Abstract
A numerical study was conducted to investigate the flow and heat transfer characteristics of a supersonic second throat exhaust diffuser for high-altitude simulations. The numerical results were satisfactorily validated by the experimental results. A subscale diffuser using nitrogen was utilized to investigate starting pressure and pressure variation in the diffuser wall. Based on the validated numerical method, the flow and heat transfer characteristics of the diffuser using burnt gas were evaluated by changing operating pressure and geometric shape. During normal diffuser operation without cooling, high-temperature regions of over 3000 K appeared, particularly near the wall and in the diffuser diverging section. After cooling, the flow and pressure distribution characteristics did not differ significantly from those of the adiabatic condition, but the temperature in the subsonic flow section decreased by more than 1000 K. Furthermore, the tendency of the heat flux from the diffuser internal flow to the wall was similar to that of the pressure variations, and it increased with operating pressure. It was confirmed that the heat fluxes of the supersonic and subsonic flows in the diffuser were proportional to the operating pressure to the 0.8 and −1.7 power, respectively. In addition, in the second throat region after separation, the heat flux could be scaled to the Mach number ratio before and after the largest oblique shock wave because the largest shock train affected the heat flux of the diffuser wall.
Highlights
Rocket engines used in launch vehicles operate in various environments, ranging from the earth’s atmosphere to outer space
The diffuser wall was considered adiabatic, and the variation in the flow characteristics of the diffuser according to the operating pressure was investigated
The heat flux exhibited the same tendency as the pressure fluctuation caused by the shock train in the second throat, and it increased gradually in the temperature recovery section, where the supersonic flow transitioned to a subsonic flow
Summary
Rocket engines used in launch vehicles operate in various environments, ranging from the earth’s atmosphere to outer space. Sung et al employed theoretical models and compared numerical simulations with experimental data to investigate the effects of major design parameters such as the area ratio of the diffuser to the rocket–motor nozzle throat zone, vacuum chamber size, rocket–motor pressure, and others to characterize startup and operational conditions [3]. Several promising flow control techniques have been examined to manipulate the shockwave–boundary layer interaction He et al performed a detailed numerical study on flow separation behavior at sea level and simulated high-altitude conditions [16]. Numerical Analysis In the present study, the STED used for high-altitude simulation was designed through use of the one-dimensional normal shock theory based on various technical and design data [18] Gk is the turbulent kinetic energy generated by the average velocity gradient, Γ is the effective diffusivity, and Y is the dissipation term [20]
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