Compressive buckling property reduction of composite plates with three simply-supported and one free edges due to a delamination

Abstract

The compressive behaviors of delaminated orthotropic plates with three simply supported and free edges, which are simplified models of stiffener plates, are studied experimentally and analytically. In the analysis, Rayleigh-Ritz approximation method is adopt,ed to obtain an estimate of buckling load. The compressive buckling load reduction is found to be significant and almost proportional to the delamination width. Local delamination buckling is found to occur when the delaminated portion is thin and large. The analytical results agree well with the experimental ones. Key Word : Composite laminate, Free edge, Interlaminar delamination, Compressive buckling. Fiber reinforced coinposite 1aminat.e~ attract intensive attentions as a prommising candidate of a future material not only in aerospace engineering but also various engineering fields owing to its remarkable properties such as high strength and high stiffness ratio, good fatigue resistance, design-ability, etc. Interlaminar delamination, which is suspected to occur owing to impact load, fatigue, etc., IS one of typical damage patterns of laminated structures. Its introduction causes sig~iificant reduct,ion of compressive strength of the plates. Therefore, a lot of works have been performed to disclose the effects of the delaminat,ion(s), particularly on one dimensional model1-lo) and have shown possibility of significant loss of compressive buckling and postbuckling properties. Works on the effect of delaminations on the compressive behaviors of two dimensional plate are limited only on special The delamination is also thought to appear at a free edge of the stiffener plate of stiffened panels and to reduce significantly its compressive strengt,h. In the present paper, the stiffener portion is modeled as a plate with three simply-supported and one free edges. The eflect of a delamination on compressive buckling properties of the orthotropic plate is analytically and experimentally studied. 'Associate Professor, Mechanical Engineering, Member AIAA 'Member AIAA 2 Theory A plat,e with three simply-supported and one free edges showu ill Fig.1 is considered. It is assumed to be uniform in thickness direction. that is, no bending-st,retchillg coupling is considered even in the delaminated area. A delamination is assunled t o locate along the free edge. The boundary conditions are where the values u , v, w denote, respectively, displacem e ~ k s ill c, y dricl : directions. The values a and b are the width and length of the plate. The term rZy is a Co~vright @ 1993 American Institute of Aeronautics and Astronautics, Inc. All rights reserved. 485 shear stress in x y plane. The values Q,, M, and My are a transverse shear force and bending moments in x and y directions. The value EO is the normalized displacement of the loading edge. When the transverse shear deformation is neglected, the conlpl~sblve bucklillg bllebb u,, of the prcvnt plate wlthout a delam~nat~on 1s where h is the plate thickness, X = a lb is aspect ratio of the panel and the coefficients DZ2, D66 are the components of bending stiffness matrix of the laminate. The component Dl l does not appear in the expression of the buckling stress of Eq.5. The Kirchhoff's hypothesis is assumed t o hold even in the neighborhood of the delamination tiofit. The continuity condition of the displacements nt Ll~e edge of the delamination ;L. = a c can be written simply ~ i t h the displacements of the midplanes of each delaminated and undelaminated portions. U I + h k ~ w o ~ ~ = uo V I + h k I w O l y = v0 (7)