Damage characteristics and residual In-plane compressive strength ...

By Willie Roberts,2014-02-10 03:07
12 views 0
Damage characteristics and residual In-plane compressive strength ...

    Impact Damage and Energy-absorbing Characteristics and Residual In-plane

    Compressive Strength of Honeycomb Sandwich Panels

     ; and M.D. Hill G. Zhou

     * Department of Aeronautical and Automotive Engineering, Loughborough University,

    Loughborough, Leicestershire, LE11 3TU, UK

     Composites Research Centre, GKN Aerospace Engineering Services, Isle of Wight

    PO32 6LR, UK

    ABSTRACT: An experimental study of the in-plane compressive behaviour of both aluminium and nomex composite sandwich panels with 8 ply carbon/epoxy skins was conducted. All sandwich panels were impact-damaged with a range of impact energies from 1 J to 55 J. Dominant damage mechanisms were found to be core crushing, skin delamination and fracture with the former two absorbing most of the impact energy. While the intact panels failed in region close to one loaded end, all the impact-damaged nomex panels failed around the mid-section region. Two thirds of the aluminium panels also failed in the mid-section region and one third failed in the loaded end region. The presence of the core played a unique role in in-plane compression with a substantial stabilising support to the skins, which counteracted the deleterious effect of impact damage. The in-plane compressive behaviour has shown the combined effects of impact damage and the core in a complex manner.

    KEY WORDS: sandwich panel, honeycomb, impact damage, CAI strength, impact damage tolerance.


    Composite sandwich constructions are widely used in aerospace structures due to their light weight, high strength-to-weight and stiffness-to-weight ratios, good buckling resistance, good energy-absorbing capacity and design versatility. A major concern with the mechanical performance of these sandwich structures is their susceptibility to localised manufacturing defect and/or impact damage; the latter could be caused by tool dropping, hail stones, or runway debris. As a result, a multitude of damage mechanisms such as skin-core debonding, core crushing and/or shear failure in addition to skin delamination and fibre fracture could occur. To maintain the aforementioned advantages in the case of any of these impact events or manufacturing defect, both mechanical properties from skins and core and geometric properties must be tailored in a design analysis such that the sandwich structures are damage-tolerant. This requires a thorough understanding of how the damage mechanisms and their energy-absorbing characteristics affect a reduction of the in-plane compressive strength (commonly known as compression-after-impact (CAI) strength) of the damaged sandwich structures and of their damage tolerance performance.

    A great deal of research has been carried out to investigate CAI strength in damaged sandwich structures as reviewed in [1-3], and thus our overview of the past research will be restricted to those for aerospace applications only. Since a standard test method for CAI strengths of sandwich panels, equivalent to the CAI standard ASTM-D 7137/D 7137M for monolithic

    laminate plates [4], is yet to be developed, over the years very different approaches of evaluating CAI strengths have been employed, involving 4-point beam bending [5-6], wide coupon compression [7], wide column compression [8-16] and panel in-plane compression [17-

     ; Author to whom any correspondence should be addressed.


    20]. Clearly most of research works adopted the wide column compression approach, in which the unloaded edges of sandwich specimens were left free. This approach was relatively simple and often produced reasonably well defined failure mechanisms in the test specimens. The employment of the panel in-plane compression approach, of present interest, was relatively recent. With the unloaded edges of sandwich specimens being simply supported, their compressive behaviour was much more complex. Although the compressive failure mechanisms induced by these two approaches might look similar in some way, the respective contributions of the skins and core to the reduction of CAI strengths could be very different. At present, there is little understanding of these differences, which is vitally important to the assessment of impact damage tolerance. McGowan and Ambur [17] investigated impact damage using an impactor of 12.7 mm diameter and CAI strength of graphite/epoxy skinned Korex core sandwich panels with the dimensions of 254 mm by 127 mm. They found that the global compressive response of the panels was not affected by the presence of impact damage with an area of less than 8 mm in diameter. However, the residual compressive loads measured showed interestingly a steady reduction with the increase of impact energy. Tomblin and his associates [18-20] conducted a series of impact and CAI tests using various sandwich configurations. They found that an impactor of 76.2 mm diameter resulted in lower CAI strengths than that of 25.4 mm diameter at the given energy level. They also observed that the sequence of failure modes was dependent on the skin stiffness and the through-the-thickness normal properties of the core.

