# Comparative Aerodynamic Performance Analysis of Camber Morphing and Conventional Airfoils
## 4 Results  
CFD simulations were run and the \(C_{L}\) and \(C_{D}\) computed for various morphed and deflected airfoil configurations, over a range of \(\alpha\) from 0\({}^{\circ}\) to 15\({}^{\circ}\). The plots of the \(C_{L}\) of different morphed cases are presented in Figure 7 and the deflected cases in Figure 8. For clarity, only those configurations that match the \(L\) profiles of morphing configurations or contribute to the overlapping regions were included in the conventional airfoil \(L\) plots in Figures 8 and 10. Figure 5: CFD setup validation case study results (NACA0012, \(Re=6\) million). (**A**) Drag polar based on NASA Langley Research Center validation case data. (**B**) Drag polar generated using the proposed CFD methodology with 2 different turbulence models. (**C**) Comparison of simulated results (**B**) with validation case (**A**). Figure 6: A schematic diagram of (**a**) a conventional airfoil with a deflecting flap (purple) and (**b**) a morphing airfoil (yellow). From Figures 7 and 8, it is noted that morphing airfoils stall starting at a higher \(\alpha\) than the conventional, deflecting airfoils. Typically, the stall region for airfoils in the 50,000-100,000 \(Re\) regions is between \(10^{\circ}\) to \(14^{\circ}\)[48] and our results confirm that. The approximated stall angles extrapolated from the plots or identified based on the readings are presented in Table 1.  
Figure 8: Conventional airfoils \(C_{L}\) vs. \(\alpha\) varying the flap deflection angle. (\(Re\) = 76,630). Figure 7: Morphing airfoils \(C_{L}\) vs. \(\alpha\) varying the maximum camber rate. (\(Re\) = 76,630). For the camber morphing, the value of \(\alpha\) at maximum \(L\) increases as the camber rate increases until it reaches a peak at \(\alpha\) = 13\({}^{\circ}\) between camber rate of 3% and 6%. However, at 7% and greater (7-9%) rates, the angle starts to decrease, and the maximum \(C_{L}\) is obtained at \(\alpha\) < 13\({}^{\circ}\). This trend of a decreasing stall angle as the camber rate increases is to be expected since as the camber of the airfoil increases, the geometric change creates a suction peak on the upper surface near the leading edge that leads to an easier onset of boundary layer separation. NACA 6412 airfoil stalls at around 13\({}^{\circ}\) but a 9% cambered airfoil--NACA 9412--stalls at around 11\({}^{\circ}\). Pressure contours of 6% and 9% morphed airfoils at their respective stall angles are shown in Figure 9 and the contours for both configurations are very similar, which is consistent with the trend and the lift values of Figure 7, where the amount of lift generated by these configurations is around a similar range. On the other hand, the conventional airfoils stall early. The stall angle corresponding to maximum \(C_{L}\) shifts to a lower value as the deflection angle increases. The highest value of stall angle obtained (\(\alpha=11^{\circ}\)) is for the airfoil configuration deflected \(1.5^{\circ}\), which is the same \(\alpha\) for which the baseline airfoil stalls. Whereas all of the morphed airfoil configurations stall at \(\alpha\geq 11^{\circ}\), none of the conventional airfoils stall at \(\alpha>11^{\circ}\). This implies that camber morphing wings would improve the manoeuvrability, agility, and stability of the aircraft. In fact, camber morphing wings flight-tested on a drone demonstrated the achievement of superior manoeuvrability by generating higher roll rates exclusively through morphing [27,31]. Amongst all cases, the greatest \(L\) is generated by NACA 9412 (9% camber rate) at \(\alpha=11^{\circ}\), which is only slightly greater than the \(L\) generated by the NACA 8412 airfoil for the same \(\alpha\). Due to decreasing stall angles as the camber rate increases, the comparative enhancement in \(L\) with increasing camber rate starts to overlap and curb, starting at 6% camber rate at \(\alpha=13^{\circ}\).