Fatigue and damage tolerance issues of Glare in aircraft structures

Airworthiness regulations require for most structural parts of the aircraft a damage tolerance design philosophy. The fatigue and damage tolerance issues related to the application of the fibre metal laminate Glare in the upper fuselage skin structure are discussed in this paper. As a result of the laminate lay-up (thin aluminium layers with in-between fibre/epoxy layers), the actual stresses in the aluminium layers are higher compared to the applied laminate stresses, resulting in a reduction of crack initiation life compared to monolithic aluminium. However, the crack propagation life is significantly longer than monolithic aluminium due to the bridging by the intact fibres over the fatigue crack. It is discussed that the prediction methods developed for monolithic aluminium can be modified for Glare. Durability issues with respect to moisture absorption and temperature effects are discussed as well as the residual strength of Glare. The through the thickness effects of fatigue in Glare joints are explained and compared to the monolithic aluminium behaviour. Because of the increased fatigue crack propagation life and the higher residual strength of Glare compared to aluminium, higher design allowable stresses are possible, increasing the inspection intervals.