Progressive Failure Analysis in Open-Hole Tensile Composite Laminates of Airplane Stringers Based on Tests and Simulations

Jian Shi,Congjie Ye,Chuwei Zhou,Mingbo Tong,Xindong Wang

doi:10.3390/app11010185

Abstract

The failure types and ultimate loads for eight carbon-epoxy laminate specimens with a central circular hole subjected to tensile load were tested experimentally and simulated using two different progressive failure analysis (PFA) methodologies. The first model used a lamina level modeling based on the Hashin criterion and the Camanho stiffness degradation theory to predict the damage of the fiber and matrix. The second model implemented a micromechanical analysis technique coined the generalized method of cells (GMC), where the 3D Tsai–Hill failure criterion was used to govern matrix failure, and the fiber failure was dictated by the maximum stress criterion. The progressive failure methodology was implemented using the UMAT subroutine within the ABAQUS/implicit solver. Results of load versus displacement and failure types from the two different models were compared against experimental data for the open hole laminates subjected to tensile displacement load. The results obtained from the numerical simulation and experiments showed good agreement. Failure paths and accurate damage contours for the tested specimens were also predicted.

Highlights

Utilizing composite laminates in aircraft structures plays quite an important role owing to their low weight, high strength, the capability to design varieties of geometries and other advantages
The analysis model with failure criteria for progressive failure analysis (PFA) was accomplished using a user-defined material subroutine (UMAT) in the ABAQUS finite element analysis code, and its implicit solver was utilized in the simulation models
Test results showed that the load versus displacement curve of the specimen was mainly linear before reaching the ultimate load, which may be due to the fact that 0◦ fiber lamination was the main load-bearing object under tensile load, and the fiber had typical brittleness characteristics before the failures occurred

Summary

Introduction

Utilizing composite laminates in aircraft structures plays quite an important role owing to their low weight, high strength, the capability to design varieties of geometries and other advantages. The ultimate load and failure type of composite structures are significant parameters to satisfy the engineering requirement, especially those discontinuous structural components such as laminates with notches or open holes that may cause complex stress states and damages [1,2,3,4]. The breakage occurring around the open hole induces the redistribution of the stresses in the composite laminates that contributes to causing the catastrophic failure [9]. Fiber breakage, delamination, etc., observed in an open hole laminate subjected to tensile loads are common failure mechanisms [10,11,12]. Many analytical and numerical models have been used to predict the failure strength of composite laminates with open holes [13,14]. Many criteria distinguish the failure of fiber and matrix and take progressive failure modeling into account [20,21,22], they are not available to capture the interaction between the constituents at the microscale level

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Conclusion