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Andrew J. Lombardi A closed-loop plasma actuated control scheme which relies on the ability to detect incipient flow separation is presented. It is known that the receptivity of the boundary layer increases with the onset of separation. The detection scheme therefore hinges on the ability to measure disturbances produced by the unsteady pulsing of a plasma actuator via an on-board pressure sensor and analog circuitry. The actuator introduces the disturbances into the flow at a low powered "Sense State." An on-board pressure sensor measures the real-time pressure signal at x/c = 0.08, which is then amplified and bandpass filtered around the actuator's pulsing frequency. An integrated circuit which measures the RMS of the signal serves as a means to measure the spectral energy at the pulsing frequency. When the signal reaches a sufficiently high value, set by an adjustable threshold, the actuator is triggered to a high powered "Control State" with more flow control authority. The method not only provides a precursor for flow separation, but also an indicator when conditions exist where active re-attachment control is no longer needed. This is demonstrated on a periodically oscillating airfoil which is pitched into light, medium, and deep dynamic stall.

Wind tunnel experiments were conducted to quantify the e↵ectiveness of alternating current dielectric barrier discharge flow control actuators to suppress leading-edge stall on a NASA energy e cient transport airfoil at compressible freestream speeds. The objective of this research was to increase lift, reduce drag, and improve the stall characteristics of the supercritical airfoil near stall by flow reattachment at relatively high Mach and Reynolds numbers. In addition, the e↵ect of unsteady (or duty cycle) operation on these aerodynamic quantities was also investigated. The experiments were conducted at the University of Notre Dame Mach 0.6 Wind Tunnel for a range of Mach numbers between 0.1 and 0.4 with an airfoil model of chord 30.48 cm at atmospheric conditions corresponding to a Reynolds number range of 560, 000 through 2, 260, 000. Lift and drag forces, as well as the quarter chord moments were measured directly by a sting which reacted on load cells and torque sensors on the outside of the 0.91⇥0.91 m wind tunnel test section. Two leading-edges of the airfoil were fabricated. The first was covered in a Kapton dielectric film of 0.127 mm and had a 7 µm copper electrode, and the second was a thick-dielectric Macor with a copper tape exposed (76 µm thick) electrode. A high voltage AC signal was applied to electrodes for the flow control case. The results show that the plasma actuators were e↵ective at reattaching the leading-edge separated flow as evidenced by the increase in maximum lift coe cient and stall angle. In the post stalled regime, the lift was dramatically increased, by as much as 90%. Drag in the stalled regime was reduced by as much as 28% and the nose down pitching moment was reduced by as much as 40%. Pressure taps on the suction surface confirmed flow reattachment as evidenced by the return of a pressure peak near the leading-edge and better pressure recovery aft of the leading-edge when the active flow control was enabled. Time-averaged PIV confirmed the airflow following the airfoil surface closely. The experiment also showed that lift was increased the most in deep stall when the plasma actuator was operated unsteady with a reduced frequency of unity, whereas in light stall steady operation was preferred. Overall, both AC DBD plasma actuator designs were able to increase the maximum lift coe cient and stall angle of attack for the full range of Mach numbers, with the thick-dielectric Macor leading-edge performing better at Mach 0.4.

Single dielectric barrier discharge plasma actuators have been used to control flow separation in a large number of applications. An often used configuration involves spanwise-oriented asymmetric electrodes that are arranged to induce a tangential wall jet in the mean flow direction. For the best effect, the plasma actuator is placed just upstream of where the flow separation will occur. This approach is generally more effective when the plasma actuator is periodically pulsed at a frequency that scales with the streamwise length of the separation zone and the free-stream velocity. The optimum frequency produces two coherent spanwise vortices within the separation zone. It has been recently shown that this periodic pulsing of the plasma actuator could be sensed by a surface pressure sensor only when the boundary layer was about to separate, and therefore could provide a flow separation indicator that could be used for feedback control. The paper demonstrates this approach on an aerofoil that is slowly increasing its angle of attack, and on a sinusoidally pitching aerofoil undergoing dynamic stall. Short-time spectral analysis of time series from a static pressure sensor on the aerofoil is used to determine the separation state that ranges from attached, to imminent separation, to fully separated. A feedback control approach is then proposed, and demonstrated on the aerofoil with the slow angle of attack motion.