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Buckling behaviour and delamination growth in impacted composite specimens under fatigue load: an experimental study
Composites Science and Technology 61 (2001) 1841–1852 www.elsevier.com/locate/compscitech
Buckling behaviour and delamination growth in impacted composite specimens under fatigue load: an experimental study L. Gunnar Melin*, Joakim Scho¨n Swedish Defence Research Agency, SE-172 90 Stockholm, Sweden Received 5 January 2001; received in revised form 22 May 2001; accepted 7 June 2001
Abstract Impact-damaged carbon-fibre/epoxy composite laminates have been tested in fatigue under constant-amplitude loading with emphasis on studying mechanisms leading to delamination growth and fatigue failure. The shapes of the buckles have been measured with an optical whole-field measurement technique, while the extension and depth of delaminations was measured several times by ultrasonic C-scanning during testing. It was observed that the delamination growth occurs mainly in transverse direction to the load and that buckles on the backside usually have the same shape as some of the delaminations. This indicates that the buckling which takes place during the compressive part of the load cycle drives delamination growth. In specimens run to the fatigue limit, buckling and delamination growth occur only in the outer 2 or 3 layers on the backside while a specimen that failed before the fatigue limit buckled through the whole specimen. # 2001 Elsevier Science Ltd. All rights reserved. Keywords: B. Fatigue, B. Impact behaviour, C. Buckling, C. Delamination, D. Ultrasonics
1. Introduction Composite structures are increasingly being used in aircraft. During the aircraft operational life those structures will often suffer impact damage. It is important, therefore, to have reliable criteria for predicting the effects of such impact damage during fatigue. Before such criteria can be developed a good understanding of the damage mechanisms during fatigue loading is necessary. Several investigators have studied the way in which different parameters affect the fatigue life of impacted carbon-fibre/epoxy composites. Specimens made of T300/976 carbon/epoxy composites with different thicknesses have been fatigue loaded after impact [1]. It was found that the thicker specimens performed better in fatigue than the thinner ones. Comparing fatigue properties of undamaged and impact damaged composite laminates loaded in tension/tension loading at a stress min ratio, R=0.1, where R ¼ max ; it was observed that in low-cycle fatigue the fatigue life was slightly shorter for impact-damaged specimens than for undamaged specimens. However, in high-cycle fatigue the fatigue life was * Corresponding author. Tel.: +46-8-5550-4257; fax: +46-8258919.
significantly shorter for impact-damaged specimens than for undamaged specimens [2]. In a fatigue study [3] of impacted AS/3501-6 composites it was found that specimens loaded in compression/compression, R ¼ 1; have almost the same fatigue life as specimens loaded in tension-compression, R=1. Specimens loaded in tension/tension, R=0 had the longest fatigue life. Similar results were obtained in Ref. [4] where fatigue testing of 156 mm wide specimens of HTA/6376 composite was performed. Testing at two R values, R=1 and R=5, did not result in any significant difference in fatigue life. From this it could be concluded that the fatigue life is determined by the compressive part of the loading and not by the tensile part. Two different laminates were investigated, one quasi-isotropic (QI) and one dominated by fibres in the zero degree direction (ZD). When fatigue life is plotted versus stress, the latter laminate was found to have a higher fatigue resistance, although when plotted against strain instead the QI laminate was found to have the better performance. Damage growth during fatigue loading has been looked at in a large number of investigations. The conclusions sometimes seem to be contradictory, which might be due to differences in material, specimen design, impact energy and tested load cases. Several authors [5–10] have
0266-3538/01/$ - see front matter # 2001 Elsevier Science Ltd. All rights reserved. PII: S0266-3538(01)00085-9
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reported that most of the damage growth occurs in the final part of the fatigue life. The increasing damage can be monitored, for example by studying the decreasing stiffness [5], the increasing crack density [5] or the growth of the delamination envelope [7]. Gerharz et al. [8] identified two stages during the fatigue damage growth, ‘phase 1’ and ‘phase 2’. During phase 1 a small increase in buckling deformation and compliance was observed. No delamination growth out of the damage envelope was observed and the number of delaminations within the original damage envelope increased. The phase 2, which occurs during the last 10% of the fatigue life, is characterised by rapid damage growth. Buckling deformation and compliance increase and delaminations grow outside the original damage envelope. Delamination growth occurring during a significant part of the fatigue life has also been reported [11–13]. Mitrovic et al. [13] found for 38 mm wide specimens impacted with 2.1 J that the delamination growth rate decreases with delamination length. In Refs. [3,4,11,12], damages were reported to grow mainly in transverse direction to the load while in reference [13] it was found that most of the delamination growth occurred in the load direction. One way to study the damage growth is by using optical whole field measurement techniques such as shadow moire´ [11,12,14] or digital speckle photography [4] (DSP) to determine the shape of the buckles. Melin et al. [4] reported that the main buckles go inward on the impacted side and outwards on the backside of the laminate. Therefore all plies through the specimen have to buckle. The buckles on the backside had both larger size and amplitude. To the authors’ knowledge, one study using ultrasonic C-scanning to detect which delaminations through the specimen that grow during fatigue loading has been made [14]. It was found that most of the delamination growth occurred in the ply layers close to the back face of the specimens where the largest delaminations had been produced by the impact. They were not able to resolve the delaminations on a ply level. In a fractographic study [15] of damage growth during fatigue of impacted composites, it was found that delaminations at neighboring interfaces interact with each other. Ply cracks were found to establish boundaries for some delaminations. Delaminations were believed to form in fatigue at crossover points of intra-ply cracks. Before results from fatigue loading of impacted coupons can be used in design, impacted structures need to be fatigue loaded to study the transferability of coupon data to structures. Two sections of a horizontal stabiliser made of CF/PEEK (APC-2) with impact damage have been tested in fatigue. No damage growth was observed and the structures failed in quasi-static loading after the fatigue loading had been finished [16]. During certification of C-130 Hercules composite flaps barely visible impact damage, BVID, and visible impact damage,
VID, were introduced to a component. During fatigue testing, the VID did not grow while the BVID grew from a diameter of 11 mm to 19 mm [17]. Poe [18] has discussed if fatigue of impacted aeronautical composite structures are important. He argues that constant amplitude compression fatigue tests suggest that fatigue lives for nonvisible impact damage are adequate but those for detectable damage may need to be verified by spectrum tests. Although a large number of investigations have studied damage growth during fatigue after impact no investigation has studied delamination growth on a ply level and the buckling growth has also not been investigated in detail. The objective of this investigation is to study delamination growth on a ply level with ultrasonic C-scan together with detailed measurements of the buckling shape and growth using DSP. The delamination growth and hence the damage growth is expected to be related to the fatigue life of the specimens in compressive load.
2. Experimental 2.1. Specimen The specimens were prepared from Hexcel HTA/6376 carbon fibre/epoxy matrix prepreg with a nominal fibre volume fraction of 65%. The specimens had a size of 450mm156 mm. Two different lay-ups have been used, one dominated by fibres in 0 direction, ZD, and one quasi-isotropic, QI. The ZD lay-up has the sequence [45,135,0,90,45,135,02,(45,135,0, 90)2,45,135,02,45,135,0, 90]s, giving totally 48 plies with 16 in 0 , 8 in 90 and 12 each in 45 and 135 directions. The QI lay-up has the sequence [(90,135,45,0)s,(0,45,135,90)s]3, totally 48 plies with 12 in each of the 0 , 90 , 45 and 135 directions. Loading is applied along the 0 direction. The QI lay-up results in a specimen that is quasi-isotropic both in the plane and in bending [19]. The nominal specimen ply thickness is 0.13 mm; hence, the cured specimens were 6.24 mm thick. 2.2. Impacting The specimens have been impacted with energies of 27.6 0.4 J. Most of the impact threats will have impact energy less than 28 J [20]. The impacts were conducted in a conventional drop-weight rig using a drop height of 0.5 m. The velocity just before impact was measured with a photo transistor device. The impactor had a tip radius of 7.5 mm. During impact the specimens were clamped in gripping zones of the specimens short sides, along the long sides the specimens were supported. The impact causes on the frontside a permanent indentation 0.05– 0.1 mm deep and about 15 mm in diameter; similarly on
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Fig. 1. Loading arrangements including CCD cameras for DSP measurement. The solid specimen-boundary indicates the permanent indentation and the dashed lines outlines the buckling shape.
