Durability of Buckled Composites Using Virtual Crack Closure Technique Fatigue R-Curve Implementation

Abstract

Composite structures have become popular in modern aircraft because they help reduce weight and increase durability. In addition, hat-stiffened panels provide the stability that the airframe skin needs. However, they can be subject to delamination in the postbuckling regime. Progressive damage analysis methods can help predict interlaminar and intralaminar failure events. Many aircraft structures are subject to cyclic loading in the postbuckling regime. Hence, fatigue life prediction becomes essential for design and sustainment purposes. Under the NASA Advanced Composites Project, composite panels stiffened with single- and multiple-hat stringers were subject to a cyclic loading sequence from a prebuckling state to a postbuckling state. This work uses the Abaqus virtual crack closure technique with enhanced capabilities through an empirical method to integrate fatigue [Formula: see text] effects (different from the [Formula: see text]) into the Paris law. This was accomplished via a user-defined subroutine to simulate the fatigue response of these panels. This novel method captured the fatigue life predictions within 5% of the test results.

Influence of Delamination Size and Depth on the Compression Fatigue Behaviour of a Stiffened Aerospace Composite Panel

“…The effect of multiple delaminations on postbuckling properties has been studied by using the virtual crack closure technique. The study in [16] employed Abaqus, enhancing fatigue analysis by integrating R-curve effects into the Paris law through an empirical method. Implemented via a user-defined subroutine, the novel method accurately predicts fatigue life within 5% of the test results.…”