Finite Element Analysis of transversely tapered [0/90] non-crimp fabric (NCF) representative of realistic laminate wing covers subjected to compressive loading was undertaken. The main objectives were to determine the effect on laminate stress distribution of contesting existing global taper ratio heuristic guidelines of 20:1. Positive challenge of these guidelines is seen as an enabler of mass reduction and therefore subsequent potential commercial gains. The laminates studied consisted of a 48-ply thick section and a 32-ply thin section comprised of 0.25mm thick NCF carbon/epoxy lamina. This configuration characterised a 12mm-8mm reduction in thickness creating a case which was structurally representative of the taper arrangement commonly associated with aircraft wing covers. The geometry of the laminate was created with a global taper ratio of 20:1 analysed and subsequently re-modelled and re-analysed with a global taper ratio of 10:1. A number of surveys were undertaken on the Finite Element Model both holistically across the laminate and in detail at the ply drop locations and these investigations failed to highlight significant increases in stress as a result of reducing the global taper ratio. The work undertaken suggests that a decrease in global taper ratio is possible without noticeable increase in associated stress and this could enable 50% reductions in mass in the taper region of composite structures to be realised in an ever marginalised mass receding opportune environment.
- B. Stress Concentrations
- C. Finite element analysis (FEA)