In order to improve the takeoff performance of Unmanned Aerial Vehicles (UAVs) in plateau, the studies of flow control over a wing with high aspect ratio using a single symmetrical layout of dielectric barrier discharge plasma actuator were carried out with the help of wind tunnel experiments and numerical simulation at low Reynolds number. The flow control effect was evaluated and the flow control mechanism was revealed. Interestingly, the laminar trailing-edge separation vortex at low angle of attack and the stall separation flow at high angle of attack can be suppressed by one symmetrical plasma actuator. The results indicate that the variation of wing aerodynamics with angle of attack conforms to the characteristics of low Reynolds number aerodynamics at the Reynolds number of 7.47×10⁴ before plasma actuation. The lift coefficient increases nonlinearly, while the drag coefficient first increases, then decreases, and then increases again. It should be noted that the trailing-edge separation vortex is the underlying mechanism of nonlinear aerodynamic phenomena. Regarding the aerodynamics control effectiveness of plasma, the aerodynamic performances of wing can be enhanced by the symmetrical plasma actuator over a wide range of angles of attack. At low angles of attack, the phenomena of aerodynamic nonlinearity were eliminated almost and the maximum lift to drag ratio was increased by about 15% by the plasma actuator. At high angles of attack, stall separation flow around the wing was suppressed and the stall angle is delayed by about 2°. Regarding the mechanism of plasma flow control, the induced vortices of symmetrical plasma actuator play an important role in achieving flow control over a wide range of angles of attack. The induced vortices transferred the momentum from the leading edge to the trailing edge of airfoil and squeezed the laminar separation bubble when controlling the trailing-edge separation vortex. When applied to the control of flow separation at high angle of attack, the induced vortices enhanced the mixing between high-energy mainstream and low-energy boundary layer airflow and suppressed the stall separation flow by integrating the large-scale separation vortices at high angles of attack.

Motivated by the demand of alleviating the impact of gusts during aircraft takeoff and landing, experimental investigations on an airfoil of GAW-1 gust alleviation using a symmetrically arranged dielectric barrier discharge plasma actuator were carried out in a low-speed wind tunnel with the help of force measurement, pressure measurement, and high-speed particle image velocimetry. The effect of gust alleviation using the plasma actuator was quantitatively evaluated, and the flow control mechanism of was revealed. The symmetrical plasma actuator which can be capable of producing a bi-direction quasi-wall jet with approximately equal velocities was mounted at the leading-edge of airfoil. The results indicated that the separation flow around the airfoil can be suppressed, and the stall angle of attack can be delayed by the plasma actuator under the gust environment. The stall angle of attack was delayed by 2° and the maximum lift coefficient was increased by 12% after plasma actuation. Meanwhile, the pressure oscillations caused by gusts can be suppressed by the plasma actuator, leading to the gust alleviation. Two typical flow structures, namely spanwise vortices and coherent structures are generated by the plasma actuator. The induced spanwise vortices near the leading-edge of airfoil and the induced coherent structures in the vicinity of the wall surface play an important role in the gust alleviation. The process of gust alleviation based on plasma actuation can be divided into three stages. Initially, the induced vortices can promote the mixing between low-energy airflow near the wall and the mainstream, thereby injecting momentum into the boundary layer during the first stage. In the second stage, a relatively enclosed region produced by the interaction between the induced vortices and the incoming flow was established and was able to create virtual aero-shaping, thereby changing the shape of the leading edge of the airfoil. Finally, a series of coherent structures transported induced momentum from the leading edge of the airfoil to the trailing edge in the third stage. The present results provide methodological support for establishing the gust alleviation technology for unmanned aerial vehicles based on plasma actuation.

Flow control using surface Dielectric Barrier Discharge (DBD) plasma actuators driven by a sinusoidal alternating-current power supply has gained significant attention from the aeronautic industry. The induced flow field of the plasma actuator, with the starting vortex in the wall jet, plays an important role in flow control. However, the energy consumed for producing the induced flow field is only a small fraction of the total energy utilized by the plasma actuator, and most of the total energy is used in gas heating and dielectric heating. Therefore, an in-depth analysis of the thermal characteristics of the plasma actuator is the key to develop its potential capability further. In addition, compared with the investigation on the aerodynamic characteristics of the plasma actuator, there is a relative lack of detail in the study of its thermal characteristics. Understanding the thermal characteristics of the plasma actuator is of great interest for providing a deeper insight into the underlying working principles, advancing its numerical simulation model, prolonging its life, and achieving several potential engineering applications, such as anti-icing and deicing. The present paper reviews the thermal characteristics of the plasma actuator, summarizes the influence of the dielectric film and actuation parameters on heating, and discusses the formation and transfer mechanism of the induced heating based on the discharge regimes of the plasma actuator in one cycle.