Study On Residual Strength Of Composite Laminates With Preset Damage

Carbon fiber reinforced plastics(CFRP)have been widely used in wing and fuselage skins,cabin siding and other parts of the aircraft.However,the CFRP structure is inevitably subject to impact from external objects,which often causes delaminations and fiber fracture in the interior of material,resulting in structural instability or even damage under low loads,and burying a huge security risk to the safety of the aircraft.Therefore,combined with experimental and finite element simulation methods,the bending behavior and residual strength of CFRP laminates with preset damage were studied.The main work and conclusions are as follows:Firstly,the bending behavior of T700/6240 unidirectional composite laminates with preset damage were studied experimentally,in order to investigate the effects of different damage forms and parameters on the residual strength of laminates,and to establish residual strength models.The study results show that,the residual strength of the upper delaminated sample is 7.15%higher than the intact sample,and the residual strength of intermediate and lower delaminated specimen is reduced by 23.98%and 1.43%,respectively;During the process of increasing the ratio of long to short axis length from 1:1,2:1 to 3:1,the residual bending strength decreases from 2.26%to 2.97%;the increase of delamination area causes the residual strength approximately quadratic function decreasing.As the position of fiber fracture change from the upper surface of the laminate to the lower surface of the laminate,the degree of attenuation of residual strength is reduced from 19.10%to 6.24%;the increase of the fiber fracture angle does not cause trending effect on the attenuation of the residual strength;the fiber fracture length increased from 5mm to 15mm,the residual strength decreased approximately linearly.Then,based on the ABAQUS finite element software platform,a three-point bending finite element model of the laminate with embedded damage was established.The error ratio between the finite element simulation results and the experimental values are less than 8.00%,which verifies the reliability of the numerical simulation method.On this basis,the three-point bending behavior of the laminate with delamination/fiber fracture mixed damage was also simulated,the contribution ratio of different damage forms was clarified.It is found that delamination area 3.14cm~2/fiber fracture length 11.6mm is used as the threshold,when the delamination area/fiber fracture length is small,the contribution of fiber fracture to the residual strength attenuation is larger than delaminations,and when the size of the stratified area/fiber break length exceeds the threshold,the contribution of delaminations will exceed the fiber fracture.The multi-delaminations and delamination/fiber fracture mixed damage were used to simulate the actual impact damage.The bending behavior of the laminate after impact damage was studied,and the calculated results were compared with the experimental results,which proved that the analysis method for the single damage form can calculate the residual strength of complex damage laminates and achieve comparable high precision.Then the test of the mechanical properties of the test piece is raised to the component level test of the aircraft wing.The bending behavior of the composite wing with impact damage is numerically simulated and the bending failure load is calculated for the actual impact damage,which provides a certain reference for the study of bending failure of composite wing structure.