Huoxing Liu

For horizontal take-off and landing reusable hypersonic vehicles, thermal management of aerodynamic stagnation heat at high Mach numbers constitutes the most critical thermal challenge. Usually, the researcher addresses this through an integrated thermal allocation strategy employing lightweight high-efficiency precoolers, which systematically redistribute thermal loads by reducing inlet/outlet temperatures while enhancing energy input to power components, thereby achieving superior energy utilization efficiency. However, the ultra-fine characteristic dimensions (∼1 mm scale) of modern precoolers intensify fluid-thermal coupling effects, rendering conventional design methodologies inadequate with prediction deviations exceeding 30%. To resolve this, we present an inverse design framework based on a prototype sCO2 precooler. Through 15 meticulously designed operational cases encompassing air inlet temperatures around 800 K and CO2/air flow ratios of 1–1.5, we establish a data-informed methodology that synergizes: (1) Experimental characterization of thermal-hydraulic performance (including heat transfer coefficients and pressure drop correlations). (2) Reverse determination of critical design parameters through integration with conventional heat exchanger design protocols. (3) Validation under representative hypersonic flight conditions. Experimental verification demonstrates that the proposed methodology reduces temperature prediction errors to <15% compared to empirical data, while maintaining pressure drop deviations within ±10%. This advancement provides a validated design paradigm for next-generation hypersonic thermal management systems requiring strong multi-physics coupling resolution.

To address the thermal protection challenges of multiple high-temperature components and the electrical power deficiency in hypersonic vehicles, this study proposes twelve multi-heat-source thermoelectric conversion schemes based on the sCO₂ Brayton cycle. A three-dimensional evaluation system for thermal management is established, incorporating thermal efficiency, coolant mass flow rate, and system mass as key metrics. A comprehensive parameter sensitivity analysis was conducted on the twelve dual-heat-source cycle configurations. For systematic performance comparison, the Non-dominated Sorting Genetic Algorithm II (NSGA-II) was employed for multi-objective optimization, with Pareto fronts analyzed to determine optimal configurations. The results demonstrate that appropriately increasing the minimum cycle temperature can significantly reduce coolant flow requirements. Multi-objective optimization reveals the following: (1) The pre-compressed aero-comb configuration achieves optimal performance in the efficiency-mass flow rate optimization scenario; (2) Both pre-compressed aero-comb and re-compressed comb-aero configurations show superiority in the efficiency-mass optimization scenario; (3) The pre-compressed aero-comb configuration exhibits lower system mass in low coolant flow regions for the mass flow rate-mass optimization scenario. Overall, the performance of the precompression aero-comb configuration is relatively superior. This work provides an important reference for the design of thermal management systems for hypersonic vehicles.

A wave rotor optimizes the use of energy resources by enhancing thermodynamic cycles, and further optimization of wave rotor geometry is emerging as an attractive research area. Among the geometric features, the stagger angle of channels lacks sufficient study in spite of its important effects. To address this question, this work developed and applied the velocity triangle models to modify the basic geometry of wave rotors for different stagger angles, and investigated the flow fields with two-dimensional numerical methods. Results showed that: (1) different stagger angles worked out similar unsteady pressure wave systems and kept nearly constant compression and expansion ratios of the wave rotor; (2) increased stagger angle made the inlet and outlet flows turn toward the axial direction, which was beneficial to compact and light-weighted integration of the wave rotor to a gas turbine; (3) increased stagger angle made the wave rotor consume more shaft power, but even the maximum shaft power was small. This work revealed a critical mechanism how the velocity variation across an unsteady pressure wave produced rim work in a staggered channel, and made a recommendation to comprehensive optimization of wave rotor geometry for better integration in a gas turbine and acceptable shaft power consumption.

