The results of experimental and computational studies are considered on a near surface electric discharge effect on supersonic airflow near a 15° compression surface. The tests were performed at Mach number M = 2, stagnation pressure P 0 = 1-2.8 bar, stagnation temperature T 0 = 290-600 K, and plasma power W pl = 6-12 kW. They demonstrated a significant effect of plasma on the flow structure and reduction of static pressure on the compression surface. Transient phenomena were analyzed and it was found that the pressure decrease on the ramp was as fast as t < 0.3 ms. Simulations based on 3D unsteady Navier-Stokes equations with plasma modeled as an array of lengthwise heat sources demonstrated adequacy of such simplifications. Further simulations attempted to find an optimal range of plasma power and position in terms of achievable effect, effectiveness of the method, and response time of the system to the plasma actuation. The electric discharge authority for a fast and effective control of aerodynamic forces in a compression ramp configuration is considered.
This work was performed to study the effect on flow disturbances in the corner separation zone of a compression surface with a hypersonic boundary layer caused by a weakly ionized transient plasma generated upstream. Schlieren imaging was used to distinguish the corner separation zone for 15°, 20°, and 25° compression ramps at Mach 4.5 (nozzle exit). A Shack-Hartmann wavefront sensor was used to determine the dominant frequencies of flow oscillations at different locations in the flow field and the resulting effect of repetitively pulsed plasma actuators. A significant rise in amplitudes of high-frequency (>80 kHz) flow perturbations was found when pulsing the plasma at a frequency (100 kHz) higher than the natural dominant frequency of the boundary layer (~65 kHz). The plasma effect was negligible when operated below this frequency (50 kHz). PCB pressure sensors were used to determine the dominant frequencies of pressure oscillations present at the surface on the flat plate and compression ramp inside the separation zone. This technique can potentially be used for active control of the boundary layer condition and supersonic flow structure on the compression surface.
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