In hypersonic flight the shock wave and turbulent boundary layer interaction (STBLI) sharply increases wall heat transfer that intensifies the aerodynamic heating problems. In this work the STBLI is modelled by compression ramp flow with a Mach number of 5, a Reynolds number based on momentum thickness of 4652 and a wall to recovery temperature ratio of 0.5. The aerodynamic heat generation and transport mechanisms are investigated in the interaction based on theoretical analysis and direct numerical simulation (DNS) that agrees with previous studies. A prediction correlation of wall heat flux in STBLI is deduced theoretically and validated by some representative data including the present DNS, which improves the prediction accuracy and can be applied to a wider Ma range compared with the canonical Q-P theory. The correlation indicates that the sharp increase of wall heat transfer in the STBLI can be explained by the boundary layer compression and the convection transport enhancement. Based on the DNS results, the aerodynamic heat generation and transport mechanisms are revealed in the separation, recirculation and reattachment zones in the STBLI. From this perspective, the peak heat flux can be further explained by the enhancement of near-wall turbulent energy dissipation, compression aerodynamic heat generation and the near-wall turbulent transport. The generation and transport of compression aerodynamic heat reveal the underlying mechanism of the strong correlation between the peak heat flux ratios and the pressure ratios in STBLIs.

High-speed vehicles that have long-term flight requirements in the atmosphere encounter significant aerodynamic drag and aero-heating challenges. Opposing jet as an active flow control technique has the potential to achieve efficient drag and heat reduction at the same time. The study aimed to investigate the control mechanism of the opposing jet drag and heat reduction in high-speed flows. Theoretical analysis and numerical simulations were conducted utilizing a blunt body with an opposing jet hole to identify the dominant control parameters and regulation rules of the opposing jet. The results showed that the recirculation vortex was the most critical flow structure for the drag and heat reduction of the opposing jet. Theoretical analysis of the recirculation vortex range and strength revealed five dominant control parameters influencing the drag and heat reduction characteristics of the opposing jet, including jet stagnation pressure ratio PsR, the total temperature ratio T0R, jet Mach number M-j, incoming flow Mach number M-infinity and diameter ratio D of the jet hole to the base model. Properly adjusting the dominant control parameters according to their regulation rules can efficiently reduce drag and heat while consuming less jet flow.

Due to the wall temperature difference between wind tunnel experiments and flight environment, the wall temperature effects on shock wave turbulent boundary layer interaction (STBLI) are worth investigating. Based on the verified DNS method, a 30 degree ramp is used to generate STBLI for Ma of 5 and wall to recovery temperature ratio ranging from 0.2 to 1.0. The results indicate that the separation zone decreases for cold wall conditions and quantitatively validate the wall-temperature-corrected interaction scaling theory in recent literature. The heat transfer results show that the wall temperature greatly influence the heat flux enhancement in STBLI: there exists significant negative minimum downstream the heating peak for near adiabatic walls and there exists sharp heating peak for the coldest wall. The analysis reveals that the coupling change of local velocity and temperature causes the negative heat flux for near adiabatic wall conditions. Overall, the decrease of wall temperature leads to the decrease of the peak heat flux enhancement and the fundamental mechanisms is the near-wall turbulent aerodynamic heat dissipation variation with wall temperatures.

While the opposing jet technique has the potential to achieve efficient drag and heat reduction, it can be severely affected by the incoming angle of attack. To analyze the angle-of-attack characteristics of opposing jet for improving drag and heat reduction, a three-dimensional blunt model was studied under various jet stagnation pressure ratios and angles of attack using the verified numerical method. The results showed that the enhanced reattachment shock on the windward side resulted in a higher pressure and temperature rise, which led to the deterioration of drag and heat reduction. Under the influence of the incoming angle of attack, the recirculation vortex transformed into a longitudinal vortex, resulting in a slanted U-shaped distribution of the surface pressure coefficient and Stanton number. Increasing the jet stagnation pressure ratio widened the coverage of the recirculation vortex on both the windward and leeward sides, which brought an improvement in drag and heat reduction. The interaction between the incoming angle of attack and the opposing jet caused a double-peak distribution of Stanton number due to the recirculation vortex reattachment and the compression of the incoming flow. The inclined opposing jet could reduce the peak values of pressure coefficient and Stanton number when subjected to the incoming flow with an angle of attack by spreading the recirculation vortex along the windward side. There should exist an optimal inclination angle that can effectively reduce the peak caused by the compression of the incoming flow without generating an excessive peak due to the recirculation vortex reattachment.

Gas turbine is a crucial power equipment around the world. Its inlet temperature increases year by year, bringing great challenges to the temperature resistance of turbine blades. Thus, exploring high-efficiency cooling structure is a must. In this paper, flow and heat transfer characteristics of a rotating rectangular smooth internal U-channel with orientation angles of 90 degrees called Coriolis-utilization rotating rectangular smooth cooling U-channel have been numerically investigated by comparing with non-rotation smooth U-channel and conventional rotating smooth U-channel with Reynolds-Averaging Navier-Stokes (RANS) method. The Reynolds number kept constant at 10000. The Rotation number was between 0 and 0.1. The numerical method has been validated by experiment results. The results demonstrate that the Coriolis-utilization rotating U-channel not only utilizes the Coriolis force merit on the leading and trailing walls to strengthen heat transfer, but also performs lower pressure loss than the conventional rotating U-channel. Therefore, the Coriolis-utilization rotating U-channel is a promising internal cooling structure of rotor blade for higher inlet-temperature turbine in the future.

燃气轮机和航空发动机被誉为是工业皇冠上的明珠，其研制水平是一个国家科技水平和综合国力的重要标志。随着燃气轮机和航空发动机工作效率和性能的不断提高，涡轮入口温度逐年上升，涡轮叶片暴露在更高的来流温度下。为使金属叶片在远超其熔点的温度中仍能安全运转，亟需发展高效的冷却技术。该文概述了燃气涡轮高效冷却技术及设计方法的发展趋势，提出了按照3个维度开展燃气涡轮冷却技术研究的思路，总结了本团队在冷却单元-气冷叶栅-整机多部件交互等方面的基础研究成果，搭建了基于实验数据驱动的高效高精度冷却结构设计平台，探索了以双层壁为代表的下一代冷却技术的特性和发展趋势。