Dynamic stall is observed in numerous applications, including sharply maneuvering fixed-wing aircraft, biomimetics, wind turbines, and most notably, rotorcraft. The associated unsteady loading can lead to aerodynamic flutter and mechanical failure in the system. The present work explores the ability of nanosecond pulse-driven dielectric barrier discharge plasma actuators to control dynamic stall over a NACA 0015 airfoil. The Reynolds number, reduced frequency, and excitation Strouhal number were varied over large ranges: Re = 167,000-500,000, k = 0.025-0.075, and St(e) = 0-10, respectively. Surface pressure measurements were taken for each combination of Reynolds number, reduced frequency, and excitation Strouhal number. Phase-locked particle image velocimetry measurements were acquired for select cases. It was observed that the trends of effect of St(e) were similar for all combinations of Reynolds number and reduced frequency, and three major conclusions were drawn. First, it was observed that low Strouhal number excitation (St(e) < 0.5) results in oscillatory aerodynamic loading in the stalled stage of dynamic stall. This oscillatory behavior was gradually reduced as Ste increased and was not observed beyond St(e) > 2. Second, all excitation resulted in earlier flow reattachment. Last, it was shown that excitation, especially at high St(e), resulted in reduced aerodynamic hysteresis and dynamic stall vortex strength. The decrease in the strength of the dynamic stall vortex is achieved by the formation of large-scale structures induced by the excitation that bleed the leading-edge vorticity before the ejection of the dynamic stall vortex. At sufficiently high excitation Strouhal numbers (St(e) approximate to 10), the dynamic stall vortex was completely suppressed.

Flow control experiments were performed using nanosecond dielectric-barrier-discharge plasma actuators on a NACA 0015 airfoil with flow approaching from the geometric trailing-edge side, which is a condition anticipated to occur on the retreating blade side of advanced helicopters such as slowed-rotor compound rotorcraft. This symmetric airfoil, which is not typical of those used in rotorcraft blades, was used for simplification of an otherwise very complex problem. The Reynolds number based on the chord length was fixed at 0.50106, corresponding to a freestream flow of approximately 38m/s. An angle of attack of 15deg was used. Fully separated flow on the suction side extended well beyond the airfoil with naturally shed vortices occurring at a Strouhal number of 0.19. Plasma actuation was evaluated at both the aerodynamic leading and trailing edges of the airfoil. Excitation at very low (impulse excitation) to moderate (approximate to 0.4) Strouhal numbers at the aerodynamic leading edge generated organized coherent structures in the shear layer over the separated region with a shedding Strouhal number corresponding to that of the excitation, which caused changes in the size of the wake, the separation area, lift, and drag. Excitation at higher Strouhal numbers resulted in weaker naturally shed vortices (rather than generating new vortices) that diffused quickly in the wake. The excitation caused the wake to elongate slightly and skew toward the aerodynamic trailing edge, but it still reduced the separation area and significantly reduced drag. The primary mechanism of control at the aerodynamic leading edge is excitation of instabilities associated with the leading-edge vortices; the excitation generates coherent large-scale structures over a range of excitation frequencies, increasing their entrainment abilities to bring high-momentum fluid into the separation region to reduce the separation size and increase the lift. On the other hand, excitation over a broad range of frequencies at the aerodynamic trailing edge was found to significantly reduce organization of the naturally shed large-scale wake structures.

Leading edge separation control was investigated using nanosecond pulsed DBD plasma actuators on a NACA 0015 airfoil installed in a recirculating wind tunnel such that the flow was fully-reversed over the airfoil. A 15° angle of attack at a Reynolds number of 0.50·10 was selected for detailed investigation. Fully separated flow on the suction side extended well beyond the airfoil with highly asymmetric velocity and vorticity fields and a shedding Strouhal number of 0.19 with a harmonic at 0.38. Excitation at very low (impulse excitation) to moderate (~0.4) Strouhal numbers using plasma actuators at the aerodynamic leading edge generated organized coherent structures in the shear layer over the separated region with a frequency corresponding to the excitation frequency, which caused changes in the size of the wake, the separation area, lift, and drag. Excitation near = . had the most significant effects: creating moderately sized structures that convected far downstream, reducing the separation area by 39%, increasing lift by 6%, and decreasing drag by 8%. Excitation at high Strouhal numbers modified the baseline vortices (rather than generating new vortices) and resulted in weaker vortices that diffused quickly in the wake, causing the wake to elongate slightly and skew toward the aerodynamic trailing edge, but still reducing the separation area and also reducing lift and drag. The primary mechanism of control is excitation of instabilities in the shear layer over the separated zone that generate more coherent large-scale structures over a range of excitation frequencies increasing their entrainment abilities to bring high-momentum fluid into the separation region to reduce the separation size and increase the lift.