Compressive Failure of Impacted NCF Composite ... - SAGE Journals

Compressive Failure of Impacted NCF Composite Sandwich Panels – Characterisation of the Failure Process FREDRIK EDGREN* AND LEIF E. ASP SICOMP AB, Box 104, SE-431 22 Mo¨lndal, Sweden

PETER H. BULL Department of Aeronautics Royal Institute of Technology SE-100 44 Stockholm, Sweden (Received May 23, 2003) (Revised August 28, 2003)

ABSTRACT: In the present study, non-crimp fabric (NCF) composite face sheet sandwich panels have been tested in compression after impact (CAI). Damage in the face sheets was characterised by fractography. Compression after impact loaded panels were found to fail by plastic fibre microbuckling (kinking) in the damaged face sheet. Studies of panels for which loading was interrupted prior to failure revealed extensive stable kink band formation at several positions and in numerous plies. Kink bands initiated and propagated within a wide region close to the point of impact. In addition, kink bands initiated in zones with high shear stresses, away from the impact centre line. Consequently, the fractographic results from this investigation do not support the assumption of modelling the impact damage as an equivalent hole. To achieve accurate predictions of kink band initiation, the stress field must be known. The results from this study imply that bending effects caused by remaining dent or material eccentricities in the damaged region must be considered. KEY WORDS: damage tolerance, fractography, CFRP, sandwich, CAI.

INTRODUCTION decades damage tolerance modelling of composite structures has been intense. State-of-the-art damage tolerance models when structures are loaded in compression either consider delamination buckling and propagation (see e.g. [1–4]) or compressive failure caused by stress concentrations in the vicinity of the damaged region treated as an open hole or a ‘‘soft inclusion’’ (see e.g. [5–8]).

I

N THE LAST two

*Author to whom correspondence should be addressed. E-mail: [email protected]

Journal of COMPOSITE MATERIALS, Vol. 38, No. 6/2004 0021-9983/04/06 0495–20 $10.00/0 DOI: 10.1177/0021998304040559 ß 2004 Sage Publications

495

496

F. EDGREN ET AL.

Analysis of delamination buckling and growth in composite panels is very complex and requires a nonlinear structural analysis in combination with fracture analysis. The material is anisotropic and inhomogeneous and delaminated members tend to be partly in contact. The materials are, however, usually sufficiently brittle to motivate a linear elastic fracture mechanics approach. Compression failure of unnotched composite laminates is usually by plastic fibre microbuckling [9], alias kinking. The kinking mechanism is a plastic shear instability where initially misaligned fibres rotate within a band, under increased applied load, and form a kink band. The compressive kinking strength is governed primarily by the shear yield strength and the fibre misalignment angle [10,11]. Fleck and Budiansky [12] also showed that the critical compressive stress for initiation of microbuckling is severely reduced by the presence of remote shear stresses. Compression failure caused by stress concentrations in the vicinity of the damaged region has been treated by an equivalent damage size concept [6,8]. The approach taken is to treat damage as an open hole or a ‘‘soft inclusion’’ where the equivalent damage area is based on the size of the damaged zone measured by ultrasonic C-scan. Compressive failure of fibre composite laminates from pre-impregnated tapes (prepregs) with an open hole is by initiation and propagation of fibre microbuckle from the edge of the hole [13]. One of the first models for prediction of the open hole compression (OHC) strength is the maximum stress criterion which assumes failure to occur when the maximum stress equals the unnotched compressive strength of the material. Another criterion in this spirit, the net section stress criterion, assumes failure to occur when the net section stress equals the unnotched strength. In addition, the point stress criterion (PSC) and the average stress criterion (ASC) have also been used to predict the OHC strength [14,15]. These criteria assume fracture to occur when the stress at a characteristic distance away from the hole edge equals the unnotched compressive strength of the laminate. Soutis et al. [16] developed a large-scale crack-bridging model for the initiation and propagation of compressive damage emanating from the edge of a hole or a notch. In this model the damage is simulated by a compressive mode I crack with a cohesive zone behind the tip. The model requires knowledge of the unnotched strength and the in-plane compressive fracture energy. This model has also been successfully used for predictions of the compression after impact (CAI) strength, where the impact damage has been treated as an equivalent hole [6,17]. An ‘equivalent damage’ approach taken by Nyman [8] is treating impact damage as a ‘‘soft inclusion’’. In the work by Nyman, Lekhnitskii’s complex variable technique [18] was employed. This technique makes it possible to calculate stresses and strains not only for notched plates but also for regions comprising different elastic moduli, so called ‘‘soft inclusions’’. Nyman assumed the damaged region to have a smooth stiffness degradation, from the edge of the damage zone to its centre, and used a maximum strain criterion (based on strength of the unnotched laminate) to predict compressive failure. Much of the research on compression failure mechanisms and damage tolerance of composites has been performed on laminates manufactured from prepregs. As described above, models for prediction of effects of damage in composite materials based on prepregs are now well established. This is not the case for non-crimp fabric (NCF) based composites. NCFs are textile preforms with stacked layers of unidirectional oriented fibre bundles stitched together through the thickness. Even though the name of the preform, non-crimp, implies otherwise, there is a certain waviness of the fibre bundles in the fabric. This will thus affect the compressive strength of the NCF laminates. The heterogeneous