the underside a buckle 0.15–0.2 mm high and up to 30 mm in diameter is created. During certification of C-130 Hercules composite flaps, impact damage was considered barely visible when the depth of the impact size was a minimum of 0.25 mm [17]. Therefore, the impact damage in this program should be considered non-visible. 2.3. Fatigue testing Fig. 1 shows the specimen during compressive loading. The fatigue testing has been conducted in a servohydraulic MTS loading machine with a maximum load of 1000 kN. An anti-buckling device was used to prevent global buckling in compression. The anti-buckling device was made of steel and had a window of 100mm100 mm on each side of the specimen centred on the impact point. To minimise friction and allow movement between the
anti-buckling device and the specimen, a layer of a low friction polymer separated the specimen and the steel. In Fig. 1 is also indicated a permanent indentation/buckle caused by the impact (solid lines) and a typical out-ofplane deformation (dotted lines). The testing was done in constant amplitude loading using either of two R values: R=1 and 5. A frequency of 0.5 Hz was used, which resulted in negligible temperature increase in the damaged area. The frequency also gave time for buckling of the damaged region. All tests were performed at room temperature and in load-controlled mode. In Table 1, the number of specimens tested with the two whole field measurement techniques are tabulated. By ultrasonic C-scan using the system Krautkramer USPC 2100, the areal shape and depth of the delaminations are measured through the thickness. Numerical values of the delamination depth are available from the C-scan system. At each measurement point not more than one delamination, the one closest to the surface, can be detected. Consequently, delaminations deeper down in the laminate are hidden. From the C-scan maps it could be reconstructed between which plies each delamination occurred. All specimens were C-scanned after impact, before starting the fatigue test. To perform a C-scan during the test, the loading machine had to be stopped and the specimen dismounted; for this reason a limited amount of C-scans were made during the fatigue tests. Those tests are outlined in Table 2. The successive delamination growth was analysed in the image-editing program Corel PHOTO-PAINT 7. With the optical non-contact measurement technique digital speckle photography (DSP) [21–23] the displacement field was measured in the window of the antibuckling devise. The method is also known as Digital Image Correlation, Electronic Speckle Photography and the Grating Method. The measurements were made with the system Aramis from the German company GOM mbH. In DSP, a random pattern is applied on the specimen surface, here dotted patterns were created using white matt spray paint. Images are acquired by CCD cameras of the undeformed and deformed states and evaluated numerically to determine the displacement field. By using two cameras imaging the specimen surface from different angles, all three displacement components can be determined. Here, a total of four cameras were available, making it possible to perform DSP measurements on both sides of the specimens as
Table 1 Number of specimens tested in constant amplitude fatigue Lay-up
Total number of specimens
Specimens with ultrasonic C-scan during test
Specimens with DSP during test
Specimens with DSP and C-scan during test
ZD QI
11 15
3 5
4 6
2 3
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Table 2 Specimens measured with ultrasonic C-scanning and DSP during fatigue test Specimen
Minimum stress (MPa)
R value
Life (number of cycles)
C-scan after N cycles
ZD14 ZD15 QI5 QI13 QI14
206 247 185 185 226
5 5 1 5 1
117 140 120 75 973 Stop at 1 020 000 230
1000, 7000, 44 000, 92 000, 115 000, 117 000 51, 114 70 000, 73 600 1 020 000 100, 200
shown in Fig. 1. Each measurement area covered the window area of the anti-buckling device. The images were acquired after different intervals of cycling. During image acquisition the specimens were kept under constant load giving a hold time of about 10 s at the maximum load. Saunders and van Blaricum [14] reported that hold times of 10 s do not have any discernible effect on the fatigue life. All DSP registrations were made at the minimum compressive load except for one specimen for which registrations were made during complete load cycles. That enabled one to see during which part of the load-cycle buckling occur and if there is any hysteresis.
3. Results A selection of the experimental data is presented in this paragraph. The focus is on the specimens with both C-scan and DSP measurement (see Table 2), where the areal shape of delaminations and buckles can be compared. The results from the specimen on which complete load cycles were measured are also presented. The fatigue results plotted against compressive stress for constant amplitude loading of the two lay-ups can be seen in Fig. 2. The stress has been calculated using the total width, thickness and elastic constants of the
Fig. 2. Minimum stress vs number of cycles for all specimens tested in constant amplitude loading.
undamaged specimens. The fatigue results are taken from Ref. [4]. 3.1. Delamination growth analysis 3.1.1. Definitions Four specimens are presented. Two had typical buckling shapes for a long and a short fatigue life: ZD14 and QI14 with long and short fatigue life, respectively. Of the other two specimens, QI13 was run up to the fatigue limit and QI5, which had a rather long fatigue life, exhibited a particular buckling shape. C-scan and DSP measurements have been made from both sides of the specimen. The result maps are presented with the same absolute specimen orientation using the co-ordinate system defined in Fig. 1. In Fig. 3 the laminate directions 0–360 in the laminate plane are defined together with a co-ordinate system with the same orientation as in the result maps. The 90 and 270 directions will be referred to as sideward directions. The out-of-plane DSP measurements of displacement are presented as iso-contour maps superimposed on Cscan maps. A 0.1 mm contour interval is used. The displacement is for the minimum load during each cycle relative zero load before the first cycle. The zero-displacement contours are labelled with a ‘0’, further, the buckles larger than 0.1 mm are labelled with a ‘+’ or ‘’ depending on if they buckles in positive or negative z-direction. For example, the buckle in Fig. 1 would be labelled with a ‘’ on both the frontside and the backside. The displacement measurements were made during the last loading before performing the C-scan except at
Fig. 3. Definitions: laminate directions 0–360 relative co-ordinate system.
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Fig. 4(b). Results for ZD14. Extension and growth of delaminations in sideward directions.
Fig. 4(a). Results for ZD14. DSP measurements of z-displacement superposed to C-scan maps. The displacement is plotted as iso-contour with 0.1 mm spacing. Left column: frontside; right column: backside.
the first cycle when the C-scanning was made first. The positioning of the contour maps relative the C-scan images has an uncertainty of 1–2 mm. When the C-scan map is missing on the result maps, it is because no Cscan was made at that occasion. Observe that the Cscan maps are sometimes a bit noisy. The colour range of the C-scan maps is shown versus distance from the front surface in Fig. 4(a). The C-scan equipment is not able to resolve delaminations in the outer 3–5 plies, these delaminations are seen as a single delamination on the C-scan maps. The C-scan analysis shows that the delaminations created during impact with few exceptions follow the fibre direction of the adjacent ply closest to the backside of the specimen. The initial distribution and the growth of the delaminations are outlined in diagrams, where the delamination shape has been projected onto two directions: 90 and 270 . These directions were the dominating growth directions. Of all observed growing delaminations only 10–20% grew along one of the 45 or 135 directions and almost no growth occurred along the 0 or 180 directions. In other words, these diagrams show the largest extension along the x-axis of the visible delaminations at different fatigue lives. In a few cases, where the largest extension of the delaminations was not at the place for the delamination growth, then the diagram describes the part of the delamination that grew. Observe that
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because only the outer delamination can be detected at each measurement point, it is probable that other delaminations are hidden. The growth of these hidden delaminations could not be measured. Some delaminations are not visible in the C-scan maps until after some load cycle. In the diagrams, it is indicated as default that these delaminations have started to grow from a width of zero though it is not known from where the growth started. Observe also that the delaminations appear at interfaces between two plies and, contrary to what the delamination diagrams indicate, has no extension in the thickness direction. The zero displacement level in the DSP contour maps has been set by subtracting the average of measurement points close to the edge of the measurement area. In compressive load, besides the buckling around the damage area, usually an out-of-plane slope was observed. This gave a difference of up to 1.0 mm in out-of-plane displacement over the buckling window. This out-of-plane rotation is caused by imperfections in the loading machine. To emphasise the local buckling at the damage area, the rotation has been eliminated from the displacement maps by a co-ordinate transformation that adjusts the displacement field to a plane. It is not expected that the slope had any significant impact on the buckling behaviour and delamination growth. 3.1.2. ZD14 (Fig. 4) This specimen was tested at a load close to the fatigue limit at a minimum stress of 206 MPa and an R value of 5 (Fig. 2). From the start the buckles on both sides extend about half the diameter of the delamination envelope and have relatively small amplitudes. (Observe that since the outer contours of the buckles are at 0.1 mm displacement, the buckles extend some distance outside that contour.) Not until after 92 000 cycles does the inward frontside buckle begin to grow in area. The outward buckle on the backside has a rounded shape. This buckle has a slow, steady growth in both amplitude and area up to 92 000 cycles. Regarding the delaminations, minor growth is observed on the outer plies on both sides for the first 1000 cycles, after that no growth is observed until after 44 000 cycles where again some delaminations near the surfaces have grown in area. The growth near the frontside correlates to the increasing width of the main inward buckle and the development of an outward buckle above it. No growth occurs near the envelope. Closer to failure, the specimen broke after 117 140 cycles; the trend is that a successively larger number of delaminations grow at positions deeper down from the surfaces. The delaminations that have started to grow also continue to grow. Between 92 000 and 115 000 cycles, all delamination growth takes place at positions about 10 mm inside the envelope. On the backside the sideward increase of the buckle area correlates to the
