https://iopscience.iop.org/article/10.1088/0963-0252/25/6/063001/pdf

Dynamics of near-surface electric discharges and mechanisms of their interaction with the airflow

Towards In-Flight Applications? A Review on Dielectric Barrier Discharge-Based Boundary-Layer Control

A pulsed-plasma jet actuator is used to control the unsteady motion of the separation shock of a shock wave/boundary layer interaction formed by a compression ramp in a Mach 3 flow. The actuator is based on a plasma-generated synthetic jet and is configured as an array of three jets that can be injected normal to the cross-flow, pitched, or pitched and skewed. The typical peak jet exit velocity of the actuators is about 300 m/s and the pulsing frequencies are a few kilohertz. A study of the interaction between the pulsed-plasma jets and the shock/boundary layer interaction was performed in a time-resolved manner using 10 kHz schlieren imaging. When the actuator, pulsed at StL ≈ 0.04 (f = 2 kHz), was injected into the upstream boundary layer, the separation shock responded to the plasma jet by executing a rapid upstream motion followed by a gradual downstream recovery motion. Schlieren movies of the interaction showed that the separation shock unsteadiness was locked to the pulsing frequency of the actuator, with amplitude of about one boundary layer thickness. Wall-pressure measurements made under the intermittent region showed about a 30% decrease in the overall magnitude of the pressure fluctuations in the low-frequency band associated with unsteady large-scale motion of the separated flow. Furthermore, by increasing the pulsing frequency to 3.3 kHz, the amplitude of the separation shock oscillation was reduced to less than half the boundary layer thickness. Investigation into the effect of the actuator location on the shock wave/boundary layer interaction (SWBLI) showed qualitatively and quantitatively that the actuator placed upstream of the separation shock caused significant modification to the SWBLI unsteadiness, whereas injection from inside the separation bubble did not cause a noticeable effect.

Control of unsteadiness of a shock wave/turbulent boundary layer interaction by using a pulsed-plasma-jet actuator

Large-eddy simulations have been conducted to investigate the mechanisms of separated-flow control using a dielectric barrier discharge plasma actuator at a low Reynolds number. In the present study, the mechanisms are classified according to the means of momentum injection to the boundary layer. The separated flow around the NACA 0015 airfoil at a Reynolds number of 63 000 is used as the base flow for separation control. Both normal and burst mode actuations are adopted in separation control. The burst frequency non-dimensionalized by the freestream velocity and the chord length (F+) is varied from 0.25 to 25, and we discuss the control mechanism through the comparison of the aerodynamic performance and controlled flow-fields in each normal and burst case. Lift and drag coefficients are significantly improved for the cases of F+ = 1, 5, and 15 due to flow reattachment associated with a laminar-separation bubble. Frequency and linear stability analyses indicate that the F+ = 5 and 15 cases effectively exc...

Mechanisms for laminar separated-flow control using dielectric-barrier-discharge plasma actuator at low Reynolds number

Particle image velocimetry measurements in close proximity to dielectric-barrier discharge plasma actuators are conducted to quantify the momentum transfer of the plasma to the surrounding air flow. Based on these data a comparative analysis of six existing approaches to estimate the induced body force is presented. Integral methods calculate an integral value for the actuator force based on the momentum-balance equation. Insight into the spatial distribution of the body force is provided by differential methods, which are based either on the Navier–Stokes equations or on the vorticity equation. It is demonstrated that the intensity as well as the domain of the force increase with increasing operating power levels. Emphasis is also placed on the issue of self-induced drag. It is shown that 30% of the induced momentum is consumed by wall friction. All results are validated with previously obtained balance force data and luminosity analysis of identical actuators.

http://iopscience.iop.org/article/10.1088/0022-3727/46/5/055202/pdf

Velocity-information-based force-term estimation of dielectric-barrier discharge plasma actuators

This paper reports experimental results on using steady and unsteady plasma aerodynamic actuation to control the corner separation, which forms over the suction surface and end wall corner of a compressor cascade blade passage. Total pressure recovery coefficient distribution was adopted to evaluate the corner separation. Corner separation causes significant total pressure loss even when the angle of attack is 0°. Both steady and unsteady plasma aerodynamic actuations suppress the corner separation effectively. The control effect obtained by the electrode pair at 25% chord length is as effective as that obtained by all four electrode pairs. Increasing the applied voltage improves the control effect while it augments the power requirement. Increasing the Reynolds number or the angle of attack makes the corner separation more difficult to control. The unsteady actuation is much more effective and requires less power due to the coupling between the unsteady actuation and the separated flow. Duty cycle and excitation frequency are key parameters in unsteady plasma flow control. There are thresholds in both the duty cycle and the excitation frequency, above which the control effect saturates. The maximum relative reduction in total pressure loss coefficient achieved is up to 28% at 70% blade span. The obvious difference between steady and unsteady actuation may be that wall jet governs the flow control effect of steady actuation, while much more vortex induced by unsteady actuation is the reason for better control effect.

Control of the corner separation in a compressor cascade by steady and unsteady plasma aerodynamic actuation

Wedge oblique shock wave control using an arc discharge plasma aerodynamic actuator was investigated both experimentally and theoretically. Schlieren photography measurements in a small-scale short-duration supersonic wind tunnel indicated that the shock wave angle decreased and its start point shifted upstream with the plasma aerodynamic actuation. Also the shock wave intensity weakened, as shown by the decrease in the gas static pressure ratio of flow downstream and upstream of the shock wave. Moreover, the shock wave control effect was intensified when a static magnetic field was applied. Under test conditions of Mach 2.2, magnetic control and input voltage 3 kV, the start point of the shock wave shifted 4 mm upstream, while its angle and intensity decreased 8.6% and 8.8%, respectively. A thermal choking model was proposed to deduce the change laws of oblique shock wave control by surface arc discharge. The theoretical result was consistent with the experimental result, which demonstrated that the thermal choking model can effectively forecast the effect of plasma actuation on an oblique shock wave in a cold supersonic flow.

https://iopscience.iop.org/article/10.1088/0022-3727/42/16/165503/pdf

Changbing Su

Papers

Control of the corner separation in a compressor cascade by steady and unsteady plasma aerodynamic actuation

Effects of plasma aerodynamic actuation on oblique shock wave in a cold supersonic flow

Experimental Investigation on Shock Wave Control by Plasma Aerodynamic Actuation

Plasma Aerodynamic Actuation Based Corner Separation Control in a Compressor Cascade