497

Compressive Failure of Composite Sandwich Panels

structure of the NCFs, with fibre bundles, waviness and through-the-thickness stitches, will also affect the impact induced damage in the laminates. The objective of this paper is to identify the mechanisms that control failure of impact damaged NCF face sheet sandwich panels for naval applications loaded in compression. Detailed knowledge of mechanisms initiating and governing failure will provide a means to judge the influence of damage features formed at impact (i.e. matrix cracks, delaminations, remaining dent and fibre fractures) on the residual strength of the panels. Thus, fundamental understanding of the interaction of damage features present in the material and the identified failure mechanism will provide a possibility to develop efficient, physically based models to predict the CAI strength of this type of panels.

EXPERIMENTAL Materials and Manufacture All sandwich panels tested in this study were manufactured by Kockums AB Karlskronavarvet. Two material configurations were investigated. The first type with a 60-mm thick Divinycell H80 core (density: 80 kg/m3) and CFRP face sheets. The face sheet laminates were manufactured from carbon fibre NCFs 0 /90 and 45 . The sandwich panels were manufactured by vacuum infusion of the resin into the stacked core material and the NCF-reinforcement. Vinylester was used for the matrix. The face sheet lay-up was: [0/90/45/45]S, with an approximate thickness of 1.8 mm. This first sandwich configuration will hereafter be referred to as ‘‘type A’’. The second material configuration, ‘‘type B’’, was manufactured from a 50-mm thick Divinycell H200 core (density: 200 kg/m3). The face sheets on this panel type were manufactured from the same materials as the type A-panels. The face sheet lay-up was [0/90/45/45]S3, with a thickness of approximately 5.4 mm. The fibre volume fraction of all composite face sheets was approximately 50%. The panels were cured at 60 C for 24 h. A summary of the materials used and the dimensions of the two panel types are presented in Table 1. Specimens were saw cut from panels to a final dimension of approximately 300 300 mm2. Tabs were bonded to the laminates to assure a safe load introduction. The two tabbed sides of the panels (where the load were to be introduced) were made plane and parallel by machining. A sketch of the panels is shown in Figure 1. In addition, panels with the dimension 150 150 mm2 were manufactured from the type B sandwich material. In one of the skins of each of these panels a hole or notch was cut. The diameter of the centrally placed hole was 50 mm. The notch, with a length of 50 mm, was oriented in the 90 -direction and placed central in the skin. The height of the notch was approximately 1.5 mm.

Table 1. Material and thickness of sandwich panels. Panel Type A B

Core Material

Core Thickness (mm)

Face Sheet Lay-up

Face Sheet Thickness (mm)

Divinycell H80 Divinycell H200

60 50

[0/90/45/45]s [0/90/45/45]s3

1.8 5.4

498

F. EDGREN ET AL. 50–60 mm

~ 300 mm Loading direction

~40 mm 1.8–5.4 mm

0° Impact point 90° ~ 300 mm Strain gauges

50 mm 50 mm

Dial gauge

~40 mm Figure 1. Dimensions of sandwich panels and positions of strain gauges and dial gauge.

Table 2. Details of the impact tests. Panel Group

Panel Type

Impact Energy (J)

Impact Mass (kg)

Drop Height (m)

Number of Panels

ABVID AVID A90BVID BBVID BVID BHOLE BNOTCH

A A A B B B B

30 50 30 100 250 – –

15.287 15.287 15.287 15.287 15.287 – –

0.200 0.333 0.200 0.667 1.667 – –

4 4 3 4 4 2 2

Mechanical Testing Impact damage was inflicted to the large panels using a conventional drop-weight rig. The impact energy was varied by altering the drop height. Tip radius of the impactor was 12.5 mm. A laser optic sensor was used to measure the displacement of the impactor during the test. During the test, the sandwich panels were placed on a rigid steel plate below the impactor. After the impact, the impactor was caught by hand on the way up to prevent it from falling down onto the specimen a second time. Four different impact energies were used. The energy levels were chosen to correspond to barely visible impact damage (BVID), and visible impact damage (VID) on the surface of the two types of panels. Totally 19 panels were impacted divided into five groups called ABVID, AVID, A90BVID, BBVID and BVID. The groups ABVID, AVID, BBVID and BVID each contained four panels, whereas the group A90BVID contained three panels. All panels within each group were impacted with the same energy level. The energy levels, impactor mass and drop heights are presented in Table 2. The impact position was central on one face sheet only per panel.


