The compressive behaviour of bonded patch repaired composite laminates is examined. A non-linear stress analysis is performed on a double-lap joint in order to identify critical joint parameters and design an efficient external patch repair. It is found that oversized patches not only increase the structure's weight but also increase the stress concentrations in the repaired region which can cause premature failure. Reducing the patch thickness near the edges of the overlap and increasing the local adhesive thickness decreases the stress concentration in both shear and peel stresses. A three- dimensional finite element analysis is then performed to determine the stresses in the optimum repaired configuration and is used with a stress failure criterion to predict the ultimate failure load. Experimental measurements show that carefully designed bonded patch repairs can recover almost 80 per cent of the undamaged laminate strength.

As a first step in developing a health monitoring system, the effect of delamination on the modal frequencies of laminated composite beams has been investigated. A piezoceramic patch driven with a linear rapid frequency sweep was used to induce vibrations on the structure and its response registered via piezoelectric sensors. Modal frequencies were obtained using concepts of resonant ultrasound spectroscopy (RUS). Changes of the modal frequencies after delamination initiation, compared to those of a non-delaminated specimen, gave a good indication of the degree of damage, demonstrating the feasibility of using measured changes in the vibration characteristics to detect damage.

The present work examines the effect of resin ductility (varied as a function of temperature) on the compressive strength of unidirectional T800/924C carbon fibre-epoxy laminates. Tests are conducted in a screw-driven machine between room temperature and 100°C. Untabbed straight-sided specimens are used in a modified Celanese test rig; conventional serrated grip faces of the Celanese jig are replaced by spark-eroded inserts to eliminate adhesively bonded tabs on the specimen ends and minimize gripping region failures. Test results show that at approximately 80°C the failure mode switches from in-plane to out-of-plane fibre microbuckling. As the test temperature increases, the shear strength/stiffness of the resin is considerably reduced; this decreases the amount of side support for the fibres and reduces the strain level at which fibre buckling occurs. Recent fracture models are used to predict the compressive strength of the T800/924C system; agreement between theory and experiment is acceptable.

Mechanically fastened joints are the most common method of connecting structural components in aerospace structures. The skinto-spar/rib connections in a wing structure and the wing-to-fuselage connection are typical examples of bolted joints in aircraft primary structures. It is well-recognised that bolt fasteners can clamp joint parts together well and show a good load carrying capability. In this respect, a number of authors (1)(2)(3)(4) have detailed the design methods for bolted joints mostly under static loading conditions. However, drilling fastener holes in members inherently introduces a stress concentration near the hole and reduces the load carrying cross sectional area. A drilling process may also cause a rough surface finish in the bore of the fastener hole which is prone to fatigue crack initiations under cyclic loads.Aircraft structures are primarily constructed from high strength light alloys and composites as their low density provides optimum strength-to-weight ratio aerospace materials. Since safety is of paramount importance in aerospace vehicles, several investigations have been conducted with the aim to optimise the design of structural bolted joints (metallic and composites) so that catastrophic failure during the flight can be prevented (5)(6)(7)(8)(9)(10)(11) .A bolted joint is most commonly preloaded through an initial torque. When the torque is applied to the nut, the bolt is axially ABSTRACT Accurate stress and strain analysis in bolted joints is of considerable interest in order to design more efficient and safer aerospace structural elements. In this paper, a finite element modelling of aluminium alloy 7075-T6 bolted plates, which are extensively used in aircraft structures, is discussed. The ANSYS Finite Element (FE) package was used for modelling the joint and estimating the stresses and strains created in the joint due to initial clamping forces and subsequent longitudinal tensile loadings. A double-lap bolted joint with a single bolt and nut was considered in the study. A three-dimensional (3D) finite element model of the joint was generated, and then subjected to three different simulated clamping forces followed by different levels of longitudinal tensile load. 3D surface-to-surface contact elements were employed to model the contact between the various components of the bolted joint. Friction effects were considered in the numerical analysis; and moreover, the clearance between the bolt and the plates was simulated in the model. FE results revealed beneficial compressive stresses near the hole edge as a result of applying the clamping. It was found that a higher clamping force can significantly decrease the magnitude of the resultant tensile stress at the hole edge and also bearing stress in the joint when subjected to the longitudinal tensile load.

No abstract