• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.93-96
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• 2011

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Mass flow rate to the main thrust chamber, mass flow rate to the gas generator and chamber pressure have been selected as design variables. The target engine is the open gas generator cycle using the LO2/RP-1 propellant. The objective function of design optimization is to maximize the specific impulse with condition of energy balance between the pump and the turbine. The properties of the combustion chamber have been obtained from CEA2. Pump & turbine efficiencies and properties of the gas generator have been modeled mathematically from reference data. The result shows 3~4% errors for the specific impulse and 2~6% errors for the pump power of the gas generator cycle compared to references.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.193-196
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• 2003

The KSR-III developed by KARI is the first rocket vehicle which is adopting the liquid propellant rocket engine system in Korea and its flight test was successfully done last year. KSR-III is a sounding rocket class launch vehicle, but there is a sense to accomplish design, manufacture, performance test and finally its flight test by domestic technology. In this paper, the authors are intended to introduce the multi-purpose test facility(PTA-II Test Facility) which is constructed for the variety of tests on KSR-III feeding system(single component tests, verification tests, cold flow tests and combustion tests) and its test results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.10
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• pp.1026-1030
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• 2010

This paper summarizes the activities performed for the development of a BLDC(Brushless Direct Current) motor driven cryogenic thrust control valve with application to KSLV-II rocket engine. The developed thrust control valve can modulate the flow rate of liquid oxygen under cryogenic temperature of 90K and high pressure of 113.2 bar with the help of electro-mechanical actuator driven by a BLDC motor. This valve can be applied to an engine combustion test after minor change because all development certification test have been performed successfully.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.62-70
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• 2015

A liquid rocket engine system can cause rapid pressure and temperature variations during the startup period. Thus the startup analysis is required to reduce time and expense for successful development of liquid rocket engine through the startup prediction. In this study, a startup analysis simulator is developed for a staged combustion cycle engine powerpack. This simulator calculates propellant flow rates using pressure and flow rate balances. In addition, a rotational speed of turbopump is obtained as a function of time by mathematical modeling. A startup analysis result shows that the time to reach a steady-state and a rotational speed at the steady-state are 1.3 sec and 27,500 rpm, respectively. Moreover it can indicate proper startup sequences for stable operation.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.3
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• pp.38-44
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• 2000

In general liquid rocket nozzles are protected from hot combustion gas by regenerative cooling techniques. But due to the complexity of the cooling system, it causes increase of system cost and frequently source of the system malfunction. Recently, instead of regenerative cooing, ablative material are used to protect combustion chamber wall and nozzle. To determine the nozzle material erosion rate and erosion shape, more than 500 hot fire test were performed by using 100 lb thrust experimental liquid rocket. Test variable were propellant feed sequence, injector, position of igniter and liquid oxygen supply temperature.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.1
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• pp.18-27
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• 2003

To perform combined tests with propellant feeding system and engine, which were developed for KSR-III launcher, vertical test stand was organized and a series of hot-fire combustion tests were carried out with engines of several injector faceplate types. In hot-fire tests in vertical installation, combustion instabilities occurred right after ignition with an engine without baffle, and such combustion instabilities did not occur at ignition add during mainstage operation for an engine with STS or composite baffle. 1.regular and temporary pressure pulsations(popping) were detected during steady operation with a baffle engine, however a development to combustion instabilities with resonant mode was highly suppressed by baffle. With a series of tests, it was confirmed that the last developed engine, which has composite baffle, was operated successfully in KSR-III flight propulsion system.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.5
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• pp.718-724
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• 2000

A model for depicting the rocket engine combustion process is presented and several experiments near a design point are provided with a FOOF type of unlike impinging injector for a propellant combination of Jet A-1 fuel and liquid-oxygen. The model is based on the assumption that the vaporization is the rate-controlling combustion process. The effects of initial drop size and initial drop velocity are systematically shown and discussed. It is seen that in the midst of considered parameters the change of initial drop size is more sensitive to the performance. The proposed model describes qualitative trends of combustion process well despite of its simplicity.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.11
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• pp.874-882
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• 2013

Kerosene(Jet A-1), one of the propellants for each stage's engine of the Korea Space Launch Vehicle-II(KSLV-II), functions as coolant at the same time as it flows inside the cooling jacket of the combustion chambers and is injected through the film cooling holes. A physical surrogate mixture model to reproduce the thermophysical characteristics of Jet A-1 has been selected and the thermodynamic/transport properties of the model fuel under high pressure including supercritical conditions have been estimated using SUPERTRAPP(NIST SRD4). Comparisons with the measured properties suggest that proposed database can be used to extract properties of Jet A-1 for conjugate heat transfer analysis of liquid propellant rocket engine thrust chambers. Predicted combustion/cooling performance of regeneratively cooled thrust chambers shall be validated through comparisons with upcoming firing test results.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.203-207
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the thrust chamber of liquid rocket using LOx and Kerosene as propellant. The heat fluxes were obtained from the measured wall temperature to the axial direction of thrust chamber for different type of coolant, the various O/F ratio, mass flow rate and the location of the film cooling injector. A thin wall combustion chamber and nozzle were used to obtain the heat flux.