499

The groups of panels with holes and notches, called BHOLE and BNOTCH, each contained two panels. A summary of all panel groups can be seen in Table 2. Impact damaged, holed and notched panels were then tested in compression. All panels, except the A90BVID group, were tested in uniaxial in-plane compression in the 0 -direction (see Figure 1). Two of the panels in the group A90BVID were tested in uniaxial in-plane compression in the 90 -direction. During the compression test the panels were placed between two rigid and parallel steel plates in a 2 MN press. The two unloaded edges of the panel were unsupported and free to move. To study failure initiation, the loading of one panel in each group was stopped prior to failure. Out-of-plane displacement of the skin at the impact point was monitored on all impacted panels using a dial gauge. The strain in the panels compressed until failure was monitored by strain gauges mounted onto the face sheets. Two strain gauges on each side of the impacted panels were attached. On the panels in the BHOLE and BNOTCH groups, strain gauges were attached only on the cut skin. A sketch of the panels showing the positions of the strain gauges is presented in Figure 1. All panels were loaded at a constant cross-head speed of 1 mm/min. Fractographic Methods All impacted panels were examined after impact by means of ultrasonic C-scan. To allow for successful ultrasonic evaluation of the damage, the surface of the damaged skins was smoothened by application of a thin layer of vinylester. This layer was ground to accomplish the smooth surface. The panels were C-scanned also after the compression tests. In addition, the impact damage distribution in the laminates was studied by optical microscopy. The studied specimens were sectioned and polished. Polishing was carried out in several steps, ending with a 4000 grade paper. The polished surfaces were studied in an optical stereo microscope. The observed damage was first drawn by hand onto a paper, and later manually transferred to a computer picture. From the panels where loading was interrupted prior to failure, a small area close to the point of impact was cut out. This laminate piece was ground and polished stepwise in the thickness direction to identify locations where failure initiated. After each step, the surface was studied under the microscope and fibre fracture was registered, if noted. The shape of the remaining dent on the surface of two of the impacted panels was measured. The panels were mounted on a milling table, where the position was monitored in x- and y-coordinates. A digital dial gauge was used to measure the z-position (out-ofplane direction) of the laminate surface. The z-coordinate was registered along lines in the x- and y-direction, through the point of impact. The in-plane size (diameter) of the dent was taken as the distance between the two points were the measured z-level of the dent intersected the measured mean level of the undamaged laminate surface in the region surrounding the dent. RESULTS AND DISCUSSION Impact Damage Characteristics Impact damage in composite face sheet laminates may consist of matrix cracks, delaminations and fibre fracture [19, 20]. In addition, impact can cause damage in the

500

F. EDGREN ET AL.

sandwich core material and a remaining dent to the face sheet. This study is focused on the damage in the sandwich face sheets. The amount of damage in a laminate face sheet and the distribution of damage between the different damage modes depend on the impact mass, impact velocity, impactor geometry and boundary conditions of the panel at impact [19, 20]. This section presents characterisation of damage in the face sheets caused by impact to the studied panels. All impacted panels were investigated by means of ultrasonic C-scan. In Figure 2(a), a C-scan image of a type A panel with BVID is presented. Delaminations extending in the 0 - and 45 -directions are clearly seen. The projected damage has an elliptical shape rather than circular. The individual delaminations, however, are more rectangular in shape, with the long boundaries of the delaminations straight. This is in contrast to the peanut shaped delaminations often found in prepreg based laminates [21]. The elliptical shape of the projected damage is found also for the higher impact energy on the same type of panel, see Figure 2(b). The impact damage in the type B panels has a more circular shaped projection, shown in Figure 3. The rectangular shape of the individual delaminations can, however, be distinguished in Figure 3(a). The in-plane dimensions of the impact damage have been measured in the 0 - and 90 directions. The average results are presented in Table 3. In the table, dimension of the overall damage is denoted ‘‘projected damage’’. The ‘‘overlapping damage’’ refers to the dimensions of the region where the delaminations in the different directions overlap. The damage size in the 0 - and 90 -direction is denoted ‘‘2b’’ and ‘‘2a’’ respectively. It was not possible to distinguish the dimensions of the ‘‘overlapping damage’’ for the panels with VID from the C-scan images due to the large amount of fibre fracture in this region. After impact, a dent remains on the surface of the impacted face sheet laminate. The 50 and 250 J impacts, corresponding to VID, leave severe damage at the point of impact. A hole is punched in the laminate with a large amount of fibre fracture. This type of damage has not been characterised further in this study. The impacts corresponding to

0º 90º

(a)

(b)

Figure 2. Ultrasonic C-scan images of impact damaged type A panels: (a) ABVID; (b) AVID.

501


0º 90º

(a)

(b)

Figure 3. Ultrasonic C-scan images of impact damaged type B panels: (a) BBVID; (b) BVID.

Table 3. Characteristics of impact damage (standard deviations within brackets). Projected Damage

Overlapping Damage

Panel type

2a (mm)

2b (mm)

2a (mm)

2b (mm)

ABVID AVID BBVID BVID

32 36 63 69

54 (3.8) 69 (22.8) 68 (1.8) 72 (1.1)

11 (1.3) – 17 (1.3) –

13 (0.6) – 17 (1.1) –

(1.3) (0.5) (5.8) (5.6)

