【作者】 杨眉；

【导师】 汪海；

【作者基本信息】 上海交通大学 ， 飞行器设计， 2011， 硕士

【摘要】 先进复合材料由于高比强度、比刚度,耐疲劳等优点在航空航天领域中得到广泛应用,但复合材料结构在加工和使用中不可避免产生缺陷,其中分层损伤是复合材料的主要失效之一。本文对飞机复合材料结构分层损伤进行研究。本文基于裂纹闭合技术和裂纹尖端单元分析方法,分别建立3D和双板有限元模型计算作为复合材料分层扩展判据的应变能量释放率,比较其预测精确度,在此基础上分析了层压板的铺层顺序对应变能量释放率的影响。介绍基于内聚力模型的界面元复合材料分层形成及扩展分析方法,利用界面元对混合型弯曲分层扩展试验进行模拟并与试验结果对比,预测层合板的层间性能及裂纹尖端网格密度对层压板混合模态下分层扩展的极限载荷影响。本文研究了层压板的极限强度和破坏失效模式,采用基于细观力学的失效准则,逐个单刚度退化,从而对层压板进行渐进失效分析。基于Hashin准则和LaRC03准则以及Chang的材料参数退化模型,利用渐进失效模型对四种铺层含圆孔的层合板的拉伸破坏进行了失效扩展过程的数值仿真分析,预测其破坏模式、损伤起始、损伤扩展以及拉伸极限载荷,并通过4种碳纤维复合材料开孔层合板拉伸强度试验对仿真结果进行验证。基于界面元,运用分层起始和分层扩展准则,对蜂窝夹层内嵌件结构进行脱胶分析,同时基于复合材料渐进失效对蜂窝夹层面板进行失效分析。利用Abaqus软件,通过fortran用户子程序对含内嵌件的蜂窝夹层结构进行整体失效分析,并通过试验验证,证明整体失效分析模型能较好预测含内嵌件蜂窝夹层结构拉脱极限强度和各部分损伤机理,类型以及演化过程。更多还原

【Abstract】 Advanced composites due to high strength and stiffness to weight and fatigue resistance have been widely used in the field of aeronautics and astronautics. However, composite structures are inevitable to have defects in the processing and use. Delamination damage is one of main failure forms. This paper is to study the delamination damage of aircraft structure composite laminates.A three-dimensional finite element model and a double-plate finite element model for predicting delamination damage growth in Laminated Composites and calculating the strain energy release rates are built based on the virtual crack closure technique (VCCT) and the crack tip element (CTE). The calculated results are compared to determine accuracy. The influences of stacking sequence and ratio of laminate on the strain energy release rate are analyzed with these methods.The analysis method of delamination formation and growth based on interface element of cohesive zone model (CZM) was introduced. The delamination in composite laminate is often under mixed-mode load. Mixed mode combines mode I under interlaminar tension load and Mode II under interlaminar sliding shear load. Mixed-mode bending (MMB) test is simulated based on interface element. The simulated result is compared with the test result to predict the interface shear properties. The element density influence on the carrying capacity of the delamination growth is considered.This paper has studied the ultimate strength and failure patterns of CFRP laminates. A progressive damage model has been developed based on micromechanics theory and the rigidity deterioration of failure element. The model and the program are utilized to simulate the procedure of failure and predict ultimate strength of open hole CFRP laminates. Hashin failure criteria and LaRC03 failure criteria are applied to determine the failure patterns of the material in the element. The material properties are degraded according to their failure patterns that are fiber breakage, matrix cracking and fiber-matrix shearing failure. Four different types of laminates with holes under uniaxial load are used to demonstrate the failure patterns, initial failure, failure progression and ultimate failure load in the progressive damage analysis. Finally, the four kind laminates with open hole are tested to verify the simulation model and program, and the testing results show satisfying simulation and prediction capacity of the proposed model and developed program in this research.A sandwich panel containing preformed inserts with debonding between the face plates and honeycomb core is considered. The debonding failure is simulated based on interface element with damage initiation criteria and damage evolution criteria. A progressive damage model is utilized to simulate the procedure of failure the panel of the sandwich. The failure of integrated sandwich panel is analyzed on the finite element software ABAQUS and fortran subroutine. The test is carried to verify the simulation analysis and the test show the simulation can exactly predict the ultimate capacity and damage mechanism, types and failure propagation of the sandwich panel containing preformed inserts.更多还原