured segments to be separated and to reveal the complete web failure EFGH.
METHOD OF FAILURE ANALYSIS
Fig. 3
In the fractographic analysis which follows, each mode of failure which occurred in the beam will be described and illustrated by a typical area taken from the specimen, together with an outline of other areas which suffered the same mode of failure. From this evidence, a diagram of the relative movements of the fractured segments of the structure during failure will be drawn and the detailed failure sequence of the complete beam will be determined, thus revealing the site and mode of critical fracture and cause of ultimate failure.
Failed segments of I-beam E
Fig. 4
44
Failure of web
Interlaminar shear Delamination of a multidirectional CFRP laminate has been shown ~° to occur preferentially at the boundary of plies of different orientation, one of which approximates to the direction of shear. Most of the delamination occurring in the I-beam took place in the compression cap and thus we may expect to find most interlaminar failures between the 0 ° and 45 ° plies. For example, the delamination between plies 23 and 24 (+ 45°/0 °, Figs 1 and 2) is shown in Fig. 5. However, the delamination between plies 27 and 28 (+ 450/- 45°), illustrated in Fig. 6, is a typical exception as will be seen later. The most significant features in a delamination are those in the matrix. The presence of misaligned fibres such as S (Fig. 5) and T (Fig. 6) causes the formation of resin rich areas and hence larger fracture features in the matrix. Also, by comparing the orientation of matrix features adjacent to misaligned
COMPOSITES. JANUARY 1984
ultimate failure, the crack spacing decreasing with increasing 0 ° stress - - a similar p h e n o m e n o n occurs in tension. For a given laminate the crack spacing may be used to provide an indication of the stress sustained. Where the cracked 45 ° plies subsequently delaminate, they frequently leave residual fibres (Fig. 8) indicating the crack spacing which was present at the time of delamination. Hence an estimate of the stress at which delamination occurred may be made. In this case, the crack spacing indicates that the delamination occurred at a 0 ° compressive stress approaching the strength of the material, after the in-plane 0 ° shear failure described later. Fig. 5
Fig. 6
Interlaminar shear failure, 0*/45* (x 20)
Interlaminar shear failure, + 4 5 0 / - 4 5 * (x 20)
Using these features and plotting the shear directions over the whole area of delaminations, it is possible to deduce the following. •
The cap delaminated between plies 15 and 16 before the web, flange and remaining cap failed at E since, apart from minor post-failure bruising, there is no evidence of the failure at E on the surface of ply 16 (see Figs 1 and 2).
•
Delamination between plies 15 and 16 commenced around C2, radiated to Bz, BI, C1 and towards PI P2 where the delamination changed across the pre-cracked + 45 ° ply (15) to plies 14/15, continued to D~D2 and subsequently to AtA2.
•
Delamination between plies 23 and 24 commenced at a split along the 0° plies MN, spread initially at + 45 °, propagated to C1 Cz and then extended to cover the area D~ D2 K M R C2 Cv (Note that ply 24 stops 120 m m from the cap centre. Delamination is then between the taper and ply 33.) Some delamination over the area C~ R L Q also occurred between plies 19/20/21.
fibres with the general features, the effect of ply orientation may be assessed. Looking in detail at Fig. 5, the misaligned fibre itself has been removed during the failure process but the fractured resin, particularly to the left of the fibre, S, shows clearly the cusps formed during a shear failure. ~a (Cusps are sometimes referred to as hackles? ,8.~2) These cusps may be observed over the whole area of both surfaces and are characteristic of a shear failure in which the mating surface moved in the direction arrowed. It is important to note that in assessing the overall failure directions, m a n y sites must be inspected since local perturbations may easily give rise to erroneous conclusions if they are based on only one or two observations. Comparing Figs 5 and 6, important differences may be observed in addition to the fibre orientations. Apart from the matrix associated with the misaligned fibres S and T, the general texture of the failure surface in Fig. 5 is much smoother than that shown in Fig. 6. Also, although it is not possible to illustrate, the colouring of the surface in Fig. 5 is amber whereas that in Fig. 6 is white. The amber colouration is that of the matrix and the surface of Fig. 5 is the result of the thicker, less finely textured matrix. These differences indicate that in the case of the 00/45 ° failure (Fig. 5), either the interlaminar bond was relatively weak or there was an interlaminar tensile component to the failing stress. Where post-failure movement has caused abrasion of the mating surfaces, the fractures appear dull, comparatively featureless and covered with debris (Figs 7 and 8). Both figures show part of the 00/45 ° shear failure between plies 15 and 16 (Figs 1 and 2). In Fig. 7 it is possible to detect cracks in the 45 ° plies at about 0.5 m m spacing. These cracks are due to the 0 ° compressive stress and start to form well before
COMPOSITES. JANUARY 1984
Fig. 7
45 ° cracking in interlaminar shear failure (x 7)
Fig. 8
45 ° crack remnants in interlaminar shear failure (x 7)
45
Shear
J peel
u
0~.~ 'ly .~8
_~7
F4
C1 Fig. 9
•
Comparison of shear and peel surfaces (x 0.3)
Delamination between plies 28 and 29 was a 0 ° interlaminar shear failure following the split along LR.
