An Integrated Procedure to Optimize the Design of a Composite Wing

Carbon Fiber Reinforced Plastics (CFRP) composites have demonstrated to be particularly suitable for aerospace structural applications due to their high specific strength and stiffness. The development of innovative methods for the structural optimization of aerospace composite components is a topic of interest for research and industry. In particular newer and fast methodologies are desirable in order to support the preliminary design of aerospace composite structure. This work proposes an integrated multilevel procedure aimed at analyzing and optimizing the design of a composite wing box by means of linear analysis. The target is to make available in a preliminary design phase a reliable structural sizing of a composite (or sandwich) wing by means of a rational and computationally efficient approach without using FEA. Assigned the external loads acting on the wing, its geometry and the design material values, this procedure attempts to achieve the objective of weight reduction by optimizing the rib spacing, the skin/web thicknesses, the layups, the sandwich core height, the spar cap area as well as the stringer one according to a no-strength failure and no buckling onset criteria. The proposed methodology is based on the use of in-house developed codes integrated with commercial codes; among the commercial codes modeFrontier® and Hypersizer® are used. The first one is a multidisciplinary and multi-objective software, while the second one is a design, analysis, and optimization code for composite structures. The main theoretical assumptions of the design methodology foresee the wing conceived as concentrated elements: the skin is sized by bending and shear; the spar webs by shear and the caps by bending. Elementary theories are also applied to determine the internal loads acting on the different structural wing box components, both for stiffened and sandwich panels. The methodology allows an automated design process to be executed with very short computational times offering the possibility to perform sensitivity analyses to study the influence of several factors (wing thicknesses, panel structural concept, stacking sequences) on the structural weight. In this way, the use of iterative FE analyses, which are unsuitable for preliminary design, is avoided. The proposed methodology is a very fast tool to optimize the ply thickness and the final laminate thickness according to minimum weight requirement and minimum gauge constraints that can influence the final material selection and its form (prepreg, fabric). The preliminary wing, sized and optimized with this procedure, represents a reliable sized structure able to withstand the assigned external loads; it has been then modeled and analyzed by FEM analysis in a second stage of the design process: the FEA results confirmed that the final FE structure is very close to the preliminary wing structure. In other words, the proposed preliminary design procedure gave accurate results in a very cost effective way if compared with standard optimizations approaches where only general purpose software are considered. The future work of this research activity is to extend this procedure also to other structural items like a fuselage or a vertical fin.