• International Journal of Aerospace System Engineering
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• v.5 no.2
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• pp.16-22
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• 2018

An analysis has been made on the performance variation due to pressure drop change at propellant supply pipes of liquid rocket engine. The objective is to compare the effectiveness of control variables to tune the liquid rocket engine performance. The mode analysis program has been used to estimate the engine performance for different modes which is realized by controlling the flow rate of propellant. The oxidizer of combustion chamber, the fuel of combustion chamber, the oxidizer of gas generator and the fuel of gas generator are the independent variables to control engine thrust, engine mixture ratio and temperature of gas generator product gas. The analysis program is validated by comparing with the powerpack test results. The error range of compared variables is order of 4%. After comparison of tuning effectiveness it is turned out that the pressure drop at oxidizer pipe of gas generator and pressure drop at combustion chamber fuel pipe and the pressure drop at the fuel pipe of gas generator can effectively tune the thrust of engine, mixture ratio of engine and temperature of product gas from gas generator respectively.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.139-143
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• 2003

Study on the ignition characteristics of combustor and gas generator for LOx-kerosene liquid rocket engine was performed experimentally through a series of combustion tests of sub-scale engine combustor and gas generator. Characteristic of gas-torch ignitor based on gaseous methane and gaseous oxygen was compared with hypergolic ignition using propellant tri-ethyl-aluminium. Gas-torch ignitor showed good performance on igniting sub-scale liquid rocket engine combustor and gas generator. It was observed that the ignition delay is also affected by the extent of nitrogen in the combustion chamber.

• 한국연소학회:학술대회논문집
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• 2012.11a
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• pp.363-366
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• 2012

Detonative combustion is considered as a promising combustion mechanism for improving thermodynamic efficiency of power generation systems as a PGC, as well as high-speed propulsion systems. Among the various types of detonative combustion, RDE is fascinated by many researchers because of the simplicity and continuos operation characteristics. Present paper is an introduction to the physical and operational concept of RDE with a brief history of RDE researches and recent development activities. Additional discussions will devoted to the relevance to the tangential mode instabilities in liquid rocket engines and improvement of liquid rocket performance.

• 한국연소학회:학술대회논문집
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• 2003.05a
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• pp.41-50
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• 2003

Numerical analysis was carried out to investigate performance and combustion characteristics of KSR-III liquid rocket engine with several types of baffle. To evaluate the change of performance and combustion characteristics with several types of baffle, the first numerical calculations were performed about baffle tab, radial blade baffle, and hub-and-spoke baffle. Then radial blade and hub-and-spoke baffle were determined to design two types of the KSR-III engine with baffles. Also to investigate the effect of injector arrangements and baffle positions, two types of radial blade baffle were calculated then numerical calculations were carried out with changing axial length of radial blade I, II and hub-and-spoke baffle. While axial length of baffle effected to performance very small, injector arrangement effected to performance largely through calculations of radial blade I, II. From the viewpoint of combustion instability, hub-and-spoke baffle controlled combustion instability effectively and there was the performance of hub-and-spoke baffle between radial blade I and II.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.93-97
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• 2010

The energy balance program which can balance the relations among energy, mass flow, pressure in the staged-combustion cycle of the liquid rocket engine has been developed. The modular approach has been chosen for the analysis; the engine cycle consists of the elements from the predefined component analysis program. The engine with the staged-combustion cycle has been decomposed into several principal component modules, such as a thruster chamber, turbopumps, turbines, supply system components and a pre-burner. The program has been verified with comparison of the results to the selected data of the space shuttle main engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.574-577
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• 2012

Thrust measuring is one of the crucial factor to decide the performance of a liquid rocket engine when the engine development test, especially for the combustion chamber, is implemented. Calculating the thrust from a combustion pressure is used when direct measuring the thrust is impossible, but direct measuring the thrust is necessary and various methods for doing it more precisely should be considered. This paper introduces the preliminary design concept about the new thrust measurement system for the vertical firing test stand, which is introduced domestically for the first time, of a liquid rocket engine combustion chamber.

• Proceedings of the KSME Conference
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• 2008.11b
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• pp.3154-3158
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• 2008

A liquid rocket engine performance has been analyzed as a function of combustion pressure with LOx/RP-1R. The present method is verified by comparing the specific impulse for various combustion pressure with given pump head model. The optimal combustion pressure is between 150 bar and 200 bar for given efficiencies. Both the optimal combustion pressure and the specific impulse increase for increased turbine efficiency. The optimal combustion pressure decreases and the specific impulse increases for increased combustion efficiency. The pump efficiency and the turbine inlet temperature have the same qualitative effect as the turbine efficiency.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.54-58
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• 2017

A combustion instability has been one of the most serious problems in the development of combustion devices including rocket engine and gas turbine. In particular, a high-frequency combustion instability generated by resonant coupling between combustion phenomena and acoustic oscillations within thrust chamber causes severe damage to the hardware. Because it is accompanied by high amplitude pressure oscillations and excessive heat flux to the chamber wall. Therefore, combustion instability is one of the difficult problems that must be resolved in developing liquid rocket engine. This paper describes the cases of combustion instability encounted during the development of thrust chamber for KSR-III and KSLV-II.

• Aerospace Engineering and Technology
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• v.12 no.1
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• pp.200-206
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• 2013

A performance sensitivity of liquid rocket engine to propellant density or supply pressure change was studied. The analysis program was verified to have 1% error comparing with the measured data of a turbopump-gas generator system. The engine combustion pressure decreases as fuel supply pressure increases due to decreased mixture ratio which reduces the turbine power. The engine combustion pressure increases as fuel density increases because the total propellant flow rate is increased substantially even though mixture ratio is slightly decreased. The engine combustion pressure increases when the oxidizer density or supply pressure increases.

• Aerospace Engineering and Technology
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• v.7 no.2
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• pp.157-161
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• 2008

This study introduce a newly developed program to calculate the combustion process of combustion chamber and gas generator of liquid rocket engine by use of Gibbs free energy minimization method based on chemical equilibrium. The simulation results of the new program and CEA code of NASA were compared and showed good agreement, thus proving the validity of the newly developed in-house program for combustion analysis.