•
The 45 ° cracking occurred only over the area C1 R L Q. This cracking and subsequent delaminations occurred after the split MN which caused an increase in the cap compressive stress on the C~ D~ side of the web.
•
The delaminations between plies 15/16 and 23/24 showed evidence of either the influence of interlaminar tensile stresses or low interlaminar strength.
Interlaminar shear failure also occurred between the web and central reinforcement pads and was such that close to the line EF (Fig. 1), the left half of the web moved in the direction E ---' F relative to the pads.
Peel The delamination of a multidirectional CFRP laminate due to peel forces normally takes place within a ply orientated close to the direction of peel and near to the boundary of that ply with one of different orientation? 3 The failure surface consists largely of fibres (or fibre imprints in the matrix), may contain m a n y loose fibres and suffers little abrasion. It is therefore highly reflective as seen in Fig. 9, where the interlaminar shear failure between plies 23 and 24 may be compared with the subsequent peeling. Also clearly visible are the imprints of the serrated terminations of plies 3 to 16. The only peel failure to occur during breakage of the I-beam took place from C2 towards D2 in ply 24. The remaining peel failures are those produced by separation of the broken segments during analysis (including Fig. 9), their distinctive appearance facilitating the identification of such artefacts.
In-plane shear Failure due to in-plane shear stress takes place preferentially in a plane parallel to a fibre direction, l° Fig. 10 shows the failure containing plies 24 to 28 along MN (Figs 1 and 2). Considering first the 0 ° fibres (plies 24-27), the broken fibres, u, failed in compression indicating that the mating surface moved in the direction arrowed. In a 0 ° + 45 ° laminate, adjacent 45 ° layers subjected to stresses due to 0 ° in-plane shear prefer to fail along the 0° line in the layer in which the resolved 45 ° component is compression rather than tension? 6 The compression failure of ply 28 ( + 45 °) is also in the direction arrowed (Fig. 10); it is thus consistent with the generalization above and shows that the 0 ° failure preceded that of the 45 ° ply. Fig. 11 illustrates part of the mating surface to that shown in Fig. 10 in which several features may be noted. There is no delamination between the -I- 45 ° and - 45 ° plies (28 and 29) where they meet, thus their delamination took place after the 0 ° shear failure. The breakage steps
46
Fig. 10
In-plane shear failure (x 20)
Fig. 1 1 + 4 5 ° compression along LR (x 20)
in the + 45 ° ply failure (Fig. 11) are not perpendicular to the laminate plane. Since it has already been shown that the failure started in the 0 ° layers, the angle suggests that the in-plane fracture progressed from right to left (R-L, Figs 1 and 2), ie also in the direction arrowed. Hence the failure is likely to have originated at the right hand termination of ply 28 (R, Figs 1 and 2). The peel failure described earlier (plies 23/24) took place across the width dictated by the 0° shear crack and hence the in-plane failure (plies 24-27) must have occurred before the delamination (plies 23/24). It is important to note also that the shear failure along LR closely coincides with the length and thickness of the central web reinforcement. Failure of the web along EF (Fig. 1) was by in-plane shear, the left hand segment moving from E to F relative to the right hand segment.