BVID cause a circular shaped dent on the surface of the laminate, with little or no fibre fracture. An example illustrating the shape of the dent in two perpendicular directions is shown in Figure 4. From this figure, it is possible to estimate the dent depth and the dimensions of the dent in the 0 - and 90 -directions. The average dimensions of the dent are presented in Table 4. The damage distribution in two of the panels with BVID was studied under a microscope. A section in the 90 -direction through the damaged laminate of an ABVID panel is presented in Figure 5. The amount of matrix cracks observed in this laminate is smaller than what is usually seen in prepreg tape based laminates [21]. The delaminations seen in the ultrasonic C-scan image, Figure 2(a), are found also in the sectioned laminate in Figure 5. A section through the impact damage of a BBVID panel, Figure 6, also shows a small amount of matrix cracks. The distribution of matrix cracks in the thickness direction tends to be cylindrical, rather than conical, in shape. From Figures 5 and 6 it can be seen that the large delaminations primarily have developed between layers with a difference angle of 90 , i.e., at 0 /90 and þ45 /45 interfaces. This is consistent with what have been found in prepreg based laminates [21]. This difference in angle of 90 between two adjacent layers is only found within fabrics, i.e., between stitched layers, in the laminates. This result implies that the stitching thread used in the NCF does not prevent delaminations from developing within the fabric.

502

F. EDGREN ET AL. 1.40 z [mm] 1.20 1.00 0.80 0.60

90˚-direction 0˚-direction

0.40 0.20

Length coordinate [mm]

0.00 −40

−30

−20

−10

0

10

20

30

40

(a) 0.60 z [mm] 0.50 0.40 0.30 90˚-direction 0˚-direction

0.20 0.10

Length coordinate [mm] 0.00 −40

−30

−20

−10

0

10

20

30

40

(b) Figure 4. Measurements of remaining dent on surface of: (a) ABVID panel and (b) BBVID panel.

Table 4. In-plane dimensions of remaining dent on laminate surface (standard deviations within brackets). Panel Type ABVID BBVID

Dent Diameter (mm)

Dent Depth (mm)

28 (0.8) 16 (3.3)

0.8 (0.1) 0.4 (0.04)

Results from Compression Tests The results from the CAI tests reveal that there are small differences in strain to failure between panels exposed to different impact energies for a specified panel type, see Table 5. Thus, the CAI strength depends very little on the impact energy levels used in this study, and hence, the amount of fibre fracture in the impact damaged zone. The data presented for each panel type in Table 5 are mean values from the three CAI tested panels. The strain levels are the average of strain measured on the front (impacted) side of the panels. Each

503


Impact point

0˚ 90˚ 45˚ −45˚ −45˚ 45˚ 90˚ 0˚

−20 −15 −10

−5

0

5

10

15

90°-direction (mm)

20

Figure 5. Panel ABVID, damage from 30 J impact.

Impact point

0˚

90˚ 45˚ −45˚ −45˚ 45˚ 90˚ 0˚ 0˚ 90˚ 45˚ −45˚ −45˚ 45˚ 90˚ 0˚ 0˚ 90˚ 45˚ −45˚ −45˚ 45˚ 90˚ 0˚

−20 −15 −10 −5 0

5

10 15 20

Figure 6. Panel BBVID, damage from 100 J impact.

Table 5. Strain to failure for compression loaded panels (standard deviations within brackets). Panel Type ABVID AVID BBVID BVID BHOLE BNOTCH AUNNOTCHED

Failure Strain (%) 0.58 (0.02) 0.60 (0.07) 0.48 (0.01) 0.49 (0.04) 0.49 (–) 0.42 (–) 0.90 (0.05)

90˚-direction (mm)

504

F. EDGREN ET AL.

panel was instrumented with two strain gauges on the front side and the back side respectively. The recorded load–strain curves indicated that there were small bending effects in the type A panels. The type B panels generally had larger bending effects. The reason for this behaviour is probably that the upper and lower surfaces of the panels, where the load are introduced, were not perfectly parallel. The stiffness of the two skins of each panel were, however, the same, indicating that the stiffness reduction due to the damage in the impacted skins is negligible. An example of the load–strain plots can be seen in Figure 7. From the group A90BVID only one panel was tested in compression until failure. The strain to failure for this panel was 0.47%. This value is not considered representative due to the fact that only one panel was tested. Compression tests were also performed on notched and holed panels. The strain to failure for these panels is also presented in Table 5. In addition, the table presents results from compression tests of unnotched monolithic laminates, of the same type as in the face sheets of the type A sandwich panels. The tests were performed according to ASTM Standard D 3410-95 [22]. The strain to failure values in Table 5 for holed/notched panels and CAI panels are not directly comparable because of the different panel dimensions and the dimensions of the hole/notch and impact damage. The main purpose with the holed tests were to compare damage initiation and propagation mechanisms with those observed in CAI tests. The out-of-plane displacement of the skin at the point of impact was measured on the CAI tested panels during the compressive loading. On all panels the skin tended to move inwards, towards the core material, with increased load. However, immediately prior to failure the behaviour differed between the type A and type B panels. The skin of the type A panels continued to move towards the core when reaching the maximum load (see Figure 8(a)). The skin of the type B panels, however, showed a tendency to move outwards when reaching the maximum load (see Figure 8(b)). No evidence of in-plane dent growth was observed in the panels. The described behaviour of the two panel types was consistent for both BVID and VID. These results imply that delaminations in the impacted region of type A panels did not open (out-of-plane) during compression loading. Delaminations in the type B panels may however have opened prior to failure. To facilitate delamination growth in the direction perpendicular to the loading direction in this type of panels, opening of delaminations is required [3]. Consequently, this result implies that delamination growth is not the mechanism controlling failure in the Type A panels.

