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Towards a predictive design methodology based on the physical modelling of the fracture of fiber composite
No full textA predictive design methodology based on modelling the fracture stress (notched tensile strength) and post-fatigue residual strength of laminated fiber composites is presented. The approach is based explicitly on the development of models of the physical processes by which damage accumulates at a notch-tip and the application of these models to cross-ply laminates for a variety of material systems, including thermosetting and thermoplastic matrices containing carbon, glass and Kevlar fiber reinforcements. The effects of temperature and humidity on composite fracture can also be examined in the context of this modelling strategy.
A pre-requisite of the model is that it has to be calibrated for each material system by performing tensile tests on notched and unnotched cross-ply laminate. From this initial calibration, which takes relatively little time, it is possible to apply the model to a prediction of the dependence of fracture stress on notch size; to an understanding of the effects of laminate stacking sequence (within the same cross-ply family) on fracture stress; and to provide insight into the effects of thermal or load cycling history on fatigue damage-growth and residual or fatigue strength.
The advantages and deficiencies of this modelling strategy are assessed, as well as the applicability of such a physical modelling approach to the predictive design and failure of composite materials in general
Fatigue damage mechanics of composite materials III: prediction of post-fatigue strength
No full textA damage-based model for post-fatigue notch strength is presented. The model is an extension of a method developed previously to predict the notch strength of laminated composites. A simple finite element representation of the notch tip damage zone is used to obtain the localized damage-modified stress distribution. A uniaxial tensile stress failure criterion is applied to the 0° plies from which the laminate strength is evaluated. In conjunction with the fatigue damage growth law described in Part II, residual strength is calculated as a function of the applied loading conditions, specimen geometry and lay-up for (90/0)s, (90/0)2s and (902/02)s T300/914C carbon-fibre/epoxy laminates subjected to tension-tension fatigue cycling
Fatigue damage mechanics of notched graphite: epoxy laminates
No full textA modeling approach is presented that recognizes that the residual properties of composite laminates after any form of loading depend on the damage state. Therefore, in the case of cyclic loading, it is necessary to first derive a damage growth law and then relate the residual properties to the accumulated damag
The fatigue damage mechanics of notched carbon fibre/PEEK laminates
No full textA model is presented for the strength, post-fatigue residual strength and damage propagation in notched, cross-ply carbon fibre/polyetheretherketone (PEEK) laminates. Fracture mechanics principles are used to predict quasi-static damage growth, and the application of a Paris law permits extension to fatigue damage. Strength is predicted by applying a failure criterion based on the tensile stress distribution in the 0° plies, as modified by damage (either quasi-static or fatigue). The volume dependence of strength is included by using a simple Weibull distribution. The parameters of the model are determined from independent experiments. Good agreement with experimental results is obtained. Comparisons are made with previous results from carbon fibre/epoxy laminates. The behaviour of the carbon fibre/PEEK is similar, although the extent of delamination and matrix cracking is reduced owing to the higher inherent toughness of the matrix
Fatigue damage mechanics of composite materials II: a damage growth model
No full textA fatigue model has been developed for damage growth at a notch tip in carbon fibre/epoxy laminates. The damage is modelled as a series of interacting matrix cracks in various forms: splitting, delamination and transverse ply cracking. The extent of fatigue damage can be successfully predicted for a family of (90i/0j)ns and (90/±45/0)s laminates. In this second paper of four, a damage-based model is proposed which can explain the effect of cyclic tensile loading on the post-fatigue strength and stiffness of a notched laminate
Fatigue damage mechanics of composite materials part IV: prediction of post-fatigue stiffness
No full textA model has been developed for predicting the stiffness of cross-ply carbon fibre composite laminates containing a notch from which damage, in the form of matrix cracks, splits and delaminations, has grown. A combination of experimental and theoretical results have been used to deduce appropriate degraded stiffness properties for the damaged regions of the laminate. These degraded stiffnesses have then been incorporated into a finite element representation of the notched laminate to predict the overall stiffness. Agreement with experimental data is satisfactory
Modelling the compressive response behaviour of monolithic and sandwich composite structures
No full textFailure in compression of fiber composite laminates with an open hole is by the initiation and growth of a microbuckle from the edge of the hole. The geometric inhomogeneity induces fiber rotation under increasing applied load; deformation localizes within a band and a microbuckle is initiated. The microbuckle then propagates in a stable manner for 2–3 mm and the component fails at a higher load than the initiation load. This process in carbon-epoxy and carbon-PEEK laminates was compared to an equivalent crack containing cohesive stresses. It is a crack bridging analysis that predicts the size of the buckled region as a function of the applied load, with the local stress supported by the buckled fibers decreasing linearly with the closing displacement of the microbuckle. The model is able to predict successfully the effects of hole size and lay-up upon the compressive strength and has been incorporated into a user-friendly computer program. Tests had been performed to obtain both the laminate un-notched strength and the compressive energy release rate associated with fiber microbuckling, which are required as the model's input. From a design point of view, it is desirable to predict these laminate properties from the mechanical properties of the fibers and the matrix and from the lay-up geometry
Design and performance of novel aircraft structures with folded composite cores
No full textThis chapter focuses on novel sandwich structures with open cellular composite cores manufactured by folding thin sheet base materials into a three-dimensional structure. Folded composite cores or ‘foldcore’ manufactured from resin impregnated aramid paper have similar densities and mechanical properties to Nomex honeycomb. Furthermore, they allow ventilation to prevent moisture build-up due to their open cell design and can be manufactured cost-efficiently in a continuous process. In aircraft fuselages, a sandwich design concept could yield significant weight savings compared to an aluminium reference fuselage, through increasing frame spacing and elimination of stringers.
The foldcore concept is described with typical base sheet materials, core geometries, and fabrication technology suitable for continuous manufacturing processes. Folded core properties and design are discussed with test methods used for measuring thin base sheet properties and determining core through-thickness compression and shear failure modes and strength properties. Core design is based on micromechanics cell models used in FE methods for simulating progressive damage and collapse mechanisms to provide core properties for use in sandwich structural analysis and design of sandwich structures. Finally, the damage tolerance of sandwich panels with carbon fibre/ epoxy skins and aramid fibre/phenolic folded core is discussed, based on experimental investigations by drop tower and gas gun impact tests. Comparison between observed and computed failure behaviour for a range of impact load cases shows good agreement, indicating that the FE methods could provide the basis for design and certification of these advanced aircraft sandwich structures
2.13 Multiaxial Fatigue of Composites: Experimental Evidences and Life Prediction Methodology
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