Series Solution-Based Approach for the Interlaminar Stress Analysis of Smart Composites under Thermo-Electro-Mechanical Loading

This paper introduces a new loading condition considering the combined thermo-electro-mechanical coupling effect in a series solution-based approach to analyze the free-edge interlaminar stresses in smart composite laminates. The governing equations are developed using the principle of complementary virtual work. The assumed stress fields satisfy the traction-free and free-edge boundary conditions. The accurate stress states of the composite structures are acquired through the procedure of generalized eigenvalue problems. The uniform temperature is employed throughout the laminate, and the electric field loading is applied to the symmetric piezo-bonded actuators to examine the combined effect of thermal and electrical stresses on the overall deformation of smart composite laminates. It was observed that the magnitude of the peeling stresses generated by mechanical loading was reduced by the combined thermal and electric excitation loading (up to 25.3%), which in turn resulted in expanding the service life of the smart composite structures. The proposed approach is implemented on three different layup configurations. The efficiency of the current methodology is confirmed by comparing the results with the 3D finite element (FEM) solution computed by ABAQUS.

Treatise on Sustainable Infrastructure Construction: Green Composites, Cross Laminated/Mass Timber, Wood Truss Connectors, Nondestructive Technologies, Health Assessment and Monitoring: Utility Poles and Geofoam

The understanding of the engineering performance of green laminated composites is necessary to the design of load bearing components in building and infrastructure construction, and packaging applications. These components are made of outer thin laminae called skins or faces and a thick inner layer called core. The use of bonding is unavoidable in the assembling of these composite products. Like all materials, the bonding materials have finite mechanical properties, e.g. stiffness, but when used in the literature, they are assumed perfectly rigid. That is an unrealistic assumption. Our analytical solutions change this assumption by using the real properties of bonding. In general, the analytical formulations are based on the equilibrium equations of forces, the compatibility of interlaminar stresses and deformation, and the geometrical conditions of the panels. Once solutions are obtained, the next step is to evaluate them. The numerical evaluations proved that perfect rigid bonding in laminated composites greatly underestimates the true performance. At low values of adhesive stiffness, the serviceability is multiple orders of magnitude of that at high values. The logical question is thus: what constitutes perfect bonding? The answer to this question lies in the core-to-adhesive stiffness. The lower the ration is the higher the error in using the rigid-bond theories. It is worth noting that green-composites in this chapter refer to components made of traditional materials such as wood, in addition to newly developed bio-based and bio-degradable and bio-based composites, made of renewable resources. In addition, bonding and adhesive are used interchangeably.

Weight minimization of fiber laminated composite beam for aircraft wing construction using exhaustive enumeration algorithm and numerical modeling

Purpose This paper aims to focus on calculating the number of layers of composite laminates required to take the applied load made up of graphite/epoxy (AS4/3501-6) which can be used in many industrial applications. Optimization for minimization of weight by variation in the mechanical properties is possible by using different combinations of fiber angle, number of plies and their stacking sequence. Design/methodology/approach Lots of research studies have been put forth by aerospace industry experts to improve the performance of aircraft wings with weight constraints. The orthotropic nature of the laminated composites and their ability to characterize as per various performance requirements of aerospace industry make them the most suitable material. This leads to necessity of implementing most appropriate optimization technique for selecting appropriate parameter sets and material configurations. Findings In this work, exhaustive enumeration algorithm has been applied for weight minimization of fiber laminated composite beam subjected to two different loading conditions by computing overall possible stacking sequences and material properties using classical laminate theory. This combinatorial type optimization technique enumerates all possible solutions with an assurance of getting global optimum solution. Stacking sequences are filtered through Tsai-Wu failure criteria. Originality/value Finally, through the outcome of this optimization framework, eight different combinations of stacking sequences and 24-ply symmetric layup have been obtained. Furthermore, this 24-ply layup weighing 0.468 kg has been validated using finite element solver for given boundary conditions. Interlaminar stresses at top and bottom of the optimized ply layup were validated with Autodesk’s Helius composites solver.

Experimental and numerical investigation of interface damage in composite L-angle sections under four-point bending

Curved laminates in aero-structures, such as the L-angle sections where webs and flanges meet, are prone to delamination due to high interlaminar stresses in these regions. Some efforts to investigate delamination in these structures can be found in the literature but commonly structures are limited to unidirectional layups or modelling approaches are constrained to the cohesive element based methods. In this work, multi-directional L-angle laminates were manufactured using unidirectional prepregs and tested under four-point bending load conditions to examine the interface damage. Acoustic emission technique was used to assist the capture of damage initiation and propagation. Three interface modelling strategies for predicting delamination, namely cohesive element, cohesive surface and perfectly bonded interface were used in the numerical study. The interface damage behaviour was successfully predicted by the simulation methods and differences among the strategies were compared.

The Hybrid Joints between an FRP Panel and a Steel Panel through Tubular Reinforcements: A Methodology for Interlaminar Stress Calculations

The advantages of laminates in terms of the chemical properties and mechanical properties/weight relationship have motivated several applications of fiber-reinforced plastic (FRP) composites in naval constructions due to the reduction in structural weight. This weight advantage has motivated multiple investigations dedicated to dissimilar material joints. We present a methodology for the interlaminar stress calculations of a tubular hybrid joint between an FRP panel and a steel panel through tubular reinforcements. The proposed formulas allow the estimation of the shear and normal stresses on the adhesive, which are generated in the bonding angle of the tubular hybrid joint. The stresses generated at the adhesive bonding ends influence on the adherent’s adjacent layer. A failure criterion is shown to check the accomplishment of the resulting stresses in the adherent laminate. Finally, the proposed formulas are validated using the finite element method and compared with the obtained interlaminar stresses.

Multiscale analysis of interlaminar stresses near a free-edge in a [±45/0/90]s laminate

A multiscale model for a [±45/0/90]s tape laminate under uniaxial extension was used to investigate the effect of modeling the heterogeneous microstructure near a free-edge. A random fiber arrangement was used for the 0° and 90° plies and homogenized properties for the 45° and −45° plies. The predicted interlaminar normal stress was compared to the prediction using a classical homogeneous model. When fibers and matrix were modeled discretely, the local stress state was shown to be sensitive to the proximity of fibers, which caused a complex stress distribution along the 0–90 ply interface. Next, the effect of reducing the size of region modeled at the microscale was investigated, since this would significantly reduce the computational effort. Reducing the region modeled at the microscale in the direction normal to the 0–90 ply interface to a size that was 25% and 10% of the ply thickness only changed the peak stresses by 3% and 8%, respectively. Reducing the microscale region in the direction normal to the free-edge to be one and two ply thicknesses in size did not have a significant effect on the predicted interlaminar normal stress at points within 75% of a ply thickness of the free-edge.