• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.5
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• pp.91-96
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• 2004

An optimal design of the gas generator for Liquid Rocket Engine (LRE) was conducted. A fuel-rich gas generator in open cycle turbopump system was designed for 10ton in thrust with RP-1/LOx propellant. The optimal design was done for maximizing specific impulse of thrust chamber with constraints of combustion temperature and for matching the power requirement of turbopump system. Design variables are total mass flow rate to gas generator, O/F ratio in gas generator, turbine injection angle, partial admission ratio, and turbine rotational speed. Results of optimal design provide length, diameter, and contraction ratio of gas generator. And the operational condition predicted by design code with resulting configuration was found to maximize the objective function and to meet the design constraints. The results of optimal design will be tested and verified with combustion experiments.

• 한국연소학회:학술대회논문집
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• 2004.11a
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• pp.141-145
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess length have been tested for the comparative study of CH flame structure and dynamic characteristics. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a full-scale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korea Institute of Military Science and Technology
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• v.6 no.4
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• pp.113-123
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• 2003

Both numerical analysis and experiment of cold and hot tests were performed to obtain basic design data for the swirl coaxial type Injector and to predict the combustion performance. Mass distribution, mixing distribution, mixing efficiency, characteristic velocity efficiency were measured by the cold tests and numerical analysis using the commercial thermo-hydraulic program. Test and analysis variables were recess, pressure drop, velocity ratio, mixing spray, mixture ratio. Hot tests were performed for the Uni-element injector to compare the performance with the cold test results, and, hot tests for Multi-element injector were performed to compare the performance with Uni-element injector. Designed thrust of the Uni-element injector liquid rocket was 35kgf at sea level and combustion chamber pressure, 20bar. Kerosene and Lox were used as a propellant.

• 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.14 no.5
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• pp.8-14
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• 2010

For the development of a liquid rocket engine, hot-firing tests of a regeneratively cooled thrust chamber were performed at chamber pressures of approximately 30 and 60 bars. In the paper, pressure fluctuation data, which were obtained from the dynamic pressure transducers installed in propellant manifolds and combustion chamber, were analyzed. Compared to the data at chamber pressure of 60 bar, the results at chamber pressure of 30 bar showed low-frequency oscillations around 150 Hz in the combustion chamber. The low-frequency waves in the combustion chamber were coupled with those in the manifolds. However, the RMS values of the chamber pressure fluctuations at chamber pressure of 30 bar were only 0.8% of the chamber pressures. Thus, it can be inferred that the thrust chamber operates in the stability boundary even at low chamber pressure.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.15 no.6
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• pp.3403-3409
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• 2014

Hot-firing tests were carried out using a mixing head with 19 swirl coaxial injectors and a combustion chamber with internal cooling channels. The propellants of liquid oxygen and kerosene(Jet A-1) were burned in a range of chamber pressures (59~82 bar) and mixture ratios (2.0~3.0). The temperature of water used as the cooling fluid was measured at the inlet and outlet of the cooling channels, and the heat flux was calculated. The aim of this study was to examine the effect of combustion instability on heat transfer in a subscale thrust chamber, and detect the temperature variation of cooling water. During several hot-firing tests, combustion instability was encountered which caused a 5~20% increase in heat flux. The peak heat flux took place in the initial stages of combustion instability.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.26-37
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• 2011

A 75-tonf-class LRE(liquid rocket engine) thrust chamber with ablative material for technology demonstration was manufactured on the basis of development technologies of 30-tonf-class LRE. Hydraulic characteristics of the thrust chamber were examined through cold flow test and ignition test of low flow condition. Test result showed that hydraulic function was good. Side ignition method with igniter ring also showed a fine function of ignition in operating ways of static condition. But a close review is required to understand the phenomena of generation and extinction of specific frequencies showed in dynamic characteristics ways. To achieve these, a large combustion test facility which is capable of performing combustion test at design condition of the 75-tonf-class thrust chamber should be constructed as soon as possible.

• Journal of Aerospace System Engineering
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• v.16 no.5
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• pp.17-25
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• 2022

The pintle injector has many advantages in the key characteristics of a liquid rocket engine, such as combustion stability, combustion efficiency, and wide range of comprehensive thrust control, design and manufacture, and test fired under supercritical conditions. The pintle injector is manufactured with a rectangular, single-row orifice for thrust control and production considerations. In order to verify the combustion performance of the pintle injector and its potential as a commercial injector, the combustion characteristics were analyzed by varying the TMR (Total Momentum Ratio) and BF (Blockage Factor). The result of the hot firing test showed that the heat flux increased as TMR increased, and it confirmed that the characteristic velocity efficiency was more affected by BF than TMR. Suppose a single-row pintle injector with efficiency characteristics insensitive to changes in TMR can achieve high efficiency at low fuel differential pressure conditions. In that case, the variable pintle injector's design flexibility can be increase.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.130-133
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• 2009

The sequence of the propellant supply is very important for the reliable and safe operation of a LRE combustion test. So combustion performance tests were performed to find an optimum test sequence by changing supply time of propellants and purge gas in the moment of ignition and extinction. The liquid rocket engine consisted of a catalytic ignitor and six swirl-coaxial injectors which used hydrogen peroxide and kerosene. Conclusively, an optimum sequence was found for stable combustion in the moment of ignition and extinction.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.258-261
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• 2003

The optimal design and combustion analysis of the gas generator for Liquid Rocket Engine (LRE) were performed. A fuel-rich gas generator in open cycle turbopump system was designed for 101on1 in thrust with RP-1/LOx combination. The optimal design was done for maximizing specific impulse of main combustion chamber with constraints of combustion temperature and power matching in turbopump system. Results of optimal design show the dimension of length, diameter, and contraction ratio of gas generator. The configuration of the gas generator and the condition for performance which can maximize the objective function were determined and found to meet the design constraints. Also, the combustion analysis was conducted to evaluate the performance of designed chamber and injector of gas generator. And the effect of the turbulence ring was investigated on the mixing enhancement in the chamber.