• 한국연소학회:학술대회논문집
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• 2012.11a
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• pp.291-294
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• 2012

The research and development status of combustion chamber of liquid rocjet engine for Korea Space Launch Vehicle(KSLV-II) are briefly described. The cold and hot firing tests of uni-element injector, the performance/heat flux measurement/hot firing tests of subscale combustion chamber and the performance/stability rating/regenerative cooling/hot fire tests of 30ton-class combustion chamber were successfully performed. Based on these results, the research and development of combustion chamber for 75ton-class liquid rocket engine are underway.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.168-172
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• 2008

Hot firing tests of sub-scale combustors were carried out to study the characteristic velocity according to the variation of propellant mass flow and injector arrangement. Test results show that there exists an effective range of relative flow-rate density on the condition of similar combustion pressure and mixture ratio. Numerical analysis has also revealed that the increase of the distance between the outermost injector array and the cylindrical chamber wall with film cooling increases the region of low mixture ratio near combustion chamber wall and it decreases the characteristic velocity of the combustor. Thus, it was confirmed that these two factors play an important part in improving the performance of LRE combustor on a predetermined chamber pressure.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.614-615
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• 2010

The propellants distribution of an oxidizer rich preburner was predicted by a simplified physical approach. The Mixing head is composed of honeycomb array of 7 fuel injectors and 24 oxidizer injectors. The OF ratio of the mixing head is 15. As results, the OF ratio of the central area is about 9 and the OF ratio of the wall area is about 30.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.56-62
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• 2011

An experimental study on design of a catalytic ignitor was performed to use an ignition source for a small bi-propellant liquid rocket engine which use hydrogen peroxide and kerosene as propellants. In the catalytic ignitor, hot gas of hydrogen peroxide which was decomposed by a catalyst induced autoignition of kerosene. Mass flow rate and O/F ratio for the ignitor were calculated by CEA code. A combustion chamber which had a quartz window and thermocouples was manufactured to determine whether the ignition is successful. Ignition performance was investigated according to exit area of fixed rings and mixture ratio. Results showed that reliable ignition performance was achieved at non-choking exit area of fixed ring and O/F ratio of 6~8.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.101-106
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• 2019

Carbon deposition on the turbine nozzle throat of a LOx/kerosene gas generator cycle(open cycle) engine causes performance reduction of the engine. Estimation methods for a turbine nozzle throat area are proposed. The discharge coefficient of the turbine nozzle was estimated with the turbine gas properties such as gas constant, specific heat ratio, and temperatures. The pressure ratio and temperature ratio of the turbine nozzle throat, was utilized to estimate the discharge coefficient also. Estimated discharge coefficient of turbine nozzle throat of KSLV-II 1st stage engine shows the carbon deposition effects on the turbine nozzle throat of a LOx/kerosene open cycle engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.16-19
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• 2011

An experimental study on the spray characteristics of Gas-centered swirl coaxial injectors(GCSCI) for high-performance staged combustion rocket engines has been carried out using cold flow tests. In this study, water and gaseous nitrogen are used as working fluids and a back-lit photography technique with image processing for the measurements of spray characteristics. Our study is focused on the effect of injector geometries like as gap thickness of liquid nozzle and gas nozzle and momentum flux ratio for fundamental understanding of the injectors.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.1
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• pp.50-57
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• 2014

A ground hot-firing test (HFT) was conducted to take out the optimal design configurations for the thrust chamber of 70 N-class liquid rocket engine under development. Monopropellant grade (purity: ${\geq}98.5%$) hydrazine was adopted as a propellant for the HFT, and three kinds of thrust chambers having characteristic lengths ($L^*$) of 2.79, 2.95, and 3.13 m were selected for their performance evaluation. It is revealed through the test and evaluation that the increase of the $L^*$ leads to a performance degradation in the test condition specified, and pulse response performance of the development model shows superior characteristics to commercialized hydrazine thrusters.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.85-90
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• 2003

To control spacecraft including satellite, we should understand precisely the performance of propulsion system and the program logic with appropriate format for satellite operations. In this study, the thruster performance functions was generated by using the best curve fitting for performance data from bi-propellant thrusters. Detailed thruster performance data are, in general, company proprietary information, therefore real firing tests were performed to understand the basic characteristics of the performance curve. Experimental rocket motor utilize liquid oxygen and kerosine as propellant and designed average thrust was 100 pound.

• Journal of ILASS-Korea
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• v.15 no.4
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• pp.189-194
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• 2010

This study has been mainly motivated to numerically model the transcritical mixing and reacting flow processes encountered in the liquid propellant rocket engines. In the present approach, turbulence is represented by the extended k-$\varepsilon$ turbulence model. To account for the real fluid effects, the propellant mixture properties are calculated by using SRK (Souve-Redlich-Kwong) equation of state model. In order to realistically represent the turbulence-chemistry interaction in the turbulent non-premixed flames, the flamelet approach based on the real fluid flamelet library has been adopted. Based on numerical results, the detailed discussions are made for the real fluid effects and the precise structure of the transcritical cryogenic liquid nitrogen jet and gaseous hydrogen/liquid oxygen coaxial jet flame.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.24-33
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• 2020

In this study, model combustion tests were conducted to investigate the combustion instability characteristics of swirl coaxial injectors for a liquid rocket engine. To examine the effects of the combustion chamber resonant frequency and the injector mixing conditions, pressure fluctuations in the combustion chamber were measured by changing the combustion chamber length, injector recess length, and propellant mixture ratio. From the test results, the variation in the pressure fluctuations for each experimental condition was confirmed and the combustion stability was evaluated by stability mapping. It was found that the longitudinal mode and Kelvin-Helmholtz instabilities occurred due to the change in the combustion chamber and recess lengths.