Oblique breakdown in a Mach 2.0 supersonic boundary layer controlled by a local cooling strip with a temperature jump is investigated using direct numerical simulations and linear stability theory. The effect of temperature on the stability of the fundamental oblique waves is first studied by linear stability theory. It is shown that the growth rate of fundamental oblique waves will decrease monotonically as the temperature decreases. However, the results of the direct numerical simulations indicate that transition reversal will occur as the growth rate of the fundamental oblique waves of cooled case becomes faster compared with that of baseline case downstream of the cooling strip. When the cooling strip is in the linear region, the transition is delayed due to the suppression effect of the cooling strip on the fundamental oblique waves. When the cooling strip is located in the early nonlinear region, the fundamental oblique waves will be suppressed by higher spanwise wavenumber steady modes generated by the mutual and self-interaction between the fundamental oblique waves and harmonic modes, which is first called the self-suppression effect (SSE) in the present study. Further research indicated that the meanflow distortion generated by steady modes plays an important role in the SSE. Compared with the stabilization effect of the cooling strip, the SSE is more effective. Moreover, the SSE might provide a new idea on the instability control, as it is observed that the SSE works three times leading to the growth rate of fundamental oblique waves slowing down at three different regions, respectively.

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Complex shock interactions and severe aerothermal loads are often encountered on the lips of three-dimensional inward-turning inlets, which presents significant challenges to the performance and safety of hypersonic flight vehicles. However, there have been few investigations on reducing the heat flux of the lips, especially when considering real gas effects. It is therefore necessary to investigate flow control methods that are suitable for the lips under real gas effects. Three flow control methods are implemented in this work: a passive method with shock control bump and stagnation bulge, an active method with counterflow jet, and a combined method. The lip is simplified as a V-shaped blunt leading edge to eliminate the influence of other structures. Numerical simulations are performed at freestream Mach numbers ranging from 6.0 to 12.0. The principles and abilities of different flow control methods for reducing heat flux are compared and analyzed. Although the passive and active methods can reduce the heat flux by more than 40% at low Mach numbers, they have an apparent deficiency under strong real gas effects at high Mach numbers. Moreover, the active method causes new heat flux peaks near the nozzle and at the reattachment position of the flow separation zone. Therefore, a combined method is proposed for further reducing the heat flux. By coupling the passive and active methods, the combined method can reduce the heat flux by nearly 60%. In general, the flow control methods investigated in this work can achieve satisfactory heat flux reduction abilities.

Hypersonic flow on a V-shaped blunt leading edge (VSBLE) at Mach 12 is numerically investigated. A series of self-induced shock–shock interactions and shock wave/boundary layer interactions (SWBLIs) cause extremely high heat flux peaks on the crotch of the VSBLE. Based on these shock structures, a simplified model that divides the flow field into inviscid and viscous areas is constructed. This model can quickly solve the shock interactions and predict the heating/pressure peaks with high accuracy. Furthermore, a shock control bump (SCB) is placed on the SWBLI region to split the strong incident shock into a weaker multi-wave system. The mechanism study of the SCB shows that the inviscid effect significantly reduces the heating/pressure peaks, and the viscous effect suppresses the SWBLI-induced separation. Finally, a VSBLE with SCBs is numerically investigated. The heat flux peak is reduced by 66 % compared to that without the SCBs. The robustness of the SCB under various working conditions is also evaluated. This paper provides an idea for the simplified solution of complex shock interactions and extends the application of the SCB as a thermal protection device in hypersonic flow for the first time.