    The present work reports results of an investigation, using the panel in-plane compression approach, into the in-plane compressive behaviour of intact and impact-damaged composite sandwich panels with both aluminium and nomex honeycombs. In early reports [21-22], damage mechanisms induced in both aluminium and nomex honeycomb sandwich panels were identified; the effects of skin thickness, core density and material, indenter nose shape, panel diameter and support condition on the damage characteristics were studied. The energy-absorbing characteristics of the identified damage mechanisms were examined. This paper reports the results of a systematic investigation of how those characteristics affect the in-plane compressive behaviour of impact-damaged composite honeycomb sandwich panels.


    Composite skins were made of unidirectional carbon/epoxy T700/LTM45 prepreg with a

     was selected due nominal ply thickness of 0.128 mm. A symmetric cross-ply lay-up of (0/90)2s

    primarily to strength considerations. Although four skin thicknesses, each with 4 plies, 8 plies, 12 plies, or 16 plies, have been made, the only 8 ply skins were used in fabricating in-plane compression panels. Two types of honeycomb cores used were 5052 aluminium with a density 33of 70 kg/m (4.4-3/16-15) and nomex with a density of 64 kg/m (HRH-3/16-4.0). The core

    depth of both aluminium and nomex honeycombs was 12.7 mm and a nominal panel thickness was 14.7 mm. The nomex honeycomb was dried in an oven overnight at 65?C before being bonded to the skins to remove any moisture absorbed from the atmosphere. Adhesive VTA260 was selected for interfacial bonding.

Skin laminates of 300 × 200 mm were laid up and cured first in an autoclave at 60C under a

    pressure of 0.62 MPa (90 psi) for 18 hours. To aid skin-core adhesion, the cured skin laminates were degreased before bonding. The designated 0? direction of carbon fibres within the skins was aligned with the ribbon direction of honeycomb core. For aluminium panels, each skin was separately bonded to the core in an oven at 80C for 5 hours under a pressure of 0.1 MPa (15

    psi). For nomex panels, skins were bonded to the core in an autoclave under the same conditions. Because of the condensation of nomex cells (a couple of cells around the panel edges), the pieces of nomex core used were slightly larger than the skins so that the condensed


    cells could be trimmed off after bonding. Each sandwich panel was then cut into two nominal 200 mm × 150 mm specimens with the longer side aligned with the direction of compressive loading. Back-to-back strain gauges were bonded on the panel surfaces at selected locations in both the longitudinal and transverse directions (shown in Figure 2(a)) to monitor mean (or membrane) and panel bending strains. These strain data allowed both the local and global behaviour of the panels to be examined.


Drop-weight Low-velocity Impact Tests

    Low-velocity impact tests were carried out on a purpose-built instrumented drop-weight impact rig in Figure 1(a) by using a hemispherical (HS) impactor of 20 mm diameter with a 1.5 kg mass. Desired impact energies were regulated by adjusting drop height and they ranged from 1 J to 55 J in this investigation. Each rectangular carbon/epoxy plate of 200 mm by 150 mm with a circular testing area of 100 mm in diameter was clamped by using a clamping device, as illustrated in Figure 1(b). A pair of photodiodes fixed over a known distance and two timers were used to record respective times that were used to calculate both impact and rebound velocities. This allowed absorbed energies to be calculated directly. For high impact energies, rebound velocities could be slightly underestimated due to the substantial panel deflections, thereby overestimating the absorbed energies. Both impact force and strain gauge responses could be recorded by a Microlink 4000 data acquisition system with a sampling rate of 50 s.

    At each selected impact energy level, three impact tests were conducted with one being diametrically cut up for examination of damage mechanisms and the remaining two being reserved for subsequent in-plane compression test. The absorbed energy could also be estimated via the closed area under load-displacement curves. While surface dent was measured after each impact test, crushed core depth was measured over the cut cross section. All impact test results along with the extents of crushed core and skin delaminations are summarised in Table 1.

