An experimental study of flow separation control with a nanosecond pulse plasma actuator was performed in wind-tunnel experiments. The discharge used had a pulse width of 12 ns and rising time of 3 ns with voltage up to 12 kV. Repetition frequency was adjustable up to 10 kHz. The first series of experiments was to measure integral effects of the actuator on lift and drag. Three different airfoil models were used, NACA-0015 with the chord of 20 cm, NLF-MOD22A with the chord of 60 cm and NACA 63-618 with the chord of 20 cm. Different geometries of the actuator were tested at flow speeds up to 80 m/s. In stall conditions the significant lift increase up to 20% accompanied by drag reduction (up to 3 times) was observed. The critical angle of attack shifted up to 5—7 degrees. The relation of the optimal discharge frequency to the chord length and flow velocity was proven. The dependence of the effect on the position of the actuator on the wing was studied, showing that the most effective position of the actuator is on the leading edge in case of leading edge separation. In order to study the mechanism of the nanosecond plasma actuation experiments using schlieren imaging were carried out. It shown the shock wave propagation and formation of large-scale vortex structure in the separation zone, which led to separation elimination. PIV diagnostics technique was used to investigate velocity field and quantitative properties of vortex formation. In flat-plate still air experiments small-scale actuator effects were investigated. Measured speed of flow generated by actuator was found to be of order of 0.1 m/s and a span-wise non-uniformity was observed. The experimental work is supported by numerical simulations of the phenomena. The formation of vortex similar to that observed in experiments was simulated in the case of laminar leading edge separation. Model simulations of free shear layer shown intensification of shear layer instabilities due to shock wave to shear layer interaction.
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机译：在风洞实验中进行了用纳秒脉冲等离子体致动器进行流分离控制的实验研究。所用放电的脉冲宽度为12 ns，上升时间为3 ns，电压最高为12 kV。重复频率可调至10 kHz。第一系列实验是测量执行器对升力和阻力的整体影响。使用了三种不同的机翼模型，弦长为20厘米的NACA-0015，弦长为60厘米的NLF-MOD22A和弦长为20厘米的NACA 63-618。在高达80 m / s的流速下测试了执行器的不同几何形状。在失速条件下，观察到明显的升力增加高达20％，同时减阻（最多3倍）。临界攻角最多可移动5-7度。证明了最佳放电频率与弦长和流速之间的关系。研究了作用对机翼上致动器位置的依赖性，表明在前缘分离的情况下，致动器的最有效位置在前缘上。为了研究席利伦成像的纳秒等离子体驱动机理，进行了实验。结果表明，冲击波在分离区传播并形成大尺度的涡旋结构，导致分离消失。 PIV诊断技术用于研究涡流形成的速度场和定量特性。在平板静止空气实验中，研究了小型促动器效应。发现由执行器产生的测得的流速约为0.1 m / s，并且观察到跨度方向的不均匀性。现象的数值模拟为实验工作提供了支持。在层流前沿分离的情况下，模拟了与实验中观察到的相似的涡旋形成。自由剪力层的模型仿真表明，由于冲击波与剪力层的相互作用，剪力层的不稳定性加剧。
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