Experience with the finite element modelling of a full-scale test of a composite aircraft control surface

Abstract

A full-scale, co-cured, carbon fibre composite control surface representative of those found on mid-size, jet transport aircraft has been designed and tested. It was designed as a postbuckling blade-stiffened structure to reduce weight and improve operational performance compared with a honeycomb sandwich panel design traditionally used in such structures. The purpose of the tests was to demonstrate the validity of the design methodology by applying static limit and ultimate loads to the structure. The control surface, manufactured from prepreg tape, was successfully loaded to the ultimate design load without evidence of failure. Buckling initiated at approximately 48% of ultimate load with significant out-of-plane displacements observed. The global behaviour predicted by the finite element (FE) model of the test arrangement was in close agreement with the experimental results. Good agreement was also demonstrated with the local behaviour, as evidenced by the strain and buckling results. The inability of the FE analysis to capture complex snap-through mode change behaviour at 89% of ultimate load was identified as a limitation. The success of the testing program demonstrated the suitability of the design methodology for this type of structure.