delamination growth at the interfaces between ply 4-5 and 5-6. Between 115 000 and 117 000 cycles, large delamination growth is observed at many interfaces. The envelope is extended by three delaminations, both in the middle of the laminate and closer to the backside. Many delaminations now have almost the same extension in sideward direction. The backside buckle has grown significantly sidewards. Its areal shape follows mainly the shape of the delaminations from ply 6-7 and out. On the frontside the inward buckle gets a narrow shape between two minor outward buckles. 3.1.3. QI5 (Fig. 5) This test was performed at a relatively low load with a minimum stress of 185 MPa and an R value of 1 (Fig. 2). Its specific buckling behaviour was not observed in any other specimen. In the first cycle, only a small inward buckle appear on the frontside. The buckle on the backside has its highest amplitude over a delamination in the outer plies beside the impact point. After 70 000 cycles, there are both an inward and below (in negative y-direction) an outward buckle of similar size and areal shape at the frontside. The largest delamination growth has occurred at the transition between the two buckles. After 73 600 cycles, the outward buckle on the frontside has increased both in area and amplitude. On its upper side, there is a sharp transition to the inward buckle and on the other sides its edge is similar to the delaminations in the six outer plies. On the backside the buckle has grown sidewards, reaching the envelope. Its maximum amplitude is 5 mm in the negative y-direction from the impact point. Until 73 600 cycles when the last C-scanning was made on both sides, almost all delamination growth occurred in the outer five interfaces. Some delaminations near the frontside at the position of the outward buckle grew partly in the 180 direction. The DSP measurement just before failure at 75 900 cycles, the specimen failed after 75 973 cycles, shows that the outward buckles on both sides have grown along the 90 direction; the place for the growth coincided though the backside buckle has a larger extension. The inward buckle on the frontside has got its amplitude decreased and can be described as being pressed upwards by the dominating outward buckle. 3.1.4. QI13 (Fig. 6) The test was run up to the fatigue limit at a minimum stress of 185 MPa and an R value of 5 (Fig. 2); under the same loading conditions, specimen QI9 failed after 21 115 cycles. Specimen QI5 was loaded at the same minimum load but with another R value. No delamination diagram is shown; the only delamination growth observed was near the backside at the two outer interfaces. This
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Fig. 5(b). Results for QI5. Extension and growth of delaminations in sideward directions.
point. The largest buckling appears in a 5 mm wide stripe in the outer ply of 0 direction, this stripe could be observed visually during testing.
Fig. 5(a). Results for QI5. DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside, right column: backside.
growth, which was up to 15 mm in length, was directed away from the impact point and the delaminations were beside the impact point. From the start, no detectable buckle emerge on the frontside. After 1020 000 cycles a small inward buckle is indicated around the impact point. On the backside, an outward buckle appears during the first cycle with maximum amplitude nearly 10 mm away from the impact point. The buckle has the same areal shape as the outer delaminations. At the fatigue limit, the buckle has increased both in amplitude and size. Its main part is still beside the impact point and the enlargement has been directed away from the impact
3.1.5. QI14 (Fig. 7) A high fatigue load of 226 MPa and an R value of 1 (Fig. 2) was applied, causing relatively large buckles on both sides from start. The backside buckle has similar sideward extension as the delamination envelope. Delamination growth is seen after 100 cycles at the interfaces close to the surfaces. Nearly all delamination growth is in sideward directions, but not only at delaminations with a 90 ply adjacent towards the backside. The envelope is enlarged at one delamination. The backside buckle grows in 270 direction together with the extended envelope. On the frontside, the inward buckle has not increased; a small outward buckle has appeared above. After 200 cycles, delaminations through the whole specimen have grown, all within the extension of the backside buckle. The backside buckle has continued to grow sidewards while a small inward buckle is detected above the main buckle. On the frontside, the main buckle is almost unchanged and the delamination growth near the frontside has partly appeared outside this buckle. The specimen failed after 230 cycles.
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Fig. 6. Results for QI13: DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside; right column: backside.
3.2. Fatigue cycle monitoring On one specimen, QI53, some specific cycles were monitored by making totally 10 DSP registrations per cycle. The specimen was loaded at 226 MPa and R=1 which is the same load as for specimen QI14. Specimen QI53 had, for this load, a rather long fatigue life of 2356 cycles. Four specific cycles are presented: the first cycle, after 300, 900 and 2200 cycles. The C-scan maps before loading are shown in Fig. 8(a) together with the buckling shapes at the minimum compressive load. The damage zone is slightly smaller than for the other presented QI specimens. The z-displacement during the cycles is shown in Fig. 8(b–d). Finally, the strain histories along the loading direction are presented in Fig. 8(e–f). The strain has been taken as the average strain over an 83 mm long section where the damage zone is included. The buckling shapes in Fig. 8(a) have a similar appearance to the previously discussed specimens. On the backside, the highest buckling occur outside the impact point and for this reason the buckling amplitude is presented both at the impact point in Fig. 8(c) and at the position with highest buckling amplitude in Fig. 8(d). During the first cycle, the buckling is, on both sides, clearly smaller during the compressive loading phase than during the unloading phase. During the fatigue life the load/buckling relationships get increasingly linear though some hysteresis effect remain on the backside. The small displacement registered during tensile loading is probably caused by the permanent buckle/indentation getting straightened out. The strain plots show a linear behaviour in tension and a slightly lower stiffness during
compression. There is almost no hysteresis and a small weakening occur in compression during the fatigue life.
4. Discussion In this section, the main features of the results and what can be deduced from them are discussed. The conclusions regarding the buckling behaviour are also supported by the other DSP-measured specimens not presented here. Two of those were presented in Ref. [4]. 4.1. Delamination growth and correlation to buckling The initial delamination shape follow with few exceptions the direction of the adjacent ply closest to the backside of the specimen, a ply that also have big influence on the delamination growth direction during fatigue while the ply closest to the frontside has almost no influence on the growth direction. In the outer interfaces during fatigue loading, delaminations can grow in any direction following a single stripe but deeper down in the specimen delaminations grow either in the direction of the adjacent ply towards the backside or in the direction transverse to the load. The dominating delamination growth direction is transverse to the load, hence most delaminations that grow have a 90 ply adjacent towards the backside though some delaminations adjacent to 45 or 135 plies also grow. Only in the outer layers is growth observed at delaminations with a 0 ply adjacent towards the backside. The longest growth takes place in the outer delaminations though a trend with increasing number of cycles
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Fig. 7(b). Results for QI14. Extension and growth of delaminations in sideward directions.
Fig. 7(a). Results for QI14. DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside, right column: backside.
is that delaminations deeper into the laminate start to grow. A delamination that has started to grow usually continues to grow. For short fatigue lives, the growth rate is approximately constant whereas for longer lives most growth occurs just before failure. The delamination envelope is usually not extended until near failure, then delaminations between the middle of the laminates and the backside grow out of the envelope. Comparing DSP measurements with the C-scan maps, it is seen that the outward buckle on the backside usually has a similar areal shape as some delaminations, contrary to the inward frontside buckle. The agreement in areal shape on the backside can from start be with the outer delaminations and later in fatigue life to delaminations deeper down from the surface; an example of this is specimen QI5. For higher loads, though, also from start the sideward extension might be similar to the envelope as for specimen QI14. This good shape agreement gives a strong support of the idea that the delamination growth is driven by buckling. A possible mechanism controlling the delamination
growth could then be as outlined in Fig. 9. It shows the buckling in a cross-section transverse to the load around the damage zone, both at the beginning and at the end of the fatigue life. Since all plies and hence all delaminations buckle in the same direction and they are buckling into the specimen on the frontside, all plies in the thickness direction have to buckle at least as much in area and amplitude as is observed on the frontside surface since otherwise the plies would have to penetrate each other, which is impossible. This discussion can be applied to each delamination starting from the frontside. Therefore, the buckle on the backside has to have the same or larger amplitude and area as the frontside buckle. This was also observed in the experiments. As a delamination buckles, all three fracture modes will be introduced along the edge of the delamination. Since the critical energy-release rate for mode I, GI,C, is approximately 240 J/m2 and for mode II, GII,C, is about 800 J/m2 [24–27], it can be assumed that the delaminations grow mainly in mode I. The value for mode III is not available but can be assumed to be in the order of GII,C. Calculations of the mode separated energy release rate of delaminations in the post buckling state have found that GI has a similar or larger value than GII and GIII [28,29]. The ratio of GI to the total energy-release rate has a large influence on the delamination-growth rate in fatigue [30].
