This work continues an ongoing effort aimed at development and use of dielectric barrier discharge (DBD) plasma actuators driven by repetitive nanosecond pulses for high Reynolds number aerodynamic flow control. These actuators are believed to influence the flow via a thermal mechanism which is fundamentally different from more commonly studied AC-DBD actuators. Leading edge separation control on an 8-inch chord NACA 0015 airfoil is demonstrated at various post-stall angles of attack for Mach numbers up to 0.26 (free stream velocity up to 93 m/s) and Reynolds numbers up to 1.15 X 106. The nanosecond (NS) pulse driven DBD is shown to extend the stall angle at low Reynolds numbers by functioning as an active trip. At post-stall angles of attack, the device is shown to excite shear layer instabilities and generate coherent spanwise vortices that transfer momentum from the freestream to the separated region, thus reattaching the flow. This is observed for all high Reynolds numbers and Mach numbers spanning the speed range of the subsonic tunnel used in this work. A comparison of leading edge separation control between NS-DBD and AC-DBD plasma actuation demonstrates the increased control authority of NS-DBD plasma at higher flow speeds. The NS-DBD actuator is also integrated into a feedback control system with a stagnation-line-sensing hot film. A simple on/off type controller is developed that operates based on a threshold of the power dissipated by the hot film. An extremum seeking controller is also investigated for dynamically varying Re. Several challenges typically associated with the integration of DBD plasma actuators into a feedback control system have been overcome. The most important of these is the demonstration of control authority at typical aircraft takeoff and landing Mach numbers.
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