Composite sandwich materials are an interesting alternative to the use of conventional composite structures for aircraft components. Current research at DLR addresses the question, as to where sandwich composites can be beneficially employed in aircraft wingbox designs.
Conventional composites structures (hereby denoted as monolithic composites) are manufactured by stacking several plies together forming a laminate. In aircraft primary structures, carbon fibre as the material basis together with a suitable resin, constitutes this ply.
Sandwich composites can be visualized as splitting a monolithic composite laminate in the middle and inserting a low-density material such as a honeycomb core in the middle. This construction efficiently increases the bending stability of the structure on account of a larger thickness, with very low increase in weight. Consequently, sandwich composites have been used extensively in aerospace applications, primarily in regions where the design loads are low, for instance, in control surfaces, engine cowlings and belly fairings. In the sailplane and satellite industry, they are also used in the primary load-bearing components.
The present research focuses on the use of sandwich composites in primary aircraft structures such as the wingbox. This is motivated by trends in aircraft wing development that show buckling as the driving design factor in large sections of the wing, when compared to strength requirements.
In order to better understand the weight comparisons between sandwich and monolithic composite designs, optimization studies are performed on isolated skin panels of the wing in a first step. A skin panel is defined as the intersection between two ribs and spars. The NASA Common Research Model (CRM) which is a representative long-range aircraft model is used as a test case. The failure modes particular to sandwich composites are addressed using conservative empirical formulae. Altogether, material failure through AML (Angle Minus Laminate in a strain-based approach), buckling instability using a finite element (FE) – based linear solver, facesheet wrinkling, shear crimping, facesheet dimpling and core-shear failure using empirical relations are accounted for.
Depending on the span-wise section considered, skin panels with sandwich composite design exhibit weight savings of up to 30% when compared to its classic monolithic counterpart. The trends show that when the skin panel is driven by buckling as a design requirement, sandwich composites outperform monolithic designs on account of their higher bending stiffness to weight ratio. When material or strain-dominated criteria drive the design, sandwich composites offer a similar performance to monolithic skin panels. The former is observed near the wing-tip while the latter is seen from the skin panels at the mid-section onwards towards the root, in the case of this reference wing.
Apart from savings in weight, sandwich composites in the lightly-loaded regions near the wing-tip allow for a larger stringer pitch without penalizing structural weight. This could potentially raise discussions on savings in manufacturing, maintenance and inspection costs.
Practical experience shows that the benefits of sandwich composites come alongside a number of engineering challenges, which must be carefully considered in realistic designs: damage tolerance and its characterization, manufacturing complexities at joints and ramp-downs, complex identification and carrying out of repairs, to name a few. However, on-going research in several university and research groups aim to tackle these very challenges on account of the potential benefits.
In the next steps of our research, performance comparisons on different wing configurations will be studied with the aim of providing a motivation for further research and development in sandwich composites and its application.
Author: Yasser M. Meddaikar, Department Loads Analysis and Design, DLR Institute of Aeroelasticity