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The architecture of 3D-printed parts made through fused deposition modelling (FDM) with raster infill resembles that of composite laminates. Classical lamination theory (CLT), the simplest model for composite laminates, has been proved successful for describing the stiffness properties of FDM parts, while strength modeling so far has been limited to unidirectional lay-ups. The aim of this paper is to show that CLT can be used to predict also FDM part failure.

Wood flour-filled polyester has been chosen as a model material. Unidirectional specimens oriented at 0°, 90° and ± 45° have been first characterized in simple tension to obtain the properties of the single layer. Next, two quasi-isotropic lay-ups, possessing different layer sequences, have been tested again in simple tension for CLT validation.

The measured properties are in good agreement with theoretical predictions, both for stiffness and strength, and an even better agreement can be achieved if a correction for taking the contour lines into account is implemented.

The paper shows that also the tensile strength of FDM parts can be predicted by using a mathematical model based on CLT. This opens up the possibility of using CLT for studying optimization of raster filled lay-ups, for example in terms of the best raster angles sequence, to better resist applied external loads.

A linear and geometrically non‐linear computation of a laminated composite torispherical shell subjected to internal pressure was performed by using the layered finite element whose formulation is based on degeneration principle. Geometric non‐linearity in terms of large deformations with total Lagrangian formulation was taken into account. The effect of the lamination schemes on geometric non‐linear behaviour and stress resultant distributions was analysed. The fibre directions have not a great influence on the shape of the load‐displacement curves. In contrast to the hoop stress resultant distribution, the moment distribution is significantly influenced by the lamination schemes. The influence of the lamination schemes on bending moments is greater in non‐linear than in linear computations. Likewise, the effect of the fibre orientation is greater on the hoop than on the meridional moment distribution. In unsymmetric laminated shells the values of the hoop moments exceed those of the meridional moments which is a considerable difference from metallic isotropic shells.

Reports on the MSc group design project of students at the College of Aeronautics, aerospace vehicle design in 1995. The students worked on advanced short take‐off and vertical landing of a combat aircraft. Details the project showing aircraft dimensions and design. Full assessment of the results is pending, but outlines a number of problems faced by the students.

A model for a symmetric three‐layer configuration is developed. This model refers to Cu/Invar/Cu (CIC) laminates. Calculated values for the coefficient of thermal expansion (CTE) are compared with literature values. The model is then extended to symmetric CIC‐Metalcore boards and its prediction is compared with experimental results. Shearing of FR‐4 is discussed.

Presents a simple heuristic optimization algorithm to determine weight‐minimal laminate structures. The algorithm is based on an inverted form of growth strategy, where material is removed in areas with low stresses. In the presented algorithm the removal of material is performed in a layerwise manner in areas with low stress or in areas where the layer orientation angle differs significantly from the principal stress direction. In order to get a production‐adapted structure, the layer orientation angles of the available individual plies are not modified and the material is removed only locally in the respective plies. Contrary to a more formal mathematical optimization no sensitivity analyses are needed by the procedure outlined so far and this keeps the corresponding numerical effort reasonably low. The structural analyses have been performed by the Finite Element Program ANSYS and the outlined heuristic algorithm has been implemented by ANSYS‐macros. Several examples of rectangular laminate plates show the effectiveness of the present algorithm.

This paper aims to present a systematic performance-oriented procedure to predict structural responses of composite layered structures. The procedure has a direct application in the preliminary design of aerospace composite structures evaluating the right and most effective material.

The aforementioned procedure is based upon the definition of stiffness invariants. In the paper, the authors briefly recall the definition and the physical explanation of the invariants, i.e. the trace; then they present the scaling procedure for the selection of the best material for a fixed geometrical shape.

The authors report the basic principles of the scaling procedure and several examples pertaining typical responses sought in the preliminary design of aeronautic structures

Typically, during early stages, engineers had to perform the daunting task of balancing among functional requirements and constraints and give the optimum solution in terms of structural concept and material selection. Moreover, preliminary design activities require evaluating different responses as a function of as less as possible parameters, ensuring medium to high fidelity. The importance of incorporating as much physics and understanding of the problem as early as possible in the preliminary design stages is therefore fundamental. A robust and systematic procedure is necessary.

The time/effort reduction in the preliminary design of composite structures can increase the overall quality of the configuration chosen.

Reduction in design costs and time.

In spite of the well-known invariant properties of composites, the application and extension to the preliminary design of composite structures by means of a scaling rule is new and original.

Layered manufacturing is an evolution of rapid prototyping (RP) techniques where the part is built in layers. While most of the previous applications focused on building “prototypes”, recent developments in this field enabled some of the prototyping methods to achieve an agile fabrication technology to produce the final product directly. A shift from prototyping to manufacturing of the final product necessitates broadening of the material choice, improvement of the surface quality, dimensional stability, and achieving the necessary mechanical properties to meet the performance criteria. The current study is part of an ongoing project to adapt fused deposition modeling to fabrication of ceramic and multi‐functional components. This paper presents a methodology of the mechanical characterization of products fabricated using fused deposition modeling.