Load [kN] 250.0

Load [kN] 600.0

200.0

500.0 400.0

150.0 Front side Back side

100.0 50.0 0.0 0.00

Strain [%] 0.10

0.20

0.30

(a)

0.40

0.50

0.60

300.0

Front side

200.0

Back side

100.0 0.0 0.00

Strain [%] 0.10

0.20

0.30

0.40

(b)

Figure 7. Typical load–strain plot for: (a) A ABVID panel and (b) A BBVID panel.

0.50

0.60

505


Displacement [mm] 0.2 0.1 0.0 −0.1 0.0 50.0 −0.2 −0.3 −0.4 −0.5 −0.6 −0.7 −0.8

Load [kN] 100.0

150.0

200.0

250.0

(a) Displacement [mm] 0.00 0.0 100.0 −0.05

Load [kN] 200.0

300.0

400.0

500.0

600.0

−0.10 −0.15 −0.20 −0.25 −0.30

(b) Figure 8. Out-of-plane displacement vs. applied load for: (a) a ABVID panel and (b) a BBVID panel. Positive value corresponds to movement outwards, away from the core material.

However, the results indicate that delamination growth in the Type B panels can not be excluded at this point. This will be discussed further in the section to follow.

Identification of Mechanisms Governing Failure The main objective of this study is to identify the mechanisms that control failure of this type of impact damaged sandwich panels. All panels loaded until failure in this investigation failed by compression failure of the impacted skin. Failure occurred along a line through the impact damage, perpendicular to the load direction. Fractographic studies of the failed panels revealed that fibre microbuckling was the mechanism controlling failure of the skins and hence the sandwich panels. In these studies, no fractographic evidence of delamination growth in the Type A nor the Type B panels was found. No evidence of delamination growth could be observed with neither ultrasonic

506

F. EDGREN ET AL.

C-scan nor sectioning. Delamination growth was therefore concluded not to govern failure of the panels. The holed and notched panels failed in the same manner as the impact damaged panels, i.e., by development and growth of fibre microbuckle kink bands. There are two questions to be answered at this point. Where do kink band formation occur and how do they propagate after initiation? KINK BAND FORMATION IN HOLED/NOTCHED PANELS The loading of one panel from each group was interrupted prior to failure, as described above. The loading of the holed and notched panels was interrupted at approximately 90–95% of the failure load. The fractographic studies of the holed and notched panels showed that fibre microbuckling initiated at the edge of the hole or notch. Regions with kink bands were found in both panel types. Since these particular panels were not loaded until failure, the kink bands found have propagated in a stable manner. In the notched panel, with the skin lay-up [0/90/45/45]S3, kink bands were found in all four 0 -plies of the laminate. The kink bands in the two middle 0 -plies have propagated approximately 4 mm. A sketch of the kink bands found in the notched panel is presented in Figure 9. It should be noted that all of these kink bands were located in the same inplane position. The fractographic studies also revealed that the extension of the kink band zone in the loading direction was larger closer to the notch edge than at the crack tip. At the notch edge several kink bands at different angles to the laminate midplane normal was connected to each other, forming a wide damage zone. Closer to the tip of the damage, away from the notch edge, the kink band zone was smaller and formed by one single kink band. This behaviour is illustrated in Figure 10. This phenomenon is consistent with what has been described by Fleck [5] for prepreg based composites. Figure 10 also shows that the fibres have buckled out-of-plane. In fact, all kink bands found in the investigated panels had buckled out-of-plane. Stable growth of kink bands was found also in the holed panel. In this panel, kink bands were found only in the 0 -ply at the outer surface. The length of the damaged zone was approximately 1 mm. A sketch of the damage found in the holed panel is presented in Figure 11. KINK BAND FORMATION IN CAI TESTED PANELS Loading of the impact damaged panels was interrupted at approximately 80–90% of the failure load. After testing, these panels were subjected to fractographic investigations. Kinkband

[0/90/45/-45]s3

Notch Notch edge Core 0

A

A

A-A

1

2

3

4

5

[mm]

Distance from notch edge

Figure 9. Sketch of kink bands found in notched panel after interrupted loading.


507

0º-direction

(a)

(b) Figure 10. Micrographs of kink bands found in notched panel BNOTCH : (a) Section 1.8 mm from notch edge; (b) Section 3.6 mm from notch edge (interrupted loading).

In these panels, several kink bands were found in the region close to the impact damage. The length of the kink band zones varied from a few mm to 22 mm. This extension of individual kink bands is larger than what is usually found in prepreg based laminates [17]. More importantly, these individual kink bands are significantly longer than those found in the holed/notched panels, see above. Kink bands were primarily found in plies with the fibre direction parallel to the load direction. Some kink bands were, however, found also in the 45 plies. In general, the area (in the x–y plane) of the region, experiencing kink band formation in a CAI test is significantly larger than that in the holed or notched panels.