Compression Fig. 12 illustrates part of the cap 0 ° compression failure C1 C2 (Fig. 1) near C2. The upper half of the picture shows the peeled surface which, it should be noted, contains evidence of the compression failure. It may thus be deduced that the peel failure occurred at the time of, or after the compression failure. In comparison, Fig. 13 shows the 0 ° compression fracture close to C1 and it is clear that, apart from some post-failure bruising, the delaminated surface contains no evidence of the compression failure. Hence, in this area, the delamination occurred before the compressive fracture. Fig. 14 illustrates part of the mating surface to that shown in Fig. 13. It can be seen that the 0 ° plies have, in fact, failed in two distinct modes - - c o m p r e s s i o n of the outer plies and tension on the innermost? ° Hence the 0 ° plies in this area failed due to combined compression and flexure: this confirms that delamination must have occurred prior to the compressive fracture and allowed buckling to take place. Similarly, it may be deduced that the cap failure at D~ was a combined compression/flexural failure and took place at the time of the adjacent delaminations between plies 23/24 and 14/15/16. In the region of D2, only partial compression failure occurred since plies 17 and 18 terminated at D 2.
COMPOSITES. JANUARY 1 9 8 4
Fig. 12
0 ° compression failure before delamination (x 1 5)
Fig. 1 3
0 ° compression failure after delamination (x 1 5)
Fig. 15
Web compression failure (x 8)
Fig. 16
Compression/bond failure of web (x 3)
Compression
Tension Fig. 14
0 ° compression/flexure failure (x 1 5)
Part of the _+ 45 ° web failure between F and G is shown in Fig. 15. (The plane surface to the right of the picture is the cut along x-x). This failure is due to compression perpendicular to the + 45 ° axis FG and illustrates the problem of abrasion between the failed surfaces which so often follows compressive failure. The angle of failure seen on the cut surface is typical of a pure compressive failure and also shows that no delamination of the web occurred in this area. Similar compressive failures took place along EH and H G due to in-plane forces approximately perpendicular to those lines. Compressive failure of the
remainder of the cap and web flanges occurred at E.
Cap~flange bond Part of the failure of the cap and web at E (Fig. 1) took place at the bond between the web flanges and cap as shown in Fig. 16. Normally failure at a good bond happens just within the CFRP as at V. ~3 The white honeycomb pattern in the lower flange is produced by the adhesive and carrier, very little area being bonded. However, the portions of the cap compression failure, W, remaining attached to the flange indicate that the cap failure took place before that of the bond despite its poor quality.
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Movement diagram
COMPOSITES. JANUARY 1 9 8 4
47
MOVEMENT DIAGRAM AND FAILURE SEQUENCE The relative movements of the various segments of the fractured I-beam as deduced above are shown in Fig. 17. F r o m this diagram and the preceding analysisl the following failure sequence may be obtained. Ultimate failure originated in the compression cap. A 0 ° in-plane shear failure occurred in plies 24-28 along MN, originating at R (the termination of the 45 ° plies 28 and 29) and causing interlaminar shear failure between plies 28 and 29 over the area LROQ. Delamination between plies 23 and 24, also originating along MN, then occurred and propagated towards C allowing the outer plies to buckle. When this delamination reached C, where 0 ° plies 25 and 26 terminated, compressive/flexural failure of plies 23, 33 and 34 commenced at Ct and propagated to the 0° split RN, prior delamination having occurred between C~ and N. Some delamination of plies 19/20/21 also took place over a similar area. Delamination between plies 15 and 16 c o m m e n c e d at C~ C2 at this stage and spread towards D, together with an extension of the delamination between plies 23 and 24. Plies 15-23 buckled and failed in compression/flexure when the dela- • minations on both sides reached D~. Simultaneously the compressive/flexural fracture at C continued to C2, followed by a peel failure in ply 24 which extended, together with the 14/15 delamination, to the termination of plies 17 and 18 at D2 due to the lower stress and greater stiffness in this less damaged cross-section between M N and side 2. The buckling described above accounts for the indications of interlaminar tensile forces mentioned earlier, although other work has shown the interlaminar strength to be low. At this stage less than half the thickness of the compression cap remained intact over the central area of the beam, the area of m a x i m u m 0 ° load. The remaining cap and web flanges thus failed at E, the boundary of the central reinforcements where the poor bond was also located. The intact tension cap then bent at G causing the web to finally fail as shown in Fig. 17.
CONCLUDING REMARKS The analysis has illustrated the techniques of fracture characterization and failure sequencing using lowpower optical microscopy. Using these techniques, the initial cause of ultimate failure of a typical aerospace structural element has been successfully determined. Critical failure of the I-beam occurred in the compression cap due to an in-plane shear stress concentration. This stress concentration was located at the termination of the ___ 45 ° plies which coincide closely with the boundaries of one of the central reinforcements.
