以提升高原无人机起飞性能为背景，以大展弦比无人机机翼模型为研究对象，以单个对称布局介质阻挡放电等离子体激励器为控制手段，采用风洞试验和数值计算相结合的方法开展了低雷诺数下机翼绕流流场控制研究，评估了流动控制效果，揭示了流动控制机理，通过单个对称布局激励器同时实现了小迎角层流分离泡控制和大迎角失速控制。结果表明，在施加激励前的气动特性方面，当雷诺数为7.47×10⁴时，机翼气动力随迎角的变化规律符合低雷诺数气动力变化特征：升力系数非线性增加，阻力系数先增加后减小再增加；后缘层流分离泡是导致气动力出现非线性现象的主要原因。在等离子体气动力控制效果方面，单个对称布局等离子体激励器能够在较大迎角范围内提升机翼气动性能。在小迎角下，等离子体激励能够基本消除气动力非线性现象，最大升阻比提高15%；在大迎角下，等离子体能够抑制机翼失速分离，施加激励后，失速角推迟2°。在等离子体流动控制机理方面，对称布局等离子体激励器诱导涡是在较大迎角范围内实现流动控制的关键。当针对后缘层流分离泡进行控制时，等离子体诱导涡将诱导动量从翼型前缘传递到后缘，通过“排挤”层流分离泡的方式，实现流动控制；当针对大迎角分离流进行控制时，等离子体诱导涡增强了高能量主流与低能量边界层气流之间的掺混，通过“融入”大尺度分离涡的方式，抑制失速分离。

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.