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Sliding discharge, as a new type of electrical discharge, is being gradually applied in plasma active flow control in recent years. In this work, the particle image velocimetry (PIV) experiments were conducted to investigate the airflow characteristics induced by the sliding discharge plasma actuator at varied voltage signals. The results show that the integral thrust produced by the negative alternating current (AC) power is stronger than normal AC power under the same voltage magnitude. The induced airflow direction changes along with the changing of two power voltage signals. Furthermore, the angle of the induced airflow is mainly linear with the loading voltage.

Plasma-based flow control is one of the most promising techniques for aerodynamic problems, such as delaying the boundary layer transition. The boundary layer’s characteristics induced by AC-DBD plasma actuators and applied by the actuators to delay the boundary layer transition on airfoil at Ma = 0.3 were experimentally investigated. The PIV measurement was used to study the boundary layer’s characteristics induced by the plasma actuators. The measurement plane, which was parallel to the surface of the actuators and 1 mm above the surface, was involved in the test, including the perpendicular plane. The instantaneous results showed that the induced flow field consisted of many small size unsteady vortices which were eliminated by the time average. The subsequent oil-film interferometry skin friction measurement was conducted on a NASA SC(2)-0712 airfoil at Ma = 0.3. The coefficient of skin friction demonstrates that the plasma actuators successfully delay the boundary layer transition and the efficiency is better at higher driven voltage.

As combat aircraft becomes more and more maneuverable, the need to understand the unsteady behavior of aircraft in dynamic flow fields becomes more important. Usually researchers pay more attention to the effects on the changes of AOA, but ignore the effects of velocity variations. It is known that the velocity of aircraft changes greatly when the aircraft undergoes a high angle of attack maneuver, like "cobra" maneuver. To completely simulate and study the effect of rapid changes in both free stream velocity and angle of attack, a pitching motion setup is developed in the unsteady wind tunnel of NUAA. By measuring unsteady loads, unsteady pressure distribution and flow visualization, the unsteady aerodynamic behavior of a pitching isolated delta wing and the pitching delta wing coupled with unsteady free stream are investigated. It is found that the oscillating free stream velocity affects the hysteresis characteristics of the pitching delta wing further. The pressure distribution and flow visualization measurements show that the changes in the structure of the leading-edge vortices are the main reason. These studies conclude that a good understanding of the unsteady aerodynamics is vitally important in the design of super-maneuverable aircraft.