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A conceptual design method for composite material stiffened panels used in aircraft tail structures and unmanned aircraft has been developed to bear compression and shear loads.

The method is based on classical laminated theory to fulfil the requirement of building a fast design tool, necessary for this preliminary stage. The design criterion is local and global buckling happen at the same time. In addition, it is considered that the panel does not fail due to crippling, stiffeners column buckling or other manufacturing restrictions. The final geometry is determined by minimising the area and, consequently, the weight of the panel.

The results obtained are compared with a classical method for sizing stiffened panels in aluminium. The weight prediction is validated by weight reductions in aircraft structures when comparing composite and aluminium alloys.

The work is framed in conceptual design field, so hypotheses like material or stiffeners geometry shall be taken a priori. These hypotheses can be modified if it is necessary, but even so, the methodology continues being applicable.

The procedure presented in this paper allows designers to know composite structure weight of aircraft tails in commercial aviation or any lifting surface in unmanned aircraft field, even for unconventional configurations, in early stages of the design, which is an aid for them.

The contribution of this paper is the development of a new rapid methodology for conceptual design of composite panels and the feasible application to aircraft tails and also to unmanned aircraft.

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.

This paper's aim is to focus on the design, manufacture and test of a stiffened panel in composite material with integrated longitudinal foam‐filled stiffeners, spar and rib caps, using one‐shot liquid infusion (LI) process, reducing weight and number of subparts respect to metallic reference baseline P180 Avanti vertical fin.

Extensive activities in computational applications in order to improve the efficiency of the design process finite element analysis/structural sizing codes have led to an optimised engineering design process that resulted in a successful stiffened carbon fibre reinforced polymer panel design in terms of weight and number of parts with respect to the metallic baseline.

The composite panel has fulfilled all the design requirements (reduction of mass and number of parts with respect to the metallic reference baseline) overcoming the certification static test, and confirming the reliability of the theoretical analyses.

The composite aircraft components, conceived as unitized structure by one‐shot process, guarantee not only a mass reduction, compared to aluminium components, but assure also the reduction of the number of subparts and of the assembly process cycle time. On the other hand, the LI technology implies the development of more specific and advanced techniques to control the manufacturing and the weight.

The stiffened panel is the most used component in the aircraft structures; the solution shown in this work can find applications in many parts of an aircraft.

The results obtained in this work can be useful to those who work in aeronautical structural departments with the aim to reduce weight and subparts of the airframe.

Traditionally, an armor-grade composite is based on a two-dimensional (2D) architecture of its fiber reinforcements. However, various experimental investigations have shown that armor-grade composites based on 2D-reinforcement architectures tend to display inferior through-the-thickness mechanical properties, compromising their ballistic performance. To overcome this problem, armor-grade composites based on three-dimensional (3D) fiber-reinforcement architectures have recently been investigated experimentally. The paper aims to discuss these issues.

In the present work, continuum-level material models are derived, parameterized and validated for armor-grade composite materials, having four (two 2D and two 3D) prototypical reinforcement architectures based on oriented ultra-high molecular-weight polyethylene fibers. To properly and accurately account for the effect of the reinforcement architecture, the appropriate unit cells (within which the constituent materials and their morphologies are represented explicitly) are constructed and subjected to a series of virtual mechanical tests (VMTs). The results obtained are used within a post-processing analysis to derive and parameterize the corresponding homogenized-material models. One of these models (specifically, the one for 0°/90° cross-collimated fiber architecture) was directly validated by comparing its predictions with the experimental counterparts. The other models are validated by examining their physical soundness and details of their predictions. Lastly, the models are integrated as user-material subroutines, and linked with a commercial finite-element package, in order to carry out a transient non-linear dynamics analysis of ballistic transverse impact of armor-grade composite-material panels with different reinforcement architectures.

The results obtained clearly revealed the role the reinforcement architecture plays in the overall ballistic limit of the armor panel, as well as in its structural and damage/failure response.

To the authors’ knowledge, the present work is the first reported attempt to assess, computationally, the utility and effectiveness of 3D fiber-reinforcement architectures for ballistic-impact applications.

The maintenance of aircraft structure with lower cost is one of the prime concerns to regulatory authorities. The carbon fiber-reinforced polymer (CFRP) patches are widely used to repair the cracked structure. The demands and application of CFRP compel its price to increase in the near future. A distinct perspective of repairing the cracked aluminum panel with the hybrid composite patch is presented in this paper. The purpose of this paper is to propose an alternative patch material in the form of a hybrid composite patch which can provide economical repair solution.

