Laminated composite structures are widely used for structural applications, owing to their tailorable mechanical response. These structures are made of tens or hundreds of plies of various materials, each having a unique orientation, called the layup. The layup is a primary parameter in the design of composite structures. The numerous design variables involved in defining the layup of laminated structures are decided based on expert intuition, often resulting in overweight designs, as analysing all possible combinations may be time-consuming and sometimes impossible. Many researchers have worked on optimising the fibre orientations; however, these were limited to the problems whose analytical equations were available. Analytical models representing the behaviour of complex problems are often unavailable, such as the wing of a typical unmanned aerial vehicle. Such structures can be efficiently modelled using numerical methods such as finite element (FE) modelling, and these FE models can be further used for optimisation. The Design Optimisation module of general-purpose ANSYS has the capability of optimising such problems. However, this module is designed for the optimisation of systems made from isotropic material. This paper presents a two-step strategy for the weight optimisation of an aircraft wing fully made of composite materials. A code was developed which can cater for the range of wing span, number of ribs, spar width, and location of spars. The optimisation design variables considered were the number of composite layers, layer orientation, and their stacking sequence. The number of layers in each structural component was optimised, followed by the optimisation of layer sequence and orientation of each lamina. The weight of the specific wing was reduced by 45%, and the maximum stress values also lowered from 421 MPa to 330 MPa, meeting the material strength limits. The method presented is easily implementable for a wide range of problems, materials, and loading conditions.