Analysis of the effect of impact damage on the repairability of CFRP composite laminates.

Abstract

Polymer composite materials are common in the aerospace application such as aircraft structures including primary and secondary structures. Therefore, there has been an increasing demand for composites in both the military and civilian aircraft industry. At least 50% of the next generation of military and civil aircraft structures are likely to be made from composites. The most important properties for composite materials in aircraft application was the high strength-to-weight ratios, stiffness-to-weight ratios and easy to repair. However, the composite materials have low resistance for impact damage. Impact can lead to significant strength reduction in aircraft structure about 40% to 60% of an undamaged composite laminate strength. Therefore, establish a numerical methodology to defined the optimum repair joint to restore sufficient strength of damaged aircraft composite structures during some operations and exercise activities with limited resources which will be the main contributions to knowledge in this thesis. To achieve this contribution need to understanding of the behaviour of Carbon Fibre Reinforced Plastic (CFRP) composite laminates subject to high velocity impact and the unrepaired composite laminates and repaired (stepped joint) subject to compression after impact test. Therefore, this study consists of two parts:- first, part a combined of numerical simulation and experimental investigation have been used to evaluate the woven CFRP laminate subject high velocity impact. The selected impact velocities were (140m/s, 183m/s, 200m/s, 225m/s, 226m/s, 236m/s, 270m/s, 305m/s, 354m/s and 368m/s) in order to evaluate the induced impact damage in three different thickness of CFRP composite laminates (6 mm, 4.125 mm and 2.625 mm) these velocities were selected according the gas gun limitation. The woven composite laminate made of Hexcel G0926 Carbon Fabric 5 harness 6K, Areal Weight 370 gsm. The resin used was Hexcel RTM 6, cured for 1 hour 40 minutes at 180° C at a pressure of 100 psi, with an average thickness of 0.375mm. The laminates were comprised of 16 layers, using the following stacking sequence: [(0/90); (±45); (±45); (0/90); (±45); (±45); (0/90); (0/90); (±45); (0/90); (±45); (0/90); (±45); (0/90); (±450); (0/90)], 11 layers, using the following stacking sequence: 0/90; ± 45; 0/90; ± 45; 0/90; ±45; 0/90; ± 45; 0/90; ±45; 0/90 and 7 layers, using the following stacking sequence: ± 45; 0/90; 0/90; ±45; 0/90; 0/90; ±45. The density of woven CFRP laminates was 1.512e-3 ±1e-6 grm/mmÖ³. The penetration process and also change of kinetic energy absorption characteristics have been used to validate the finite element results. The experimental and numerical method in this study show a significant damage occurs, including delamination, compression through thickness failure, out-of-plane shear failure and in-plane tensile failure of the fibres located at the rear surface when the projectile penetrates the laminate. The penetration mechanism of the projectile had a “plugging-type” (shear) failure and the hole that was formed after impact was conical in shape were shown in experimental and also verified in the numerical model. The residual kinetic energy in numerical model is 5.0 % larger than experimental data which is significantly matched in all simulated cases. In part two a finite element model is established to optimise the repair joint to restore sufficient strength of damaged composite laminate and used compression after impact test to compare the compression failure load of the sample. In order to achieve this an optimised repair models of stepped lap joints with variable parameters such as number of steps and length of steps have been experiment the undamaged composite laminate and composite laminate subject to high velocity impact and also created a numerical model for these experimental. The experimental CAI failure load of undamaged 7 Plies CFRP composite laminate higher than the failure load of damaged specimens by approximately 23%. The undamaged 11 Plies CFRP composite laminate failed at approximately 40% higher than the damaged specimens. Moreover, the difference between the experimental and numerical results of above tests was about 10%. The numerical model of repaired composite laminate show the damage initiated at the end of overlap and the average compression failure load of the stepped lap joint increased with the increasing of the number of step and length of step. The 85% and 90% of compressive failure load has been restored.