13th International Conference on Fracture June 16–21, 2013, Beijing, China -1- Application of a novel finite element method to design of splices in a fiber metal subjected to coupled thermo-mechanical loading XueCheng Ping1,*, MengCheng Chen2, BingBing Zheng1, YiHua Xiao1 1 School of Mechanical and Electronical Engineering, East China JiaoTong University, 330013, China 2 School of Civil Engineering, East China JiaoTong University, 330013, China * Corresponding author: xuecheng_ping@yahoo.com.cn Abstract The splice concept has been developed in fiber metal laminates (FMLs) due to the limited dimensions of pretreatment facility, the autoclave curing facility and C-scan facility. Failure always initiates in the splice opening between the metal splice edges before delamination occurs in the loading direction. In this paper, a novel finite element method for obtaining stress intensities is introduced in the estimation of the failure strength of splices in FMLs. A super wedge tip element for application to bi-material wedge is developed utilizing the numerical stress and displacement field solutions based on an ad hoc finite element eigenanalysis method, from which singular stress fields near apex of arbitrary bi-material wedges under coupled thermo-mechanical loading can be obtained. Failure of double spliced FMLs with varying spliced width and fiber-layer thickness were investigated, and fracture criterion based on stress intensity factors are presented to predict splice failure. Keywords Thermo-mechanical loading, Fracture criterion, fiber metal laminate, Splice, Super wedge tip element 1. Introduction Fiber metal laminates, such as Arall (Aramid reinforced aluminum laminates) and Glare (S2-glass fiber reinforced aluminum laminates), were developed at Delft University of Technology as a family of structural aerospace sheet materials. They take the advantages of metal alloys and fiber-reinforced composites providing superior mechanical properties [1], and have been applied to fuselage and leading edges in aircraft structures as a replacement of high-strength aluminum alloys due to their light weight, high strength, and excellent fatigue resistance. The dimensions of a FML are only limited by the width of the aluminum layers and not by their length. The available prepreg dimensions are also not limited. The maximum dimensions of a Glare skin panel are further limited by the dimensions of the pretreatment facility, the autoclave curing facility and C-scan facility. Also the handling and transportation of the panel may be a limiting factor. Currently only aluminium sheet with widths up to 1524 mm (60 inch) can be manufactured to the required accuracy with the necessary nominal thicknesses between 0.3 and 0.5 mm. This limitation would imply the necessity of applying many costly mechanical longitudinal or circumferential joints (at a maximum distance of 1524 mm) in an aircraft fuselage. To avoid this disadvantage, the internal splicing concept has been developed to increase the maximum available sheet size [2-4]. Due to the difference in elastic properties and thermal expansion coefficients of the components joined and discontinuity of junction geometry, residual stresses exist in each layer of the FML, and most importantly, high stress concentrations occur at the aluminum splice edges under thermal environment during manufacturing and coupled thermo-mechanical loading. As a result, the structural efficiency of bonded structures depends not on the structure itself but on splice strength. Generally, the failure behavior of the spliced FML always includes two steps, i.e., the onset of
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