• 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.

• Journal of Power System Engineering
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• v.21 no.2
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• pp.89-94
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• 2017

Pneumatic systems are widely applied in various industry because it have a many advantage(low cost, high safety, etc.). Air compressors supply the working fluid to the pneumatic systems and consume a lot of electrical energy at the manufacturing site. The one of the suggested idea is to reduce the energy consumption by reducing the suction temperature of the air compressor and increasing the discharge flow rate. In this paper, the discharge flow rate and air power of the positive displacement type air compressor is simulated by changing the temperature of suction air and the relationship between the suction air temperature and the performance variation of the air compressor is analyzed. As a result, we know that as the suction temperature of air is lowered, the discharge mass flow-rate is increased, but the specific enthalpy is reduced rather than increased, which means that the power of the discharged air is not greatly increased even if lower the suction air temperature.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.1-1
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• 2005

Japan Aerospace Exploration Agency(JAXA) has been conducting research and development of the Scramjet engines and their derivative combined cycle engines as hypersonic propulsion system for space access. Its history will be introduced first, and its recent advances, focusing on the engine performance progress, will follow. Finally, future plans for a flight test of scramjet and ground test of combined cycle engine will be introduced. Two types of test facilities for testing those hypersonic engines. namely, the 'Ramjet Engine Test Facility (RJTF)' and the 'High Enthalpy Shock Tunnel (HIEST)' were designed and fabricated during 1988 through 1996. These facilities can test engines under simulated flight Mach numbers up to 8 for the former, whereas beyond 8 for the latter, respectively. Several types of hydrogen-fueled scramjet engines have been designed, fabricated and tested under flight conditions of Mach 4, 6 and 8 in the RJTF since 1996. Initial test results showed that the thrust was insufficient because of occurrence of flow separation caused by combustion in the engines. These difficulty was later eliminated by boundary-layer bleeding and staged fuel injection. Their results were compared with theory to quantify achieved engine performances. The performances with regards to combustion, net thrust are discussed. We have reached the stage where positive net thrust can be attained for all the test coditions. Results of these engine tests will be discussed. We are also intensively attempting the improvement of thrust performance at high speed condition of Mach 8 to 15 in High Enthalpy Shock Tunnel (HIEST). Critical issues for this purposemay be air/fuel mixing enhancement, and temperature control of combustion gas to avoid thermal dissociation. To overcome these issues we developed the Hypermixier engine which applies stream-wise vortices for mixing enhancement, and the M12-engines which optimizes combustor entrance temperature. Moreover, we are going to conduct the flight experiment of the Hypermixer engine by utilizing flight test infrastructure (HyShot) provided by the University of Queensland in fall of 2005 for comparison with the HIEST result. The plan of the flight experiment is also presented.

• Journal of the Korean Society of Combustion
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• v.7 no.3
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• pp.32-36
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• 2002

A numerical study was carried out to investigate combustion phenomena in a model Scramjet engine, which had been experimentally studied at the University of Tokyo using a high-enthalpy supersonic wind tunnel. The main airflow was Mach number 2.0 and the total temperature of hot flow was 1800K. Equivalence ratio was set to be 0.26 which is higher than that of experiment to investigate the effect of strong precombustion shock. The results showed that self-ignition occurred at the rear bottom wall of the combustor and combined with the shear layer flame between fuel jet and main airflow. Then, precombustion shock was generated at the step location and reversely enhanced the mixing and combustion process behind the shock. Due to the high equivalence ratio, the precombustion shock moved upstream of the step compared with that of experiment.

• 한국연소학회:학술대회논문집
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• 2001.06a
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• pp.127-132
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• 2001

A numerical study was carried out to investigate the combustion phenomena in a model Scramjet engine, which had been experimentally studied in the University of Tokyo using a high-enthalpy supersonic wind tunnel. The main airflow was 2.0 in Mach number and the total temperature of hot flow was 1800K. Equivalence ratio was set to be rather higher value of 0.26 than that of experiment to investigate the effect of strong precombustion shock. The results showed that self-ignition occurred at the rear bottom wall of the combustor and combined with the shear layer flame between fuel jet and main airflow. Then, precombustion shock was generated at the step location and reversely enhanced the mixing and combustion process behind the shock. Due to the high equivalence ratio, the precombustion shock moved upstream of the step compared with that of experiment.

• 한국연소학회:학술대회논문집
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• 2000.12a
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• pp.83-91
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• 2000

A numerical study was carried out to investigate the 'unstart' process of thermally-choked combustion in model scramjet engines. The combustion mechanism of supersonic combustor will be compared with the experimental results obtained from the T3 free-piston shock tunnel at ANU (Australian National University) and the high enthalpy supersonic wind tunnel at UT (University of Tokyo). For the numerical simulation of supersonic combustion. multi-species Navier-Stokes equations were considered. and detailed chemistry reaction mechanism of $H_2$-Air were adopted. The governing equations were solved by Roe's FDS method and LU-SGS method with MUSCL scheme. In this study. it is found that the thermal choking process could result from excessive heat release due to combustion. In detail, sufficient heat release could be generated at local region of very high temperature increased by reflection of shock waves or vortex sheets. Accordingly the flow of downstream of the combustor fell to subsonic field propagated upstream along the combustor. Sometimes the subsonic flow field propagated into isolator could generate precombustion shock waves in the isolator.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.647-650
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• 2017

As the necessity of development of supersonic vehicle increases, securing an aerodynamic data from low to high altitude is requested for flying vehicles crusing in various high-tech environment. Therefore our research team built equipment by improving heating device of Model Aerodynamic Facility(MAF) of Konkuk University to simualte a real gas environment. Guided weapon system and temperature and velocity distribution according to the flow that is produced from the pier of supersonic vehicle is planned to be researched by using this equipment.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.198-202
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• 2005

Gas generator is the equipment to produce high enthalpy gas used to generate sufficient power to operate turbine and pump system for propellant feeding in liquid rocket engine. Since the limit in operating temperature is imposed due to turbine blade, the gas generator has to be operated at the temperature far below stoichiometric maintaining fuel rich combustion. In this research, fundamental study was performed to understand the non-equilibrium combustion process with in-house code and CFD-ACE as well.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.197-200
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• 2008

The existing computer code to solve the air flow in a segmented constrictor-type arc-heated wind tunnel named ARCFLO4 is improved to accept an air-argon mixture as the working gas and to consider the separated injection of air and argon. The new version of the code is used to calculate the flows in the Aerodynamic Heating Facility of NASA Ames Research Center where argon concentration is relatively high. The calculation shows that argon tends to increase the diameter of the arc-column, increase ionization fraction, decrease thermal efficiency of the arc-heater, and push the ratio of the centerline-to-average enthalpy toward unity. The calculated operating characteristics of the arc-heater agree well with the experimental data.