This article presents a multi-objective optimization of a large aircraft composite wing subject to multiple constraints including strength, damage tolerance, and aeroelastic stability. Based on a preliminary design of the wing structure, the investigation demonstrated a multi-disciplinary design optimization process to replace the usual manual iteration of finite element analysis in the detailed design phase. For potential application, the optimization process was performed by making the full use of the commercial software MSC Nastran, which is widely employed in aerospace industry. The optimization procedure was divided into two stages according to the design objectives. This allows the designer to interact and make decision in the design process. The first stage was focused on the minimum weight optimization subject to the multiple constraints. The laminate ply thickness and orientation of the wing skins were taken as design variables. The optimized structure weight was reduced by 44.6 per cent. The second stage was focused on aeroelastic tailoring to reduce the wing gust response to the same level as the original value of the preliminary design with a minimum weight penalty. This led to the final optimized structure weight saving by 34.5 per cent. A post-process was performed to trim the optimized skin laminate layup and ply thickness under the manufacturing constraint. For a conservative design, an option of increasing the skin laminate thickness was made and led to a slight increase of the structure weight. However, the final weight saving is still over 30 per cent. A practical optimum design of the composite wing structure with significant weight saving can be achieved by a practical approach.

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A computer program for use in the conceptual stage of aircraft design has been developed. The program obtains minimum mass designs for high aspect ratio, composite wings, subject to constraints on flutter speed, divergence speed and material stress. The wing is modelled as a series of composite beam elements and both flutter speed and divergence speed are calculated using a normal mode approach. Modal analysis is carried out by applying the Wittrick-Williams algorithm to the dynamic stiffness method, whereas unsteady aerodynamic loads are calculated from strip theory, although an option which uses lifting-surface theory is also presented. A previously published example is given to validate the analysis. Single level optimisation is carried out using a sequential quadratic programming strategy combined with the modified methods of feasible directions optimizer, for which flutter sensitivities are obtained by an efficient determinant interpolation technique. Design variables include topological variables such as spar and engine positions as well as layer thicknesses, which are modelled using quadratic functions. The wing of a regional turboprop aircraft is optimized to illustrate the use of the program. The problem was modelled using 10 elements and had 43 design variables, 162 constraints and required just over 20 minutes of CPU time on a workstation. This, coupled with the fact that a full three-dimensional FE model of the same wing would require over 1000 elements, illustrates the suitability of the dynamic stiffness method to the conceptual design stage.