48
ACKNOWLEDGEMENT © Copyright Controller, HMSO, London, 1983
REFERENCES l
2 3 4
5 6 7
8 9
10 11 12
Auvinet, J. and Rouchon, J. 'The possibilities of using a
scanning electron microscope for the study of composite materials having an organic matrix' RAE Library Translation 1874 (1975) Adams,D.F. 'A scanning electron microscope study of hybrid composite impact response"JMater Sci 10 (1975) p 1591 't Hart, W.G.J. 'Scanning electron microscopy of fracture surfaces of carbon composite materials" NLR TR 76035 U (1976) Voloshin, A. and Arean, I.. 'Failure of glass-epoxy lamina --fractographic study"J Composite Mater 13 (1979) p 240 Morris, G.E. 'Determining fracture directions and fracture origins on failed graphite/epoxy surfaces' ASTM STP 696 (American Society for Testing and Materials. 1979) p 274 Miller, A.G. and Wingert, A.L. 'Fracture characterisation of commercial graphite/epoxy systems' ibid p 223 Faure, J.-P. and Ropars, M. 'Contribution a l'etude du compartement a la rupture des composites a matrice thermostable PSP a partier d'observations au microscope electronique a balayage' ONERA (1980) pp 1980-1989 Kline, R.A. and Chang, F.H. 'Composite failure surface analysis' J Composite Mater 14 (1980) p 315 Theocaris, P.S.'and Stassinakis, C.A. 'Crack propagation in fibrous composite materials studied by SEM' J Composite Mater 15 (1981) p 133 Purslow,D. "Some fundamental aspects of composites fractography" Composites 12 No 4 (October 1981) pp 241-247 Morris, G.E. and Hetter, C.M. 'Fractographic studies of graphite/epoxy fatigue specimens' ASTM STP 775 (American Society of Testing and Materials, 1982) p 27 Liechti, K.M., Masters, J.E., Ulman, D.A. and Lehman, M.W.
'SEM/TEM fractography of composite materials" AFWAL-TR-82--4085 (1982)
13 Purslow,D. 'The fractographic analysis of the failure of a carbon fibre composite wind tunnel model' RAE TR 83020 (Royal Aircraft Establishment, UK. 1983) 14 Richards-Frandsen, R. and Naerheim, Y. 'Fracture morphology of graphite/epoxy composites' J Composite Mater 17 (1983) p 105 15 Purslow,D. 'Fractographic analysis of failures in CFRP' AGARD CP 355 (1983) 16 Potter, R.T. and Purstow, D. 'The environmental degradation of notched CFRP in compression' Composites 14 No 3 (1983) pp 206-225; RAE TR 83029 (Royal Aircraft Establishment, UK. 1983) 17 Potter, R.T. 'The significance of defects and damage in composite structures' AGARD CP 335 (1983)
AUTHOR David Purslow is with the Procurement Executive, Ministry of Defence, Royal Aircraft Establishment, Materials and Structures Department, X32 Building, Farnborough, Hants GU14 6TD, UK.
COMPOSITES. JANUARY 1984
Keywords: composite materials; [-beams, fractography; optical microscopy; three-point bending; failure mechanisms.
The fractographic analysis of broken coupons and components made from carbon fibre-reinforced plastics (CFRP) is becoming an increasingly frequent and important facet of the development of successful composite aerospace structures. In the understanding of basic fracture processes and in many component failure investigations, the use of a scanning electron microscope (SEM) is indispensable. 1-16. However, extracting relevant samples, sometimes many from a large component, mounting, and coating with a conducting surface, is a time consuming occupation. Also, there are certain to be circumstances in which access to an SEM is limited or unobtainable. This paper therefore describes the successful failure analysis of an aerospace structural element using lowpower optical microscopy alone. It must be added, of course, that the ability to use this technique is based
on considerable experience in composites fractography using an SEM. n°'~3"~5~6 The component chosen to illustrate the efficacy of low-power fractography is part of a research programme to investigate the relationship between defects and performance in structural components? 7 in this case a typical I-beam loaded in three-point bending.
