• 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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• 2012.05a
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• pp.52-57
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• 2012

A liquid rocket system consists of a combustion chamber, a gas generator, a turbo pump, and a turbine, etc. Each component is connected by supply components such as valves, pipes, and orifices. Since each component has a combined effect on engine performance, preliminary analysis for overall system must be required before the conceptual design stage. Comprehensive analysis program considered the supply system has not been developed yet. In this paper, a supply component model of the liquid rocket engine has been designed after verification of each component. The gas generator cycle with supply components has been composed. The results of the cycle has been compared to those of the F-1 engine with the representative gas generator cycle.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.6
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• pp.104-110
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• 2018

Operating mode analysis of a liquid propellant rocket engine(LRE) is a crucial tool through the development of an engine. The operating mode analysis of an engine based on a collection of the acceptance tests of components shows discrepancies when compared to the test results. We propose a correction method for performance parameters to develop an engine analysis model for the gas generator cycle of an LRE. In order to simulate engine behavior, the performance parameters for the analysis model are tuned based on the test results of the 75tf engine of KSLV-II.

• 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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• 2017.05a
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• pp.1165-1168
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• 2017

This paper suggests a fault diagnosis of damage on inner liner of regeneratively-cooled combustion chamber during gas generator cycle rocket engine hot firing test. This method focuses on a phenomenon that fuel flow rate difference between two flow estimate methods changes under an inner liner damage of combustion chamber causing fuel leakage and it is expected that it contributes to detect a damage on the combustion chamber in early stage and prevent further destruction during the hot firing test.

• 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, $0.28-0.58MW{\cdot}s/kg$.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.22-30
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• 2015

A gas generator which generates turbine driving gas by burning a part of propellants is used in an open cycle liquid rocket engine and as a main component of an open cycle liquid rocket engine autonomous hot firing tests are required to investigate the combustion performance and characteristics of the gas generator. However, since the combustion gas generated by a gas generator is choked at the turbine nozzle in the turbine manifold, it is necessary to consider the internal volume of turbine manifold as well as that of the gas generator for correct investigation of the combustion performance, characteristics, and acoustic characteristics of the gas generator. Therefore, in the paper hot firing test results of a gas generator with a turbine manifold simulator are described and characteristic prediction using the autonomous test of a gas generator is explained.

• Aerospace Engineering and Technology
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• v.9 no.1
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• pp.93-97
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• 2010

The analysis of specific impulse of the liquid rocket engine employing gas generator cycle with LOx/kerosene as propellant has been performed. The relative error of performance of 300 ton level engine is 0.1%s for specific impulse and 12% for optimal combustion pressure comparing with the published data. The difference of the performance model and the material property models of gas generator product gas are the presumed major reason of discrepancy. The optimal condition of 30 ton level engine is combustion pressure of 68 bar and mixture ratio of 2.2 for maximum specific impulse. This optimal condition can be changed by performance models.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.151.2-151.2
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• 2012

터보펌프 구동에 사용된 가스발생기 생성가스를 연소기로 공급하여 주추력 발생에 사용하는 다단연소 사이클 로켓엔진은 고추력을 요하는 우주 발사체에 널리 사용되고 있다. 다단연소 사이클 로켓엔진에 사용되는 가스발생기를 예연소기라 부르며 케로신과 액체산소를 추진제로 하는 다단연소 사이클 로켓엔진에는 산화제 과잉 예연소기가 사용된다. 예연소기는 터보펌프 구동을 목적으로 하기 때문에 예연소기 생성가스의 횡단면 온도분포는 터빈에 의해 제한되는 온도범위 내에서 균일하여야 하며 넓은 운전영역에서 안정적인 연소가 이루어져야 한다. 산화제 과잉 예연소기는 모든 추진제가 혼합헤드를 통해 분사되는 방식과 추진제를 혼합헤드와 연소실로 나누어 공급하는 방식이 있다. 기술검증을 위해 산화제 일부와 연료를 혼합헤드를 통해 연소실에 공급하여 1차 연소시키고 나머지 산화제를 연소실 냉각채널을 거쳐 연소실 중앙의 분사공을 통해 연소실로 주입하여 기화시키는 형태로 최종적으로 연소압 20MPa, 혼합비 60에서 작동하는 산화제 과잉 예연소기를 설계하여 연소시험을 수행하였다. 혼합헤드에는 별도의 점화용 분사기 없이 전체 연료 분사기를 통해 점화용 연료인 TEA/TEB 혼합물을 분사하여 점화하였다. 추진제를 2단으로 공급할 수 있도록 고안된 가압식 연소시험 설비에서 10회, 누적 60초 이상의 연소시험이 성공적으로 수행되었다. 연소시험결과 넓은 작동영역에서 안정적 연소특성과 생성가스 온도 분포의 균일성을 확인할 수 있었다. 고온 고압의 산화제 과잉 예연소기 기술 확보를 통해 케로신/액체산소 다단연소 사이클 로켓엔진 개발을 위한 기술적 기반을 마련하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.135-138
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• 2006

We developed an engine system design program by balancing the pressure-mass-power relation which can be acquired from each component's specification. In gas generator type open-cycle rocket engine system it is possible to distinguish the variables into two categories, which are input variables and requirement variables. We define 11 design variables corresponding to the 11 balance equations as functions of pressure, mass and power of target engine system. We solved these equations by Newton method. As an example we designed gas generator cycle engine system and finally we could conclude that this developed program is well suited to the engine system design.