Numerical and Experimental Study of Defects Evolution in a Composites Wing Box Under Compressive Loads

Due to their high specific strength and stiffness, Carbon Fibres Reinforced Plastics (CFRP) are commonly considered suitable for aerospace structural applications. However, their failure mechanisms are not completely predictable and this is the main reason why the CFRP integration in the aerospace industry has been generally slow in the last twenty years. Indeed, the lack of robust numerical tools, able to take into account the damage tolerance of composite structures, has led to over-conservative designs, not fully realising the promised economic benefits of composites materials. The Project MACMES (Damage Management of Aircraft Composite Structures Monitored by Embedded Sensors), funded by the General Defence Secretariat/National Armaments Directorate, of the Italian Ministry of Defence, in the framework of the National Military Research Plan (PNRM), addresses this issue by suggesting an integrated approach consisting in a synergy between numerical predictive tools (FEM models) and experimental ones (embedded optic fibres) for the damage management of aircraft composite structure.Within the MACMES project, such integrated approach was applied to a composite wing-box in order to monitor the buckling and the internal damage evolution under compressive loads. The experimental compressive test was performed by ALENIA under displacement control. Two defects were artificially included into the wing box: a skin/stringer debonding and an embedded circular bay delamination. Position and size of the delamination and skinstringer debonding were chosen as a result of a sensitivity analysis performed during the preliminary design phase and aimed at determining the best configuration (in terms of defects size and position) which guarantees a satisfactory experimental measurement of the damage growth from initiation to the global buckling load and well before the failure load of the wing box. The non-linear post-buckling behaviour of the damaged composite structure was simulated by developing appropriate FE numerical models. The adopted numerical models use the Virtual Crack Closure Technique to simulate the inter-laminar damage evolution and the numerical analyses have been performed by means of the FEM code ABAQUS and B2000++. The obtained numerical results have been assessed and compared to each other in terms of delaminated area evolution, delamination growth initiation load and strain distributions in order to investigate the effectiveness of the adopted numerical platforms in predicting the evolution of inter-laminar damages. Comparisons with experimental data, in terms of load displacement curves and strains in the dedonding area, are presented to assess the accuracy of the numerical simulations