In-plane Compression Test

    As part of specimen preparations for compression test, the core at the panel ends intended for compressive load application was removed to a depth of about 5 mm (slightly more than one cell size). Epoxy end pots were cast between the two skins to prevent an end-brooming failure and after curing the two potted ends were machined to parallel. In each in-plane compression test, a panel was placed in a purpose-built support jig, as illustrated in Figure 2(b), such that the L or ribbon direction of the honeycomb core was aligned with the loading direction. The jig provided simple support along the unloaded edges, which were free to move in the width direction during loading. Quasi-static load was applied to the panel at the machined ends via a universal testing machine at speed of 1 mm/min. Although the loaded ends were not clamped, they were effectively close to the clamped condition but without surface clamping pressures. Load, strain and cross-head displacement in all tests were recorded through a data acquisition system at a sampling rate of 0.5 Hz. All tested panels were cut up for a study of damage mechanisms. All in-plane compression results for both intact and impact-damaged panels are summarised in Table 2. To aid an understanding of the in-plane compressive behaviour of sandwich panels, a small number of 16 ply monolithic cross-ply panels were also compression-tested.


    As an assessment of the compressive strength reduction of impacted sandwich panels required quantitative information on damage states with varying severity induced over the range of


    impact energies, the initiation and propagation of those induced damage mechanisms must be characterised in terms of impact energy, absorbed energy and damage extent. As the strain rate sensitivity of the aluminium honeycomb induced between 1 m/s and 6 m/s impact velocity was considered to be very small, carbon/epoxy skin and nomex core were generally strain-rate insensitive within this impact velocity range. As a result, this characterisation effort was significantly facilitated in combination with much more cost-effective quasi-static bending and indentation tests and with systematic microscopic inspections of cut cross sections in [21-22]. Unless expressly stated, the damage characteristics as discussed below were restricted to those associated with sandwich panels with the 8 ply skins, either impacted by the HS impactor or loaded by the HS indenter.

    For the aluminium sandwich panels, the initial damage threshold that was indicated by either a load drop or a significant loss of the slope of a responsive curve occurred at the average force of 0.51?0.03 kN (measured quasi-statically). Such indication could be seen in Figure 3 from two sample load-displacement curves from both quasi-static and impact testing. Under the impact condition, the initial damage threshold was exceeded even at the lowest impact energy of 1.43 J. As the stabilised compressive and crush strengths of the core from manufacturer are usually obtained using a flat-ended (FE) indenter, they could not be used directly to check to see if the threshold was associated with core crushing here. Nevertheless, further diagnostic tests in conjunction with a systematic cross-sectioning of tested specimens revealed that this incipient damage was due to a combination of core crushing and small delaminations in the loaded (or top) skin in the shape of a cone towards the skin-core interface. These two damage mechanisms did not seem to be related. The estimated contact pressures of the HS indenter/impactor ranged from 2.54 MPa at the apex to 2.11 MPa on the annular region of about 5 mm radius [23] and they were greater than the core crush strength of 1.71 MPa but significantly less than the stabilised compressive strength of 4.1 MPa. This discrepancy could be attributed to the fact that the core crushing in the sandwich panels in bending induced by the HS indenter involved substantial honeycomb cell rotations, tearing and oblique crushing in addition to normal plastic folding [21], which was the primary crushing mechanism in the measurement of the crush strength. As impact energy was increased, the damage mechanisms in the sandwich panels were characterised by the continued core crushing and by propagation of the top skin delaminations. A cross-sectional view of impact-damaged aluminium panel (Al 13J 2) is shown in Figure 4. At about 25 J, top skin fracture occurred and the extent of damage in both the skin and core was very significant as shown in Figure 5. The extent of crushed core was the same as the extent of skin delamination at the low impact energies (see Figure 8(a-b)). The gap between the two started to grow when impact energy was greater than 12 J and became greater for the given higher impact energy. The surface dent depth (in Table 1) was also greater with the greater impact energy, as expected. There was no local skin-core interface debonding found in all tests before the ultimate load. The maximum depth of the crushed cells at an impact energy of 25 J reached only about the middle of the core thicknesses and the bottom skin remained intact on all occasions.

