Abstract:The three鄄dimensional finite element model of foam core sandwich panel with non鄄penetrating damage
of SR20 airplane is established. The convergence of the model is analyzed and the effective mesh density is given.
Stress analysis of the sandwich panel with non鄄penetrating damage under uniaxial and twin鄄axial tensile loading is car鄄
ried out. The material principle direction stresses distribution of the sandwich panel with no defects and repaired panel
are given. The uniaxial and twin鄄axial tensile strengths of the intact and repaired panel are calculated based on the
strength criterion of maximum stress. It is shown that under uniaxial tensile loading, the initial damage mode of the re鄄
paired panel is shear failure. The failure occurs on the motherboard which is beside the boundary of the repair area
and about 30毅to the x symmetry axis of the panel. Under twin鄄axial tensile loading, the initial damage mode of re鄄
paired panel is the fracture of the reinforce fiber. The failure occurs on the motherboard which is near the boundary of
the repair area and is about 45毅to the x symmetry axis of the panel. In the ideal repair status, the strength recovery
coefficient of the repaired panel under uniaxial tensile loading is 85. 8%, and 96. 7% for the same repaired panel un鄄
der twin鄄axial tensile loading. The strength of the repaired sandwich panel decreases due to the stress concentration
caused by the material discontinuity of the repair area. The strength of the repaired sandwich panel decreases with the
increase of the number of surface patches. This happens thanks to the increase of local stiffness caused by the addi鄄
tional surface patches.
