Crash Simulation Of A Composite Unmanned Aerial Vehicle Fuselage

In this research the results of experimental works and numerical simulation works pertaining to the crash behavior and crashworthiness characteristic of the upper part of the composite unmanned aerial vehicle (UAV) fuselage sections that were subjected to quasi-static transverse compressive loading are presented in detail. The influence of varying angles of lamina and special cases of laminates is thoroughly analyzed. The fuselage sections were made of 8 plies of C-glass/epoxy in a [45/-45/90/0]s layup. Two types of density of C-glass/epoxy, 200 g/m2 and 600 g/m2, were used with a total thickness of 0.00224 m and 0.004 m respectively for the 8-plies. Each ply has a thickness of 0.00028 m for C-glass/epoxy 200 g/m2 and 0.0005 m for C-glass/epoxy 600 g/m2. The C-glass/epoxy fuselage section was compressed using MTS machine of 250 kN loading capacity at very low-strain rate typical for static testing. The experimental data are correlated with predictions from a finite element model developed using the ABAQUS/Standard with user subroutine. The simulation of the composite fuselage sections was carried out, refined several times and validated with the experimental results. The ABAQUS analysis results for both the C-glass/epoxy 200 g/m2 and C-glass/epoxy 600 g/m2 fuselage sections agreed well with the experimental data. ABAQUS analyses predicted the location of progressive damage to the sections using three failure theories, Maximum Stress Failure Theory, Tsai-Hill Failure Theory and Tsai-Wu Failure Theory. Tsai-Hill Failure Theory is found to have the least error percentage compared to the other two failure theories used. Finally, the finite element model was then used to study the influence of varying angles of lamina and special cases of laminates. 15o angle of lamina and cross-ply laminate is found to have the most energy absorption.

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