The purpose of this study is to manufacture test boards for re-enacting plant or field situations where vacuum chamber for expelling gas bubbles and autoclave equipment would not be accessible. This research focuses on the examination and enhancement of tensile strength for the nanocomposites consisting of uniaxial glass fiber mats, nanoclay (NC) and epoxy.

The parameters considered are the weight content of Cloisite 15A NC, the volume of glass fiber (V_gf) and the direction of glass fibers (θ). The composites are made by hand lay-up technique and tested according to ASTM D 638 standard. Taguchi L9 orthogonal array is used to design the experiments.

The results imply that the orientation of fibers exhibited high significance with a p-value of 0.001 for the upgrade of strength. NC percentage and the volume of fiber have a low effect as the p-values obtained were 0.375 and 0.294. Confirmation tests were performed at the optimal levels of parameters and the outcomes were in the permissible range of the anticipated values of S/N ratio and mean tensile strength. The negligible effect of nanoclay is due to the lack of infusion of resin into the d-spacing of clay layers due to the low configuration settings of mixing conditions which was confirmed by XRD studies. The negligible effect of glass fiber volume is due to the void content and lack of stress transfer between fibers uniformly due to the void content and improper mixing of nanoclay.

The limitation of this study is that a low-speed mechanical stirrer was used to mix NC in the epoxy and the mixture was not subjected to vacuum and ultrasonication for degassing and deagglomeration.

These composites can be used as substitute materials in place of metallic parts in the aerospace and automobile sector. These composites can be used in civil structures instead of steel and concrete, which have low strength-to-weight ratio and where the requirement of strength is in the range of 60 to 390 MPa.

The composites can be used in a variety of applications, for example, structural works, automotive panels and low-cost housing.

This research gives an idea about the combined contribution of NC, V_gf and “θ” to the improvement of tensile strength of the glass-epoxy composite.

Among various efforts pursued to produce fuel efficient vehicles, light weight engineering (i.e. the use of low‐density structurally‐efficient materials, the application of advanced manufacturing and joining technologies and the design of highly‐integrated, multi‐functional components/sub‐assemblies) plays a prominent role. In the present work, a multi‐disciplinary design optimization methodology has been presented and subsequently applied to the development of a light composite vehicle door (more specifically, to an inner door panel). The door design has been optimized with respect to its weight while meeting the requirements /constraints pertaining to the structural and NVH performances, crashworthiness, durability and manufacturability. In the optimization procedure, the number and orientation of the composite plies, the local laminate thickness and the shape of different door panel segments (each characterized by a given composite‐lay‐up architecture and uniform ply thicknesses) are used as design variables. The methodology developed in the present work is subsequently used to carry out weight optimization of the front door on Ford Taurus, model year 2001. The emphasis in the present work is placed on highlighting the scientific and engineering issues accompanying multidisciplinary design optimization and less on the outcome of the optimization analysis and the computational resources/architecture needed to support such activity.

The purpose of this paper is to develop a general mathematical model for laminated curved structure of different geometries using higher-order shear deformation theory to evaluate in-plane and out of plane shear stress and strains correctly. Subsequently, the model has to be validated by comparing the responses with developed simulation model (ANSYS) as well as available published literature. It is also proposed to analyse thermal buckling load parameter of laminated structures using Green–Lagrange type non-linear strains for excess thermal distortion under uniform temperature loading.

Laminated structures known for their flexibility as compared to conventional material and the deformation behaviour are greatly affected due to combined thermal/aerodynamic environment. The vibration/buckling behaviour of shell structures are very different than that of the plate structures due to their curvature effect. To model the exact behaviour of laminated structures mathematically, a general mathematical model is developed for laminated shell geometries. The responses are evaluated numerically using a finite element model-based computer code developed in MATLAB environment. Subsequently, a simulation model has been developed in ANSYS using ANSYS parametric design language code to evaluate the responses.

Vibration and thermal buckling responses of laminated composite curved panels have been obtained based on proposed model through a customised computer code in MATLAB environment and ANSYS simulation model using ANSYS parametric design language code. The convergence behaviour are tested and compared with those available in published literature and ANSYS results. Finally, the investigation has been extended to examine the effect of different parameters (thickness ratios, curvature ratios, modular ratios, number of layers and support conditions) on the free vibration and thermal buckling responses of laminated curved structures.

The present paper intends to give sufficient amount of numerical experimentation, which may lead to help in designing of finished product made up of laminated composites. Most of the aerospace, space research and defence organisation intend to develop low cost and high durable products for real hazard conditions by taking combined loading and environmental conditions. Further, case studies might lead to a lighter design of the laminated composite panels used in high-performance systems, where the weight reduction is the major parameter, such as aerospace, space craft and missile structures.

In this analysis, the geometrical distortion due to temperature is being introduced through Green–Lagrange sense in the framework of higher-order shear deformation theory for different types of laminated shells (cylindrical/spherical/hyperboloid/elliptical). A simulation-based model is developed using ANSYS parametric design language in ANSYS environment for different geometries and loading condition and compared with the numerical model.