    The damage characteristics of the nomex panels were more or less the same as the aluminium ones as a cross-sectional view of impact-damaged nomex panel (Nom 16J 2) shows in Figure 6, though nomex cells were fractured with a lesser degree of cell folding and showed a substantial spring-back upon unloading. Furthermore, the initial flexural rigidity of the nomex sandwich panels (within the elastic range) was significantly lower than those of the aluminium ones due to the fact that the shear properties of the nomex cores were significantly lower. In addition, the local indentation in the nomex panels was found to be very dominant. This is interesting in considering the fact that the impacted nomex panels exhibited much less surface indent with the maximum depth of crushed cells reaching not even the middle of the core thickness from the 23-J impact test, as can be seen in Figure 7.



    The energy absorption characteristics of both aluminium and nomex sandwich panels were established respectively over three regions shown in Figures 8(a-b) and 9(a-b) in terms of damage extent and absorbed energy. In Figure 8(a-b), both extents of skin delamination and crushed core exhibit distinctive transitions at around 5 J and 23 J, respectively, thereby defining those three regions. In Region 1, the minimum impact energy required to induce skin delamination and core crushing can be extrapolated to be around 0.75 J for the current sandwich material systems. The subsequent trends for these two damage mechanisms seem nonlinear, irrespective of the type of core materials. The magnitudes of both extents at any given impact energy are roughly the same and this feature continues up to the middle of Region 2, which approximately corresponded to about 12 J. These characteristics seem to suggest that the respective propagations of core crushing and skin delamination were collectively responsible for absorbing most of the impact energies. Nevertheless, among the two damage mechanisms, impacted skin delaminations and core crushing, the latter was believed to be the major damage mechanism that absorbed impact energy significantly more than the former. As the interlaminar shear (ILS) strength of cross-ply skins in the present sandwich panels is around 62 MPa [22] and the core crush strengths are 4.02 MPa and 3.93 MPa, respectively, it was thus much easier for the impact energy to be dissipated throughout the continued core crushing in the impact direction than driving the local ILS stress level up to above the ILS strength in the annular region, which was substantially away from the impactor in the direction perpendicular to the impact direction.

    From the middle of Region 2 onwards, the respective extents of skin delamination and crushed core start separating, with the former being substantially less than the latter. These diverging trends have continued into Region 3 when fibre fracture in the impacted skin started occurring. While the diverging trends of both damage mechanisms in Region 2 are linear for the aluminium sandwich panels, the core crushing extent for the nomex sandwich panels seems to escalate with an increase in impact energy. This observation seems to suggest that the nomex honeycomb absorbed impact energy through the extra extension of core crushing in the plane of the sandwich rather than through the continuous cell fracturing in the loading direction. In Region 3, while both extents start decreasing for the aluminium sandwich panels and thus indicate localised penetration of the impactor into the sandwich panels, they seem to level off for the nomex sandwich panels. It is noted that the gap between the two extents for the nomex panels is much wider due likely to the lack of the plastic folding capability of nomex in addition to its relatively low flexural rigidity associated with low through-the-thickness shear properties. The maximum crushed core depth at an impact energy of 25 J reached only the middle of the core thickness in both types of panels.

    In Figure 9(a-b), it is interesting to note that the rate of energy absorption in the damaged panels can be seen to be linear and remain constant throughout Regions 1 and 2 for both types of panels. This constant level of energy absorption is about 67% for the aluminium sandwich panels and about 59% for the nomex sandwich panels. This strongly suggests that the greater percentage of energy absorption occurred in the through-the-thickness direction in Region 2. This trend continues up to either an impact energy of about 23 J or Region 3. This finding is in accordance with the early observations about the nonlinear trends and divergence of the two damage extents. At the beginning of Region 3, fibre fracture of the impacted skin set in and the amount of energy absorption went up to over 94%, irrespective of the type of core materials. This just confirms that this final similar level of energy absorption was attributed to multiple ply fractures in the impacted skin, which were identical for both types of panels.


