Abstract
This study considers the effect of an electric discharge on the flow structure near a 19.4° compression ramp in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. The stagnation temperature and pressure were varied in a range of 294–600 K and 1–3 bar, respectively, to attain various Reynolds numbers ranging from 5.3 × 105 to 3.4 × 106 based on the distance between the exit of the Mach-2 nozzle and the leading edge of the ramp. Surface pressure measurements, schlieren visualization, discharge voltage and current measurements, and plasma imaging with a high-speed camera were used to evaluate the plasma control authority on the ramp pressure distribution. The plasma being generated in front of the compression ramp shifted the shock position from the ramp corner to the electrode location, forming a flow separation zone ahead of the ramp. It was found that the pressure on the compression surface reduced almost linearly with the plasma power. The ratio of pressure change to flow stagnation pressure was also an increasing function of the ratio of plasma power to enthalpy flux, indicating that the task-related plasma control effectiveness ranged from 17.5 to 25.
Highlights
Active flow control is one of the most promising methodologies for improving the aerodynamic characteristics of supersonic transport vehicles
A supersonic flow field is dominated by shock waves, and their formation and interaction with the boundary layer is an essential aspect of flow control to attain a desirable flow field around the vehicle [1]
Supersonic flow over a compression ramp geometry has been studied for decades [4,5] because the ramp geometry is employed in aerodynamic shapes
Summary
Active flow control is one of the most promising methodologies for improving the aerodynamic characteristics of supersonic transport vehicles. A micro vortex generator, or a micro ramp composed of triangular obstacles with a forward-facing compression ramp, has been studied both at supersonic [2] and hypersonic speeds [3] as a flow modification method. It provides a means of controlling boundary layer separation and shock-induced separation by forming an adverse pressure gradient along the surface. Supersonic flow over a compression ramp geometry has been studied for decades [4,5] because the ramp geometry is employed in aerodynamic shapes. The compression ramp can be regarded as a geometry composed of a supersonic vehicle body with its flap deflected outwards
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