## 3 Design Setting of Optimum Design in UAV Main Wing Skin
### Factor and Design Variables for the Genetic Algorithm  
The more constraints there are, the smaller the generation can find an optimization value. In this paper, only the constraints for the first failure mode and the buckling failure mode were used, and optimization was carried out based on the values used in the actual work. Positive margin of safety (M.S.) values used in actual work are generally determined by the experience of development projects or engineers, and it is recommended to set the minimum safety rate within the range of 0 < M.S. < 0.5. If the minimum safety rate is 0.5 or more, it becomes too conservative, which does not mean much of the structural optimization work. As shown in Equation (1), the Tsai-wu breakage theory used in the initial breakage mode showed the interaction of flat stress on the right side, and the failure \begin{table}
\begin{tabular}{c c} \hline \hline \multicolumn{2}{c}{**Blade-Type Stiffener Shape Dimensions**} \\ \hline length (Skin), L [mm] & 250 \\ width (Skin), W [mm] & 160 \\ location (Stiffener), s [mm] & 30 \\ width (Flange), F [mm] & 24 \\ width (Web), W [mm] & 25 \\ Laminate Stacking (Skin) & [0\({}_{4}\)/-45\({}_{4}\)/+45\({}_{4}\)/90\({}_{4}\)]s \\ Laminate Stacking (Stiffener) & [0\({}_{2}\)/-45\({}_{2}\)/+45\({}_{2}\)/90\({}_{2}\)]s \\ Weight [kg] & 0.30 \\ \hline \hline \end{tabular}
\end{table} 
Table 1: Geometry information and blade-type stiffener shape dimensions. \begin{table}
\begin{tabular}{c c c c c c} \hline \hline \multicolumn{1}{c}{**E\({}_{11}\) (GPa)**} & \multicolumn{1}{c}{**E\({}_{22}\) (GPa)**} & \multicolumn{1}{c}{**G\({}_{12}\) (GPa)**} & \multicolumn{1}{c}{**G\({}_{13}\) (GPa)**} & \multicolumn{1}{c}{**G\({}_{23}\) (GPa)**} & \multicolumn{1}{c}{**v\({}_{12}\)**} \\ \hline
146.00 & 9.00 & 3.40 & 3.40 & 2.60 & 0.33 \\ \hline \multicolumn{1}{c}{**X\({}_{\text{Tension}}\) (MPa)**} & \multicolumn{1}{c}{**X\({}_{\text{Comp}}\) (MPa)**} & \multicolumn{1}{c}{**Y\({}_{\text{Tension}}\) (MPa)**} & \multicolumn{1}{c}{**Y\({}_{\text{Comp}}\) (MPa)**} & \multicolumn{1}{c}{**Shear (MPa)**} & \multicolumn{1}{c}{**t-ply (mm)**} \\ \hline
2251.00 & 1078.00 & 44.00 & 200.00 & 69.00 & 0.19 \\ \hline \hline \end{tabular}
\end{table} 
Table 2: Composite material properties, wing skin material. Figure 7: The geometry of the blade stiffened the panel. index on the left side determines that breakage occurs if it is equal to or greater than 1. In this paper, the convergence value was set to 0.8. \[F_{1}\,\sigma_{1}+F_{2}\,\sigma_{2}+F_{11}\sigma_{1}{}^{2}+F_{22}\sigma_{2}{}^{2 }+F_{66}\tau_{12}{}^{2}+2F_{12}\sigma_{1}\sigma_{2}=1, \tag{1}\]  
where \(F_{1}=\frac{1}{\chi^{2}}-\frac{1}{\chi^{2}}\), \(F_{2}=\frac{1}{\chi^{2}}-\frac{1}{\chi^{2}}\), \(F_{11}=\frac{1}{\chi^{2}\chi^{2}}\), \(F_{66}=\frac{1}{5^{2}}\), \(F_{12}=\frac{(F_{11}F_{22})^{0.5}}{2}\). In the buckling failure mode, the safety rate was calculated using the interaction between the compression load and the shear load to verify the structural safety of the wing skin. In this paper, the convergence value was set to 0.05. In Equation (2), \(N_{x,cr}\), \(N_{xy,cr}\) of the compression buckling critical load of the simple support structure is as follows:  
\[N_{x,cr}=(\frac{\pi}{b})^{2}\left[D_{11}m^{2}(\frac{b}{a})^{2}+2(D_{12}+2D_{66 })n^{2}+D_{22}(\frac{n^{4}}{m^{2}})(\frac{a}{b})^{2}\right], \tag{2}\]  
\[N_{xy,cr}=(\frac{\pi}{b})^{2}(D_{11}\times{D_{22}}^{3})^{1/4}\left[8.125+\frac {5.05(D_{12}+2D_{66})}{(D_{11}\times{D_{22}})^{0.5}}\right], \tag{3}\]  
Here, \(a\) and \(b\) are the shape dimensions, \(D_{11}\), \(D_{12}\), \(D_{22}\), and \(D_{66}\) are the reduced stiffness \(D\) matrix, and \(n\) and \(m\) are 1, 2, or 3. The values of the factors used in this optimum design issue are shown in Table 3. The main design variables for performing the optimization analysis of the reinforced composite laminate are the number of laminations and the lamination angle.