    In Figure 10, the surface dent depths produced by impact in both aluminium and nomex sandwich panels are plotted against impact energy. The BVID criterion from a perspective of maintenance inspections requires that an 8 J impact produces a dent depth of no more than 0.5 mm, which is regarded as ‘barely visible’ under normal lighting condition from a distance of about 1.5 m. In the figure, the aluminium sandwich data indicate the BVID energy threshold of only about 3 J. In other words, the 8 J impact would produce a surface dent depth of around 1.3 3 core. On the contrary, the nomex mm in the current panels with 8 ply skins and 70 kg/m

    sandwich panels seem to meet the BVID criterion with ease with the BVID energy threshold of being about 25 J. This is of course deceptive, as the extents of core crushing and skin delamination in both types of sandwich panels were similar up to a level of 10 J. Thus such low dent depths in the nomex sandwich panels were due primarily to the nomex core’s capability of

    storing part of the impact energy elastically, which were released during unloading in the form of spring-back.


In-plane Compressive Behaviour

    In addition to the damage characteristics as discussed above, two other factors were considered also to contribute to the complexity of the in-plane compressive behaviour of the damaged sandwich panels. One was that the two skins during in-plane compression were locally stabilised and supported, to some extent, by the core in the through-the-thickness direction. Unless it was very extensive with respect to panel width, the impact damage alone may not have been sufficient to weaken the local stability of the skins in the mid-section region. As a result, the stabilising and supporting effect of the core could significantly reduce the effect of the impact damage. This could particularly be the case if the shear rigidity and compressive strength of the core was relatively high for the given flexural rigidity of the skins and if the impact damage was relatively small. Thus, much of the in-plane compressive characteristics established from monolithic composite panels [24-25] could not necessarily be the same as those from the sandwich panels. The other was that the presence of the impact damage resulted in through-the-thickness local asymmetry with respect to the compression loading and supporting conditions. Therefore, the in-plane compressive strength reduction of the damaged sandwich panels was attributed not only to the effect of impact damage but also to asymmetry and the relative magnitude of the shear rigidity and compressive strength of the core.

    For the intact sandwich panels with no damage and thereby with no local asymmetry, the first aforementioned factor could have much greater influence over the compressive behaviour. Consequently, the likelihood of failure at one of panel ends increased significantly. From the compressive responses of the intact aluminium sandwich panels as examined in [26], a tilting of one panel end could be seen at the early stage of loading via the bending strain response of the far-field strain gauges (SG1) as shown in Figure 11. As the local strain gauges (SG2) on the mid-section exhibited very small bending strain with large mean strain in the figure, these responses provided no indication of local buckling and thus the panel failed in the region close to the loaded end, as shown in Figure 12. The compressive response of the intact nomex panels was very similar in terms of both strain response and failure characteristic.

    Once impact damaged, about two thirds of the aluminium sandwich panels failed around the mid-section region, as one example from specimen ‘Al 13J 2’ shows in Figure 13, and all the

    impact damaged nomex sandwich panels failed in the same way. The remaining third of the impacted aluminium sandwich panels failed in a similar way to the control panels with the failure being close to one loaded end. One of such panels’ strain responses are shown in Figure

    14, which can be seen to be similar to those of the intact panel in Figure 11. Moreover, this


    phenomenon did not appear to be affected by the severity of the impact damage, as the impact energies applied to those four panels spread from 3.7 J to 25 J. In the last case with the impact energy of 25 J, the respective extents of delamination in the impacted skin and core crushing were close to a half of the panel width. A further examination of the aluminium sandwich panels revealed that there was little difference in CAI strength between those with the mid-section failure and those with the end failure. Nevertheless, the majority of the impact damaged nomex sandwich panels not only failed around the mid-section region but showed also classic strain reversal, as one exhibits in Figure 15, which immediately triggered catastrophic failure. As the aluminium core had substantially greater through-the-thickness shear and compressive properties than the nomex, these differences increased the probability of the aluminium sandwich panels failing at one of panels ends for the given impact damage.

