For hypersonic flow, it was found that the most effective plasma actuator is derived from an electromagnetic perturbation. An experimental study was performed between hypersonic flow and plasma aerodynamic actuation interaction in a hypersonic shock tunnel, in which a Mach number of 7 was reached. The plasma discharging characteristic was acquired in static flows. In a hypersonic flow, the flow field can affect the plasma discharging characteristics. DC discharging without magnetic force is unstable, and the discharge channel cannot be maintained. When there is a magnetic field, the energy consumption of the plasma source is approximately three to four times larger than that without a magnetic field, and at the same time plasma discharge can also affect the hypersonic flow field. Through schlieren pictures and pressure measurement, it was found that plasma discharging could induce shockwaves and change the total pressure and wall pressure of the flow field.
The attenuation of electromagnetic （EM） waves in unmagnetized plasma generated by an inductively coupled plasma （ICP） actuator has been investigated both theoretically and experimentally. A numerical study is conducted to investigate the propagation of EM waves in multilayer plasma structures which cover a square fiat plate. Experimentally, an ICP actuator with dimensions of 20 cm×20 cm×4 cm is designed to produce a steady plasma slab. The attenuation of EM waves in the plasma generated by the ICP actuator is measured by a reflectivity arch test method at incident waves of 2.3 GHz and 10.1 GHz, respectively. A contrastive analysis of calculated and measured results of these incident wave frequencies is presented, which suggests that the experiment accords well with our theory. As expected, the plasma slab generated by the ICP actuator can effectively attenuate the EM waves, which may have great potential application prospects in aircraft stealth.
This study demonstrates the potential for shock wave-boundary layer interaction control in air by plasma aerodynamic actuation.Experimental investigations on shock wave-boundary layer interactions control by plasma aerodynamic actuation are conducted in a Mach 3 in-draft air tunnel.Schlieren imaging shows that the discharges cause the oblique shock to move forward.Schlieren imaging and static pressure probes also show that separation phenomenon shifts backward and the size of separation is enlarged when plasma aerodynamic actuation is applied.The intensity of shock wave is weakened through wall pressure probe.Furthermore,numerical investigations on shock wave-boundary layer interactions control are conducted with plasma aerodynamic actuation.The discharge is modeled as a steady volumetric heat source which is integrated into the energy equation.The input energy level is about 7 kW through discharge process.Results show that the separation phenomenon shifts backward and the intensity of shock is reduced with plasma actuation.These numerical results are consistent with the experimental results.
SUN QuanLI YingHongCUI WeiCHENG BangQinLI JunDAI Hui
Experimental investigation of aerodynamic control on a 35° swept flying wing by means of nanosecond dielectric barrier discharge （NS-DBD） plasma was carried out at subsonic flow speed of 20-40 m/s, corresponding to Reynolds number of 3.1 × 10^5-6.2× 10^5. In control condition, the plasma actuator was installed symmetrically on the leading edge of the wing. Lift coefficient, drag coefficient, lift-to-drag ratio and pitching moment coefficient were tested with and without control for a range of angles of attack. The tested results indicate that an increase of 14.5% in maximum lift coefficient, a decrease of 34.2% in drag coefficient, an increase of 22.4% in maximum lift-to-drag ratio and an increase of 2° at stall angle of attack could be achieved compared with the baseline case. The effects of pulsed frequency, amplitude and chord Reynolds number were also investigated. And the results revealed that control efficiency demonstrated strong dependence on pulsed fre- quency. Moreover, the results of pitching moment coefficient indicated that the breakdown of lead- ing edge vortices could be delayed by plasma actuator at low pulsed frequencies.
Han MenghuLi JunNiu ZhongguoLiang HuaZhao GuangyinHua Weizhuo
According to the mechanism of the arc plasma heating effect,and from a phenomenological perspective of view,the plasma actuation was simplified as heating energy injected into the supersonic flow field for the numerical research on controlling detached shock of the blunt body in non-center symmetrical positions.Besides,experimental research on the form and strength of detached shock wave control by plasma aerodynamic actuation in non-center symmetrical positions was conducted in a high-speed shock tunnel(M=2).The results showed that the detached distance of shock wave increased and the strength of normal shock wave ahead of the detached shock wave reduced when plasma actuation was applied.The control effect was greatly improved after the magnetic field was applied and the effect of upwind-direction flow was the best one.When the upwind-direction flow was applied with 1000 V voltage actuation,the distance of detached shock wave would increase from 3.4 to 7.6 mm and the time average strength of normal shock wave was weaken by 5.5%.At last,the mechanism of plasma actuation on controlling the detached shock wave was briefly analyzed.
SUN QuanCHENG BangQinLI YingHongKONG WeiSongLI JunZHU YiFeiJIN Di
Experimental investigation of active flow control on the aerodynamic performance of a flying wing is conducted. Subsonic wind tunnel tests are performed using a model of a 35° swept flying wing with an nanosecond dielectric barrier discharge （NS-DBD） plasma actuator, which is installed symmetrically on the wing leading edge. The lift and drag coefficient, lift-to- drag ratio and pitching moment coefficient are tested by a six-component force balance for a range of angles of attack. The results indicate that a 44.5% increase in the lift coefficient, a 34.2% decrease in the drag coefficient and a 22.4% increase in the maximum lift-to-drag ratio can be achieved as compared with the baseline case. The effects of several actuation parameters are also investigated, and the results show that control efficiency demonstrates a strong dependence on actuation location and frequency. Furthermore, we highlight the use of distributed plasma actuators at the leading edge to enhance the aerodynamic performance, giving insight into the different mechanism of separation control and vortex control, which shows tremendous potential in practical flow control for a broad range of angles of attack.
