Experiments and Analyses on Post-Buckling Behavior of Stringer-Stiffened Laminated Composite Helicopter Tailplane

The present paper describes the results of an experimental and numerical investigation on a composite laminated helicopter tailplane, whose lower panels are subjected to buckling phenomena during the life envelope. The structure is made of carbon fabric, carbon unidirectional and honeycomb. Two different types of lower laminate composite panels are investigated: Z stringer-stiffened panels and L stringer-stiffened panels. During the tests, the applied load is measured using a load cell, the displacement of the tailplane ends using potentiometers and the panels strains using strain gauges. Moire' fringes are installed to highlight the panels deformation in the lower part of the structure. At the beginning three Z stringer-stiffened and three L stringer-stiffened laminate composite lower panels are tested until 85% of the target load. Then the last L stringer-stiffened panel is tested until collapse. The buckling load happens at 63% of the target load, while the collapse happens at 90%. At the same time, the finite element analyses of the tailplane structures with both Z stringer-stiffened and L stringer-stiffened panels are performed, simulating the dynamic of a slow compression test, using LS-DYNA. The finite element results are compared with the experimental data, obtaining an excellent numerical-experimental correlation. Indeed the collapse of the structure is predicted at 93% of the target load and experimentally happened at 90%.