In this study, an electrical parameter test system and a high-speed Schlieren system are used to study the control of a cylindrical detached shock wave through high-energy surface arc plasma excitation. The results show that, when plasma excitation is not applied, the bow shockwave angle around the cylinder is 52 •. After the plasma excitation is applied, the arc discharge releases a large amount of heat within a short time, generating a shockwave and a control gas bulb (CGB). As a result, the bow shockwave angle first decreases and then increases, the pressure ratio before and after the shockwave decreases, and the intensity of the bow shockwave weakens. At t = 280 μs, the bow shockwave angle is reduced to a minimum of 46 •. The effective interference time of high-energy surface arc plasma excitation on a cylindrical detached shockwave is 820 μs. A high temperature is used to control the heating effect of the bubbles, which will increase the local sound velocity near the wall, reduce the local Mach number, cause the sound velocity to move online, and eventually push the bow shockwave away from the cylinder. Concurrently, heating will accelerate the gas flow, reduce the pressure, and cause the mass flow in the unit flow area of the heated area to decrease, resulting in a strong compression effect, which deforms the bow shockwave. The high-energy surface arc plasma excitation will provide a potential technical means for high-speed aircraft detached shockwave control.

Pulsed arc plasma excitation is characterized by strong local heating effect and wide disturbance range, and it has a broad application prospect in supersonic flow control. In this paper, by using electrical parameter measurement system and high speed schlieren technique, we study the electrical and flow field characteristics of pulsed arc plasma excitation under the condition of <i>Ma</i> = 3 incoming flow. The nano-particle planar laser scattering (NPLS) is used to investigate the flow structure of the supersonic flat boundary layer, and the transition characteristics of the boundary layer at different plasma excitation frequencies are studied. The experimental results show that in the flow field with <i>Ma</i> = 3 and the total incoming pressure <i>P</i>₀ = 1 atm (1 atm = 1.01 × 10⁵ Pa), the peak voltage of the pulsed arc plasma actuator discharge is 6 kV, the peak current is 70 A, the time scale of the discharge is about 300 ns, the single discharge energy is 70 mJ; the pulsed arc discharge will produce the precursor shock wave with higher velocity and the thermal deposition zone with higher temperature, which will exert continuously disturbance on the boundary layer. The pulsed arc plasma excitation with perturbation can promote the transition of supersonic plate boundary layer. Moreover, the high-frequency impact effect of pulsed discharge can promote the transition to occur ahead of time, and the higher the frequency, the better the effect is. As the excitation frequency increases, the transition position of the boundary layer of the supersonic flat plate moves forward, and the length of the transition area of the boundary layer becomes shorter as the excitation frequency increases. When the excitation frequency is 60 kHz, the length of transition zone is 0 and the thickness of turbulent boundary layer is 25 mm. When a high frequency is applied (<i>f</i> = 40, 60 kHz), the transition path of the boundary layer is that the shock wave generated by the plasma excitation triggers the unstable wave, and the development of unstable waves directly skips the linear growth stage, passes through the bypass and transitions into turbulent flow. The pulsed arc plasma excitation can be used to promote supersonic boundary layer transition.

In order to solve the problem of single arc plasma actuator’s failure to suppress the boundary layer separation, the effectiveness of the array surface arc plasma actuator to enhance the excitation intensity is verified by experiment. In this study, an electrical parameter measurement system and high-speed schlieren technology were adopted to delve into the electrical, flow field, and excitation characteristics of the high-energy array surface arc plasma actuator under low ambient pressure. The high-energy array surface arc discharge released considerable heat rapidly; as a result, two characteristic structures were generated, i.e., the precursor shock wave and thermal deposition area. The duration increased with the increase in environmental pressure. The lower the pressure, the wider the thermal deposition area’s influence range. The precursor shock wave exhibited a higher propagation speed at the initial phase of discharge; it tended to decay over time and finally remained at 340 m/s. The lower the environmental pressure, the higher the speed would be at the initial phase. High-energy array surface arc plasma actuator can be employed to achieve effective high-speed aircraft flow control.

The symmetric flying wing has a simple structure and a high lift-to-drag ratio. Due to its complicated surface design, the flow field flowing through its surface is also complex and variable, and the three-dimensional effect is obvious. In order to verify the effect of microsecond pulse plasma flow control on the symmetric flying wing, two different sizes of scaling models were selected. The discharge energy was analyzed, and the force and moment characteristics of the two flying wings and the particle image velocimetry (PIV) results on their surface flow field were compared to obtain the following conclusions. The microsecond pulse surface dielectric barrier discharge energy density is independent of the actuator length but increases with the actuation voltage. After actuation, the stall angle of attack of the small flying wing is delayed by 4°, the maximum lift coefficient is increased by 30.9%, and the drag coefficient can be reduced by 17.3%. After the large flying wing is actuated, the stall angle of attack is delayed by 4°, the maximum lift coefficient is increased by 15.1%, but the drag coefficient is increased. The test results of PIV in the flow field of different sections indicate that the stall separation on the surface of the symmetric flying wing starts first from the outer side, and then the separation area begins to appear on the inner side as the angle of attack increases.