Numerical evaluation of failure mechanisms on composite specimens subjected to impact loading

Composite panels are in common use, especially in aeronautic and aerospace structures due to their high strength/weight and stiffness/weight ratio. The out-or-plane impact loading is considered potentially dangerous mainly because the damage may be left undetected and because the loading itself acts in the through-the-thickness direction of the laminated composite panel. This direction is the weakest in the composite since no fibres are present in that direction. The impact loading can lead to damage involving three modes of failure: matrix cracking, delamination and eventually fibre: breakage for higher impact energies. Even when no visible impact damage is observed at the surface on the point of impact, matrix cracking and delamination can occur, and the residual strength of the composite is considerably reduced. The objective of this study is to determine the mechanisms of the damage growth of impacted composite laminates when subjected to impact loading. Fur this purpose a series of impact tests were carried out on composite laminates made of carbon fibre reinforced epoxy resin matrix. An instrumented drop-weight-testing was used together with a C-scan ultrasonic device for the damage identification. Two stacking sequences of two different epoxy resins and carbon fibres, representative of four different elastic behaviours with a different number of interfaces were used. A numerical evaluation of these specimens was also carried out, using static analysis only. Results showed that the delaminated area due to impact lending depends on the number of interfaces between plies. Two failure mechanisms due to impact were identified, which are influenced by the stacking sequence and by the thickness of the panels. (C) 2000 Elsevier Science Ltd. All rights reserved.