SPECIMENDESIGNAND TESTING The I-beam is illustrated in Fig. 1. Two back-to-back channel sections of -t- 45° CFRP formed the web which was strengthened in the loading areas Z by tapered 90 ° + 45 ° reinforcement pads on both sides (pad layup: [(04 °. + 45 °. - 45°)3]0. The web and reinforcement pads were co-cured: pre-cured 0° + 45 ° tapered caps were then bonded to the web flanges. Details of materials and fibre lay-up of the caps are given in Fig. 2.
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+
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....... ............ ......
B2
Schematic of failed I-beam
O010-4361/84/010043-06 $03.00 © 1984 Butterworth 8- Co (Publishers) Ltd COMPOSITES. VOLUME 15. NO 1. JANUARY 1984
43
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Fig. 2
3 0 0 mm
Lamination of tapered caps: material XAS-914
A series of beams with various types and positions of artificial defects were tested in three-point bending both statically and in fatigueJ 7 The specimen described in this paper was manufactured as a 'non-defective' control and was loaded to failure statically.
D2
rc1
GENERAL DESCRIPTION OF FAILURE The broken specimen is shown in Figs 1, 3 and 4. (The lettering is consistent between all of the figures). The compression cap delaminated and split in several places between A and B and broke at positions C, D and E. The undelaminated cap plies and web flanges failed at E and one central reinforcement pad broke away from the web. The tensile flanges and cap remained intact and the web was therefore cut along xx to allow the fractured segments to be separated and to reveal the complete web failure EFGH.
METHOD OF FAILURE ANALYSIS
Fig. 3
In the fractographic analysis which follows, each mode of failure which occurred in the beam will be described and illustrated by a typical area taken from the specimen, together with an outline of other areas which suffered the same mode of failure. From this evidence, a diagram of the relative movements of the fractured segments of the structure during failure will be drawn and the detailed failure sequence of the complete beam will be determined, thus revealing the site and mode of critical fracture and cause of ultimate failure.
Failed segments of I-beam E
Fig. 4
44
Failure of web
Interlaminar shear Delamination of a multidirectional CFRP laminate has been shown ~° to occur preferentially at the boundary of plies of different orientation, one of which approximates to the direction of shear. Most of the delamination occurring in the I-beam took place in the compression cap and thus we may expect to find most interlaminar failures between the 0 ° and 45 ° plies. For example, the delamination between plies 23 and 24 (+ 45°/0 °, Figs 1 and 2) is shown in Fig. 5. However, the delamination between plies 27 and 28 (+ 450/- 45°), illustrated in Fig. 6, is a typical exception as will be seen later. The most significant features in a delamination are those in the matrix. The presence of misaligned fibres such as S (Fig. 5) and T (Fig. 6) causes the formation of resin rich areas and hence larger fracture features in the matrix. Also, by comparing the orientation of matrix features adjacent to misaligned
COMPOSITES. JANUARY 1984
ultimate failure, the crack spacing decreasing with increasing 0 ° stress - - a similar p h e n o m e n o n occurs in tension. For a given laminate the crack spacing may be used to provide an indication of the stress sustained. Where the cracked 45 ° plies subsequently delaminate, they frequently leave residual fibres (Fig. 8) indicating the crack spacing which was present at the time of delamination. Hence an estimate of the stress at which delamination occurred may be made. In this case, the crack spacing indicates that the delamination occurred at a 0 ° compressive stress approaching the strength of the material, after the in-plane 0 ° shear failure described later. Fig. 5
Fig. 6
Interlaminar shear failure, 0*/45* (x 20)
Interlaminar shear failure, + 4 5 0 / - 4 5 * (x 20)
Using these features and plotting the shear directions over the whole area of delaminations, it is possible to deduce the following. •
The cap delaminated between plies 15 and 16 before the web, flange and remaining cap failed at E since, apart from minor post-failure bruising, there is no evidence of the failure at E on the surface of ply 16 (see Figs 1 and 2).
•
Delamination between plies 15 and 16 commenced around C2, radiated to Bz, BI, C1 and towards PI P2 where the delamination changed across the pre-cracked + 45 ° ply (15) to plies 14/15, continued to D~D2 and subsequently to AtA2.
