## 3 Design of the Skin and the Stringers
### Material Selection and Design of the Skin
It yields the inverse reserve factor (IRF), which assumes values equal to or greater than zero. If the IRF reaches or exceeds 1, failure of the ply is assumed. The investigations described above show that laminates made of unidirectional plies reinforced by high-strength or intermediate-modulus carbon fibres provide the necessary combination of properties. However, since the \"typical\" materials mentioned in the database usually do not directly correspond to commercially available materials, a material type with similar properties then had to be identified that can actually be obtained from suppliers. This led to the selection of IM/8552 carbon fibre [28] for the design of the MLE. This is a \"prepregmaterial, i.e., the fibres of the semi-finished product stacked to make the laminate are already impregnated (pre-impregnated) by a specific amount of resin. Apart from being a unidirectional intermediate-modulus carbon fibre material which is certified for aviation use, its material properties were readily available in the literature. All these aspects were relevant with respect to its selection. By applying the [PERSON]'s failure criterion as described above and using the available material properties, it could easily be verified that the achievable combinations of bending stiffnesses and strengths of laminates made of the selected material are suitable for the present application. Tailoring the stiffness of the MLE skin along the contour was achieved by suitable composite laminate lay-ups, varying the skin thickness by ply drop-offs in the laminate. These were determined in another iterative process based on finite element analysis (FEA) of the structure. Models consisted of linear shell elements (ABAQUS element type S4R, i.e., with reduced integration) with displacement boundary conditions imposed at the top edges of the stringers. The upper edge (approximately parallel to the span-wise direction) of the skin-contour, where it would be fastened to the wing's front spar, was fixed in all directions (translational and rotational), and the aerodynamic pressure distribution for the respective flight condition (cruise or high-lift) was applied to the outer surface of the skin. The solutions were obtained using an implicit solver and considering geometric non-linearity. This way, the reaction forces occurring at the locations of the displacement boundary conditions are equivalent to the forces applied by an actuation system holding the skin in the required shape under the respective aerodynamic pressure load. Lay-ups were manually modified, and the resulting stresses due to morphing and aerodynamic pressure loads were analysed by the FEM. Subsequently, the IRF according to [PERSON]'s failure criterion as well as the deviation from the target shapes were determined from the FE-solution. Based on these, further modifications of the lay-ups were made, and the process repeated until a suitable solution had been found. A simplified flow-chart of the process for defining the skin laminate is shown in Figure 7.  
Tailoring is essential, because the bending strain occurring in the composite plies on the inside and outside surfaces of the laminate is proportional to the laminate thickness as well as the local change of curvature between S4 and S5. Assuming that the target shapes are approximated as intended, increasing the stiffness by adding plies thus also increases the ply stresses, which may then lead to failure. The iterative process described above  
Figure 7: Simplified flow-chart of the process resulting in definition of the skin laminate. considers the IRF (strength related; becomes more critical as thickness increases because surface strains increase) as well as the approximation of the target shapes (stiffness related; becomes less critical as thickness increases because aerodynamic loads cause less deviation from the target shapes) so that both aspects are addressed adequately.