Aerothermal characteristics of bleed slot in hypersonic flows

Abstract

Two types of flow configurations with bleed in two-dimensional hypersonic flows are numerically examined to investigate their aerodynamic thermal loads and related flow structures at choked conditions. One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer interactions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.

Characteristics and mixing enhancement of a self-throttling system in a supersonic flow with transverse injections

“…One is that there is no recirculation zone upstream of the barrel shock due to the upstream three throttling orifices which remove the low momentum portion of boundary layer. Besides, a shock generated by the upstream throttling hole [40] can decrease the freestream Mach number and weaken the interaction between the injection and freestream, leading to a lower adverse pressure gradient upstream the jet at z/D = 0 plane in Case6. It is also can be seen that separation region R4 is much smaller than R1.…”