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Fig. 8(b)–(f). Results for QI53. (b) z-Displacement on frontside at the impact point during specific load cycles. (c) z-Displacement on backside at the impact point. (d) z-Displacement on backside at the point with maximum buckling amplitude. (e) Average strain in loading direction on frontside. (f) Average strain in loading direction on backside. Fig. 8(a). Results for QI53. (a) DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside; right column: backside.
In the first cycle in Fig. 9, buckling occur through the whole laminate but since the buckle is smaller than the delamination envelope, it is only at the outer delamination that a mode I component appear. The delaminations that do not buckle at the crack tip will be closed. As a consequence GI will be small for those delaminations and no growth will occur. As the outer delaminations grow and the buckling area gets larger at the same rate, the edge of the buckle will reach the delamination envelope. Hence, the delaminations deeper down in the specimen will also be loaded in mode I hence promoting delamination growth. This has also been observed in the experiments (Figs. 4 and 5). For specimen ZD14, delamination growth occurred deep inside the specimen, which gives a larger delamination width than is the case close to the backside. This would suggest that those delaminations buckle all the way to the crack tip. Smaller diameter buckles are then necessary to open the crack tips of the delaminations close to the backside. This means that several delaminations buckle separately and that the buckle
observed on the surface is the results of several buckles superimposed on each other. Support for this can be found from Fig. 4(a) at the backside after 117 000 cycles; the slope of the buckle is small over the delaminations 2–3 mm into the specimen whereas at the edge of the outer delaminations the slope of the buckle increases significantly. Close to failure, minor buckles usually appear above and under the main buckles. The minor buckles typically have an amplitude of 0.2 mm, buckle in opposite direction to the main buckles and do not cause any delamination growth. These buckles probably do not have any significant influence on the fatigue life. The specimen QI5 is an exception where an outward buckle on the frontside becomes dominant over the inward buckle. Its areal shape is similar to the delaminations within the outer five layers why at least these plies buckle. These delaminations grow in area as the outward buckle appears. It is not known whether this buckling behaviour had any effect on the fatigue life. No differences between the two R values 1 and 5 regarding the buckling behaviour and delamination growth could be distinguished. The corresponding conclusion from the measured fatigue life was presented by
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Fig. 9. Outline of proposed mechanism for buckling-induced delamination growth.
Melin et al. [4]. This suggests that the tensile part of the loading has a small influence on the delamination growth, whereas the compressive load is the harmful part. From Fig. 8(b–d), it can be concluded that buckling occurred only during compressive loading. The load/strain plots in Fig. 8(e–f) show that the specimens were less stiff during the compressive loading than in tension. This is reasonable since a damage zone can take less load as it buckles in compression than when it is stretched in tension. 4.2. Fatigue limit In most QI specimens, the delaminations nearest the backside was outside the impact point. From start at a low load usually the largest buckling occur there—see specimen QI5 or QI53. Specimen QI13, which was run up to the fatigue limit, had two particular buckling features compared to the failed specimens: the backside buckle was fully outside the impact point and almost no inward buckling occurred on the frontside. During testing, the only evolution observed was that the backside buckle grew in a direction away from the impact point. It is probable that only the outer two plies buckled— one indication of this is that all observed delamination growth occurred there. Also, in the other QI specimens run to the fatigue limit, the only observed delamination growth was in the outer 2–3 plies. This indicates that buckling in only the outer 2–3 plies is not sufficient to cause failure. Failure demands that buckling also occurs deeper down in the specimen. This suggests that a fatigue threshold exists; the specimen can only fail if the buckling goes at least 5–8 plies deep on the backside. For this, a minimum compressive
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load has to be exceeded during the load cycle, a threshold load that is individual for each impact damaged specimen. The tests indicate that when the threshold is exceeded, the buckling at the impact point has an amplitude of at least 0.15 mm on both sides. This could be one way to measure the fatigue threshold though the number of tests is too few to claim a rule. No DSP testing has been made on a ZD specimen run to fatigue limit, making it difficult to make statements regarding buckling behaviour at the fatigue limit. A difference from the QI specimens is that the delamination in the outer 45 ply on the backside covers the impact point, this ply is the first to buckle at low loads. This leads to different buckling shapes at low loads for the two laminates. To know whether a ZD specimen is expected to fail it is probably not enough to know if the impact point on the backside buckles, one also has to conclude from the buckle shape if only a stripe in the outer ply buckles or if plies deeper down buckles, giving it a more circular shape.
5. Summary and conclusions Impact-damaged carbon fibre/epoxy composite laminates have been tested in fatigue at constant amplitude loading. The objective has been to study the mechanisms that leads to delamination growth and in the end fatigue failure. To achieve this, the shapes of the buckles has been measured with an optical whole field measurement technique and the laminates have been studied by ultrasonic C-scan several times during testing. Both quasi-isotropic and zero-degree dominated laminates have been tested. The observations indicate that the fatigue life is controlled by delamination growth in the direction transverse to the load, though the final failure mechanism can be another one. More than half of the delaminations that grow are along plies in transverse direction. The delamination growth is driven by buckling that occurs during the compressive part of the load cycles; an indication of this is that the buckle on the backside usually had the same areal shape as some delamination. In specimens run to fatigue limit, buckling and delamination growth occur only in the outer 2–3 layers on the backside. For the quasiisotropic laminate, it is suggested that a threshold compressive load exists, a threshold which is individual for each specimen. At loads higher than the threshold, the specimen buckles through the whole thickness and fatigue failure is probable.
Acknowledgements The authors wish to express their thanks to the Swedish National Research Program (NFFP336+345) and to the Swedish Defence Material Administration (FMV).
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References [1] Tai NH, Ma CCM, Lin JM, Wu GY. Effects of thickness on the fatigue-behavior of quasi-isotropic carbon/epoxy composites before and after low energy impacts. Comp Sci Technol 1999; 59(11):1753–62. [2] Tai NH, Yip MC, Lin JL. Effects of low-energy impact on the fatigue behavior of carbon/epoxy composites. Comp Sci Technol 1998;58(1):1–8. [3] Ramkumar RL. Effect of low-velocity impact damage on the fatigue behavior of graphite/epoxy laminates. ASTM STP 813, 1983, 116–135. [4] Melin LG, Scho¨n J, Nyman T. Fatigue testing and buckling characteristics of impacted composite specimens. Int J Fatigue (in press). [5] Park SW, Cao C, Zhou M, Talreja R. Post-impact fatigue behavior of an IM7/PETI-5 cross-ply laminate. Fatigue 99, China, vol. 3. 1999, A99-43943, p. 1711–7. [6] Caprile C, Sala G, Buzzi A. Fatigue of low-energy impacted and moisture aged composites, ICAF 95, vol. 1, Melbourne, Australia, 1995. p. 365–83. [7] Symons DD, Davis G. Fatigue testing of impact-damaged T300/ 914 carbon-fibre-reinforced plastic. Comp Sci Tech 2000;60:379–89. [8] Gerharz JJ, Idelberger H, Huth H. Impact damage in fatigue loaded composite structures. In: Proceedings of the 15th Symposium of the International Committee on Aeronautical Fatigue, Israel, 1989. p. 497–516. [9] Beheshty MH, Harris B. Post-impact fatigue behaviour of CFRP and the growth of low-velocity impact damage during fatigue. In: Proc. Int. Conf. on Fatigue of Composites, France, 1997. p. 355–62. [10] Symons DD, Davis G. Fatigue testing of impact-damaged T300/ 914 carbon-fibre-reinforced plastics. Composite Science and Technology 2000;60:379–89. [11] Blaricum TJ, Saunders DS, Clark G, Preuss TE. Damage tolerance of impact damaged carbon fibre composite wing skin laminates. In: Proceedings of the 14th Symposium of the International Committee on Aeronautical Fatigue, Canada, 1987. p. 537–56. [12] Clark G, Van Blaricum TJ. Load spectrum modification effects on fatigue of impact-damaged carbon fibre composite coupons. Composites 1987;18(3):243–51. [13] Mitrovic M, Hahn HT, Carman GP, Shyprykevich P. Effect of loading parameters on the fatigue behavior of impact damaged composite laminates. Comp Sci Tech 1999;59:2059–78. [14] Saunders DS, van Blaricum TJ. Effects of load duration on the fatigue behaviour of graphite/epoxy laminates containing delaminations. Composites 1988;19:217–28. [15] Clark G, Saunders DS. Morphology of impact damage growth by fatigue in carbon fibre composite laminates. Materials Forum 1991;15:333–42.
