A hybrid structural health monitoring (SHM) system, consisting of a piezoelectric transducer and fiber optic sensors (FOS) for generating and monitoring Lamb waves, was investigated to determine their potential for damage detection and localization in composite aerospace structures. As part of this study, the proposed hybrid SHM system, together with an in-house developed algorithm, was evaluated to detect and localize two types of damage: a through thickness damage (hole of 2 mm in diameter) and a surface damage (2 mm diameter bore hole with a depth of 0.65 mm) located on the backside of the plate. The experiments were performed using an aircraft representative composite plate skin, manufactured from carbon fiber reinforced polymer (CFRP).

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The paper presents the results of an experimental study that used digital image correlation to measure the residual strains created by hole cold expansion both before and after insertion of an interference fit fastener. The study used 7075-T6 aluminium test coupons with multiple fastener holes to more closely simulate a real aircraft structure where multiple fastener holes are often cold expanded sequentially and where the interactions between the holes are a function of both the hole-to-free-edge distance as well as the pitch between the fastener holes. Both hole-to-free-edge distance and coupon thickness were varied to measure their effects on residual strain after hole cold expansion and after hole cold expansion and interference fit fastener insertion. The results showed that as the edge distance was decreased, the tensile strains at the low edge side of the coupon increased exponentially for both thick (6.35 mm) and thin (1.59 mm) coupons.

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