Lamination-Parameter-Based Mass Minimization of the Common Research Model Wing Using Sandwich Composites

Fiber-reinforced composites are widely used in primary aircraft structures on account of their superior performance when compared to metallic structures. When buckling is a dominant driver of the structural design, the use of sandwich composites could potentially yield more efficient designs. This paper applies a recently developed approach for optimizing practical commercial-scale aircraft wings using sandwich composites in a preliminary design stage to perform design studies using the NASA Common Research Model (CRM) as a reference. The approach uses lamination parameters as design variables in a continuous optimization step. Structural constraints for classic composite laminate design, such as material failure and buckling, and for sandwich design, such as crimping, wrinkling, dimpling, and core shear failure, are accounted for using industrial-standard and empirical methods driven by finite element analyses. The optimization studies present comparisons in structural weight between sandwich composite designs and their monolithic counterparts. The studies present several cases where sandwich composites offer superior structural performance, as well as potential cost savings by affording a lesser number of stringers in the design.