• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.2
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• pp.10-17
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• 2004

The results of the combustion performance tests of gas generator which supplies hot gas into the turbine of turbo-pump for liquid rocket engine and uses LOx and kerosene as propellant are described. The gas generator consists of a injector head with F-O-F impinging injector, a water cooled combustion chamber, a gas torch igniter, a turbulence ring and an instrument ring. The effect of turbulence ring and combustion chamber length on performance of gas generator are investigated. The ignition and combustion at design point are stable and the pressure and gas temperature at gas generator exit meets the target. The turbulence ring installed at middle of chamber effectively mixes hot gas with cold gas and the effect of residence time of hot gas in gas generator on combustion efficiency is small. Test results show that the main parameter controlling the gas temperature at gas generator exit is overall O/F ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.4
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• pp.38-45
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• 2007

A liquid rocket engine fuel-rich gas generator has been developed for the first time in the country, which can produce combustion gas over the rate of 4 kg/s at 900 K and 58 bar. The gas is not only for driving a turbopump but also for providing heat source for propellant supply tanks. The final design of the gas generator had been fixed based on the concept and preliminary development tests, and was validated through structure and heat transfer analysis. The manufacturing involved precision machining, surface finish, and special welding technique. The final assessment on the characteristics of ignition and combustion had been carried out for two different versions of injector heads. This concluded that the present product satisfies the development requirements such as spatial temperature distribution and the development has been successful.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.220-223
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• 2005

System design parameters of gas generator cycle liquid rocket engines were investigated and compared in the present study. Characteristic velocity of combustor, pressure drop of combustor injector, exit pressure of pump, pump efficiency and specific power of turbine were considered as a system design parameter. The result shows the characteristic velocity is in the range of 1700-1770 m/s, pressure drop of combustor injector, 4-10 bar, pump exit pressure ratio to combustion pressure, 120-230%, pump efficiency, 60-80%, specific power of turbine, $0.28-0.58MW{\cdot}s/kg$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.87-91
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• 2004

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of launch vehicle successfully. We performed the dispersion analysis of gas generator cycle LRE (liquid rocket engine) accompanied with ANASYN. As a result, the vacuum thrust dispersion of the engine was $+5.34\%,\;-5.27\%$ and the mixture ratio deviated $+9.07\%,\;-9.82\%$ from the nominal value due to the errors of components and engine inlet condition of propellants. By applying the gas generator regulator only, the dispersion of the engine performance increases. Error in turbine efficiency is the most influential factor to the dispersion of engine performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.447-450
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• 2012

An analysis of chill-down process was performed for 30 tonf Turbopump-Gas generator coupled tests. The chill-down process must be fulfilled before liquid rocket engine test using cryogenic propellant. Cavitation, damage and/or combustion instability due to bubble of propellant must be eliminated by chill-down process in a test specimen, especially cryogenic pump. The analysis of test data obtained by 30 tonf TP-GG coupled tests was performed in order to be based on the test process of KSLV-II liquid propellant rocket engine which will be developed. To macroscopically understand the process of chill-down from the viewpoint of test procedure the temperatures of important part and total time of chill-down process were analyzed.

• Bulletin of the Korean Space Science Society
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• 2003.10a
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• pp.70-70
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• 2003

액체 로켓 엔진에 있어 극저온 추진제인 액체 산소를 사용하는 경우에는 He을 가압제로 사용하는 것이 가장 바람직하지만, 기체인 헬륨은 발사 대기시, 선가압시, 비행중에 액체산소에 서서히 녹게 된다. 일정량 이상의 He이 용해되어 있는 LOX가 엔진에 공급되는 경우에는 터보펌프의 이상 작동 또는 연소 불안정을 야기하게 되므로, 추진기관이 작동하는 동안에 용해되어 있던 He이 액체 산소에서 분해되어 가스로 발생되는지 여부를 판단하고, 이는 엔진의 연소 시험을 통해서 검증되어야 한다. 본 연구에서는 가상의 작동 상태에 대해 최대로 용해될 수 있는 러e의 양을 계산하고, 현재 사용되는 발사체의 경우와 비교를 하여 추진시스템 운용 조건을 적절히 조절하는 방안을 제시하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.863-866
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• 2011

Recently environmental issue is more and more emphasized and 'Green Growth' became on of the key words of this Government. Based on this trend, the exhaust gases out of the gas generator cycle and the staged combustion cycle LRE whose propellants are kerosene and LOx were compared. For this purse, 8 tonf class of each cycle engine was designed and the amount and the components of the gases were investigated by using CEA. As expected, the staged combustion cycle engine generates less pollutants than the other cycle. In addition, the graphite that is generated by the gas generator can be reacted with the oxygen in the atmosphere creating additional pollutants.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.125-128
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• 2008

As a interstage of the 30tonf level LOx/kerosene liquid rocket engine development, turbopump-gas generator open-loop coupled tests are performed. Test schematic and test results of open-loop coupled tests are presented. In engine system operation environment simulating combustion chamber by flow control orifice, chill-down procedure, startup characteristics, nominal operability of turbopump+gas generator open-loop coupled Test Plant are confirmed The results of open-loop coupled test were used for the preparation on turbopump+gas generator closed-loop test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.129-132
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• 2008

For the development of the 30tonf level LOx/kerosene liquid rocket engine, turbopump-gas generator closed-loop coupled tests are performed. To simulate engine operation conditions, combustion chamber was substituted by flow control orifices. In simulated engine system operation environment, chill-down procedure, startup characteristics, nominal operability of turbopump+gas generator coupled Test Plant are confirmed. Turbopump and gas generator are confirmed to operate well in simulated engine environment. The control system for regulating power and mixture ratio of Test Plant are also successfully confirmed.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.4
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• pp.443-448
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• 2010

In this paper, the changes in performance parameters, e.g., the combustor pressure, turbine power, engine mixture ratio, temperature of gas generator, and product gas, of a liquid rocket engine employing gas generator cycle with the variations in propellant-supply pressure have been described. Engine performance is numerically calculated using the 13 major system-level variables of the rocket engine. The combustor pressure and turbine power increase with an increase in the oxidizer-supply pressure and decrease with an increase in fuel-supply pressure. The lower mixture ratio of gas generator for increased fuel mass flow rate decreases the gas generator gas temperature and deteriorates the gas material properties as the turbine working fluid. The turbine power decreases with an increase in fuel-supply pressure; this results in a decrease in the main-combustor pressure, which is directly proportional to engine thrust.