[16] Ishikawa T, Hayashi Y, Sugimoto S, Matsushima M, Amaoka K. Development and test results of full CF/PEEK (APC-2) horizontal stabilizer models-basis for sst structure. AIAA 95-3931. [17] Ayling J. C-130 Hercules composite flaps fatigue test program. In: 44th International, SAMPE Symposium, 1999. p. 1442–8. [18] Poe Jr., CC Residual strength of composite aircraft structures with damage. ASM handbook, vol. 19: fatigue and fracture (USA). ASM International, 1996. p. 920–35. [19] Olsson R. Analytical prediction of large mass impact damage in composite laminates. Euromech Colloquium 400, 1999. p. 110– 17. [20] Kan HP, Whitehead RS, Kautz E. Damage tolerance certification methodology for composite structures. Eighth DOD/NASA/ FAA Conference on Fibrous Composites in Structural Design, part, 2, NASA Langley Research Center, 1990. p. 479–98. [21] Bergmann D, Ritter R. 3D deformation measurement in small areas based on grating method and photogrammetry. SPIE 2782; 1996:212–23. [22] Melin LG. Optical whole field measurement techniques for mechanical testing—a review. FFA TN 1999–35. The Aeronautical Research Institute of Sweden, 1999. [23] Sjo¨dahl M, Synnergren P, Johnson P. Applications of digital speckle photography in experimental mechanics. In: Lasers and optics in manufacturing, Proceedings, SPIE, vol. 3098, 1997. p. 195–203. [24] Zhao S, Ga¨dke M, Prinz R. Mixed-mode delamination behavior of carbon/epoxy composites. J of Reinforced Plastics and Composites 1995;14:804–26. [25] Asp LE, Sjo¨gren A, Greenhalgh ES. Delamination growth in a carbon/epoxy composite under fatigue loading. J of Composites Technology and Research 2001;23:55–68. [26] Scho¨n J, Nyman T, Blom A, Ansell H. A numerical and experimental investigation of delamination behaviour in the DCB specimen. Comp Sci Tech 2000;60:173–84. [27] Scho¨n J, Nyman T, Blom A, Ansell H. Numerical, and experimental investigation of a composite ENF-specimen. Engineering Fracture Mechanics 2000;65:405–33. [28] Kru¨ger R, Ko¨nig M, Albinger J, Ha¨nsel C. Combined, experimental-numerical approach for the determination of mixed-mode energy release rates at delamination growth. 35th AIAA/ASME/ ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, part, 2, Hilton Head, SC, 1994. AIAA Paper 941460. [29] Suemasu H, Wakabayashi H. Failure mechanism of composite laminates with different size multiple delaminations subjected to compressive load. 41st AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, Atlanta, GA, 2000. AIAA Paper 2000-1402. [30] Scho¨n J. A model of fatigue delamination in composites. Comp Sci Tech 2000;60:553–8.
Buckling behaviour and delamination growth in impacted composite specimens under fatigue load: an experimental study L. Gunnar Melin*, Joakim Scho¨n Swedish Defence Research Agency, SE-172 90 Stockholm, Sweden Received 5 January 2001; received in revised form 22 May 2001; accepted 7 June 2001
Abstract Impact-damaged carbon-fibre/epoxy composite laminates have been tested in fatigue under constant-amplitude loading with emphasis on studying mechanisms leading to delamination growth and fatigue failure. The shapes of the buckles have been measured with an optical whole-field measurement technique, while the extension and depth of delaminations was measured several times by ultrasonic C-scanning during testing. It was observed that the delamination growth occurs mainly in transverse direction to the load and that buckles on the backside usually have the same shape as some of the delaminations. This indicates that the buckling which takes place during the compressive part of the load cycle drives delamination growth. In specimens run to the fatigue limit, buckling and delamination growth occur only in the outer 2 or 3 layers on the backside while a specimen that failed before the fatigue limit buckled through the whole specimen. # 2001 Elsevier Science Ltd. All rights reserved. Keywords: B. Fatigue, B. Impact behaviour, C. Buckling, C. Delamination, D. Ultrasonics
1. Introduction Composite structures are increasingly being used in aircraft. During the aircraft operational life those structures will often suffer impact damage. It is important, therefore, to have reliable criteria for predicting the effects of such impact damage during fatigue. Before such criteria can be developed a good understanding of the damage mechanisms during fatigue loading is necessary. Several investigators have studied the way in which different parameters affect the fatigue life of impacted carbon-fibre/epoxy composites. Specimens made of T300/976 carbon/epoxy composites with different thicknesses have been fatigue loaded after impact [1]. It was found that the thicker specimens performed better in fatigue than the thinner ones. Comparing fatigue properties of undamaged and impact damaged composite laminates loaded in tension/tension loading at a stress min ratio, R=0.1, where R ¼ max ; it was observed that in low-cycle fatigue the fatigue life was slightly shorter for impact-damaged specimens than for undamaged specimens. However, in high-cycle fatigue the fatigue life was * Corresponding author. Tel.: +46-8-5550-4257; fax: +46-8258919.
significantly shorter for impact-damaged specimens than for undamaged specimens [2]. In a fatigue study [3] of impacted AS/3501-6 composites it was found that specimens loaded in compression/compression, R ¼ 1; have almost the same fatigue life as specimens loaded in tension-compression, R=1. Specimens loaded in tension/tension, R=0 had the longest fatigue life. Similar results were obtained in Ref. [4] where fatigue testing of 156 mm wide specimens of HTA/6376 composite was performed. Testing at two R values, R=1 and R=5, did not result in any significant difference in fatigue life. From this it could be concluded that the fatigue life is determined by the compressive part of the loading and not by the tensile part. Two different laminates were investigated, one quasi-isotropic (QI) and one dominated by fibres in the zero degree direction (ZD). When fatigue life is plotted versus stress, the latter laminate was found to have a higher fatigue resistance, although when plotted against strain instead the QI laminate was found to have the better performance. Damage growth during fatigue loading has been looked at in a large number of investigations. The conclusions sometimes seem to be contradictory, which might be due to differences in material, specimen design, impact energy and tested load cases. Several authors [5–10] have
0266-3538/01/$ - see front matter # 2001 Elsevier Science Ltd. All rights reserved. PII: S0266-3538(01)00085-9
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reported that most of the damage growth occurs in the final part of the fatigue life. The increasing damage can be monitored, for example by studying the decreasing stiffness [5], the increasing crack density [5] or the growth of the delamination envelope [7]. Gerharz et al. [8] identified two stages during the fatigue damage growth, ‘phase 1’ and ‘phase 2’. During phase 1 a small increase in buckling deformation and compliance was observed. No delamination growth out of the damage envelope was observed and the number of delaminations within the original damage envelope increased. The phase 2, which occurs during the last 10% of the fatigue life, is characterised by rapid damage growth. Buckling deformation and compliance increase and delaminations grow outside the original damage envelope. Delamination growth occurring during a significant part of the fatigue life has also been reported [11–13]. Mitrovic et al. [13] found for 38 mm wide specimens impacted with 2.1 J that the delamination growth rate decreases with delamination length. In Refs. [3,4,11,12], damages were reported to grow mainly in transverse direction to the load while in reference [13] it was found that most of the delamination growth occurred in the load direction. One way to study the damage growth is by using optical whole field measurement techniques such as shadow moire´ [11,12,14] or digital speckle photography [4] (DSP) to determine the shape of the buckles. Melin et al. [4] reported that the main buckles go inward on the impacted side and outwards on the backside of the laminate. Therefore all plies through the specimen have to buckle. The buckles on the backside had both larger size and amplitude. To the authors’ knowledge, one study using ultrasonic C-scanning to detect which delaminations through the specimen that grow during fatigue loading has been made [14]. It was found that most of the delamination growth occurred in the ply layers close to the back face of the specimens where the largest delaminations had been produced by the impact. They were not able to resolve the delaminations on a ply level. In a fractographic study [15] of damage growth during fatigue of impacted composites, it was found that delaminations at neighboring interfaces interact with each other. Ply cracks were found to establish boundaries for some delaminations. Delaminations were believed to form in fatigue at crossover points of intra-ply cracks. Before results from fatigue loading of impacted coupons can be used in design, impacted structures need to be fatigue loaded to study the transferability of coupon data to structures. Two sections of a horizontal stabiliser made of CF/PEEK (APC-2) with impact damage have been tested in fatigue. No damage growth was observed and the structures failed in quasi-static loading after the fatigue loading had been finished [16]. During certification of C-130 Hercules composite flaps barely visible impact damage, BVID, and visible impact damage,
VID, were introduced to a component. During fatigue testing, the VID did not grow while the BVID grew from a diameter of 11 mm to 19 mm [17]. Poe [18] has discussed if fatigue of impacted aeronautical composite structures are important. He argues that constant amplitude compression fatigue tests suggest that fatigue lives for nonvisible impact damage are adequate but those for detectable damage may need to be verified by spectrum tests. Although a large number of investigations have studied damage growth during fatigue after impact no investigation has studied delamination growth on a ply level and the buckling growth has also not been investigated in detail. The objective of this investigation is to study delamination growth on a ply level with ultrasonic C-scan together with detailed measurements of the buckling shape and growth using DSP. The delamination growth and hence the damage growth is expected to be related to the fatigue life of the specimens in compressive load.