Contribution of Honeycomb Core

    The early discussion of the in-plane compression test results clearly suggests that the presence of the core between the two skins during in-plane compression played a complicated role in the compressive behaviour of the sandwich panels, irrespective of whether or not they were impact damaged. The intact sandwich panels had the average panel width-to-thickness ratio of 10 and the significant longitudinal flexural rigidity. With the core providing a stabilising support to the skins in the through-the-thickness direction through both transverse shear and normal compression properties, thus both skins of the panels in the mid-section region were constrained from deforming and a tendency of local buckling in the mid-section region was very small. Moreover, with the two skins being apart, the effective longitudinal load transmission was limited. As a result, all the intact panels failed at the location of being close to one of the panel ends.

    For the impact damaged panels, in some cases, their in-plane compressive resistance was not weakened sufficiently by even the significant impact damage in terms of damage extent (e.g. ‘Al 9.5J 2’ and ‘Al 25J 2’) so that those panels did not fail respectively in the impact damaged (or mid-section) region. In those cases, the impact damaged skin was able to provide a level of the compression resistance similar to the intact skin as in the case of the intact sandwich panels. Although the imbalance of local bending moments induced by the shift of the mid-plane towards the intact skin in the mid-section region could facilitate local buckling, the stabilising effect of the core was greater so that the impact damage in the impacted skin could neither propagate transversely nor facilitate local buckling.

    For the remaining impacted damaged panels which failed in the mid-section region, it was their stabilised compressive strengths that affected the final failure characteristics of the compressive behaviour. As the adhesive’s tensile strength of 5.5 MPa was substantially greater than the

    stabilised compressive strength of 4.2 MPa for aluminium (and 4.0 MPa for nomex), the undamaged (distal) skin could have only bent inwards, thereby crushing the core, as the photographs in Figure 13 show that the zone of the core in the aluminium panels crushed during in-plane compression was close to the distal skin. As for the damaged skin in the dished shape, whether it compressed the already crushed core further (see Figure 13) or buckled outward as shown in Figure 16 depended on, among others, a state of local interfacial bonding, dent depth and surface curvature. Although both inward and outward local skin buckling were observed, the latter could be in favour if local debond around the curved region occurred during impact. This seems to suggest that the denser core and/or core with greater transverse shear stiffness provides the sandwich panels with stronger compressive resistance, which seems in agreement with other results in [27]. The only additional feature of the nomex sandwich panels was that the spring-back of the fractured cells led to a lesser degree of local dishing in the impacted area and thereby the less dent depths (see Figures 6 and 7), as the fractured cells were close to the


    mid-plane of the panels. Nevertheless, this lesser local curvature did not seem to affect the aforementioned characteristics.


    The impact damage tolerance of both aluminium and nomex panels can be assessed by plotting the variation of either CAI strengths or residual compressive (mean) strains against either damage extent (in the width direction) or impact energy in Figures 17-19. The actual reduction of their CAI strengths is seen more clearly when CAI strengths were normalised with the baseline values of the sandwich panels in Figure 18. Although a damage area could also be used as a damage measure [24-25], the damage extent is an equally useful indicator as demonstrated here.

    In Figure 17, once the panels were impact-damaged, there seems an immediate but moderate reduction in compressive strength. From 3.7 J to 13 J on the impact energy axis, however, the variation of CAI strengths is limited, as a state of damage in these panels was dominated by a steady growth of core crushing and skin delaminations. Although the growth of the extents exceeded over 50% of the panel width, they seemed insufficient to trigger local buckling and/or sideway propagation of skin delaminations due likely to the aforementioned contribution of the core for stabilising the mid-section region. This seems to suggest that much of the effect of impact damage was cancelled out’ so that it could not lead to a further reduction of CAI

    strengths. At the BVID threshold of 8 J, the reduction of CAI strengths was unfortunately substantial. When the impact energy level reached over 20 J, a further reduction of CAI strengths became noticeable because of the fracture of the impacted skins. At highest impact energies (about 26 J for the aluminium sandwich panels and 28 J for the nomex sandwich panels), the reduction of compressive strength was about 50% for both sandwich panels. In addition, it is noted that the different failure locations among the CAI tested panels are not distinguishable over the values of CAI strength. Nevertheless, this does imply that the relative magnitude of the shear rigidity of the aluminium honeycomb was at the critical state such that a reduction of the shear properties could lead to the consistent mid-section failure, or vice versa.

