Kevin D Potter

In-situ measurement of chemical shrinkage of MY750 epoxy resin by a novel gravimetric method

A novel approach has been developed to measure in-situ chemical shrinkage of epoxy resins at the temperature of cure, during which the epoxy resins pass through liquid, rubbery and glassy states. A small sample of MY750/HY917/DY073 epoxy resin system, sealed in a thin-walled silicone bag, was suspended in a pot of silicone fluid and weighed independently of the silicone bath. The buoyancy of the sample was monitored as its density increases with respect to the constant density fluid during isothermal cures at three different temperatures. The relationship between the chemical shrinkage and degree of cure was deduced from a cure kinetics model for the resin. The good match of the results for the three different cure cycles suggests that chemical shrinkage is only a function of degree of cure regardless of time and temperature. A bi-linear relationship was fitted to the experimental data. The break point is at a degree of cure of about 28%, with corresponding chemical shrinkage of 3%. This point is linked to the gel point of the resin, which was measured as approximately 25% degree of cure in previous work. Total cure shrinkage of 6.9% for the MY750 resin system was obtained by extrapolating the results to a degree of cure of 100%. The method is sensitive, reliable and keeps the resin stress-free; therefore it should be applicable to a wide range of materials.

Development of curvature during the cure of AS4/8552 [0/90] unsymmetric composite plates

Abstract The development of residual stresses in [0/90] unsymmetric flat laminates (AS4/8552 composite system) was monitored by stopping the cure cycle at pre-determined points and evaluating the related level of deformation. Cylindrical shapes of consistent curvature are obtained and the “cured” curvature can be eliminated by re-heating samples up to a certain temperature, designated as the stress free temperature, T sf . The stress free temperature has also been measured after the interrupted cure cycles: the evolution of the stress free temperature during the cure cycle allowed the vitrification point to be estimated. Both curvature and T sf increase with increasing cure but level off after the vitrification point, attaining a constant value. The stress free temperature of samples cured beyond vitrification is systematically higher than their cure temperature. It is evident that a certain percentage of non-thermoelastic stress is present in the structure, possibly due to resin chemical shrinkage. The “cured” curvature is mostly driven by the transverse coefficient of thermal expansion, α 2 . From curvature–temperature profiles, it is found that α 2 is almost constant during the cycle, below T g . Post-cure of laminates that have not been fully cured tends to increase their “residual” curvature, but the effect of the post-cure is relatively small for specimens cured beyond the vitrification point.

Composite corrugated structures for morphing wing skin applications

Composite corrugated structures are known for their anisotropic properties. They exhibit relatively high stiffness parallel (longitudinal) to the corrugation direction and are relatively compliant in the direction perpendicular (transverse) to the corrugation. Thus, they offer a potential solution for morphing skin panels (MSPs) in the trailing edge region of a wing as a morphing control surface. In this paper, an overview of the work carried out by the present authors over the last few years on corrugated structures for morphing skin applications is first given. The second part of the paper presents recent work on the application of corrugated sandwich structures. Panels made from multiple unit cells of corrugated sandwich structures are used as MSPs in the trailing edge region of a scaled morphing aerofoil section. The aerofoil section features an internal actuation mechanism that allows chordwise length and camber change of the trailing edge region (aft 35% chord). Wind tunnel testing was carried out to demonstrate the MSP concept but also to explore its limitations. Suggestions for improvements arising from this study were deduced, one of which includes an investigation of a segmented skin. The overall results of this study show that the MSP concept exploiting corrugated sandwich structures offers a potential solution for local morphing wing skins for low speed and small air vehicles.

Loss of bifurcation and multiple shapes of thin [0/90] unsymmetric composite plates subject to thermal stress

