Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

Abstract

The present study analyses an aircraft composite fuselage structure manufactured by the Liquid Resin Infusion (LRI) process and subjected to a compressive load. LRI is based on the moulding of high performance composite parts by infusing liquid resin on dry fibres instead of prepreg fabrics or Resin Transfer Moulding (RTM). Actual industrial projects face composite integrated structure issues as a number of structures (stiffeners, …) are more and more integrated onto the skins of aircraft fuselage.

A post-buckling test of a composite fuselage representative panel is set up, from numerical results available in previous works. Two stereo Digital Image Correlation (DIC) systems are positioned on each side of the panel, that are aimed at correlating numerical and experimental out-of-plane displacements (corresponding to the skin local buckling displacements of the panel). First, the experimental approach and the test facility are presented. A post-mortem failure analysis is then performed with the help of Non-Destructive Techniques (NDT). X-ray Computed Tomography (CT) measurements and ultrasonic testing (US) techniques are able to explain the failure mechanisms that occured during this post-buckling test. Numerical results are validated by the experimental results.

Introduction

A representative test of a panel can be specified in terms of a compressive [1], [2] or shear loading state [3]. A pressurised test at the panel level is more difficult to assess since it needs a combined loading [4]. Several studies can be found concerning a compressive post-buckling test of co-cured or co-bonded stiffened panels, manufactured in prepregs. It is known that opposite buckling waves – symmetric or antisymmetric – are the areas, where delamination initiates [5], [6], leading to the loss of connection between skin and stiffeners [1]. In the case of excellent bonding, severe stress concentrations may cause delaminations and fracture in the skin or in the stiffener flanges at these locations [1].

This paper is the experimental investigation of a previous work which was the numerical modelling of this panel [7]. It aims to show the particular behaviour of a panel manufactured by a resin infusion process [8]. Numerical models consider the failure initiation anywhere through the thickness, rather than the damage initiates at the interface. A conventional testing machine can highlight the decohesion phenomenon in the skin–stiffener interface, by a uni-axial compressive test that makes the panel buckles. No strain gages are used because the displacement fields serve at correlating the numerical and experimental results through the use of stereo-DIC techniques, being tried and tested at the laboratory for more than 10 years.

Before the test, a newly made model that takes into account a layup modification of the stiffeners has led to redefine the load at collapse F_max = 127.026 kN for a shortening UX = 2.40 mm. This model is defined as the reference model. The load at collapse will be used to set-up the test. Geometric imperfections of the skins are not considered since they have an almost negligible effect on the load at collapse [9]. The side stiffeners provide enough rigidity so that the panel borders do not buckle [7].

The critical buckling load F_cr (corresponding to skin local buckling) has been recalculated by the linear buckling analysis [7] on the reference model. It equals 90.4 kN, which is close to the end of the linear behaviour in the non linear analysis, where F_cr is approximately 90 kN. Now, numerical results are concordant with the buckling mechanism [10]: from the critical buckling load, the skin local buckling causes a load redistribution in the stiffeners until collapse, and the behaviour becomes nonlinear. Thus linear models cannot be used anymore.

These mechanisms given by numerical results need to be validated by experiment. First of all, the experimental test is set-up. The panel and the tool placement is carefully checked so that the panel is properly lined up with the load direction of the testing machine. Then numerical and experimental results are correlated through the out-of-plane displacement field (UZ) measured by two stereo-DIC measurement systems that are positioned on each side of the panel. X-ray CT and US techniques are used to analyse the failure zones on the panel surface and through the thickness of the stiffener flanges, allowing numerical and experimental results to be corroborated.