A Structure-cell Correlation Model For Composites Mechanical Damage Analysis With Experimental Verification Under Multiple Loading Cases

With the extensive use of composite materials in military and civil aviation structures,especially in the new commercial aircraft in our country,progressive damage analysis of composite materials own great significance in achieving the dual indicators of structural safety and weight control.In this paper,the mechanical damage responses of various laminated composites with different laying sequences under four loading conditions were analyzed through theoretical,experimental and numerical methods.These analyses included both standardized composite specimen and real aviation composite structures,and the loading cases ranged from quasi-static tensile / compressive load to dynamic impact load.Based on the Representative Damage-trackable Cell(RDC),a structure-cell correlation theory for composite structures mechanical responses analysis under various loading cases was proposed.On the basis of this theory,the decoupling and reconstruction of the relationships among external-load field,material properties,mechanical response distribution and damage states field were realized.By adopting this theory,different numerical models were established and the corresponding simulations were conducted.In addition,by the implementation of more than one hundred verification tests,a high reliability test database was established.By comparing the prediction and experimental results on fourteen kinds of mechanical damage response parameters,the rationality of the proposed correlation analysis theory was verified.Through the above work,some conclusions about the mechanical responses under multiple loading cases could be drawn as follows.(1)Under uniaxial tensile load,the damage on 90-Degree promoted the damage initiation on its adjacent layer.Therefore,layers with same fiber direction in symmetric lamiantes own different damage initiation pierod.Among several damage modes,the IFF damage was the first one to initiate.With the growth of IFF damage,the crack invaded into the interface,which also promoted the generation of delamination damage and the damage in adjacent layers.In the fracture evolution process of nothed plates,this coupling among different damage modes and various layers was always the main reason that caused the final failure of composite plates.(2)Under conditional impact load,the damage of each sublayer within stadardrized composite plates generated peanut-shape damage distribution.The corresponding damage within different sublayers in laminates own different directions,which leads to the projection area of whole lamiantes’ impact damage form a circular distribution area.Under the compression load after impact,the composite plate generated cross-sectional fracture plane which acrossed the pre-impact-damage area located in the geometric center.(3)Under the edge impact load,the obvious matrix conquassation defects were found on the composite stringers’ blade edge.In addition,a wider damage range of delamination generated within the interface cohesive layers.The delamination damage was in elliptical shape,whose long axis was located at the blades’ edge.The length of this long axis determined the size of the BVID.Under the compression load after the edge impact,the blade failure at the damage location due to the evlotion of both IFF and DEA damage.And this phenomenon finally promoted the total fracutre of composite stringers.