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The background-oriented schlieren technique was used to visualize the blade-tip vortices of a Eurocopter AS532UL Cougar helicopter in maneuvering flight. The test program covered a large part of the flight envelope, including maneuvers such as hover flight, fast forward flight, flare maneuvers, and high-speed turns. For selected flight conditions, the aerodynamic results are presented here. It is shown that, with the reference-free background-oriented schlieren method, the detection of vortex filaments up to vortex ages of ψ v 540 deg is possible. The visualization of the vortex system is used to detect aerodynamic phenomena such as blade-vortex interactions, vortex-airframe interactions, and the occurrence of smooth sinuous disturbances. A detailed description of the applied reference-free background-oriented schlieren setup is given, and the suitability of different natural backgrounds for the background-oriented schlieren method is analyzed. Nomenclatureof vortex above tip path plane, m I = intensity count K = Gladstone-Dale constant, m 3 · kg −1 M = optical magnification factor Ma = Mach number n = refractive index of air N b = number of rotor blades R = rotor radius, m T = rotor thrust, N t = time, s u, v, Δy = image displacements, m x, y, z = coordinates, m Z B= distance background -camera lens, m Z D = distance background-phase object, m Z i = distance camera lens-camera sensor, m ϵ y = deflection angle, rad ρ = density of air, kg · m −3 σ = solidity; N b · c · πR −1 Ψ = azimuth; Ωt, rad ψ v = vortex age, rad Ω = rotor rotational frequency, rad · s −1 ω z = rotation of the image displacement; ∂v∕∂x − ∂u∕∂y

A back-flow flap attached to the suction side of an airfoil is investigated in both passively and actively actuated modes for the control of dynamic stall. This method of dynamic stall control has low power requirements and no parasitic drag when not actuated. Experiments in a low-speed wind tunnel at 50 m/s were used to characterize the reduction in dynamic stall hysteresis using pressure measurements on the midline airfoil section. It was found that the pitching moment peak is reduced by an average of 25% for all deep stall test cases for active actuation of the flap, while for passive actuation the pitching moment peak is reduced by 19%. In each case the maximum lift remained the same, while the peak drag increased by an average of 2.5% for the active flap, and by 0.9% for the passive flap. With the flap closed at low angles of attack, the reference values of the airfoil are retained. KeywordsHelicopter blade Á Wind tunnel Á Flow control Á Dynamic stall Á Back-flow flap Á Experiment List of symbols a; a max Angle of attack; maximum ( ) c Airfoil model chord (m) C D ; C D max Drag coefficient; peak C L ; C L max Lift coefficient; peak C M ; C M min Pitching moment coefficient; peak C P Pressure coefficient f Frequency of pitching (Hz) M Mach number k Reduced frequency: k ¼ pfc=v 1 Re Reynolds number based on c t Time (s) T Period (s) u, v Velocity: in x direction, in z direction (m/s) v 1Freestream velocity (m/s) x, zCoordinates: chord, upward (m)This paper is based on a presentation at the AHS 72nd Annual Forum,