508

F. EDGREN ET AL. Kinkband

[0/90/45/-45]s3

Hole

2

Hole edge

1

Core 0

A

A

1

2

A-A

3

4

5

[mm]

Distance from hole edge

Figure 11. Sketch of kink band found in holed panel after interrupted loading.

10 mm = Impact point (a)

(b)

(c) Load direction

(d)

(e)

Figure 12. Projection of damage: (a) ABVID; (b) AVID; (c) A90BVID; (d) BBVID; and (e) BVID.

The kink bands were found to initiate typically 5–15 mm from the point of impact. A projection of the kink bands found in each panel is presented in Figure 12. The type A panels have been investigated on one side of the impact point only. The effect of the delaminations on the position of kink band initiation was studied on an A90BVID panel. The projection of the impact-induced delaminations on a type A panel is elliptical in shape (see Figure 2). By loading this panel in the 90 -direction, with the major axis of the ellipse perpendicular to the loading direction, it was possible to study the influence of one single delamination on the position of the kink band initiation. Figure 12(c) shows that kink bands have initiated in the region close to the point of impact. From this result it can be concluded that a single delamination, extending outside the central part of the impact damage, does not promote kink band initiation as much as does the damage closer to the point of impact. An example of the kink band distribution in the thickness direction of a BBVID panel is presented in Figure 13. It can be seen that kink bands have developed in the 0 -layers and in some of the 45 -plies. In two of the 45 -plies, ply 11 and 19, the kink bands have propagated in the 90 -direction. Microscopy revealed that these kink bands have propagated step-wise. Within each growth increment, the kink band has grown in a direction almost perpendicular to the fibre direction, see Figure 14 (note the step-wise

509


Figure 13. Kink bands found in interrupted compression loaded BBVID panel. Ply 1 is the surface layer.

Fibre direction

Kink bands

0º 90º Figure 14. Kink band formation in a 45 -layer.

growth also within the fibre bundles). In Figure 13 it can be seen that the kink bands developed in the 45 -plies, ply 3, 11 and 19, are located almost in the same in-plane position with respect to the point of impact. This implies that there is an effect that promotes kink band development in that in-plane position. This effect is thus larger and has a stronger influence on the initiation position than the effect of fibre waviness. It should also be noticed that no kink bands have formed in the 0 -plies between the damaged 45 -plies at this position. The fractographic studies show that the fibres have primarily buckled out-of-plane. There are, however, indications of in-plane shear stresses affecting the kink band formation in the 45 -plies. In Figure 14 transverse cracks, in the fibre direction, can be seen between each kink band. These cracks develop prior to the kink bands and indicate the presence of in-plane shear stresses [23]. In Figure 15 the distribution of kink bands in an interrupted compression loaded BVID panel is presented. A 21-mm long kink band has formed in ply 11 (45 -ply, lower left corner). Another kink band has formed in plies 8 and 9 (0 -plies), in the same in-plane

510

F. EDGREN ET AL.

Figure 15. Kink bands found in interrupted compression loaded BVID panel. Ply 1 is the surface layer.

position as the kink band found in ply 11. However, the 0 -ply kink band has propagated less than the 45 -ply kink band. In this study it has not been possible to determine at which end a certain kink band has initiated. It may however be reasonable to assume that the kink bands initiate at the point closest to the impact point and propagate away from it. The distance of stable kink band propagation found in the panels described in this study obviously depends on the level at which the loading is interrupted. It is not possible, by use of the method described in this paper, to find the maximum length of stable growth. It is, however, possible to verify stable growth and to get an estimation of the magnitude of the stable propagation length.

Concerns for Modelling The results from this investigation show that compression loaded impact damaged face sheet laminates, of the type studied here, fail by plastic fibre microbuckling. Kink bands develop at several positions and in different plies before failure. The length of the kink bands developed prior to unstable crack growth and final failure is substantial. To the authors knowledge this behaviour has not been found in prepreg based laminates. This behaviour is probably connected with the heterogeneous structure of the NCF reinforcement, with inherent in-plane and out-of-plane fibre waviness. In the holed panels kink bands initiated and propagated at load levels only slightly lower than the failure loads. Failure occurred at the positions of maximum compressive stresses, points 1 and 2 in Figure 11. In the impact damaged panels, kink bands developed already at approximately 50–70% of the failure load. Also, kink bands initiated and propagated within a wide region and were not concentrated to a narrow zone, as for the holed panels. Several kink bands were found in the regions were the shear stresses around a circular shaped impact damage could be expected to have their maximum, see Figure 12. It was also found that the effect of one single delamination, extending outside the central part of the impact damage, on kink band initiation position is negligible compared to the effect of the damage close to the point of impact. Models based on linear elastic fracture mechanics have been developed that successfully predict the behaviour of compression loaded laminates with a hole [16]. These models have been successfully used for predictions of the CAI strength of prepreg based composite


