Study On The Delamination Damage Of Aircraft Structure Composite Laminates

Advanced composites due to high strength and stiffness to weight and fatigue resistance have been widely used in the field of aeronautics and astronautics. However, composite structures are inevitable to have defects in the processing and use. Delamination damage is one of main failure forms. This paper is to study the delamination damage of aircraft structure composite laminates.A three-dimensional finite element model and a double-plate finite element model for predicting delamination damage growth in Laminated Composites and calculating the strain energy release rates are built based on the virtual crack closure technique (VCCT) and the crack tip element (CTE). The calculated results are compared to determine accuracy. The influences of stacking sequence and ratio of laminate on the strain energy release rate are analyzed with these methods.The analysis method of delamination formation and growth based on interface element of cohesive zone model (CZM) was introduced. The delamination in composite laminate is often under mixed-mode load. Mixed mode combines mode I under interlaminar tension load and Mode II under interlaminar sliding shear load. Mixed-mode bending (MMB) test is simulated based on interface element. The simulated result is compared with the test result to predict the interface shear properties. The element density influence on the carrying capacity of the delamination growth is considered.This paper has studied the ultimate strength and failure patterns of CFRP laminates. A progressive damage model has been developed based on micromechanics theory and the rigidity deterioration of failure element. The model and the program are utilized to simulate the procedure of failure and predict ultimate strength of open hole CFRP laminates. Hashin failure criteria and LaRC03 failure criteria are applied to determine the failure patterns of the material in the element. The material properties are degraded according to their failure patterns that are fiber breakage, matrix cracking and fiber-matrix shearing failure. Four different types of laminates with holes under uniaxial load are used to demonstrate the failure patterns, initial failure, failure progression and ultimate failure load in the progressive damage analysis. Finally, the four kind laminates with open hole are tested to verify the simulation model and program, and the testing results show satisfying simulation and prediction capacity of the proposed model and developed program in this research.A sandwich panel containing preformed inserts with debonding between the face plates and honeycomb core is considered. The debonding failure is simulated based on interface element with damage initiation criteria and damage evolution criteria. A progressive damage model is utilized to simulate the procedure of failure the panel of the sandwich. The failure of integrated sandwich panel is analyzed on the finite element software ABAQUS and fortran subroutine. The test is carried to verify the simulation analysis and the test show the simulation can exactly predict the ultimate capacity and damage mechanism, types and failure propagation of the sandwich panel containing preformed inserts.