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Fig. 7 Effects of a–c) deflecting aileron, d–f) flap, and g–i) both flap and aileron by 1, 3, and 5 deg, respectively, on the lift coefficient for 0 to 25 deg. a) 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 Nomenclature Example: F/A 0/1 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 b) 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 h) Flap deflection = 0 deg, Aileron deflection = 1 deg 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 c) 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 f) 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 i) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) that occur at high angles of attack explains the lack of effectiveness of the trailing-edge ailerons and flaps, documented in Fig. 7. In particular, the surface streaks show a large degree of flow separation and crossflow near the trailing edge, especially at the wingtips. This further explains why the ailerons were less effective than the inboard flaps. The effect of the angle of attack on the LEV are further evidenced in Fig. 9, which presents photographs from laser-fluorescence flow visualization obtained at 5, 10, and 15 deg (Figs. 9a–9c, Figs. 9d–9f, and Figs. 9g–9i, respectively) at multiple streamwise locations. These images focus on the large-scale vortex that scales with the wing half-span. They demonstrate that even after the VBD associated with the leading-edge separation vortex has occurred, the outer flow maintains a (weak) spiral behavior with an expanding core as it moves downstream in the wake of the wing. Figure 10 shows surface flow visualization results from a parallel study conducted on a half-span wing model, demonstrating similar LEV and VBD patterns to those in Fig. 8 on the leeward surface for different angles of attack. These surface visualizations capture more of the fine detail of the flow modules associated with the leading-edge flow separation and reattachment. At a 10-deg angle of attack, for example, they particularly highlight the primary vortex core and vortex at the wingtip. At 16 deg, the improvement in the visual information is substantially better than that shown in Fig. 8; this shows that the location of the breakdown of the primary vortex has moved inboard to approximately half of the wing span. The vortex at the wingtip is observed to have expanded to cover all of the region that would be occupied by an aileron. The results from different plasma actuator experiments are presented in Figs. 11–15. Figures 11–13 highlight the results from the plasma actuator experiments conducted on the half-span wing model for achieving control at high angles of attack. The details of the specific plasma actuator configurations examined (e.g., P9, P10, P15, P18, and P19) are included in the captions for Figs. 11–13. In general, the letter P indicates a case with the plasma actuator on, and the number following P indicates a specific plasma actuator configuration, details of which are provided in the corresponding figure captions. Figures 11a–11f show the effects of P9, P10, and P15 Baseline F/A 0/1 PATEL ET AL. 1269 Baseline F/A 0/3 Baseline F/A 0/5 d) e) g) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) 0 5 10 15 20 25 Angle of Attack (deg) Baseline F/A 5/0 Baseline F/A 1/0 Baseline F/A 3/0 Baseline F/A 1/1 Baseline F/A 3/3 Baseline F/A 5/5 CL CL CLPDF Image | Plasma Actuators for Hingeless Aerodynamic Control of an Unmanned Air Vehicle
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