Interlaminar fatigue damage model of carbon fiber-reinforced polymer composite laminates

Abstract

Load-bearing structures made of carbon fiber-reinforced polymer (CFRP) composite laminates, such as the skin of aircraft wings, helicopter rotors, and wind turbine blades, are likely to experience time-varying loads. The fluctuating stresses could result in fatigue damage and failure of the laminates in the form of matrix cracking, fiber breakage and buckling, fiber/matrix debonding, and interface delamination. The latter is a significant damage mechanism in view of the relatively weak interlaminar bonding. In this respect, the current research has developed the interlaminar damage-based fatigue life model of fibre-reinforced polymer (FRP) composite laminates. The model incorporates the observed continuous cyclic degradation of interlaminar properties. The bi-linear traction-relative displacement softening rule for the cohesive zone model (CZM) is extended to accommodate the normalized interlaminar strength and stiffness degradation under the fatigue load cycles. The normalized fatigue life model accounts for the effect of mean stress on the observed interlaminar fatigue lives. Fatigue crack nucleation (separation) is governed by the interface's critical strain energy release rates. Hybrid finite element-experimental approach is employed to establish material parameters for the quasi static CZM. The experimental fatigue data for Mode I and Mode II of CFRP composite laminates from literature is employed to extract the residual properties. The normalized properties versus normalized fatigue life curves are then quantified based on the “wear-out” failure model. The curves are characterized by the curve fitting parameters α, β, λ, γ, μ, and ϕ for the interlaminar tensile strength, stiffness, and fracture energy. In view of the relatively large number of load cycles to capture the initiation and propagation of the interlaminar crack, the load cycle block approach is devised to improve computational efficiency. The model is coded in the UMAT Subroutine of Abaqus FE software. It is examined for interlaminar fatigue of CFRP composite laminate under Mode I , Mode II and mixed-mode loading conditions with a stress ratio, κ = 0.11, 0.15, and 0.1, respectively. The damage begins at approximately 8200 cycles and interface crack extends after accumulating 14100 applied fatigue cycles for Mode I load case. The damage begins at approximately 220000 cycles and interface crack extends after accumulating 350700 applied fatigue cycles for Mode II load case. The stress is highly concentrated at the crack front region. The FE-predicted fatigue lives are comparable with measured data and within the experimental scatter, hence validating the model. The crack tip opening and sliding displacements evolve with an initially slow rate of 2.6x10-9 and 1.85x10-10 mm/cycle respectively up to the onset of fatigue crack nucleation event at approximately 188800 cycles and then peaks at 1.5x10-7 and 7.1x10-8 mm/cycle respectively as the interface crack begins to accelerate after accumulating 284700 applied fatigue cycles for mixed mode flexure fatigue loading. The developed model will benefit various industries, including aerospace, automotive, and maritime, involved in the structural design for performance, reliability prediction, life extension and failure investigation of CFRP composite laminate structures.