•
Delamination between plies 23 and 24 commenced at a split along the 0° plies MN, spread initially at + 45 °, propagated to C1 Cz and then extended to cover the area D~ D2 K M R C2 Cv (Note that ply 24 stops 120 m m from the cap centre. Delamination is then between the taper and ply 33.) Some delamination over the area C~ R L Q also occurred between plies 19/20/21.
fibres with the general features, the effect of ply orientation may be assessed. Looking in detail at Fig. 5, the misaligned fibre itself has been removed during the failure process but the fractured resin, particularly to the left of the fibre, S, shows clearly the cusps formed during a shear failure. ~a (Cusps are sometimes referred to as hackles? ,8.~2) These cusps may be observed over the whole area of both surfaces and are characteristic of a shear failure in which the mating surface moved in the direction arrowed. It is important to note that in assessing the overall failure directions, m a n y sites must be inspected since local perturbations may easily give rise to erroneous conclusions if they are based on only one or two observations. Comparing Figs 5 and 6, important differences may be observed in addition to the fibre orientations. Apart from the matrix associated with the misaligned fibres S and T, the general texture of the failure surface in Fig. 5 is much smoother than that shown in Fig. 6. Also, although it is not possible to illustrate, the colouring of the surface in Fig. 5 is amber whereas that in Fig. 6 is white. The amber colouration is that of the matrix and the surface of Fig. 5 is the result of the thicker, less finely textured matrix. These differences indicate that in the case of the 00/45 ° failure (Fig. 5), either the interlaminar bond was relatively weak or there was an interlaminar tensile component to the failing stress. Where post-failure movement has caused abrasion of the mating surfaces, the fractures appear dull, comparatively featureless and covered with debris (Figs 7 and 8). Both figures show part of the 00/45 ° shear failure between plies 15 and 16 (Figs 1 and 2). In Fig. 7 it is possible to detect cracks in the 45 ° plies at about 0.5 m m spacing. These cracks are due to the 0 ° compressive stress and start to form well before
COMPOSITES. JANUARY 1984
Fig. 7
45 ° cracking in interlaminar shear failure (x 7)
Fig. 8
45 ° crack remnants in interlaminar shear failure (x 7)
45
Shear
J peel
u
0~.~ 'ly .~8
_~7
F4
C1 Fig. 9
•
Comparison of shear and peel surfaces (x 0.3)
Delamination between plies 28 and 29 was a 0 ° interlaminar shear failure following the split along LR.
•
The 45 ° cracking occurred only over the area C1 R L Q. This cracking and subsequent delaminations occurred after the split MN which caused an increase in the cap compressive stress on the C~ D~ side of the web.
•
The delaminations between plies 15/16 and 23/24 showed evidence of either the influence of interlaminar tensile stresses or low interlaminar strength.
Interlaminar shear failure also occurred between the web and central reinforcement pads and was such that close to the line EF (Fig. 1), the left half of the web moved in the direction E ---' F relative to the pads.
Peel The delamination of a multidirectional CFRP laminate due to peel forces normally takes place within a ply orientated close to the direction of peel and near to the boundary of that ply with one of different orientation? 3 The failure surface consists largely of fibres (or fibre imprints in the matrix), may contain m a n y loose fibres and suffers little abrasion. It is therefore highly reflective as seen in Fig. 9, where the interlaminar shear failure between plies 23 and 24 may be compared with the subsequent peeling. Also clearly visible are the imprints of the serrated terminations of plies 3 to 16. The only peel failure to occur during breakage of the I-beam took place from C2 towards D2 in ply 24. The remaining peel failures are those produced by separation of the broken segments during analysis (including Fig. 9), their distinctive appearance facilitating the identification of such artefacts.
In-plane shear Failure due to in-plane shear stress takes place preferentially in a plane parallel to a fibre direction, l° Fig. 10 shows the failure containing plies 24 to 28 along MN (Figs 1 and 2). Considering first the 0 ° fibres (plies 24-27), the broken fibres, u, failed in compression indicating that the mating surface moved in the direction arrowed. In a 0 ° + 45 ° laminate, adjacent 45 ° layers subjected to stresses due to 0 ° in-plane shear prefer to fail along the 0° line in the layer in which the resolved 45 ° component is compression rather than tension? 6 The compression failure of ply 28 ( + 45 °) is also in the direction arrowed (Fig. 10); it is thus consistent with the generalization above and shows that the 0 ° failure preceded that of the 45 ° ply. Fig. 11 illustrates part of the mating surface to that shown in Fig. 10 in which several features may be noted. There is no delamination between the -I- 45 ° and - 45 ° plies (28 and 29) where they meet, thus their delamination took place after the 0 ° shear failure. The breakage steps
46
Fig. 10
In-plane shear failure (x 20)
Fig. 1 1 + 4 5 ° compression along LR (x 20)
in the + 45 ° ply failure (Fig. 11) are not perpendicular to the laminate plane. Since it has already been shown that the failure started in the 0 ° layers, the angle suggests that the in-plane fracture progressed from right to left (R-L, Figs 1 and 2), ie also in the direction arrowed. Hence the failure is likely to have originated at the right hand termination of ply 28 (R, Figs 1 and 2). The peel failure described earlier (plies 23/24) took place across the width dictated by the 0° shear crack and hence the in-plane failure (plies 24-27) must have occurred before the delamination (plies 23/24). It is important to note also that the shear failure along LR closely coincides with the length and thickness of the central web reinforcement. Failure of the web along EF (Fig. 1) was by in-plane shear, the left hand segment moving from E to F relative to the right hand segment.
