Hybrid Structures for Concentrated Load Transmission in Fibre Composites: Initial Experiments

Abstract

This study can be considered as an element of advanced hybrid structural concepts in composite aircraft design, supporting the development of hybrid titanium composite laminates (HTCL) and its transition region for local–reinforced / fastening–optimised CFRP structures. Focus is on similarities with existing FML, model adaptation and experimental validation. Investigations on fatigue crack initiation and propagation in HTCL show promising intermediate results; Classical laminate theory (CLT) is a suitable method for residual stress calculation as initial experiments show a good correlation with the CLT results. A crack bridging mechanism seems to occur, although the delamination area is not resolvable with the used technique of Xray computer tomography. The visualisation of delamination and in-situ techniques for measurement of crack lengths are special points of attention. INTRODUCTION The current generation aircraft in development consist of +50wt% (carbon) fibre composites as these materials couple a low density with extremely high directional strength and stiffness. Additionally, fibrous composites perfectly allow in-plane M. Bos (ed.), ICAF 2009, Bridging the Gap between Theory and Operational Practice, 1069–1081. © Springer Science+Business Media B.V. 2009 P. Vermeer, R. C. Alderliesten and R. Benedictus 1070 properties to be freely specified; generally by variation of fibre type, orientation and lay-up, so also a quasi-isotropic state can be obtained. However, a concentrated transmission of load strikes on inherent violations of the composite’s basic design; structural coupling by conventional mechanical fastening provokes the composite’s notch sensitivity, its low shear and bearing strengths and the critical dependence on the laminate lay-up-. Although increasing the (quasi-isotropic) joint thickness would make the design uncritical to bearing failure, this is in conflict with the aspiration to structural efficiency. In addition unfavourable effects and increasing complexity are often induced. As fibre metal laminates (FML) are known to be suitable for mechanical fastening [1], adopting the concept of FML seems to be an elegant solution for optimising the joining area of carbon fibre reinforced polymer (CFRP) structures. Hence, a local (inverted) hybrid titanium composite laminate (HTCL) is created by insertion of thin titanium foils in the CFRP laminate, replacing those fibre layers whose orientation is contributing least to load transfer [2] (figure 1). Figure 1: Schematic lay-up of a hybrid laminate with its corresponding areas HTCL (2×Ti / 3×CFRP) with hole, Transition zone and CFRP (0°/90°) This requires investigation of the transition zone from the CFRP laminate into the hybrid area and the behaviour of the HTCL under fatigue loading. The latter has a correlation to previous studies on FML for fuselage applications, and it is likely that the different constituents in HTCL affect the laminate’s mechanical behaviour. The current paper discusses initial experiments on the fatigue crack initiation and propagation in HTCL. These investigations are part of a larger research project funded by the German Federal Ministry of Economics and Technology called: development of new materials for fuselage applications (German acronym ENWERUM). The overall scientific objective of this research is to analyse, characterise and model the mechanical behaviour of the HTCL, the joining area and the transition region. Hybrid titanium composite laminates 1071 NOMENCLATURE α coefficients of thermal expansion Ni fatigue initiation life cycles C Paris equation coefficient Nf fatigue life to failure cycles da/d N crack growth rate R stress ratio E modulus of elasticity [MPa] σp,Ti,f stress in a layer, Ti, fibre G shear modulus [MPa] σ applied stress Kt stress concentration factor S stiffness matrix ΔK(t h) (threshold) stress intensity range Sp,Ti,f stiffness of a layer, Ti, fibre Ktip stress intensity factor at crack tip τ shear stress [MPa] Kff,br SIF far field, bridging ΔT difference curing temperature and RT m Paris equation exponent ν Poisson’s ratio MODELLING Autoclave curing at elevated temperatures (commonly 120–180°C) causes a significant residual stress state in the laminate as a result of the constituent’s different coefficients of thermal expansion. The actual stress level in the separate layers is an accumulation of the laminate’s curing stress and the (externally) applied load, to be calculated with the classical laminate theory as exemplary done by Homan [3] and summarised in Eqn. 1 to 4 (assuming no off-axis loading):