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To investigate the wall cooling of the thrust chamber in an engine, two film-cooling structures, namely, a circular hole structure and a slot structure, were designed. Numerical simulations were performed to study the coupled flow and regenerative cooling heat transfer in thrust chambers with different structures. The influences of parameters such as the film mass flow rate and film hole size on wall cooling were analyzed. Experiments were conducted in a thrust chamber to validate the accuracy of the numerical calculation method. The results indicate that the slot-structured film adheres better to the wall than the circular-hole-structured film, and the film closely adhering to the wall provides better insulation against hot gas, resulting in a reduction of approximately 6% in wall temperature. When the film hole size changes, the change in circumferential wall temperature in the upstream region of the slot-structured film is more pronounced. This paper aims to provide a reference for the design of the cooling structure at the head of the thrust chamber in engineering and suggests directions for optimization and improvement.

To meet the design requirements of long-endurance hydrogen energy hybrid powered Unmanned Aerial Vehicle, the combined MATLAB and XFOIL optimization methods, based on the optimized airfoil, were used to optimize the design of the airfoil of HM114. Two aerodynamic layouts for straight wing and elliptical wing (HY1) were designed. Computational fluid dynamics was adopted to analyze the aerodynamic performance of the two layouts. The results show that the lift coefficient of the optimized airfoil is increased by 12.38%, and the lift-drag ratio is increased by 14.03%; the two layouts achieve the maximum lift-drag ratio when the angle of attack is α=4°, while the stall angle of attack is α=12°, the HY1 layout's smaller pressure difference can contribute more lift to UAV; the HY1 layout has better stall performance when the pressure boundary of the wing surface is behind; both schemes satisfy the longitudinal static stability, but the HY1 layout has better stability. The results of aerodynamic optimization and flow field characteristic analysis in this paper can provide a reference for the aerodynamic layout design of hydrogen energy long-endurance fixed-wing aircraft.

The buckling failure of the afterburner cylinder is a serious safety concern for aero-engines. To tackle this issue, the buckling simulation analysis of the afterburner cylinder was carried out by using finite element method (FEM) software to obtain the buckling mode and critical buckling loads. It was found that the afterburner cylinder was susceptible to buckling when subjected to differential pressure or the compressive force of the rear flange. Buckling would occur when the differential pressure reached 0.4 times the atmospheric pressure or when the axial compressive force on the rear flange reached 222.8 kN. Buckling was also found at the front of the cylinder under the auxiliary mount load. Additionally, under various loads on the rear flange, buckling occurred in the rear section, with the buckling mode being closely related to the load characteristics. Based on the simulation results and structural design requirements, two structural improvements were proposed, including the wall-thickening scheme and the grid reinforcement scheme. FEM simulation analysis results showed that both schemes would improve the rigidity and stability of the afterburner cylinder. For the 0.3 mm increase in the wall thickness scheme, the critical buckling load increased by 17.86% to 66.4%; for the grid reinforcement scheme, the critical buckling load increased by 169% to 619%. Therefore, the grid reinforcement scheme had a stronger anti-buckling ability and was deemed the optimal solution. The findings of this paper could provide technical support for the structural design of large-sized and thin-walled components of aero-engines.

To better understand the performance of gaseous film cooling near the injector region in the LOX/GH2 thrust chambers, a methodology is employed to simulate the coupled heat transfer of the film cooling in thrust chambers with regenerative cooling. The conjugated flow and heat transfer behaviors of film coolant, hot gas, cooling channels and regenerative coolant are numerically investigated. A three-dimensional non-adiabatic flamelet model using real gas equation of state is developed to solve the combustion and validated against the experimental data. Film cooling performance is predicted for the conditions with different geometrical parameters and mass flow rates of the film coolant. The result shows that the reverse flow zones are formed and developed in the region near the injectors The occurrence of those zones is responsible for the significant reduction of the hot-gas-side wall temperature near the head plate of the injector. The coolant mass flow rate has a great influence on the film cooling performance due to the variety of coolant momentum in the exit of orifices and vorticity in the near-injector region. An optimum mass flow rate for maximizing the averaged effectiveness exists for a given film orifice geometric configuration. The high averaged effectiveness and the uniform flux distribution of hot-gas-side wall are observed at small orifices spacing. The film cooling effectiveness in the orifice exit region is obviously enhanced when the diameter of the orifice increased in the front part of the combustion chamber. The results would be useful for the analysis and optimization design of the straight cylindrical coolant orifices in the LOX/GH2 thrust chamber.

In order to understand heat transfer mechanisms in the combustion chamber with multi-elements injector and provide benchmark data of wall heat fluxes for the CFD code validation, experiments were carried out in a GH2/GO2 heat-sink combustion chamber. To obtain the transient heat flux in the experiments, the Levenberg-Marquardt method was modified and applied to the rocket combustion chamber based on the temperature measured by single coaxial thermocouple (Named “single-point method”). The comparison between the single-point method and the two-point method with temperatures measured in two points shows that the single-point method could be used as heat flux measurements of the heat-sink combustion chamber in engineering. Heat flux distribution was obtained in experimental conditions of different work time and chamber pressure by the modified method, feasibly and efficiently. For the transient variations of the heat flux, an inverse variation trend of heat flux to time in the cylindrical segment and the nozzle segment was observed. A typical variation of the averaged heat flux was obtained under the conditions with different chamber pressure and work time. The results of wall heat flux scaled with pressure to the power 0.8 can be uniformized for all pressure levels. The heat fluxes obtained in the cylindrical chamber and nozzle section may be applied for the life cycle prediction of rocket engines.