Compression Fig. 12 illustrates part of the cap 0 ° compression failure C1 C2 (Fig. 1) near C2. The upper half of the picture shows the peeled surface which, it should be noted, contains evidence of the compression failure. It may thus be deduced that the peel failure occurred at the time of, or after the compression failure. In comparison, Fig. 13 shows the 0 ° compression fracture close to C1 and it is clear that, apart from some post-failure bruising, the delaminated surface contains no evidence of the compression failure. Hence, in this area, the delamination occurred before the compressive fracture. Fig. 14 illustrates part of the mating surface to that shown in Fig. 13. It can be seen that the 0 ° plies have, in fact, failed in two distinct modes - - c o m p r e s s i o n of the outer plies and tension on the innermost? ° Hence the 0 ° plies in this area failed due to combined compression and flexure: this confirms that delamination must have occurred prior to the compressive fracture and allowed buckling to take place. Similarly, it may be deduced that the cap failure at D~ was a combined compression/flexural failure and took place at the time of the adjacent delaminations between plies 23/24 and 14/15/16. In the region of D2, only partial compression failure occurred since plies 17 and 18 terminated at D 2.
COMPOSITES. JANUARY 1 9 8 4
Fig. 12
0 ° compression failure before delamination (x 1 5)
Fig. 1 3
0 ° compression failure after delamination (x 1 5)
Fig. 15
Web compression failure (x 8)
Fig. 16
Compression/bond failure of web (x 3)
Compression
Tension Fig. 14
0 ° compression/flexure failure (x 1 5)
Part of the _+ 45 ° web failure between F and G is shown in Fig. 15. (The plane surface to the right of the picture is the cut along x-x). This failure is due to compression perpendicular to the + 45 ° axis FG and illustrates the problem of abrasion between the failed surfaces which so often follows compressive failure. The angle of failure seen on the cut surface is typical of a pure compressive failure and also shows that no delamination of the web occurred in this area. Similar compressive failures took place along EH and H G due to in-plane forces approximately perpendicular to those lines. Compressive failure of the
remainder of the cap and web flanges occurred at E.
Cap~flange bond Part of the failure of the cap and web at E (Fig. 1) took place at the bond between the web flanges and cap as shown in Fig. 16. Normally failure at a good bond happens just within the CFRP as at V. ~3 The white honeycomb pattern in the lower flange is produced by the adhesive and carrier, very little area being bonded. However, the portions of the cap compression failure, W, remaining attached to the flange indicate that the cap failure took place before that of the bond despite its poor quality.
~
-.m---
+
1. J
Fig. 1 7
Movement diagram
COMPOSITES. JANUARY 1 9 8 4
47
MOVEMENT DIAGRAM AND FAILURE SEQUENCE The relative movements of the various segments of the fractured I-beam as deduced above are shown in Fig. 17. F r o m this diagram and the preceding analysisl the following failure sequence may be obtained. Ultimate failure originated in the compression cap. A 0 ° in-plane shear failure occurred in plies 24-28 along MN, originating at R (the termination of the 45 ° plies 28 and 29) and causing interlaminar shear failure between plies 28 and 29 over the area LROQ. Delamination between plies 23 and 24, also originating along MN, then occurred and propagated towards C allowing the outer plies to buckle. When this delamination reached C, where 0 ° plies 25 and 26 terminated, compressive/flexural failure of plies 23, 33 and 34 commenced at Ct and propagated to the 0° split RN, prior delamination having occurred between C~ and N. Some delamination of plies 19/20/21 also took place over a similar area. Delamination between plies 15 and 16 c o m m e n c e d at C~ C2 at this stage and spread towards D, together with an extension of the delamination between plies 23 and 24. Plies 15-23 buckled and failed in compression/flexure when the dela- • minations on both sides reached D~. Simultaneously the compressive/flexural fracture at C continued to C2, followed by a peel failure in ply 24 which extended, together with the 14/15 delamination, to the termination of plies 17 and 18 at D2 due to the lower stress and greater stiffness in this less damaged cross-section between M N and side 2. The buckling described above accounts for the indications of interlaminar tensile forces mentioned earlier, although other work has shown the interlaminar strength to be low. At this stage less than half the thickness of the compression cap remained intact over the central area of the beam, the area of m a x i m u m 0 ° load. The remaining cap and web flanges thus failed at E, the boundary of the central reinforcements where the poor bond was also located. The intact tension cap then bent at G causing the web to finally fail as shown in Fig. 17.
