• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.111-117
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• 2004

An analytical study was carried out to evaluate the regenerative cooling characteristics in the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel. As a supplementary cooling method, a radiative cooling was applied to the nozzle extension. It was found out from this work that the cooling system with the regenerative and radiation cooling only is not adaptable for the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel for the $2^{nd}$ stage of the space launch vehicle, with the viewpoint of the thermal and thermo-structural instability and the excessive pressure drop in the cooling channel.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.375-383
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• 2005

The present paper describes the first development of a LOX/kerosene type turbopump in Korea. The liquid rocket engine, that the turbopump can be applied to, has a 30-ton(metric) level of vacuum thrust and employs a gas generator cycle. The turbopump consists of two single-stage centrifugal pumps, that is, LOX and kerosene pumps, and one single-stage impulse turbine. Inter-propellant seal(IPS) is located between the LOX pump and the kerosene pump to avoid any interaction between the propellants. A series of component and TPU(Turbopump Unit) test has been completed in the level of simulant propellants and ready for hot firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.207-210
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• 2006

Injector is one of the most important elements in Liquid rocket Engine design, and how to arrange these injectors on the head determines the engine performance. In this study, when the swirl injectors are used for the 1st designing of injector head of 170 tonf UDMH-LOX as the propellant of LRE, a distribution relation of the mass flow rate per unit area was calculated from the function of ${\Phi}$, which is related with the mass flow rate characteristics of swirl injector. And the combustion characteristics by circumferential axis were estimated using this relation under the consideration of combustion core and film cooling area.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.150-155
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• 2010

Liquid rocket injectors play crucial roles on propulsive performance, combustion stability, and heat transfer characteristics. Nevertheless, their developments have mainly relied on empirical methods and expensive hot-firing tests due to lack of fundamental understanding of high pressure combustion phenomena in the near-injector regions. The present study was motivated by recent efforts to develop reliable modeling of liquid rocket combustion. The turbulent combustion model based on the flamelet concept has been extended to take into account real-fluid behaviors occurred at supercritical pressures, and validated against measurements for a cryogenic nitrogen injection, a non-premixed turbulent jet flame at atmospheric pressure, and a LOx/$GH_2$ coaxial shear injector at a supercritical pressure.

• Journal of the Korean Chemical Society
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• v.11 no.1
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• pp.22-27
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• 1967

The thermodynamic properties of the liquid oxygen are calculated by applying "The Transient State Theory of Significant Liquid Structure". The results show good agreement with experimental observations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.4
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• pp.333-342
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• 2016

Liquid rocket propulsion is a system that produces required thrust for satellites and space launch vehicles by using chemical reactions of a liquid fuel and a liquid oxidizer. Monomethylhydrazine/dinitrogen tetroxide, liquid hydrogen/liquid oxygen and RP-1/liquid oxygen are typical combinations of liquid propellants commonly used for the liquid rocket propulsion system. The objective of the present study is to investigate useful design and performance data of liquid rocket engine by conducting a numerical analysis of thermochemical reactions of liquid rocket propellants. For this, final products and chemical compositions of three liquid propellant combinations are calculated using equilibrium constants of major elementary equilibrium reactions when reactants remain in chemical equilibrium state after combustion process. In addition, flame temperature and specific impulse are estimated.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.3
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• pp.75-86
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• 2004

KSR-III is the first Korean sounding rocket propelled by a liquid propellant propulsion system and it has been developed over 5 years using purely domestic technologies. The propulsion system of KSR-III is a 13-ton class see-level thrust liquid rocket engine(LRE) which utilizes liquid oxygen and kerosene for its propellants and employed pressurized propellant feeding and ablative cooling system. The problem of combustion instabilities which has brought the most difficulty in the development was resolved by implementation of a baffle. Through the development of KSR-III LRE, meaningful achievements have been made in the core technologies of LRE such as design of injectors and combustion chambers and test, evaluation, and control of combustion instabilities. The acquired technologies will be applied to the development of higher performance LREs necessary for future space development programs such as Korean Small Launch Vehicles(KSLV) In this paper, the development of KRE-III LRE system is described including its design, analyses. performance tests and evaluation.

• Journal of Power System Engineering
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• v.12 no.3
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• pp.5-11
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• 2008

산업공정에서 널리 사용하는 반응기는 성질이 상이한 물질을 혼합하는 시스템으로서 본 연구에서는 이젝터(ejector)에 의한 반응기의 개발을 수행하였다. 액체-가스 이젝터는 구동유체에 의하여 기체가 흡입되면서 각종 유해가스를 제거하는 목적 또는 기체와 액체의 혼합 등 목적으로 사용된다. 본 실험에서 액체구동 가스혼합반응기의 실험 장치를 구축하고 이젝터 내부의 유동패턴과 기체용해도 자료를 도출하며 고효율 이젝터 설계를 위한 진공도 측정과 디퓨저 각도가 다른 이젝터의 실험 및 수치해석을 수행하였다. 이젝터의 성능은 흡입 측에서의 진공압력으로 평가되며 이 진공압력은 이젝터의 노즐 설계 및 유동조건에 의하여 결정되므로 이에 대한 기본적인 특성 도출이 선결되어야 한다. 순환유체의 유량이 70LPM, 80LPM 90LPM조건에서 두 가지 디퓨저에 대하여 비교실험을 수행하였다. 실험적 연구와 수치해석연구를 통하여 혼합성능과 이젝터의 내부유동특성에 대하여 고찰한 결과 디퓨저의 각도가 5.0도일 때 진공도가 더욱 높으며 구동액체의 유량이 작을 때는 진공도차이가 크지만 유량이 증가함에 따라 진공도 차이가 감소된다. 구동액체의 유량이 증가할수록 용존산소농도는 증가하며 디퓨저의 각도가 5.0도일 때는 용존산소 농도가 더 높게 나타나는 것을 확인할 수 있었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.191-195
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• 2005

Propellant remains as outage at engine shutdown contributes no useful impulse to the rocket and produces an unwanted increase in burnout weight. Minimization of outage, is therfore is a basic consideration in attaining the maximum performance capability of my bipropellant liquid rocket. This paper present the calculation procedures of outage and optimum loading propellant mass. And some control methods and measurement techniques for outage are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.781-783
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• 2011

The oxidizer filling system, ground facility of the launch complex, should accept difficult requirements from the launcher sufficiently. The launcher do not have unnecessary insulators for mass reduction and manages liquid oxygen mass fastidiously to satisfy the mission requirement. So, the ground facility should be able to accept its requirements, then we should make the operating process being adjusted. In this paper, the operating process design and verification results on the oxidizer filling ground facility for liquid rocket is demonstrated.