511

laminates [6,17]. The impact damage in these applications has been modelled as an equivalent hole. The fractographic results from this study, however, reveals that failure of CAI loaded NCF face sheet laminates do not fail in the same manner as compression loaded holed laminates. This result indicates that failure initiates not only due to high compressive stresses, as the case with a hole, but are also affected by the shear stresses. Thus, the assumption to replace the impact damage in these materials with an open hole must be questioned. The impacted panels thus have a more ‘‘ductile’’ behaviour than holed panels when loaded in compression. Damage development and growth was observed at relatively low load levels. Development of damage at low stress levels could prove fatal when a structure is subjected to, for example, fatigue. For efficient use of this type of material, there is a need for accurate predictions of the damage development. To achieve accurate predictions of the stress level and position at initiation and propagation of damage, knowledge of the stress field in the region of interest is required. In addition, knowledge of the statistical variations of the structure of the NCF reinforcement is also required. For a holed or notched panel, the stress concentration due to the presence of a hole or notch is responsible for the initiation of the kink bands at the hole/notch edge. In these panels, there is no deteriorated material or geometric out-of-plane imperfection caused by impact. In impact damaged panels, however, several mechanisms can be responsible for the initiation of kink bands. The deteriorated material, through fibre fracture, matrix cracks and delaminations, in the impacted region will cause stress redistribution and stress concentrations that may be sufficient for a kink band to form. Furthermore, there is a remaining dent in the skin of the impacted panels. When the panel is loaded in compression, this geometric imperfection causes local bending effects in the laminate and accompanying stresses. The question a modeller must answer is – which damage feature or combination of features affects the stress state to trigger kink band formation? From this study, knowledge on the effect of individual damage features in the impacted region on the initiation of kink bands has been gained. The following conclusions can be drawn regarding the importance of the individual damage features on the stress field and subsequent kink band development: . The stress-field is affected by the presence of the remaining dent in the face sheet laminate. Compressive loading of a laminate with a dent causes bending effects in the laminate and accompanying stresses. Out-of-plane kink band formation in regions with anticipated effects from local bending, out-of-plane shear and flexural stresses, suggests that the remaining dent must be considered in CAI strength predictions for this type of panels. . Local bending effects may also occur as a result from material eccentricities. Such eccentricities may be caused by multiple delaminations in combination with distributed matrix cracks. For this reason, the importance of material eccentricities on the stress field must be investigated further and cannot be excluded at this point. . The amount of fibre fracture does not affect the damage development. Kink band formation appears identical in panels with and without fractured fibres (see Figure 15). Strain to failure is marginally affected by the presence of fractured fibres when comparing test results from panels with BVID (no fractured fibres) and VID (extensive fibre fracture), see Table 5. Thus, for this type of panel, detailed modelling of fibre fractures may not be required for CAI strength predictions.

512

F. EDGREN ET AL.

. Effects of matrix cracks have not been evaluated. However, in-plane stiffness reduction caused by matrix cracking is significantly less than that caused by fractured fibres. Hence, it is reasonable to assume that matrix cracks need not be considered in CAI strength predictions for this type of panel. . The effect of a single delamination, extending outside the central part of the impact damage, on kink band position was found negligible. Detailed analysis of single delaminations extending far away from the impact point need not to be considered in CAI strength models for these panels.

To what extent the remaining dent and the eccentricity of the impact damage, caused by multiple delaminations and matrix cracks, affects the stress-field must be investigated further.

CONCLUSIONS In this study, NCF composite sandwich panels with damage have been studied under uniaxial in-plane compression. Compression after impact (CAI) tests revealed that the difference in strain to failure for panels with barely visible impact damage (BVID) and visible impact damage (VID) is very small. Fractographic investigations showed that compressive failure of the studied sandwich panels was governed by plastic fibre microbuckling in the impacted face sheet laminate, where fibre microbuckling caused formation and propagation of kink bands. The investigations showed no evidence of delamination growth during compression loading prior to failure. Formation of kink bands in the region close to the point of impact was found to initiate already at approximately 50–70% of the failure load. Kink bands developed and propagated mainly in the 0-plies. Several kink bands were, however, also found in 45-plies. Kink bands were found to develop within a wide region close to the point of impact, i.e., in the vicinity of the highly degraded material. In particular, kink bands were found to initiate in regions with anticipated high shear stresses. It was also observed that the kink band initiation position was not affected by the presence of one single delamination, extending outside the central part of the impact damage. Fractographic studies of panels for which the loading was interrupted prior to failure revealed stable propagation of kink bands up to a length of 22 mm. The failure process for impacted panels was compared to that for holed and notched panels of the same type. In contrast to impact damaged panels, studies on holed and notched panels showed that kink bands developed within a very narrow zone, at the positions of maximum compressive stress. The length of the kink bands developed prior to failure was significantly shorter than what was found in the CAI tested panels, typically 1–4 mm. The difference in damage development between holed/notched and impacted panels implies that the local stress field, and hence, the failure process is affected differently by the impact damage than by a hole. CAI strength predictions based on the assumption that the damaged region may be replaced by a hole are therefore not valid for this type of materials and damage geometry. A complex stress field is created in the vicinity of the impact damage due to the deteriorated elastic properties and the remaining dent. These features of the impact damage cause stress concentrations, bending effects and both inplane and out-of-plane shear stresses in the region close to the point of impact. These effects and the variations of the NCF structure are likely the main reasons for the