CONCLUDING REMARKS The analysis has illustrated the techniques of fracture characterization and failure sequencing using lowpower optical microscopy. Using these techniques, the initial cause of ultimate failure of a typical aerospace structural element has been successfully determined. Critical failure of the I-beam occurred in the compression cap due to an in-plane shear stress concentration. This stress concentration was located at the termination of the ___ 45 ° plies which coincide closely with the boundaries of one of the central reinforcements.
48
ACKNOWLEDGEMENT © Copyright Controller, HMSO, London, 1983
REFERENCES l
2 3 4
5 6 7
8 9
10 11 12
Auvinet, J. and Rouchon, J. 'The possibilities of using a
scanning electron microscope for the study of composite materials having an organic matrix' RAE Library Translation 1874 (1975) Adams,D.F. 'A scanning electron microscope study of hybrid composite impact response"JMater Sci 10 (1975) p 1591 't Hart, W.G.J. 'Scanning electron microscopy of fracture surfaces of carbon composite materials" NLR TR 76035 U (1976) Voloshin, A. and Arean, I.. 'Failure of glass-epoxy lamina --fractographic study"J Composite Mater 13 (1979) p 240 Morris, G.E. 'Determining fracture directions and fracture origins on failed graphite/epoxy surfaces' ASTM STP 696 (American Society for Testing and Materials. 1979) p 274 Miller, A.G. and Wingert, A.L. 'Fracture characterisation of commercial graphite/epoxy systems' ibid p 223 Faure, J.-P. and Ropars, M. 'Contribution a l'etude du compartement a la rupture des composites a matrice thermostable PSP a partier d'observations au microscope electronique a balayage' ONERA (1980) pp 1980-1989 Kline, R.A. and Chang, F.H. 'Composite failure surface analysis' J Composite Mater 14 (1980) p 315 Theocaris, P.S.'and Stassinakis, C.A. 'Crack propagation in fibrous composite materials studied by SEM' J Composite Mater 15 (1981) p 133 Purslow,D. "Some fundamental aspects of composites fractography" Composites 12 No 4 (October 1981) pp 241-247 Morris, G.E. and Hetter, C.M. 'Fractographic studies of graphite/epoxy fatigue specimens' ASTM STP 775 (American Society of Testing and Materials, 1982) p 27 Liechti, K.M., Masters, J.E., Ulman, D.A. and Lehman, M.W.
'SEM/TEM fractography of composite materials" AFWAL-TR-82--4085 (1982)
13 Purslow,D. 'The fractographic analysis of the failure of a carbon fibre composite wind tunnel model' RAE TR 83020 (Royal Aircraft Establishment, UK. 1983) 14 Richards-Frandsen, R. and Naerheim, Y. 'Fracture morphology of graphite/epoxy composites' J Composite Mater 17 (1983) p 105 15 Purslow,D. 'Fractographic analysis of failures in CFRP' AGARD CP 355 (1983) 16 Potter, R.T. and Purstow, D. 'The environmental degradation of notched CFRP in compression' Composites 14 No 3 (1983) pp 206-225; RAE TR 83029 (Royal Aircraft Establishment, UK. 1983) 17 Potter, R.T. 'The significance of defects and damage in composite structures' AGARD CP 335 (1983)
AUTHOR David Purslow is with the Procurement Executive, Ministry of Defence, Royal Aircraft Establishment, Materials and Structures Department, X32 Building, Farnborough, Hants GU14 6TD, UK.
COMPOSITES. JANUARY 1984