• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.174-180
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• 2006

The present paper describes transient thermal and mechanical analyses of a lox/kerosene type turbopump in a LRE(Liquid Rocket Engine). Turbopumps are used to pressurize propellants to achieve higher specific impulse of LRE. The turbopump under development has been designed and verified by structural analyses using finite element methods. Some parts of the turbopump operate under cryogenic environments, while the others work under ambient and high temperature environments. Therefore, numerical analysis at a turbopump system level is essential. In this study, casing assemblies of lox pump and fuel pump were analyzed to determine strength test and air-tightness test conditions. Also, some operational stress and strains of fuel pump casings were measured and analyzed. Based on these results, stress concentration of fuel pump casings during the operation could be successfully predicted.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.65-72
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• 2015

Performance tests of a turbopump for the developing 7-tonf liquid rocket engine were conducted. The performance of turbopump components and their power matching were measured and examined firstly under the LN2 and water environment. In the real propellant(LOX and kerosene) environment tests, design and off-design performances of turbopump were fully verified. During the off-design tests, turbopump running time was set the same as engine operating time and pump inlet pressure were set lower than nominal operating value in order to investigate pump suction capability. It have been verified that subject turbopump satisfies required performance - flow rate, head, suction performance and operational time - in the operating regime of developing liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.23-26
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• 2005

The method of decreasing the ecological risk for the LRE(Liquid Rocket Engine) test is developed, working on the cryogenic oxidizer and the high-boiling fuel(Kerosene). This Method is based on the study that contains a technical solution method and an accident occurrence range for decreasing of accident probability and damage. This paper contains schematic on the all risk circumstance bring to accident, block-diagram for an accident growth process in case of the propellant leakage, evaluation method of safety distance from test stand to residential area. It will be used to alternative method for the risk reduction of complex technical systems.

• Journal of the Korea Institute of Military Science and Technology
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• v.12 no.3
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• pp.396-405
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• 2009

The experimental research was conducted to setup a performance prediction logic for the regenerative cooling system on a small scale liquid rocket engine using kerosene and LOX. Total heat flux of the combustion gas side was determined for the flow rate of the coolant, combustion pressure using the calorimeter thrust chamber. Based on the experimental investigation, a performance prediction scheme for the regenerative cooling system is setup in our own way. A performance prediction logic for the regenerative cooling system has been developed by the correction scheme of the combustion gas side. The key parameters determining the temperature limitation of the coolant are the mass flow rate of the coolant and the length of the combustion chamber and the nozzle. And the parameters to control the limitation of the usable wall temperature are the number of channels and wall thickness.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.225-232
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• 2003

With an aim placed on its exploitation on practical injector design, liquid phase mixing due to unlike split triplet impinging element is experimentally investigated by a series of cold tests. Non-reacting kerosene/water spray simulates the kerosene/LOX propellant combination. Measurements of local mixture ratio distribution were made for different injection configurations and different momentum ratios. Mixing and mixing controlled characteristic velocity efficiencies are measured in terms of oxidizer/fuel jet momentum ratio from 0.5 to 8. Extent of mixing and its influence on hot performance are estimated in terms of mixing efficiency and mixing controlled characteristic velocity. Envelope of design locus for optimum mixing quality and corresponding maximum hot performance are proposed. Effects of momentum ratio, orifice diameter ratio and jet velocity ratios are also presented and discussed.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.1
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• pp.68-76
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• 2022

Upper stage of the Korea Space Launch Vehicle-II(KSLV-II) uses a 7-tons class liquid rocket engine and is an open gas generator cycle with a turbopump supply method that uses kerosene/liquid oxygen as the propellant combination. This study first provided a brief overview of the design and development process of the upper stage engine. In addition, it introduced the solutions and results applied to some of the problems that occurred during the development process of the upper stage engine.

• 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.

• 한국연소학회:학술대회논문집
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• 2004.11a
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• pp.213-216
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• 2004

The results of combustion performance test of fuel-rich gas generator(GG) using LOx and kerosene as propellant at design and off-design point are described. The parameters used in this analysis are the average exit temperature($T_{GG}$) and the characteristic velocity($C^{\ast}$). The average gas temperature at the exit of gas generator is found to be a function of propellant O/F ratio. For the gas generator having residence time of 4msec or more, the effect of flame residence time and combustion chamber pressure on the exit temperature is not significant. The exit characteristic velocity is found to be linearly proportional to the gas temperature at the exit of gas generator.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.12-12
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• 2000

우리 나라는 일본상공의 비행을 피하기 위하여 제주도와 남해안 근해로 발사장 선정이 국한되는 지정학적인 위치로 볼 때 발사장 선택에 제한이 없는 공중발사에 대한 가능성 연구가 필요한 시점에 있다. 본 연구는 우리 나라와 같은 분단 된 특수상황 그리고 지정학적 위치에서의 발사장을 고려한 우주 발사체 개발의 필요성에 따라 F-4에 장착 가능한 3단형 공중발사 로켓을 설계하고 1/3의 축소 모형을 제작하였다. 2kg의 payload를 갖는 발사체의 1단은 LRM ( Lox/kerosene )을 사용하였고 2, 3단은 SRM ( HTPB/AP/Al )을 사용하였으며 발사고도는 11-l2km 상공에서 F-4에 의해서 발사되고 31km지점에서 1단 분리가 이루어지며 62km지점에서 2단 분리와 nose fairing을 분리하게 된다. 전장은 6.85m 이며 전체 무게는 560.6kg 이고 전체 발사체 시스템의 CAD 도면은 아래 그림 1과 같이 주어져 있다. 그림 2에서는 F-4E phantom의 장착성을 검토해 본 결과 장착이 가능함을 알 수 있었으며 추진제 양의 감소로 크기를 대폭 줄일 수 있었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.203-207
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the thrust chamber of liquid rocket using LOx and Kerosene as propellant. The heat fluxes were obtained from the measured wall temperature to the axial direction of thrust chamber for different type of coolant, the various O/F ratio, mass flow rate and the location of the film cooling injector. A thin wall combustion chamber and nozzle were used to obtain the heat flux.