Abstract The shape of thermally loaded thin [0/90] composite plates predicted by the classical lamination theory is a saddle, with a value of curvature independent of the in-plane dimensions of the structure. Including the effect of geometrical nonlinearities reveals the occurrence of bifurcation of the saddle solution for a critical value of in-plane dimensions (or temperatures), depending on material properties. In the post-bifurcation regime, the saddle shape becomes unstable, while two cylindrical configurations develop on the stable branches. One cylinder can be snapped into another by applying an external force. In the present paper the behaviour of [0/90] plates under thermal stress is simulated with an ABAQUS Finite Element Model and results are compared to a Rayleigh–Ritz model. For square plates, results from the FEM model are in good agreement with the Rayleigh–Ritz model and only small discrepancies are found. The introduction of imperfections, such as asymmetry in thickness, produces a loss of bifurcation which is predicted by the Rayleigh–Ritz and by the ABAQUS model. The range of existence of the cylindrical solutions can also be predicted. For rectangular plates the FEM model predicts a loss of bifurcation which is not found by the Rayleigh–Ritz model. It is found that the length-to-width ratio (aspect ratio) plays a crucial role in the existence and stability of the cylindrical shapes. Numerical simulations are supported by experimental results on thin [0/90] unsymmetric plates with length-to-width ratio equal to 10.

Understanding and control of adhesive crack propagation in bonded joints between carbon fibre composite adherends I. Experimental

Carbon fibre composites are being widely considered for many classes of heavily loaded components. A common feature of such components is the need to introduce local or global loads into the composite structure. The use of adhesive bonding rather than mechanical fasteners offers the potential for reduced weight and cost. However, such bonded joints must be shown to behave in a predictable and reliable way. A major aspect of this is to demonstrate that the progress of cracks through the bonds is well understood. The work presented here illustrates one possibility for establishing a measure of control over that crack propagation.

The sensitivity of a Weibull failure criterion to singularity strength and local geometry variations

Previous experimental evidence has shown scale sensitivity of adhesives both in terms of ultimate stress and strain to failure; this sensitivity may be modelled using Weibull statistics. The implications of this on the sensitivity of strength predictions to mesh density around singularity points and small changes in local geometry is investigated using an idealised model of an enclosed corner in an adhesive joint. The sensitivity of two other failure criteria are also compared to the Weibull method. It is seen that there is a critical relationship between the Weibull shape parameter, m, and the singularity strength, λ. This relationship marks the transition between mesh insensitivity and mesh sensitivity, even with a singular stress (or strain) field present. If the singularity is removed by localised rounding, then the Weibull method of failure prediction shows much less sensitivity to the exact local geometry variation than other methods of failure prediction.

Bistable Prestressed Symmetric Laminates

The bistability of unsymmetric cross-ply [0n/90n] T laminates has already been investigated in much detail. In this work a new type of bistable laminate is presented which has a symmetric lay-up. Bistability derives from an unsymmetric fiber prestress applied to the laminate. The experimental procedure used to apply this fiber prestress is presented in detail. Experimental results are compared with analytical and finite element models which have the ability to model fiber prestress accurately. As well as having minimal hygrothermal variability, it is noted that the snap-through loads for a prestressed symmetric laminate can be much higher than its unstressed [0n/90n]T equivalent.

The application of thermally induced multistable composites to morphing aircraft structures

One approach to morphing aircraft is to use bistable or multistable structures that have two or more stable equilibrium configurations to define a discrete set of shapes for the morphing structure. Moving between these stable states may be achieved using an actuation system or by aerodynamic loads. This paper considers three concepts for morphing aircraft based on multistable structures, namely a variable sweep wing, bistable blended winglets and a variable camber trailing edge. The philosophy behind these concepts is outlined, and simulated and experimental results are given.

The early history of the resin transfer moulding process for aerospace applications

Abstract The resin transfer moulding (RTM) process has been the subject of a great deal of practical and theoretical development for aerospace applications since the early 1980s. This article looks at the very early developments of RTM in an aerospace setting. This development took place over a few years at the start of the 1950s. By 1956 almost all the features of RTM for aerospace applications had been introduced in a series of six patents. This achievement was made without any of the theoretical infrastructure now considered critical and was the work of a small group within a single company. The developed technology dropped from view in the general aerospace composites community and had to be redeveloped 25 years after the last patent was applied for.

Understanding and control of adhesive crack propagation in bonded joints between carbon fibre composite adherends II. Finite element analysis

Carbon fibre composites are being widely considered for many classes of heavily loaded components. A common feature of such components is the need to introduce local or global loads into the composite structure. The use of adhesive bonding rather than mechanical fasteners offers the potential for reduced weight and cost. However, such bonded joints must be shown to behave in a predictable and reliable way. A major aspect of this is to demonstrate that the progress of cracks through the bonds is well understood. The simulation work presented here complements the experimental work presented in Part I. The observed failure processes and their sequence are successfully described and modelled.