2. Experimental 2.1. Specimen The specimens were prepared from Hexcel HTA/6376 carbon fibre/epoxy matrix prepreg with a nominal fibre volume fraction of 65%. The specimens had a size of 450mm156 mm. Two different lay-ups have been used, one dominated by fibres in 0 direction, ZD, and one quasi-isotropic, QI. The ZD lay-up has the sequence [45,135,0,90,45,135,02,(45,135,0, 90)2,45,135,02,45,135,0, 90]s, giving totally 48 plies with 16 in 0 , 8 in 90 and 12 each in 45 and 135 directions. The QI lay-up has the sequence [(90,135,45,0)s,(0,45,135,90)s]3, totally 48 plies with 12 in each of the 0 , 90 , 45 and 135 directions. Loading is applied along the 0 direction. The QI lay-up results in a specimen that is quasi-isotropic both in the plane and in bending [19]. The nominal specimen ply thickness is 0.13 mm; hence, the cured specimens were 6.24 mm thick. 2.2. Impacting The specimens have been impacted with energies of 27.6 0.4 J. Most of the impact threats will have impact energy less than 28 J [20]. The impacts were conducted in a conventional drop-weight rig using a drop height of 0.5 m. The velocity just before impact was measured with a photo transistor device. The impactor had a tip radius of 7.5 mm. During impact the specimens were clamped in gripping zones of the specimens short sides, along the long sides the specimens were supported. The impact causes on the frontside a permanent indentation 0.05– 0.1 mm deep and about 15 mm in diameter; similarly on
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Fig. 1. Loading arrangements including CCD cameras for DSP measurement. The solid specimen-boundary indicates the permanent indentation and the dashed lines outlines the buckling shape.
the underside a buckle 0.15–0.2 mm high and up to 30 mm in diameter is created. During certification of C-130 Hercules composite flaps, impact damage was considered barely visible when the depth of the impact size was a minimum of 0.25 mm [17]. Therefore, the impact damage in this program should be considered non-visible. 2.3. Fatigue testing Fig. 1 shows the specimen during compressive loading. The fatigue testing has been conducted in a servohydraulic MTS loading machine with a maximum load of 1000 kN. An anti-buckling device was used to prevent global buckling in compression. The anti-buckling device was made of steel and had a window of 100mm100 mm on each side of the specimen centred on the impact point. To minimise friction and allow movement between the
anti-buckling device and the specimen, a layer of a low friction polymer separated the specimen and the steel. In Fig. 1 is also indicated a permanent indentation/buckle caused by the impact (solid lines) and a typical out-ofplane deformation (dotted lines). The testing was done in constant amplitude loading using either of two R values: R=1 and 5. A frequency of 0.5 Hz was used, which resulted in negligible temperature increase in the damaged area. The frequency also gave time for buckling of the damaged region. All tests were performed at room temperature and in load-controlled mode. In Table 1, the number of specimens tested with the two whole field measurement techniques are tabulated. By ultrasonic C-scan using the system Krautkramer USPC 2100, the areal shape and depth of the delaminations are measured through the thickness. Numerical values of the delamination depth are available from the C-scan system. At each measurement point not more than one delamination, the one closest to the surface, can be detected. Consequently, delaminations deeper down in the laminate are hidden. From the C-scan maps it could be reconstructed between which plies each delamination occurred. All specimens were C-scanned after impact, before starting the fatigue test. To perform a C-scan during the test, the loading machine had to be stopped and the specimen dismounted; for this reason a limited amount of C-scans were made during the fatigue tests. Those tests are outlined in Table 2. The successive delamination growth was analysed in the image-editing program Corel PHOTO-PAINT 7. With the optical non-contact measurement technique digital speckle photography (DSP) [21–23] the displacement field was measured in the window of the antibuckling devise. The method is also known as Digital Image Correlation, Electronic Speckle Photography and the Grating Method. The measurements were made with the system Aramis from the German company GOM mbH. In DSP, a random pattern is applied on the specimen surface, here dotted patterns were created using white matt spray paint. Images are acquired by CCD cameras of the undeformed and deformed states and evaluated numerically to determine the displacement field. By using two cameras imaging the specimen surface from different angles, all three displacement components can be determined. Here, a total of four cameras were available, making it possible to perform DSP measurements on both sides of the specimens as
Table 1 Number of specimens tested in constant amplitude fatigue Lay-up
Total number of specimens
Specimens with ultrasonic C-scan during test
Specimens with DSP during test
Specimens with DSP and C-scan during test
ZD QI
11 15
3 5
4 6
2 3
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Table 2 Specimens measured with ultrasonic C-scanning and DSP during fatigue test Specimen
Minimum stress (MPa)
R value
Life (number of cycles)
C-scan after N cycles
ZD14 ZD15 QI5 QI13 QI14
206 247 185 185 226
5 5 1 5 1
117 140 120 75 973 Stop at 1 020 000 230
1000, 7000, 44 000, 92 000, 115 000, 117 000 51, 114 70 000, 73 600 1 020 000 100, 200
shown in Fig. 1. Each measurement area covered the window area of the anti-buckling device. The images were acquired after different intervals of cycling. During image acquisition the specimens were kept under constant load giving a hold time of about 10 s at the maximum load. Saunders and van Blaricum [14] reported that hold times of 10 s do not have any discernible effect on the fatigue life. All DSP registrations were made at the minimum compressive load except for one specimen for which registrations were made during complete load cycles. That enabled one to see during which part of the load-cycle buckling occur and if there is any hysteresis.
3. Results A selection of the experimental data is presented in this paragraph. The focus is on the specimens with both C-scan and DSP measurement (see Table 2), where the areal shape of delaminations and buckles can be compared. The results from the specimen on which complete load cycles were measured are also presented. The fatigue results plotted against compressive stress for constant amplitude loading of the two lay-ups can be seen in Fig. 2. The stress has been calculated using the total width, thickness and elastic constants of the
Fig. 2. Minimum stress vs number of cycles for all specimens tested in constant amplitude loading.
undamaged specimens. The fatigue results are taken from Ref. [4]. 3.1. Delamination growth analysis 3.1.1. Definitions Four specimens are presented. Two had typical buckling shapes for a long and a short fatigue life: ZD14 and QI14 with long and short fatigue life, respectively. Of the other two specimens, QI13 was run up to the fatigue limit and QI5, which had a rather long fatigue life, exhibited a particular buckling shape. C-scan and DSP measurements have been made from both sides of the specimen. The result maps are presented with the same absolute specimen orientation using the co-ordinate system defined in Fig. 1. In Fig. 3 the laminate directions 0–360 in the laminate plane are defined together with a co-ordinate system with the same orientation as in the result maps. The 90 and 270 directions will be referred to as sideward directions. The out-of-plane DSP measurements of displacement are presented as iso-contour maps superimposed on Cscan maps. A 0.1 mm contour interval is used. The displacement is for the minimum load during each cycle relative zero load before the first cycle. The zero-displacement contours are labelled with a ‘0’, further, the buckles larger than 0.1 mm are labelled with a ‘+’ or ‘’ depending on if they buckles in positive or negative z-direction. For example, the buckle in Fig. 1 would be labelled with a ‘’ on both the frontside and the backside. The displacement measurements were made during the last loading before performing the C-scan except at
Fig. 3. Definitions: laminate directions 0–360 relative co-ordinate system.
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Fig. 4(b). Results for ZD14. Extension and growth of delaminations in sideward directions.