High bypass ratio turbofan engine is the mainstream of civil aviation engines at present. It has higher thermal efficiency and better economic performance than other types of engines under the condition of high subsonic, high altitude flight. This trend will not change in the coming decades. In a turbofan engine with high thrust, high mass flow and high bypass ratio, the fan blade

As the load of the turbine components of aircraft engines continuously increases, shock loss becomes the dominant factor of turbine stage loss and has become a hot topic. The Schlieren technique is one of the few effective experimental methods to observe and study shock wave and, thus, has been widely used. Nevertheless, limited by camera accuracy and computer image processing technology, quantitative schlieren analysis methods were difficult to achieve in engineering applications. Fortunately, several quantitative schlieren methods have been developed with the help of new digital technology. Applying the schlieren technique to the highly loaded turbine cascade test is of great significance to the study of shock wave in highly loaded turbine cascades. In this paper, the results of the quantitative density field and shock intensity and loss in the cascade are obtained by using a double-reflection-type monochrome schlieren device. The boundary condition of the density field is obtained by pressure test, and matlab software is used as image processing calculation tool. The quantitative results of this paper prove the feasibility of applying quantitative schlieren method to highly loaded turbine cascade tests. Also, the implemented image processing method and density boundary condition acquisition method are suitable and convenient for cascade flow and shock measurement tests.

With recent advancements in the development of material and manufacturing technology, the leading edge geometry of turbomachine blades has attracted widespread attention. “Sharp” leading edges always have a better aerodynamic performance, though it is prone to deformations easily. Thus, flat plates and real compressor cascades with different leading edge deformations were investigated to study the influence, which is applicable for thin blades at low speeds. Different boundary layer characteristics, including the velocity profile, transition process, and loss, are compared. The results show that there are several kinds of contradictory influence mechanisms and that the final phenomenon is closely related to the condition of the original boundary layer. In low turbulence, with large and laminar separation, the deformations can suppress separation and decrease loss. In high turbulence, with short and transitional separation, deformations can promote the transition process and increase the loss. The sensitivities of different the original leading edge shapes are also compared. This indicates that a good design always has a better robustness at low turbulence values, while it is worse at high turbulence values. The cascade experiment and simulation show that the deformation influence is similar to flat plates and that it is enlarged near the hub, which affects the corner separation.

Leakage flow between the rotor and the stator can cause serious performance degradation of wave rotors which utilize nonsteady shock waves to directly transfer energy from burned gases to precompressed air. To solve this problem, primary flow features relevant to leakage are extracted and it was found that the leakage-attributed performance degradation could be abstracted to a special initial-boundary value problem of one-dimensional Euler equations. Then, a general loss assessment method is proposed to solve the problem of nonsteady flow loss prediction. Using the above method, a reasonable physical hypothesis of the initial-boundary value problem depicting the nonsteady leakage flow process is proposed and further, a closed-form leakage loss analytical model combined with an empirical correction method for the discharge coefficient is established. Finally, with the experimentally verified CFD method, comprehensive numerical verification is conducted for the loss prediction model; it is proved that the physical hypothesis of the proposed model in this paper is reasonable and the model is capable of predicting nonsteady shock wave attenuation due to leakage exactly within the range of parameter variations of wave rotors.

A wave rotor enhances the performance of a gas turbine with its internal compression and expansion, yet the thermodynamic efficiency estimation has been troubling because the efficiency definition is unclear. This paper put forward three new thermodynamic efficiency definitions to overcome the trouble: the adiabatic efficiency, the weighted-pressure mixed efficiency, and the pressure pre-equilibrated efficiency. They were all derived from multistream control volumes. As a consequence, they could correct the efficiency values and make the values for compression and expansion independent. Moreover, the latter two incorporated new models of pre-equilibration inside a control volume, and modified the hypothetical “ideal” thermodynamic processes. Parametric analyses based on practical wave rotor data demonstrated that the trends of those efficiency values reflected the energy losses in wave rotors. Essentially, different thermodynamic efficiency definitions indicated different ideal thermal cycle that an optimal wave rotor can provide for a gas turbine, and they were recommended to application based on that essence.