等离子体冲击气动激励控制三角翼前缘涡的影响因素和机理研究

Widely used in modern fighters, delta wing exists aerodynamic problems such as vortex broke down at large attack angles. As a new type of active flow control technique, plasma aerodynamic actuation has the advantages of fast response, simple structure and the convenience for closed loop control, showing significant potential in the control of vortex flows of delta wings. The study of flow control over delta wings by AC-DBD(alternating current dielectric barrier discharge) plasma actuation is still in the initial stage and the freestream velocity is still very low(<20m/s) for effectiveness of flow control. NS-DBD(nanosecond pulsed surface dielectric barrier discharge) performs outstandingly in flow separation control over an airfoil. The application of NS-DBD actuation to the moderate swept delta wing with blunt leading edge by the applicant in early stage, has significantly improved the capability of leading-edge vortex control, with the effectively controlled velocity up to 50m/s. As the leading-edge vortex of the delta wing is greatly influenced by the sweepback angle and the leading-edge cross section, the control of leading-edge vortex of delta wing with different configurations is to be deeply researched. .Selecting multiple delta wings’ flow as the control object, this project is conducted by theoretical study, experiment test and numerical simulation. On the one hand, it is expected to reveal the control laws by parameters of actuation and geometry of delta wings, and obtain the key actuation parameters which dominates control effect. By optimizing the design of actuator and actuating under the optimal parameters, the actuation control ability will be greatly improved. And on the other hand, the corresponding mechanism can be revealed by analyzing the action process of the actuation on the leading-edge vortex, through means of flow visualization, numerical simulation, etc. This project research is helpful to provide a theoretical basis and technical reserve for solving aerodynamic problems of delta wings at high angles of attack.

三角翼被现代战机广泛采用，大迎角时存在涡破裂等气动问题。作为新型流动控制技术，等离子体激励响应快、结构简单、便于闭环控制，解决三角翼气动问题具有潜力。目前AC-DBD激励对三角翼绕流的可控速度低，鉴于NS-DBD激励控制翼型流动分离能力突出，申请人前期将NS-DBD激励应用到中等后掠钝前缘三角翼前缘涡控制，显著提升了可控速度。三角翼前缘涡受后掠角和前缘截面形状影响较大，NS-DBD激励对不同构型三角翼前缘涡的控制有待系统深入研究。针对不同三角翼的绕流，通过实验研究、数值仿真、理论分析等方法深入系统研究：有望揭示NS-DBD激励参数和三角翼几何因素对流动控制效果的影响规律；通过关键激励参数和激励器优化设计，大幅提升NS-DBD激励控制三角翼绕流的能力；结合流动显示和数值仿真，分析NS-DBD激励对不同三角翼前缘涡的作用过程，揭示流动控制机理，为解决三角翼大迎角气动问题提供理论基础和技术储备。

1 获得不同因素对等离子体冲击气动激励控制三角翼绕流的影响规律.在不同来流条件、几何模型和激励参数下，通过表面介质阻挡纳秒脉冲放电产生等离子体冲击气动激励，研究了NS-DBD激励三角翼绕流控制效果及控制规律。前缘NS-DBD激励可有效改善30°后掠和47后掠三角翼大迎角气动特性；最佳无量纲激励频率均为F+=f×c/U∞=1~2；45 m/s时最大升力系数分别提高18.3%和9.6%；对于47°后掠三角翼，钝前缘的流动控制效果好于尖前缘；后掠角增大后，流动控制效果减弱，激励未能改善60°后掠三角翼的升阻特性。.2 揭示等离子体冲击气动激励三角翼前缘涡的流动控制机理.在前缘施加NS-DBD激励唯象学模型，进行47°后掠钝前缘三角翼绕流控制仿真。在F+=1.44时，激励显著提高了失速前后的升力系数；激励在前缘分离剪切层处诱导产生流向涡，改变了前缘剪切层结构，使其向内卷吸；激励后时均流场形成了明显的负压峰值，前缘涡附着线外移，吸力面回流区减小。.锁相PIV测试表明，激励在前缘剪切层处诱导流向涡，激励频率决定了沿剪切层共存诱导涡的数量，当激励频率超过一定阈值时（F+=4），不能产生附着的前缘涡。.3 发展了阵列式表面电弧等离子体气动激励控制三角翼流动分离的方法.PSAD的瞬时脉冲能量高于典型的介质阻挡放电，可形成强的等离子体冲击气动激励。采用纹影和PIV技术测量静止空气中由激励所引起的流场变化。每次激励都会产生局部压力波，其传播速度接近声速。特别值得注意的是在纹影成像中观察到的热空气流，其热量流入周围空气中从而形成对流扰动。每个激励器放电后产生了一对弱的反向涡，最大传播速度为0.1m/s。诱导的瞬间扰动表明基于PSAD的等离子体流动控制机理是由PSAD激励产生的瞬时热释放。在30°后掠角三角翼上进行风洞实验，12路通道PSAD激励可以有效地改善三角翼的气动力性能，将迎角从14°推迟到18°，并将最大升力系数提高了12.9%。.4 根据中等后掠三角翼前缘分布式激励的测力试验结果，前缘分布式的等离子体激励可以在中等迎角下实现一定的飞行控制，也可以减缓大迎角下分离引起的机翼振动。.