This paper is categorized into two parts. (1) A frame work to design the aircraft wing structure and (2) analysis of a morphing airfoil with auxetic structure. The developed design frame work in the first part is used to arrive at the sizes of the various components of an aircraft wing structure. The strength based design is adopted, where the design loads are extracted from the aerodynamic loads. The aerodynamic loads acting on a wing structure are converted to equivalent distributed loads, which are further converted point loads to arrive at the shear forces, bending and twisting moments along the wing span. Based on the estimated shear forces, bending and twisting moments, the strength based design is employed to estimate the sizes of various sections of a composite wing structure. A three dimensional numerical model of the composite wing structure has been developed and analyzed for the extreme load conditions. Glass fiber reinforced plastic material is used in the numerical analysis. The estimated natural frequencies are observed to be in the acceptable limits. Furthermore, the discussed design principles in the first part are extended to the design of a morphing airfoil with auxetic structure. The advantages of the morphing airfoil with auxetic structure are (i) larger displacement with limited straining of the components and (ii) unique deformation characteristics, which produce a theoretical in-plane Poisson’s ratio of–1. Aluminum Alloy AL6061-T651 is considered in the design of all the structural elements. The compliance characteristics of the airfoil are investigated through a numerical model. The numerical results are observed to be in close agreement with the experimental results in the literature.

Design concepts of an aircraft wing: composite and morphing airfoil with auxetic structures

Assessment of residual strength of repaired solid laminate composite materials through mechanical testing

This paper investigates the structural behavior of the wing subjected to the aerodynamic loads during the flight using finite element analysis of wing flexure deformation. In this work, three different types of wing models are established. Material characteristics, the wing structure, and 1 design principle have been taken into account. The assembly of the wing model 2 consists of the thin skin, two spars, and the multi-ribs. The two spars consist of primary and secondary spars. For this study, NACA 23015 is chosen as the baseline airfoil as this airfoil is very similar to the customized airfoil being used in Airbus A320. Two spars mainly bear the bending moment and shear force, which are made of titanium alloy to ensure sufficient rigidity. The skin and wing ribs are made of aluminum alloy to lighten the structural weight; a static structural analysis is applied. Total deformation, equivalent elastic strain, and equivalent von Mises stress are obtained to study the wing’s structural behavior. Furthermore, the modal analysis is then applied. The natural frequencies and the modal shape of the wing for three orders are obtained through the pre-stress modal analysis. The modal analysis results help designers minimize excitation on the natural frequencies and prevent the wing from fluttering. According to the results, designers can emphasize strengthening and testing the stress concentration and large deformation area.

Structural analysis of three-dimensional wings using finite element method

This paper discusses the size optimization of the fuselage frame used in a-19-passenger amphibious aircraft using the finite element method. The size optimization was performed to obtain mass reduction against specified displacement and buckling constraints. The finite element model for the fuselage frame used shell elements and rod elements. The loads subjected to the fuselage frame were based on several landing conditions: bow landing, unsymmetric bow landing, and step landing. The mass reduction was achieved by optimizing the area of the rod elements. The constraints were maximum of 1% additional displacement than the initial design and a buckling factor of more than or equal to one. Size optimization results showed that the mass reductions were -16.06% for bow landing, -47.36% for unsymmetric bow landing, and -22.19% for step landing. The optimized design for bow landing was selected to be applied in the fuselage frame to obtain mass reduction and fulfill the constraints. 

Size optimization of the fuselage frame used in A-19-passenger amphibious aircraft using finite element method

Stiffened panel structural optimization on wing skin of “WHALE” aircraft with local and global buckling criteria

Low cost and less weight are the two primary objectives of any aircraft structure. Efficient design of aircraft components is therefore, required to reduce cost and weight of the aircraft structure. For components with compressive loadng, Ribs and stringer spacings and stringer cross-section play a major role in achieving less weight. The main objective of the present study was aimed at establishing optimum stringer and ribs spacings and stringer cross-section for minimum weight of buckling design driven components using FEM packages. The analysis of effect of plate with ribs and stringer spacing was modeled in FEM to study the effect of these on the weight of aircraft structure using aluminum as material. The analysis was carried out through linear buckling analysis of the model. The software tools used were HyperMesh as a pre and post processor and Radioss as a solver.

N. Lohith

Papers

Effect of ribs and stringer spacings on the weight of aircraft structure for aluminum material