• Bulletin of the Korean Space Science Society
• /
• 2003.10a
• /
• pp.70-70
• /
• 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
• /
• 2010.11a
• /
• pp.676-677
• /
• 2010

In Naro Space Center, Naro launch vehicle launched 2 times. Launch pad for Naro launch vehicle in Naro space center equipped propellant feeding facility for operating launch process. This paper studied development process and operating method for liquid oxygen filling system of cryogenic propellant systems in launch pad propellant feeding facility.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.10a
• /
• pp.79-82
• /
• 2004

Propellant feeding system is the system to satisfy propellant feeding requirements(mass flow rate, pressure, temperature) at engine inlet of launch vehicle. Propellant feeding test facility is being constructed for the development scheme of pressurization system, processing in tank, propellant piping system, and flow control system that are main technologies in order to develope propellant feeding system. This paper introduces the propellant feeding test facility being constructed in KARI.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.10 no.4
• /
• pp.1-10
• /
• 2006

In this paper, a calculation method of the thermodynamic parameters of cryogenic propellant is proposed when a cryogenic propellant tank is pressurized by gaseous helium(GHe) bubbling. Temperature of cryogenic propellant and mass of dissolved GHe into propellant were analyzed at the various operation of pressurization of tile liquid oxygen(LOX) and hydrogen($LH_2$) tank using helium bubbling. It was evaluated how the GHe bubbling influences to the thermodynamic parameters of LOX and $LH_2$ with results of the analysis. With the proposed calculation method, It will be able to confirm the feasibility of GHe bubbling as a pressurization system of cryogenic propellant tank and to optimize the pressurization system using GHe bubbling.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 1999.10a
• /
• pp.6-6
• /
• 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 for Aeronautical & Space Sciences
• /
• v.31 no.3
• /
• pp.58-64
• /
• 2003

Pogo is the instability resulting from the interaction between rocket structure and propulsion system of liquid propellant rocket. The coupling of structure and propulsion system can lead to severe problem in rocket. For the analysis of pogo, a time-invariant linearized mathematical model is developed for a selected flight time. Propulsion system is modeled using element representations for each components. Rocket structure is modeled using FEM. Form the results of modal analysis of structure, the behavior of structure can be represented. System equations for coupling structure and propulsion system are composed. The stability in obtained by the eigen solution of system matrix. The optimization of the design variables such as size, place of accumulator for suppressing pogo instability in carried out. This article of study can be used to determine the degree of stability, and guide the design of pogo suppression system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.7-10
• /
• 2006

In order to inject satellites into a target orbit, launch vehicles should have a precise attitude and control system capable of controlling three axises of pitch, yaw and roll. For launch vehicles, there are two types of attitude control system currently in popular use; the first one is a cold gas method, and the other is a liquid propulsion system using a single and dual property propellant. The purpose of this paper is to analyze the characteristics of thrust control system using said propellant, thereby providing for a rationale for its application to the upper stages of launch vehicles, in terms of the simplicity of the system, economics of structure and operation.

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

• Aerospace Engineering and Technology
• /
• v.3 no.1
• /
• pp.232-241
• /
• 2004

KARI(Korea Aerospace Research Institute) is achieving the KSLV program according to National Space Technology Development Program. In this paper, the authors are intend to introduce the Integrated Power Plant(IPP) test facility which will be constructed for the variety of tests on KSLV program. IPP test facility refers to comprehensive testing equipment for liquid rocket launch vehicle. Using this facility, KARI can verify the adaptedness of parts and subsystems for launch vehicle and finally can qualify the system characteristics of launch vehicle doing kinds of test including hot firing test. IPP test facility will make it possible to simulate the vehicle launching circumstances and to predict the performance of launch vehicle during its flight test.

• Bulletin of the Korean Space Science Society
• /
• 2003.10a
• /
• pp.92-92
• /
• 2003

대형 위성 발사체를 우주로 발사하기 위해서는 복잡한 추진기관시스템을 정밀하게 제어해야 하며, 이를 위해서는 로켓의 궤적에 따른 추진제 질량과 추력을 적절하게 제어해야 한다. 정확하게 계산된 비행궤도를 따라 로켓을 최종 목표 지점까지 올리는 일은 엔진의 추력과 공연비를 동시에 조절하는 엔진제어기술을 이용하여 가능하게 된다. 추력제어는 엔진시스템에 대한 정확한 이해와 이를 바탕으로 한 추진제 유량 제어를 통해 가능하기 때문에 액체로켓 엔진에 대한 엔진시스템 분석과 해석이 선행되어야 한다. 본 연구에서는 향후 연구 대상이 될 엔진시스템의 구성과 추력 및 공연비 제어시스템의 기본 제어 방법을 소개하고자 한다.