• Journal of computational fluids engineering
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• v.1 no.1
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• pp.123-141
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• 1996

In this paper the upwind flux difference splitting Navier-Stokes method has been applied to study the perfect gas and the equilibrium chemically reacting hypersonic flow over an axisymmetric sphere-cone(5°) geometry. The effective gamma(γ), enthalpy to internal energy ratio was used to couple chemistry with the fluid mechanics for equilibrium chemically reacting air. The test case condition was at altitude(30km) and Mach number(15). The equilibrium shock thickness over the blunt body region was much thinner than that of perfect gas shock. The pressure difference between perfect gas and equilibrium gas was about 3 ∼ 5 percent. The heat transfer coefficient were also calculated. The results were compared with VSL results in order to validate the current numerical analysis. The results from current method were compared well VSL results ; however, not well at near nose. The proper boundary condition and grid system will be studied to improve the solution quality.

• 한국전산유체공학회:학술대회논문집
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• 1995.10a
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• pp.203-212
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• 1995

In this paper the upwind flux difference splitting Navier-Stokes method has been applied to study the perfect gas and the equilibrium chemically reacting hypersonic flow over an axisymmetric sphere-cone($5^{\circ}$) geometry. The effective gamma($\bar{r}$), enthalpy to internal energy ratio was used to couple chemistry with the fluid mechanics for equilibrium chemically reacting air. The test case condition was at altitude(30Km) and Mach number(15). The equilibrium shock thickness over the blunt body region was much thinner than that of perfect gas shock. The pressure difference between perfect gas and equilibrium gas was about $3\sim5$ percent. The skin friction coefficient and heat transfer coefficient were also calculated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.04a
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• pp.35-38
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• 2001

The aerothermodynamic characteristics of SCRamjet engine for the airbreathing populsion system of the next generation flight vehicle are described. As the flow is passing by, combustion caused the total pressure loss and the Mach number decrease, but nozzle exit velocity is large enough to produce net thrust. To simulate supersonic combustion test, preliminary design of ground-based blowdown type supersonic combustion tunnel is attained. Minimum allowable operating pressure and mass flow rate are calculated for the design Mach number of 2.5 at the test section of a supersonic combustion tunnel.

• Journal of computational fluids engineering
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• v.18 no.3
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• pp.87-93
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• 2013

A computational study has been performed to examine the effects of catalytic walls on the stagnation region heat transfer. The boundary conditions for none, finite, and fully catalytic walls have been incorporated into a multi-block compressible Navier-Stokes solver. In the present study, both chemical and thermal non-equilibrium effects were included. The flows over a blunt body model were simulated by varying surface catalytic recombination rates. A full range of catalycities was explored in the context of a constant wall temperature assumption. Detailed information on species concentrations, temperature, and surface heat flux are presented. Comparison with available flight data of surface heat flux is also made.

• 한국전산유체공학회:학술대회논문집
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• 2002.05a
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• pp.109-114
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• 2002

The effect of nose radius on aerodynamic heating are investigated by using the Wavier-Stokes code extended to thermochemical nonequilibrium airflow. A spherical blunt body, whose radius varies from 0.003048 m to 0.6096 m, flying at Mach 25 at an altitude of 53.34 km is considered. Comparison of heat flux at stagnation point with the solution of Viscous Shock Layer and Fay-Riddell are made. Obtained result reveals that the flow chemistry for very small radius is nearly frozen, and therefore the contribution of heat flux due to chemical diffusion is smaller than that of translational energy. As the radius becomes larger, the portion of diffusion heat flux becomes greater than translational heat flux and approaches to a constant value.

• 한국연소학회:학술대회논문집
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• 2006.10a
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• pp.187-195
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• 2006

In this work, we introduce a newly constructed arc-jet device of 150 kW input power. The design of this device is a Huels type with a narrow downstream electrode. General features of this device are first described. From the measured values of electrical power input, heat discharged into cooling water, gas flow rate, and settling chamber pressure, average enthalpy was determined using the heat balance and sonic throat methods. Using the settling chamber pressure and average enthalpy values, the flow properties in the nozzle and the heat transfer rate to the stagnation point of a blunt body are calculated accounting for thermochemical nonequilibrium. The envelope of enthalpy, pressure, degree of dissociation, and heat transfer rate are presented. Stagnation temperature is predicted to be between 4630 to 6050 $^{\circ}K$, and the stagnation point heat transfer rate is predicted to be between 175 and 318 W/$cm^{2}$ for a blunt body of 3 mm nose radius. Degree of dissociation in the stagnation region of the blunt body exceeds 30%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.584-587
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• 2011

An internal flow aerodynamic test was performed for a Mach 5 scramjet engine. The test was done without fuel injection, as a preliminary test for the combustion test. Test engine is an engineering model with intake cross-section of $70mm{\times}200mm$ and total length of 1.7m. Test facility is a blowdown-type, high enthalpy, hypersonic facility. 19 pressures were measured through the holes on the model surface along the engine internal flow passage. It was found that the facility start is possible, and also supersonic flow is maintained inside the engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.6
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• pp.36-44
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• 2021

Numerical analysis was performed on the shock wave-boundary layer interaction generated in the internal flow path of the curved interstage of the scramjet engine flight test vehicle. For numerical analysis, the turbulence model k-ω SST was used in the compressibility Raynolds Averaged Navier Stokes(RANS) equation. Representatively, the separation bubbles on the upper wall of the nozzle, the interaction between the concave shock wave and the boundary layer, and the shock wave-shock wave interaction at the edge were captured. The analysis result visualizes the shock wave-boundary layer interaction of the curved internal flow path to enhance understanding and suggest design considerations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.46 no.12
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• pp.986-993
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• 2018

Lateral thrust jet has better maneuverability performance than the control surface like the conventional fin for attitude control or orbital transition of guided weapons. However, in the supersonic region, a jet interaction flow occurs due to the lateral thrust jet during flight, and a complicated flow structure is exhibited by the interaction of the shock wave, boundary layer flow, and the vortex flow. Especially, hit-to-kill interceptors require precise control and maneuvering, so it is necessary to analyze the effect of jet interaction flow. Conventional jet interaction analyses were performed under low altitude conditions, but there are not many cases in the case of medium altitude condition, which has different flow characteristics. In this study, jet interaction flow analysis is performed on the lateral jet controlled interceptor operating at medium altitude. Based on the results, the structural characteristics of the flow field and the changes of aerodynamic coefficient are analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.8-8
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• 2000

In this paper are presented main power and gasdynamic characteristics of C-l6VK hypersonic test cell of Research Test Center of CIAM. Gasdynamic adjustment of the C-l6VK test cell was carried out with the working section constructed on scheme of Ramjet/scramjet test in free stream. Gasdynamic adjustment was conducted stage by stage in tile following sequence. First, check and preparation of all technical systems and checking measuring system. Than determination of the characteristics of test cell on cold (without the heating of air at entrance) regime and determination of the characteristics of test cell on regimes with the heating of air. Finally determination of tile characteristics of test cell with the loading of the working part by object. On tile final stage of gasdynamic adjustment two experiments with tile axisymmetric Scramjet model loaded into the working part of test cell were conducted. The first experiment was conducted with the purpose of determination of flow parameters with the object leaded into the working part and verification of experiment cyclogram. The second experiment was conducted with injection of hydrogen into the combustion chamber of object, that is tile conditions on test cell simulated Scramjet flight Mach number M = 6. Such methodology of gasdynamic adjustment allows to determine influence of experimental object on flow parameters in the working part at different conditions of experiment (with the burning in combustion chamber of object and without the homing), and also to compare flow characteristics in the object duct.