513

behaviour found in the studied panels. Thus, a physically sound CAI strength model requires accurate predictions of the stress field in the region of interest and hence need to consider all of the above-mentioned effects. For this purpose, the effect of the remaining dent appears to be most important. ACKNOWLEDGEMENTS This work was part of the SaNDI (Thales JP3.23) project. The support of FMV – The Defence Material Administration of Sweden is gratefully acknowledged. The authors would also like to thank Prof. D. Zenkert for valuable discussions. Kockums AB Karlskronavarvet has supplied the sandwich panels and So¨ren Nilsson at the Aeronautical Research Institute of Sweden conducted the ultrasonic C-scans. REFERENCES 1. Chai, H., Babcock, C.D. and Knauss, W.G. (1981). One-dimensional Modelling of Failure in Laminated Plates by Delamination Buckling, Int. J. Solids Struct., 7: 1069–1083. 2. Whitcomb, J.D. (1981). Finite Element Analysis of Instability Related Delamination Growth, Journal of Composite Materials, 15: 403–426. 3. Nilsson, K-F., Asp, L.E., Alpman, J.E. and Nystedt, L. (2001). Delamination Buckling and Growth for Delaminations at Different Depths in a Slender Composite Panel, Int. J. of Solids Struct., 38: 3039–3073. 4. Asp, L.E. and Nilsson, K.-F. (2002). Delamination Criticality in Slender Compression-loaded Composite Panels, Key Engineering Materials, Trans Tech Publications, Switzerland, 221–222: 3–16. 5. Fleck, N.A. (1997). Compressive Failure of Fiber Composites, Advances in Applied Mechanics, pp. 43–117, Academic Press, New York. 6. Soutis, C. and Curtis, P.T. (1996). Prediction of the Post-impact Compressive Strength of CFRP Laminated Composites, Compos. Sci. Technol., 56: 677–684. 7. Zhuk, Y., Guz, I. and Soutis, C. (2001). Compressive Behaviour of Thin-skin Stiffened Composite Panels with a Stress Raiser, Composites Part B, 32: 697–709. 8. Nyman, T. (1999). Fatigue and Residual Strength of Composite Aircraft Structures, Doctoral Thesis, Report 99-26, Royal Institute of Technology, Stockholm, Sweden, ISSN0280-4646. 9. Jelf, P.M. and Fleck, N.A. (1992). Compression Failure Mechanisms in Unidirectional Composites, Journal of Composite Materials, 26(18): 2707–2726. 10. Argon, A.S. Fracture of Composites. In: Treatise of Materials Science and Technology, Vol. 1, Academic Press, New York. 11. Budiansky, B. (1983). Micromechanics, Computers and Structures, 16(1–4): 3–12. 12. Fleck, N.A. and Budiansky, B. (1990). Compressive Failure of Fibre Composites Due to Microbuckling. In: Dvorak, G.J. (ed.), Inelastic Deformation of Composite Materials, pp. 235–274, Springer, New York. 13. Soutis, C., Curtis, P.T. and Fleck, N.A. (1993). Compressive Failure of Notched Carbon Fibre Composites, Proc. R. Soc. Lond. A, 444: 241–256. 14. Nuismer, R.J. and Labor, J.D. (1979). Application of the Average Stress Failure Criterion: Part II: Compression, Journal of Composite Materials, 13: 49–60. 15. Rhodes, M.D. and Mikulas, M.M. and McGowan, P.E. (1984). Effects of Orthotropy and Width on the Compression Strength of Graphite/Epoxy Panels with Holes, AIAA J., 22: 1283–1292. 16. Soutis, C., Fleck, N.A. and Smith, P.A. (1984). Failure Prediction Technique for Compression Loaded Carbon Fibre-epoxy Laminate with a Single Hole, Journal of Composite Materials, 24: 1476–1498.

514

F. EDGREN ET AL.

17. Soutis, C., Smith, F.C. and Matthews, F.L. (2000). Predicting the Compressive Engineering Performance of Carbon Fibre-reinforced Plastics, Composites Part A, 31: 531–536. 18. Lekhnitskii, S.G. (1968). In: Tsai, S.W. and Cheron, T. (eds), Anisotropic Plates, 2 edn, Gordon and Breach Science Publishers, New York. 19. Abrate, S. (1991). Impact on Laminated Composite Materials, Applied Mechanical Review, 44(4): 155–190. 20. Abrate S. (1994). Impact on Laminated Composite Materials: Recent Advances, Applied Mechanical Review, 47(11): 517–544. 21. Sjo¨gren, A. (1999). Fractographic Characterization of Impact Damage in Carbon Fiber/Epoxy Laminates, FFA TN 1999-17, Bromma. 22. ASTM Standard D 3410-95. (1999). Standard Test Method for Compressive Properties of Polymer Matrix Composite Materials with Unsupported Gage Section by Shear Loading. 23. Greenhalgh, E. (2003). Private Communication.