The mechanics of interface delamination in CFRP composite laminates is examined using finite element method. For this purpose a 12-ply CFRP composite, with a total thickness of 2.4 mm and anti-symmetric ply sequence of [45/-45/45/0/-45/0/0/45/0/-45/45/-45] is simulated under three-point bend test setup. Each unidirectional composite lamina is treated as an equivalent elastic and orthotropic panel. Interface behavior is defined using damage, linear elastic constitutive model and employed to describe the initiation and progression of delamination during flexural loading. Complementary three-point bend test on CFRP composite specimen is performed at crosshead speed of 2 mm/min. The measured load-deflection response at mid-span location compares well with predicted values. Interface delamination accounts for up to 46.7 % reduction in flexural stiffness from the undamaged state. Delamination initiated at the center mid-span region for interfaces in the compressive laminates while edge delamination started in interfaces with tensile flexural stress in the laminates. Anti-symmetric distribution of the delaminated region is derived from the corresponding anti-symmetric ply sequence in the CFRP composite. The dissipation energy for edge delamination is greater than that for internal center delamination. In addition, delamination failure process in CFRP composite can be described by an exponential rate of fracture energy dissipation under monotonic three-point bend loading.

The honeycomb (HC) core of sandwich structures undergoes flexural loading and carries the normal compression and shear. The mechanical properties and deformation response of the core need to be established for the design requirements. In this respect, this article describes the development of the smallest possible representative cell (RC) models for quantifying the deformation and failure process of the Nomex polymer-based hexagonal HC core structure under the out-of-plane quasi-static loadings. While the hexagonal single and multi-cell models are suitable for the tension and compression, a six-cell model is the simplest RC model developed for shear in the transverse and ribbon direction. Hashin’s matrix and fiber damage equations are employed in simulating the failure process of the orthotropic cell walls, using the finite element (FE) analysis. The FE-calculated load–displacement curves are validated with the comparable measured responses throughout the loading to failure. The location of the fracture plane of the critical cell wall in the out-of-plane tension case is well predicted. The wrinkling of the cell walls, leading to the structural buckling of the HC core specimen in the compression test, compares well with the observed failure mechanisms. In addition, the observed localized buckling of the cell wall by the induced compressive stress during the out-of-plane shear in both the transverse and ribbon direction is explained. The mesoscale RC models of the polymer hexagonal HC core structure have adequately demonstrated the ability to predict the mechanics of deformation and the mechanisms of failure.

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Finite element method is increasingly used in the analysis of aircraft structures, including Unmanned Aerial Vehicles (UAVs). The structural model used for finite element analysis however needs to be validated in order to ensure that it correctly represents the physical behaviour of the actual structure. In this work, a case study of a straight, unswept and untapered wing structure made of composite material subjected to aerodynamic loading was modelled and analysed using finite element method. Four-noded, reduced integration shell elements were used, with structural components attached by adhesive joints modelled using tied surface constraints. For the validation process an experimental set-up of the actual wing was loaded using sandbags to simulate the aerodynamic loads. The deflection of the wing at three key locations were obtained and compared between both methods. It was found that the difference between both results ranges between 0.3% (at the tip) to 36.1% (near the root, for small deflections).

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The difficulties presented by inertial disturbances in the testing of fibre reinforced composite materials at elevated strain rates necessitated the development of a technique for the prediction of high strain rate material property data. These data have been used in a finite element analysis to simulate impact behaviour of a random glass epoxy composite. High strain rate properties obtained by extrapolating results of experiments conducted at low to intermediate strain rates were used in the finite element analysis of a simple three-point bend beam impact. Three point bend impact tests were performed on the laminates, and comparisons were made of the results predicted from this analysis and actual impact test data. The results show that the finite element model created may be used to predict the behaviour of random glass/epoxy laminates. However, the inclusion of flexible post-failure degradation rules to allow for progressive damage, will improve the accuracy of the analysis.