ISSN: 2319-8753 International Journal of Innovative Research in Science, Engineering and Technology Vol. 2, Issue 6, June 2013 Copyright to IJIRSET www.ijirset.com 2471 OPTIMIZATION OF AIRCRAFT WING WITH COMPOSITE MATERIAL Shabeer KP 1 , Murtaza M A 2 PG student, Department of Mechanical Engineering, The Oxford College of Engineering, Bangalore, Karnataka, India 1 Professor, Department Of Mechanical Engineering, the Oxford College Of Engineering, Bangalore, Karnataka, India 2 Abstract: The objective of this paper is to develop an accurate model for optimal design through design the structure of wing that combine the composite (Skins) and isotropic materials (all other structures) and compare this with the same wing made by changing the orientation of composite ply orientation in skin. The optimum design for each wing with different ply orientation can be obtained by comparing stress and displacement. Structural modelling is completed with the help of CATIA V5, each components moddeled separately and assembled using Assembly workbench of CATIAV5, this assembly is then converted to IGS file. Finite element modelling is completed in MSc Patran using the IGS file as geometry, the element type used for meshing was 2D shell elements with QUAD4 element topology and different parts are connected using RBE2 connection. Static analysis done using MSc Nastran. The finite element model obtained is analysed by applying an inertia force of 1g and then aerodynamic result (lift) is used to simulate the wing loading on the wings. Optimum design is found by tabulating stress and displacement for each ply combination Keywords: Composite Wing, Modelling in CATIA V5, Finite element Analysis in Nastran, Optimum ply orientation. І.INTRODUCTION The critical element of aircraft is the design of the wings. Several factors influence the selection of material of which strength allied to lightness is the most important. Composite materials are well known for their excellent combination of high structural stiffness and low weight. Because of higher stiffness-to-weight or strength-to-weight ratios compared to isotropic materials, composite laminates are becoming more popular. Composite structures typically consist of laminates stacked from layers with different fiber orientation angles. The layer thickness is normally fixed, and fiber orientation angles are often limited to a discrete set such as 0°, ±30°, ±45°, ±75°, and 90°. This leads to different combinations of ply orientation and among that one will gives the better results , that is the optimized design for composite structures. A unidirectional laminate is a laminate in which all fibers are oriented in the same direction, cross-ply laminate is a laminate in which the layers of unidirectional lamina are oriented at right angles to each other and quasi-isotropic laminate behaves similarly to an isotropic material; that is, the elastic properties are same in all direction. Unidirectional composite structures are acceptable only for carrying simple loads such as uniaxial tension or pure bending. In structures with complex requirements of loading and stiffness, composite structures including angle plies will be necessary. Since each laminate in the composite material can have distinct fibre orientations which may vary from the adjoining laminates, the optimum ply orientation is also obtained as a result of the parametric study conducted using NASTRAN finite element package by varying the orientation sequence in the composite. II. GEOMETRICAL CONFIGURATIONS The wing design is an iterative process and the selections or calculations are usually repeated several times. A variety of tools and software based on aerodynamics and numerical methods have been developed in the past decades, there by a reduction in the number of iterations is observed. Normally two spar construction is common in transport aircraft wing design. The spar near to the leading edge of the wing is called as front spar and the spar closer to the aft portion of the wing is called as rear spar of the wing. One end of the spar near the root of the wing is connected to the fuselage called root of wing, the other end towards the tip of the wing is a free end. This configuration is very similar to the cantilever beam arrangement in any engineering structure. Spars and Ribs are connected using L angle fittings. Figure 1 below shows the Location of Spar and Ribs from root of wing and Figure 2 shows the complete wing structure modelled in CATIA V5.