Fig. 4(a). Results for ZD14. DSP measurements of z-displacement superposed to C-scan maps. The displacement is plotted as iso-contour with 0.1 mm spacing. Left column: frontside; right column: backside.
the first cycle when the C-scanning was made first. The positioning of the contour maps relative the C-scan images has an uncertainty of 1–2 mm. When the C-scan map is missing on the result maps, it is because no Cscan was made at that occasion. Observe that the Cscan maps are sometimes a bit noisy. The colour range of the C-scan maps is shown versus distance from the front surface in Fig. 4(a). The C-scan equipment is not able to resolve delaminations in the outer 3–5 plies, these delaminations are seen as a single delamination on the C-scan maps. The C-scan analysis shows that the delaminations created during impact with few exceptions follow the fibre direction of the adjacent ply closest to the backside of the specimen. The initial distribution and the growth of the delaminations are outlined in diagrams, where the delamination shape has been projected onto two directions: 90 and 270 . These directions were the dominating growth directions. Of all observed growing delaminations only 10–20% grew along one of the 45 or 135 directions and almost no growth occurred along the 0 or 180 directions. In other words, these diagrams show the largest extension along the x-axis of the visible delaminations at different fatigue lives. In a few cases, where the largest extension of the delaminations was not at the place for the delamination growth, then the diagram describes the part of the delamination that grew. Observe that
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because only the outer delamination can be detected at each measurement point, it is probable that other delaminations are hidden. The growth of these hidden delaminations could not be measured. Some delaminations are not visible in the C-scan maps until after some load cycle. In the diagrams, it is indicated as default that these delaminations have started to grow from a width of zero though it is not known from where the growth started. Observe also that the delaminations appear at interfaces between two plies and, contrary to what the delamination diagrams indicate, has no extension in the thickness direction. The zero displacement level in the DSP contour maps has been set by subtracting the average of measurement points close to the edge of the measurement area. In compressive load, besides the buckling around the damage area, usually an out-of-plane slope was observed. This gave a difference of up to 1.0 mm in out-of-plane displacement over the buckling window. This out-of-plane rotation is caused by imperfections in the loading machine. To emphasise the local buckling at the damage area, the rotation has been eliminated from the displacement maps by a co-ordinate transformation that adjusts the displacement field to a plane. It is not expected that the slope had any significant impact on the buckling behaviour and delamination growth. 3.1.2. ZD14 (Fig. 4) This specimen was tested at a load close to the fatigue limit at a minimum stress of 206 MPa and an R value of 5 (Fig. 2). From the start the buckles on both sides extend about half the diameter of the delamination envelope and have relatively small amplitudes. (Observe that since the outer contours of the buckles are at 0.1 mm displacement, the buckles extend some distance outside that contour.) Not until after 92 000 cycles does the inward frontside buckle begin to grow in area. The outward buckle on the backside has a rounded shape. This buckle has a slow, steady growth in both amplitude and area up to 92 000 cycles. Regarding the delaminations, minor growth is observed on the outer plies on both sides for the first 1000 cycles, after that no growth is observed until after 44 000 cycles where again some delaminations near the surfaces have grown in area. The growth near the frontside correlates to the increasing width of the main inward buckle and the development of an outward buckle above it. No growth occurs near the envelope. Closer to failure, the specimen broke after 117 140 cycles; the trend is that a successively larger number of delaminations grow at positions deeper down from the surfaces. The delaminations that have started to grow also continue to grow. Between 92 000 and 115 000 cycles, all delamination growth takes place at positions about 10 mm inside the envelope. On the backside the sideward increase of the buckle area correlates to the
delamination growth at the interfaces between ply 4-5 and 5-6. Between 115 000 and 117 000 cycles, large delamination growth is observed at many interfaces. The envelope is extended by three delaminations, both in the middle of the laminate and closer to the backside. Many delaminations now have almost the same extension in sideward direction. The backside buckle has grown significantly sidewards. Its areal shape follows mainly the shape of the delaminations from ply 6-7 and out. On the frontside the inward buckle gets a narrow shape between two minor outward buckles. 3.1.3. QI5 (Fig. 5) This test was performed at a relatively low load with a minimum stress of 185 MPa and an R value of 1 (Fig. 2). Its specific buckling behaviour was not observed in any other specimen. In the first cycle, only a small inward buckle appear on the frontside. The buckle on the backside has its highest amplitude over a delamination in the outer plies beside the impact point. After 70 000 cycles, there are both an inward and below (in negative y-direction) an outward buckle of similar size and areal shape at the frontside. The largest delamination growth has occurred at the transition between the two buckles. After 73 600 cycles, the outward buckle on the frontside has increased both in area and amplitude. On its upper side, there is a sharp transition to the inward buckle and on the other sides its edge is similar to the delaminations in the six outer plies. On the backside the buckle has grown sidewards, reaching the envelope. Its maximum amplitude is 5 mm in the negative y-direction from the impact point. Until 73 600 cycles when the last C-scanning was made on both sides, almost all delamination growth occurred in the outer five interfaces. Some delaminations near the frontside at the position of the outward buckle grew partly in the 180 direction. The DSP measurement just before failure at 75 900 cycles, the specimen failed after 75 973 cycles, shows that the outward buckles on both sides have grown along the 90 direction; the place for the growth coincided though the backside buckle has a larger extension. The inward buckle on the frontside has got its amplitude decreased and can be described as being pressed upwards by the dominating outward buckle. 3.1.4. QI13 (Fig. 6) The test was run up to the fatigue limit at a minimum stress of 185 MPa and an R value of 5 (Fig. 2); under the same loading conditions, specimen QI9 failed after 21 115 cycles. Specimen QI5 was loaded at the same minimum load but with another R value. No delamination diagram is shown; the only delamination growth observed was near the backside at the two outer interfaces. This
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Fig. 5(b). Results for QI5. Extension and growth of delaminations in sideward directions.
point. The largest buckling appears in a 5 mm wide stripe in the outer ply of 0 direction, this stripe could be observed visually during testing.
Fig. 5(a). Results for QI5. DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside, right column: backside.
growth, which was up to 15 mm in length, was directed away from the impact point and the delaminations were beside the impact point. From the start, no detectable buckle emerge on the frontside. After 1020 000 cycles a small inward buckle is indicated around the impact point. On the backside, an outward buckle appears during the first cycle with maximum amplitude nearly 10 mm away from the impact point. The buckle has the same areal shape as the outer delaminations. At the fatigue limit, the buckle has increased both in amplitude and size. Its main part is still beside the impact point and the enlargement has been directed away from the impact
3.1.5. QI14 (Fig. 7) A high fatigue load of 226 MPa and an R value of 1 (Fig. 2) was applied, causing relatively large buckles on both sides from start. The backside buckle has similar sideward extension as the delamination envelope. Delamination growth is seen after 100 cycles at the interfaces close to the surfaces. Nearly all delamination growth is in sideward directions, but not only at delaminations with a 90 ply adjacent towards the backside. The envelope is enlarged at one delamination. The backside buckle grows in 270 direction together with the extended envelope. On the frontside, the inward buckle has not increased; a small outward buckle has appeared above. After 200 cycles, delaminations through the whole specimen have grown, all within the extension of the backside buckle. The backside buckle has continued to grow sidewards while a small inward buckle is detected above the main buckle. On the frontside, the main buckle is almost unchanged and the delamination growth near the frontside has partly appeared outside this buckle. The specimen failed after 230 cycles.
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Fig. 6. Results for QI13: DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside; right column: backside.
3.2. Fatigue cycle monitoring On one specimen, QI53, some specific cycles were monitored by making totally 10 DSP registrations per cycle. The specimen was loaded at 226 MPa and R=1 which is the same load as for specimen QI14. Specimen QI53 had, for this load, a rather long fatigue life of 2356 cycles. Four specific cycles are presented: the first cycle, after 300, 900 and 2200 cycles. The C-scan maps before loading are shown in Fig. 8(a) together with the buckling shapes at the minimum compressive load. The damage zone is slightly smaller than for the other presented QI specimens. The z-displacement during the cycles is shown in Fig. 8(b–d). Finally, the strain histories along the loading direction are presented in Fig. 8(e–f). The strain has been taken as the average strain over an 83 mm long section where the damage zone is included. The buckling shapes in Fig. 8(a) have a similar appearance to the previously discussed specimens. On the backside, the highest buckling occur outside the impact point and for this reason the buckling amplitude is presented both at the impact point in Fig. 8(c) and at the position with highest buckling amplitude in Fig. 8(d). During the first cycle, the buckling is, on both sides, clearly smaller during the compressive loading phase than during the unloading phase. During the fatigue life the load/buckling relationships get increasingly linear though some hysteresis effect remain on the backside. The small displacement registered during tensile loading is probably caused by the permanent buckle/indentation getting straightened out. The strain plots show a linear behaviour in tension and a slightly lower stiffness during
compression. There is almost no hysteresis and a small weakening occur in compression during the fatigue life.
