Computation of energy release rates for cracked composite panels with nonlinear deformation

The effect of geometrically nonlinear behavior on the calculation of energy release rates in a centrally cracked tensile test plate and a damaged pressurized fuselage panel is evaluated. Tensile plate test specimens are used to generate allowables for the failure prediction of damaged fuselage panels under pressure load. Both the tensile plate and the damaged fuselage panel deform nonlinearly under loading. However, the nonlinear behavior of each is different. The tensile plate behaves nonlinearly after it buckles transversely in the cracked region well before failure. while the damaged pressure loaded fuselage behaves nonlinearly due to a membrane stress stiffening effect. Deformations predicted by nonlinear analyses for the test plate and the fuselage panel are significantly different from those predicted by linear analyses. Therefore, energy release rate allowables derived from cracked flat plate tests based on linear deformation theory need to be adjusted before being applied to damaged pressurized fuselage panels. Two methods are presented to evaluate the effect of nonlinear deformations on energy release rate calculations. One is the virtual crack closure technique (VCCT) and the other is the gradient method. Energy release rates computed by both methods are in excellent agreement for the postbuckled cracked flat plate and both show that postbuckling can significantly increase energy release rates. Energy release rates predicted by both methods for the damaged fuselage panel are also in very good * Aerospace Engineer. CSBISD. Member AIAA Old Dominion University. Member AIAA 3 Branch Head, CSBISD. Member AlAA Copyright 0 1 9 9 5 by the A~nerlcan Instltutc of Aeronautics and Astronaut~cs, Inc. No copyright is asxrted in the C n i t d Statcc; under title 17. U.S. Code The U S . Gnvern~nent haa a royalty-ire license to exercise all right$ under the copyright claimed herein Hx government purposts All rights are reserved by the copyr~pht owncr. agreement and both show that the membrane stiffening effect can noticeably reduce the energy available for crack growth. Almost all of the energy release rate of the tensile flat cracked plate and the damaged fuselage panel are of the Mode-I type. This indicates that the failure modes of the flat plate and the fuselage panel are similar.