## 4 Results and Discussion
### The Second High-Speed Cruise Stage of Flying Back to the Airport  
During the second cruise stage in Figure 10, an aircraft needs smaller lift due to fuel consumption, ranging from 0.745\(W_{0}\) to 0.665 \(W_{0}\). Both VCaCh and VCa methods have little influence on lift-drag ratio at small \(L\)/\(W_{0}\). The lift-drag ratio can be improved by trailing-edge downward deflection with the VCa morphing method, but the improvement effect is small. The effect of the VCaCh morphing method is poor during this stage, and it fails to improve the lift-drag ratio. Table 5 shows that the improvement effect on lift-drag ratio of VCa_\(-\)0.8 airfoil with better aerodynamic efficiency is also very small, only 0.3%. Combined with the results of the first cruise stage, it can be concluded that with the decrease in required lift due to fuel consumption, the effect of variable camber on the lift-drag ratio is smaller under the transonic cruise stage. The results of drag decomposition show that trailing-edge downward deflection with the VCa morphing method can reduce total drag by reducing pressure drag at \(L\)/\(W_{0}\) = 0.7 during the second cruise stage. Although the pressure drag of VCaCh is smaller than VCa when the downward deflection angle is the same, the friction drag difference between them accounts for a larger proportion, resulting in smaller total drag of VCa. Therefore, the VCa morphing method is superior to VCaCh during this stage. Figure 11 shows the pressure distribution of BA airfoil under different \(L\)/\(W_{0}\) at \(Ma\) = 0.734. With the decrease in \(L\)/\(W_{0}\) (attack angle decreases), the leading-edge suction peak value decreases, the position of shock wave moves forward, and the shock wave is weakened. Furthermore, the shock wave of the BA airfoil is significantly weak when \(L\)/\(W_{0}\) = 0.7. According to the pressure coefficient distributions of morphing airfoil at \(L\)/\(W_{0}\) = 0.7 in Figure 12, the shock wave of VCaCh_\(-\)0.8 airfoil is weakened little compared with the BA airfoil, which has little influence on shock-wave drag. Therefore, the decrease in pressure drag of VCaCh_\(-\)0.8 airfoil is not enough to offset the increase in friction drag. The shock-wave strength of VCaCh_\(-\)0.8 airfoil is equivalent to that of the BA airfoil, which means its pressure drag decreases mainly because of the more obvious decrease in form drag. \begin{table}
\begin{tabular}{c c c c c c} \hline \hline
**Airfoil** & \(\mathbf{D_{\text{Irition}}}\) & \(\mathbf{D_{\text{pressure}}}\) & \(\mathbf{D_{\text{total}}}\) & \(\mathbf{L}\)/\(\mathbf{D}\) &
\begin{tabular}{c} \(L\)/\(\mathbf{D}\) \\ **Increment (\%)** \\ \end{tabular} \\ \hline BA & 27.77734 & 33.06096 & 60.8383 & 74.032 & 0 \\ VCaCh_7 & 30.77227 & 27.26645 & 58.03872 & 77.603 & 4.82 \\ VCa_7 & 29.39886 & 28.76330 & 58.16216 & 77.439 & 4.60 \\ VCh\_0.05 & 29.15994 & 31.2976 & 60.45753 & 74.465 & 0.58 \\ \hline \hline \end{tabular}
\end{table} 
Table 4: The drag decomposition and lift–drag ratio of morphing airfoil at \(Ma\) = 0.3, \(L\)/\(W_{0}\) = 0.83. Figure 10: Baseline and morphing-airfoil comparison: variation of \(L\)/\(\mathbf{D}\) with \(L\)/\(W_{0}\) at \(Ma\) = 0.734. Combining with the pressure coefficient distribution during the above-two flight stages, it is found that the morphing airfoil has an influence on the leading-edge suction peak value. When the trailing edge deflects downward at the same \(L/W_{0}\), the leading-edge suction peak value increases compared with the BA airfoil, and the VCa airfoil increases more than VCaCh when the trailing edge deflects the same angle. Instead, compared with the BA airfoil, the leading-edge suction peak value decreases when the trailing edge deflects upward, and the decrease in VCaCh airfoil is greater than VCa when the trailing edge deflects the same angle. Then, the regularity can be explained as follows.