• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.764-767
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• 2010

The high enthalpy plasma research center in Chonbuk national university is under construction for MW class plasma wind tunnel. Four types of plasma equipment will be installed in the research center. The equipments are 1set of 0.4 MW class enhanced Huels type plasma equipment, 1 set of 2.4 MW class enhanced Huels type plasma equipment, 1 set of 60 kW RF plasma equipment and 1 set of 200 kW RF plasma equipment. And electrical, water and gas utilities to assistant plasma equipments are under construction. The research center consists of experiment building, research building, power supply building, air supply building, cooling tower foundation.

• 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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• 2008.03a
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• pp.460-468
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• 2008

Recent activities on the scramjet and rocket-ramjet combined-cycle engine of Japan Aerospace Exploration Agency(JAXA) are herein presented. The scramjet engines and combined-cycle engines have been studied in the world and JAXA has also studied such the engines experimentally, numerically and conceptually. Based on the studies, 2 to 3 m long, hydrogen-fueled engine models were designed and tested at the Ramjet Engine Test Facility(RJTF) and the High Enthalpy Shock Tunnel(HIEST). A scramjet engine model was tested in Mach 10 to 14 flight condition at HIEST. A 3 m long scramjet engine model was designed to reduce a dissociation energy loss in a high temperature condition. Drag reduction by a tangential injection and two ways of a transverse fuel injection were examined. Combustor model tests at three operating modes of the combined-cycle engine were conducted, demonstrating the combustor operation and producing data for the engine design at each mode. Aerodynamic engine model tests were conducted in a transonic wind tunnel, demonstrating the engine operation in the ejector-jet mode. A 3 m long combined-cycle engine model has been tested in the ejector-jet mode and the ramjet mode since March 2007. Carbon composite material was examined for application to the engines. Production of the cooling channel on a nickel alloy plate succeeded by the electro-chemical etching.

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.4
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• pp.316-329
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• 2012

In this paper, optimum design process of ICP (Inductively Coupled Plasma) torch, which has been used widely in aerospace application, such as supersonic plasma wind tunnel, is presented. For this purpose, the behaviors of equivalent circuit parameters (equivalent resistance and inductance, coupling efficiency) were investigated according to the variations of torch design parameters (frequency, $f$, confinement tube radius, $R$ and coil turn numbers, $N$) in the basis of analytical and numerical MHD (Magneto Hydro-Dynamics) models combined with electrical circuit theory. From the results, it is found that equivalent resistance is increased with the increase of $f$ values but vice versa for equivalent inductance. For elevated values of $R$ and $N$, however, both parameters tend to increase. Based on these observations, ICP torch with a power level of 10 kW can be optimized at the design ranges of $f$=4~6 MHz, $R$=17~25 mm and $N$=3~4 to maximize the electrical coupling efficiency, which is the ratio of equivalent resistance to equivalent inductance.