The effect of curing conditions on bonded repair with boron/epoxy composite patch of aircraft skin

Abstract

Composite materials are becoming more important in the construction of aerospace structures. Bonded joints fulfil a significant role in the development of new technologies, particularly in the aerospace industry where they give new possibilities for connecting structural elements. The Composite Patch Bonded Repair is a method that can be used for repairing the metallic and composite structures. The method includes the following steps: surface preparation prior to bonding, the placing of Composite Patch and a cure cycle for the composite prepreg and the adhesive film.The primary advantages of composite materials for repair are their high strength, relatively low weight, and corrosion resistance. The purpose of this research is to define the influence of the curing time and temperature on bonded repair structures, by bonding the boron patch composite to the Aluminium 7075-T6 with the use of Structural Adhesive Film FM73K for lower residual stresses by analysis the distortion due the bending curvature. A bending behaviour is observed at the bond area of samples were made for a study of Fatigue crack growth in bonded repair of damaged aircraft structure. Analysis on metallic aircraft structures development and qualification aspect, indicated secondary bending structures curvature from both single side and double side doubler in the repair area did influence the behaviour; they recommended that would require detailed investigation. One way of minimising the level of residual stress is to cure the repair at the lowest possible temperature. FM-73K is a fracture toughened adhesive used for bonded repairs that are typically cured at 120°C for 1 hour. However, during the curing process, adhesively bonded composite/metal laminate structures that are held at elevated temperatures over 120 °C as per data sheet recommendation for material processing, very high residual stresses could build up because of the difference in coefficients of thermal expansion (CTE) for different materials. This thermal mismatch results in delamination or debonding of materials, which facilitates fatigue crack growth in the polymer/metal interface.