    The normalised CAI strengths results are presented in Figure 18 with compressive strength retention factor plotted against damage extent. The data trends are very similar to those observed from Figure 17. The residual compressive mean strains obtained from the far-field strain gauge locations can also be used in the impact damage tolerance assessment in Figure 19, with the advantage for the fact that allowable strains are often used in a composite structure design. Although the overall trends here are similar to those in Figure 17, the values of strain loss seem to be slightly worse than CAI strengths.

    Although the trends of data in the form of either CAI strengths or residual compressive strains are similar when plotted against impact energy in Figure 17, the fact that a 75% increase in the extent of skin delaminations did not lead to a further reduction in CAI strengths seem to indicate another possibility. That is, the impact energy may not be an effective damage measure, as the majority of the impact energies were actually absorbed by a crushing of the honeycomb cores, while the skins that absorbed less energy provided the major compression resistance. Unlike in damaged monolithic laminates, both impact energy and energy absorption were intimately related to the creation of damage or fracture surfaces, which in turn weakened the laminates in subsequent in-plane compression, so that their variation corresponded to the variation of CAI strengths. Therefore, they have proven to be effective damage measures along with damage area [24-25]. However, sandwich panels had a role-sharing multi-functioning characteristic and posed a significant challenge. Unless the through-the-thickness shear and normal properties of the core were low, they could significantly affect the in-plane compressive


    behaviour of the sandwich panels as in the present case. Thus, a further investigation over the role of the core in the in-plane compression in the future is essential to the impact damage tolerance assessment of sandwich panels.


    Both aluminium and nomex composite sandwich panels were impact damaged with energies ranging from 1 J to 55 J. Dominant damage mechanisms were found to be core crushing, skin delamination and skin fracture with the former two absorbing the most impact energies. As the initial damage occurred at relatively low load, energy absorption prior to the skin fracture was dominated by cell crushing in both in-plane and through-the-thickness directions in the initial region and primarily in the latter direction in the extent growth region. Different core materials with a similar density made little difference on either the damage or energy-absorbing characteristics.

    Both intact and impact damaged composite sandwich panels were tested in in-plane compression. They failed in three different modes because the presence of the core affected the way in which the panels failed. The presence of the core provided the sandwich panels with high flexural rigidity, enhanced shear rigidity and normal compressive strength such that the likelihood of local buckling around the mid-section region in these panels was substantially reduced through local stabilisation and support. Because of this, the intact panels failed in a region being close to one loaded end due to limited load transmission. Two thirds of the impact damaged aluminium panels along with all the damaged nomex panels failed around the mid-section region due to the significant effect of the impact damage. However, one third of the impact damaged aluminium panels failed in the loaded end region similar to the intact panels. This was found because the presence of the core seemed to have counteracted the deleterious effect of the impact damage in such a way that local buckling or sideway propagation of the impact damage in those panels was constrained.


1. Abrate, S. (1997). Localized impact on sandwich structures with laminated facings, Applied

    Mechanics Review, 50: 69-82.

    2. Tomblin, J.S., Lacy, T., Smith, B., Hooper, S., Vizzini, A. and Lee, S. (1999). Review of

    damage tolerance for composite sandwich airframe structures, Report DOT/FAA/AR-99/49,


    3. Hayman, B. (2007). Approaches to damage assessment and damage tolerance for FRP

    sandwich structures, J. of Sandwich Structures and Materials, 9: 571-596.

    4. ––. (2005). Standard test method for compressive strength properties of damaged polymer

    matrix composite plates, ASTM-D 7137/D 7137M 05.

    5. Gottesman, T., Bass, M. and Samuel, A. (1987). Criticality of impact damage in composite thnd ICCM and 2 ECCM, London, pp. 3.27-3.35. sandwich structures, In: Proc. of 6

    6. Charles, J.P. and Guedra-Degeorges, D. (1991). Impact damage tolerance of helicopter rdsandwich structures, In: Proc. of 23 Int. SAMPE Technical Conference, New York.

    7. Akay, M. and Hanna, R. (1990). A comparison of honeycomb-core abd foam-core carbon-

    fibre/epoxy sandwich panels, Composites, 21: 325-331.

