A nonlocal damage theory for laminated plate with application to aircraft damage tolerance

Design of commercial aircraft structure, composed of composite material, requires the prediction of failure loads given large scale damage. In particular, a fuselage of graphite/epoxy lamination was analyzed for damage tolerance given a standard large crack that severed both skin and internal structure. Upon loading, a zone of damage is known to develop in front of a crack-tip in composite laminates; and, its material behavior within the damage zone is characterized as strain softening. This investigation sought to develop a computational model that simulates progressive damage growth and predicts failure of complex laminated shell structures subject to combined tensile and flexural load conditions. This was accomplished by assuming a macroscopic definition of orthotropic damage that is allowed to vary linearly through the shell thickness. It was further proposed that nonlocal plate strain and curvature act to force damage growth according to a set of uniaxial criteria. Damage induced strain softening is exhibited by degradation of laminate stiffness. An expression for the damage reduced laminated plate stiffness was derived which assumed the familiar laminated plate (ABD) stiffness matrix format. The model was implemented in a finite element shell program for simulation of fracture and evaluation of damage tolerance. Laminates were characterized for damage resistance according to material parameters defining nonlocal strain and the damage growth criteria. These parameters were selected using an inverse method to correlate simulation with uniaxial strength and fracture test results. A novel combined-tension-plus-flexure fracture test was developed to facilitate this effort. Analysis was performed on a section of pressurized composite fuselage containing a large crack. Good agreement was found between calculations and test results.