Analytically Validated Highly Damped Repair for Acoustically - Induced Cracking on the F/A-18 Nacelle Skin

摘要

The skin of an aircraft can vibrate as a result of pressure waves caused by engine and/or aerodynamic effects. In modern fighter aircraft such as the F/A-18, overall sound pressure levels have been recorded up to 170 dB over the surface of the skin. In the F/A-18, cracking has occurred in the lower nacelle external skin, typically along the boundaries of the panel. These cracks often originate from a fastener line and grow along the boundary and then turn into the centre of the panel. In the case of the F/A-18, cracking was due to higher than expected pressure levels caused by an aerodynamic disturbance at the inlet lip. Initial design of this region was based on an overall sound pressure of 160db, which was based on initial computational fluid dynamics (CFD) work. An attempt was made to repair a panel with a boron fibre patch however this repair did not appear to significantly reduce the crack growth. The design of the repair was based on in-plane loads only. In this report the aim is to develop a patch design, based on a finite element analysis, which significantly reduces crack stress intensity factor for structures subject to intense acoustic environments. A significant reduction in stress intensity is achieved by using a highly-damped patch which incorporates a uni-directional boron/epoxy repair and constrained damping layers which consists of viscoelastic material, which provides the necessary damping, and 0/90 boron/epoxy constraining layers. Such configurations have shown that theoretical reductions in stress intensity factors of up to 10 times are achievable. This has the potential of reducing the mode I stress intensity factor (K1) to the order of K1 threshold thus substantially reducing crack growth.