    8. Kassapoglou, C., Jonas, P.J. and Abbott, R. (1988). Compressive strength of composite

    sandwich panels after impact damage: an experimental and analytical study, J. of

    Composites Technology and Research, 10: 65-73.

    9. Palm, T.E. (1991). Impact resistance and residual compression strength of composite thsandwich panels, In: Proc. of 8 ICCM, Honolulu, pp. 3-G-1-3-G-14.


    10. Tsang, P.H.W. and Lagace, P.A. (1994). Failure mechanisms of impact-damaged sandwich

    panels under axial compression, AIAA-94-1396, pp. 745-754.

    11. Kassapoglou, C. (1996). Compression strength of composite sandwich structures after

    barely visible impact damage, J. of Composites Technology and Research, 18: 274-284.

    12. McGowan, D.M. and Ambur, D.R. (1997). Damage-tolerance characteristics of composite

    fuselage sandwich structures with thick facesheets, NASA TM 110303, USA. 13. Cvitkovich, M.K. and Jackson, W.C. (1999). Compressive failure mechanisms in composite

    sandwich structures, J. of the American Helicopter Society, Oct.: 260-268.

    14. Moody, R.C., Harris, J.C. and Vizzini, A.J. (1999). Curvature effect on the damage th ICCM, Paris. tolerance iof impacted damaged composite sandwich panels, In: Proc. of 12

    15. Lagace, P.A. and Mamorini, L. (2000). Factors in the compressive strength of composite

    sandwich panels with thin facesheets, J. of Sandwich Structures and Materials, 2: 315-330.

    16. Moody, R.C., Harris, J.C. and Vizzini, A.J. (2002). Scaling and curvature on the damage

    tolerance iof impacted composite sandwich panels, J. of Sandwich Structures and Materials,

    4: 71-82.

    17. McGowan, D.M. and Ambur, D.R. (1998). Damage characteristics and residual strength of

    composite sandwich panels impacted with and without a compression loading, AIAA-98-

    1783, pp. 713-723.

    18. Tomblin, J.S., Raju, K.S., Liew, J. and Smith, B.L. (2001). Impact damage characterization

    and damage tolerance of composite sandwich airframe structures, Report DOT/FAA/AR-

    00/44, USA.

    19. Tomblin, J.S., Raju, K.S., Liew, J. and Smith, B.L. (2002). Impact damage characterization

    and damage tolerance of composite sandwich airframe structures - Phase II, Report

    DOT/FAA/AR-02/80, USA.

    20. Tomblin, J.S., Raju, K.S. and Arosteguy, G. (2004). Damage resistance and tolerance of

    composite sandwich panels - scaling effects, Report DOT/FAA/AR-03/75, USA. 21. Zhou, G., Hill, M.D., Loughlan, J. and Hookham, N. (2006). Damage characteristics of

    composite honeycomb sandwich panels in bending under quasi-static loading, J. of

    Sandwich Structures and Materials, 8: 55-90.

    22. Zhou, G., Hill, M.D. and Hookham, N. (2007). Investigation of parameters governing the

    damage and energy-absorbing characteristics of honeycomb sandwich panels, J. of

    Sandwich Structures and Materials, 9: 309-342.

    23. Zhou, G. (1996). Static behaviour and damage of thick composite laminates, Composite

    Structures, 36: 13-22.

    24. Zhou, G. and Rivera, L. (2005). Investigation for the reduction of in-plane compressive

    strength in preconditioned thin composite panels, J. of Composite Materials, 39: 391-422.

    25. Zhou, G. and Rivera, L. (2007). Investigation for the reduction of in-plane compressive

    strength in preconditioned thick composite panels, J. of Composite Materials, 41: 1961-


    26. Zhou, G. and Hill, M.D. (2007). Impact damage and residual compressive strength of thhoneycomb sandwich panels, In: Proc. of 16 ICCM, Kyoto.

    27. Nokkentved, A., Lundsgaard-Larsen, C. and Berggreen, C. (2005). Non-uniform

    compressive strength of debonded sandwich panels - I. Experimental investigation, J. of

    Sandwich Structures and Materials, 7: 461-482.


Report this document

For any questions or suggestions please email