• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.25-37
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• 2005

For pump-fed rocket propulsion system, the temperature of LOX to be supplied to turbopump inlet should be satisfied with pump inlet temperature requirement during all operating stages, as excessive temperatures can result in cavitation due to reduction in NPSH, thus either damaging the pump or adversely affecting pump performance rise. So exact estimation of LOX temperature rise is absolutely needed for developing reliable propulsion system. This paper presents systematic analysis scheme for estimating inner process of cryogenic propellant tank which is needed for LOX temperature rise. And this paper presents LOX temperature rise and thermal stratification for all rocket operating stages including cooling, filling, waiting, pre-pressurization and firing, with the application of buoyancy driven boundary layer theory.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.10a
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• pp.6-6
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• 1999

액체 로켓 엔진은 추진기관 공급 시스템으로 작동이 된다. 추진기관 공급 시스템에는 유공압장치 및 각종 배관, 필요한 압력과 유량을 연소실과 가스발생기로 공급하는 시스템, 엔진의 점화 및 정지, 발사체의 사용 목적에 따라 부과되는 기능을 수행하기 위한 장비들이 포함된다. 공급시스템은 크게 가압가스를 이용하는 방법과 터보펌프를 이용하는 방법의 두 가지로 나눌 수 있다. 잘 알려진 바와 같이 일반적으로 추력이 큰 로켓엔진의 경우에는 터보 펌프식이, 추력이 크지 않은 경우에는 가압가스 방식이 이용된다. 일반적으로 가압가스 방식은 연소실 압력이 커질수록 추진제 탱크의 압력도 커지므로, 그 두께가 두꺼워져서 비효율적이 된다. 따라서 연소실 압력이 비교적 크지 않은 추력이 약 10t 내외에서 많이 사용되고, 시스템이 터보 펌프식보다 구조가 매우 간단하므로, 작동의 신뢰도는 매우 높다.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.26-35
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• 2004

As part of thrust control technology research on turbopump-fed type liquid-propellant rocket engine system, the low frequency dynamic characteristics of turbopump was investigated. It can be described that a turbopump system has a 1st-order lag element. When the value, which was resulted by subtraction of the variation of turbine moment with respect to the variation of revolution number from the variation of pump moment with respect to the variation of revolution number, was positive, the time constant of the 1st-order lag element was positive which stood for a stable system. Increasing the above-mentioned valve within positive range leaded to the increase of response and to the decrease of controllability.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.4
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• pp.85-93
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• 2015

Liquid rocket engines of KSLV-II employ a turbopump feed system for propellants. A turbopump real-propellant test facility based on liquid oxygen and kerosene has been constructed for the experimental verification of the turbopump performance using the real media of propellants(i.e., LOX/Kerosene). The verification tests of sub-systems were performed such as LOX/kerosene feed system and alcohol burner system. Finally, the performance of the whole system was executed and verified through a sets of validation tests with the development model of the KSLV-II turbopumps. It has been confirmed that the test facility satisfies the operating conditions and time of the turbopump at the design and off-design performance test using real-propellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.107-110
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• 2009

A basic design for a Technical Development Model (TDM) of liquid oxygen lines from the turbopump exit to the oxidizer valves of the combustion chamber and the gas generator was conducted to develop a turbopump-fed liquid rocket engine. The TDM is composed of straight lines, elbows, bellows, a branch, an orifice, flanges and a heat insulator. Materials were determined by consideration of operation conditions, weight constraint and manufacturing procedures. The size and the location of each component were determined by flow analysis of the required flowrate and the pressure loss. Basic designs of the components were conducted by consideration of the operating temperature and the maximum expectation operating pressure. The safety factors were evaluated by structural analysis of design of each component.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.122-126
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• 2006

Two types of pressurization system and low weight feeding piping system are developed. With sub-system tests, ullage pressure control performance was verified for 1 step and 2 step pressurization system and the feeding performance of feeding piping system was also verified. The weight of the feeding piping system is low enough for the application of launch vehicle. In addition, LOX conditioning system is developed for avoiding geysering and LOX temperature rise. Integrated performance was verified through integrated on-board feeding system performance tests.

• Aerospace Engineering and Technology
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• v.2 no.2
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• pp.151-156
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• 2003

Liquid rocket engine system is classified into an engine of pressurization and turbo pump type by the way of fuel fed-supporting system. In the KSR-III sounding rocket, an engine of pressurization type was used, but there was lots of technical problems to be solved for a use as the first stage engine of space launch vehicle. So, an engine of turbo pump type was required to be developed to overcome the technical limitation of liquid rocket engine. In this research, the analysis of propellant of Kerosine-LOX and methane-LOX which are noticed as a future propellant was carried out for the purpose of studying the basic characteristics. And to review the basic characteristics of an engine of turbo pump type, among the sizing variant of the space launch vehicle, the ways of injecting a satellite to a direct orbit and transient orbit were discussed in this paper.

• 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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• 2011.11a
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• pp.892-897
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• 2011

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility(CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.4
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• pp.87-92
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.