The patch hybridization is performed by preparing the hybrid composite from tows of carbon fiber and glass fiber. Rule of hybrid mixture and modified Halpin–Tsai’s equation are used to evaluate the elastic constant. The stress intensity factor and interfacial stresses are determined using finite element analysis. The debonding initiation load is evaluated after testing under mode-I loading condition.

The hybrid composite patch has rendered the adequate performance for reduction of stress intensity in the cracked panel and control of interfacial stresses in the adhesive layer. The repair efficiency and repair durability of the composite patch repair was ensured by incorporation of the hybrid composite patch.

The studies involving patch hybridization for the application of composite patch repair are presently lacking. The influence of the patch stiffness, methodology to prepare the hybrid composite patch and effects of hybridization on the performance of composite patch repair is presented in this paper.

The genetic algorithm (GA) technique is widely used for the optimization of stiffened composite panels. It is based on sequential execution of a number of specific operators, including the encoding of particular design variables. For instance, in the case of a stiffened composite panel, the design variables that need to be encoded are: the number of plies and their stacking sequences in the panel skin and stiffeners. This paper aims to present a novel, implicit, heuristic approach for encoding composite laminates and, through its use, demonstrates an improvement in the optimization process.

The stiffened panel optimization has been formulated as a constrained discrete minimum-weight design problem. GAs, which use both new encoding schemes and those previously described in the literature, have been used to find near-optimal solutions to the formulated problem. The influence of the new encoding scheme on the searching capabilities of the GA has been investigated using comparative analysis of the optimization results.

The new encoding scheme allows the definition of stacking sequences in composites using shorter symbolic representations as compared with standard encoding operators and, as a result of this, a reduction in the problem design space. According to numerical experiments performed in this work, this feature enables GA to obtain near-optimal designs using smaller population sizes than those required if standard encoding schemes are used.

The approach to encoding laminates presented in this paper is based on the original heuristics. In the context of GA-based optimization of stiffened composite panels, the use of the new approach rather than the standard encoding technique can lead to a significant reduction in computational time employed.

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.

This paper seeks to address issues related to the development of a knowledge‐based engineering system for estimating manufacturing cost and weight of a composite structure at the conceptual stage of a design.

The system has been developed in the CATIA V5 knowledge environment and is applied to structures made of composite materials. At the conceptual stage of the design process, a structure is often represented by simple surfaces. The system adds the details necessary to accurately estimate weight and manufacturing cost using geometry and process‐based techniques. Knowledge captured from an expert was used to construct the knowledge base in the system.

It has been found that the system can provide continuous tracking of the weight and cost as the design evolves. Structural FEA and optimisation using MSC.NASTRAN have been integrated into the design process to enables the designer to conduct “what‐if” analyses to explore different design options involving geometry parameters such as the internal configuration of the structure.

The paper demonstrates that tools embedded in CAD systems can be expected to be able to facilitate the task of estimation of weight and manufacturing cost at the conceptual stage of the design process.

Describes preliminary structural design work on a notional uninhabited tactical aircraft (UTA), carried out at Cranfield University. UTAs are seen as an important future element of military fleets. A notional baseline requirement was derived, leading to the evolution of a design solution. The basic requirements for such a UTA are naturally highly classified but, although industry has been hesitant to comment, the baseline requirements and design solution developed herein are believed to be reasonable.

This paper aims to present a new robustness-based design strategy for thin-walled composite structures under compressive loading, which combines strength requirements in terms of the limit and ultimate load with robustness requirements evaluated from the structural energy until collapse.

In order to assess the structural energy, the area under the load-shortening curve between several characteristic points such as local buckling, global buckling, onset of degradation and collapse load is calculated. In this context, a geometrically nonlinear finite element analysis is carried out, in which the ply properties are selectively degraded by progressive failure.

The advantage of the proposed methodology is observed by analyzing unstiffened composite plates under compressive loading, wherein the lightest plate that satisfies both strength and robustness requirements can be attained.

As a practical implication, this methodology gives a new argument to accept the collapse load close to the ultimate load once robustness is ensured.

The structural energy is employed to investigate the robustness of thin-walled composite structures in postbuckling, and new energy-based robustness measures are proposed. In the design of composite structures, this innovative strategy might lead to a more robust design when compared to an approach that only accounts for the ultimate load.