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Keywords: aerospace structures, composite, buckling, delamination, exact strip, Wittrick-Williams algorithm.

Summary

Composite materials are widely used in aerospace structures. Although these materials offer many advantages over metals in terms of strength, stiffness and low mass, additional types of failure can be encountered. One of the most serious of these is delamination, which reduces the resistance to compressive and shear loads. Therefore, understanding the effects of delamination on buckling behaviour is essential for the correct design and safe use of composite materials in aircraft.

This paper uses exact stiffness analysis for the critical buckling analysis of composite plates with through-the-length delaminations. The usual discretisation of the finite element and finite strip methods is avoided by using analytical solutions of the member stiffness equations, the resulting transcendental eigenproblems being solved by the Wittrick-Williams algorithm. Accuracy is enhanced by coupling sinusoidal responses to give an infinitely long plate model.

Numerical results show that, as the delamination width is increased, there is a reduction in the critical buckling load, accompanied by a transition from a global mode to a local mode of the delaminated portion. This transition tends to be sudden for longitudinal and shear loading, and more gradual for transverse loading. The local buckling load decreases as the delamination is moved towards the surface of the plate, becoming negligibly small for wide delaminations of only a few plies. The effect of multiple through-the-length delaminations is dominated by local buckling of the thinnest portion of the delaminated region. The chosen example was insensitive to the widthwise location of the delamination, and gave qualitatively similar results for both simply supported and clamped edge conditions.

Because the method uses much smaller models than finite element analysis, solution times are very competitive. This approach is therefore suitable for parametric studies in the preliminary design of aerospace structures.

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Front Page Browse CCP CSETS CTR IJRT Other Authors Search Purchase Guide FAQ Contact us	Civil-Comp Proceedings ISSN 1759-3433 CCP: 88 PROCEEDINGS OF THE NINTH INTERNATIONAL CONFERENCE ON COMPUTATIONAL STRUCTURES TECHNOLOGY Edited by: B.H.V. Topping and M. Papadrakakis Paper 309 Critical Buckling of Delaminated Composite Plates Using Exact Stiffness Analysis M. Damghani, C.A. Featherston and D. Kennedy Cardiff School of Engineering, Cardiff University, United Kingdom doi:10.4203/ccp.88.309 purchase the full-text of this paper Full Bibliographic Reference for this paper M. Damghani, C.A. Featherston, D. Kennedy, "Critical Buckling of Delaminated Composite Plates Using Exact Stiffness Analysis", in B.H.V. Topping, M. Papadrakakis, (Editors), "Proceedings of the Ninth International Conference on Computational Structures Technology", Civil-Comp Press, Stirlingshire, UK, Paper 309, 2008. doi:10.4203/ccp.88.309 Keywords: aerospace structures, composite, buckling, delamination, exact strip, Wittrick-Williams algorithm. Summary Composite materials are widely used in aerospace structures. Although these materials offer many advantages over metals in terms of strength, stiffness and low mass, additional types of failure can be encountered. One of the most serious of these is delamination, which reduces the resistance to compressive and shear loads. Therefore, understanding the effects of delamination on buckling behaviour is essential for the correct design and safe use of composite materials in aircraft. This paper uses exact stiffness analysis for the critical buckling analysis of composite plates with through-the-length delaminations. The usual discretisation of the finite element and finite strip methods is avoided by using analytical solutions of the member stiffness equations, the resulting transcendental eigenproblems being solved by the Wittrick-Williams algorithm. Accuracy is enhanced by coupling sinusoidal responses to give an infinitely long plate model. Numerical results show that, as the delamination width is increased, there is a reduction in the critical buckling load, accompanied by a transition from a global mode to a local mode of the delaminated portion. This transition tends to be sudden for longitudinal and shear loading, and more gradual for transverse loading. The local buckling load decreases as the delamination is moved towards the surface of the plate, becoming negligibly small for wide delaminations of only a few plies. The effect of multiple through-the-length delaminations is dominated by local buckling of the thinnest portion of the delaminated region. The chosen example was insensitive to the widthwise location of the delamination, and gave qualitatively similar results for both simply supported and clamped edge conditions. Because the method uses much smaller models than finite element analysis, solution times are very competitive. This approach is therefore suitable for parametric studies in the preliminary design of aerospace structures. purchase the full-text of this paper (price £20) go to the previous paper go to the next paper return to the table of contents return to the book description purchase this book (price £145 +P&P)
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