4. Discussion In this section, the main features of the results and what can be deduced from them are discussed. The conclusions regarding the buckling behaviour are also supported by the other DSP-measured specimens not presented here. Two of those were presented in Ref. [4]. 4.1. Delamination growth and correlation to buckling The initial delamination shape follow with few exceptions the direction of the adjacent ply closest to the backside of the specimen, a ply that also have big influence on the delamination growth direction during fatigue while the ply closest to the frontside has almost no influence on the growth direction. In the outer interfaces during fatigue loading, delaminations can grow in any direction following a single stripe but deeper down in the specimen delaminations grow either in the direction of the adjacent ply towards the backside or in the direction transverse to the load. The dominating delamination growth direction is transverse to the load, hence most delaminations that grow have a 90 ply adjacent towards the backside though some delaminations adjacent to 45 or 135 plies also grow. Only in the outer layers is growth observed at delaminations with a 0 ply adjacent towards the backside. The longest growth takes place in the outer delaminations though a trend with increasing number of cycles
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Fig. 7(b). Results for QI14. Extension and growth of delaminations in sideward directions.
Fig. 7(a). Results for QI14. DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside, right column: backside.
is that delaminations deeper into the laminate start to grow. A delamination that has started to grow usually continues to grow. For short fatigue lives, the growth rate is approximately constant whereas for longer lives most growth occurs just before failure. The delamination envelope is usually not extended until near failure, then delaminations between the middle of the laminates and the backside grow out of the envelope. Comparing DSP measurements with the C-scan maps, it is seen that the outward buckle on the backside usually has a similar areal shape as some delaminations, contrary to the inward frontside buckle. The agreement in areal shape on the backside can from start be with the outer delaminations and later in fatigue life to delaminations deeper down from the surface; an example of this is specimen QI5. For higher loads, though, also from start the sideward extension might be similar to the envelope as for specimen QI14. This good shape agreement gives a strong support of the idea that the delamination growth is driven by buckling. A possible mechanism controlling the delamination
growth could then be as outlined in Fig. 9. It shows the buckling in a cross-section transverse to the load around the damage zone, both at the beginning and at the end of the fatigue life. Since all plies and hence all delaminations buckle in the same direction and they are buckling into the specimen on the frontside, all plies in the thickness direction have to buckle at least as much in area and amplitude as is observed on the frontside surface since otherwise the plies would have to penetrate each other, which is impossible. This discussion can be applied to each delamination starting from the frontside. Therefore, the buckle on the backside has to have the same or larger amplitude and area as the frontside buckle. This was also observed in the experiments. As a delamination buckles, all three fracture modes will be introduced along the edge of the delamination. Since the critical energy-release rate for mode I, GI,C, is approximately 240 J/m2 and for mode II, GII,C, is about 800 J/m2 [24–27], it can be assumed that the delaminations grow mainly in mode I. The value for mode III is not available but can be assumed to be in the order of GII,C. Calculations of the mode separated energy release rate of delaminations in the post buckling state have found that GI has a similar or larger value than GII and GIII [28,29]. The ratio of GI to the total energy-release rate has a large influence on the delamination-growth rate in fatigue [30].
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Fig. 8(b)–(f). Results for QI53. (b) z-Displacement on frontside at the impact point during specific load cycles. (c) z-Displacement on backside at the impact point. (d) z-Displacement on backside at the point with maximum buckling amplitude. (e) Average strain in loading direction on frontside. (f) Average strain in loading direction on backside. Fig. 8(a). Results for QI53. (a) DSP measurements of z-displacement superposed to C-scan maps. Left column: frontside; right column: backside.
In the first cycle in Fig. 9, buckling occur through the whole laminate but since the buckle is smaller than the delamination envelope, it is only at the outer delamination that a mode I component appear. The delaminations that do not buckle at the crack tip will be closed. As a consequence GI will be small for those delaminations and no growth will occur. As the outer delaminations grow and the buckling area gets larger at the same rate, the edge of the buckle will reach the delamination envelope. Hence, the delaminations deeper down in the specimen will also be loaded in mode I hence promoting delamination growth. This has also been observed in the experiments (Figs. 4 and 5). For specimen ZD14, delamination growth occurred deep inside the specimen, which gives a larger delamination width than is the case close to the backside. This would suggest that those delaminations buckle all the way to the crack tip. Smaller diameter buckles are then necessary to open the crack tips of the delaminations close to the backside. This means that several delaminations buckle separately and that the buckle
observed on the surface is the results of several buckles superimposed on each other. Support for this can be found from Fig. 4(a) at the backside after 117 000 cycles; the slope of the buckle is small over the delaminations 2–3 mm into the specimen whereas at the edge of the outer delaminations the slope of the buckle increases significantly. Close to failure, minor buckles usually appear above and under the main buckles. The minor buckles typically have an amplitude of 0.2 mm, buckle in opposite direction to the main buckles and do not cause any delamination growth. These buckles probably do not have any significant influence on the fatigue life. The specimen QI5 is an exception where an outward buckle on the frontside becomes dominant over the inward buckle. Its areal shape is similar to the delaminations within the outer five layers why at least these plies buckle. These delaminations grow in area as the outward buckle appears. It is not known whether this buckling behaviour had any effect on the fatigue life. No differences between the two R values 1 and 5 regarding the buckling behaviour and delamination growth could be distinguished. The corresponding conclusion from the measured fatigue life was presented by
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Fig. 9. Outline of proposed mechanism for buckling-induced delamination growth.
Melin et al. [4]. This suggests that the tensile part of the loading has a small influence on the delamination growth, whereas the compressive load is the harmful part. From Fig. 8(b–d), it can be concluded that buckling occurred only during compressive loading. The load/strain plots in Fig. 8(e–f) show that the specimens were less stiff during the compressive loading than in tension. This is reasonable since a damage zone can take less load as it buckles in compression than when it is stretched in tension. 4.2. Fatigue limit In most QI specimens, the delaminations nearest the backside was outside the impact point. From start at a low load usually the largest buckling occur there—see specimen QI5 or QI53. Specimen QI13, which was run up to the fatigue limit, had two particular buckling features compared to the failed specimens: the backside buckle was fully outside the impact point and almost no inward buckling occurred on the frontside. During testing, the only evolution observed was that the backside buckle grew in a direction away from the impact point. It is probable that only the outer two plies buckled— one indication of this is that all observed delamination growth occurred there. Also, in the other QI specimens run to the fatigue limit, the only observed delamination growth was in the outer 2–3 plies. This indicates that buckling in only the outer 2–3 plies is not sufficient to cause failure. Failure demands that buckling also occurs deeper down in the specimen. This suggests that a fatigue threshold exists; the specimen can only fail if the buckling goes at least 5–8 plies deep on the backside. For this, a minimum compressive
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load has to be exceeded during the load cycle, a threshold load that is individual for each impact damaged specimen. The tests indicate that when the threshold is exceeded, the buckling at the impact point has an amplitude of at least 0.15 mm on both sides. This could be one way to measure the fatigue threshold though the number of tests is too few to claim a rule. No DSP testing has been made on a ZD specimen run to fatigue limit, making it difficult to make statements regarding buckling behaviour at the fatigue limit. A difference from the QI specimens is that the delamination in the outer 45 ply on the backside covers the impact point, this ply is the first to buckle at low loads. This leads to different buckling shapes at low loads for the two laminates. To know whether a ZD specimen is expected to fail it is probably not enough to know if the impact point on the backside buckles, one also has to conclude from the buckle shape if only a stripe in the outer ply buckles or if plies deeper down buckles, giving it a more circular shape.
5. Summary and conclusions Impact-damaged carbon fibre/epoxy composite laminates have been tested in fatigue at constant amplitude loading. The objective has been to study the mechanisms that leads to delamination growth and in the end fatigue failure. To achieve this, the shapes of the buckles has been measured with an optical whole field measurement technique and the laminates have been studied by ultrasonic C-scan several times during testing. Both quasi-isotropic and zero-degree dominated laminates have been tested. The observations indicate that the fatigue life is controlled by delamination growth in the direction transverse to the load, though the final failure mechanism can be another one. More than half of the delaminations that grow are along plies in transverse direction. The delamination growth is driven by buckling that occurs during the compressive part of the load cycles; an indication of this is that the buckle on the backside usually had the same areal shape as some delamination. In specimens run to fatigue limit, buckling and delamination growth occur only in the outer 2–3 layers on the backside. For the quasiisotropic laminate, it is suggested that a threshold compressive load exists, a threshold which is individual for each specimen. At loads higher than the threshold, the specimen buckles through the whole thickness and fatigue failure is probable.
Acknowledgements The authors wish to express their thanks to the Swedish National Research Program (NFFP336+345) and to the Swedish Defence Material Administration (FMV).
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