• 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.

• 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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• 2012.05a
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• pp.436-442
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• 2012

In the process of liquid rocket engine design, obtaining method of the dynamic characteristics of engine should be emphasized typically to determine the control logic and algorithms of the throttle valves in the propellant feed pipeline. However, determining the dynamic characteristics of an engine through the autonomous test is very hard and laborious, so that the numerical approach is prevailing. In this study, using the previously developed dynamic analysis model of the engine around the steady state, we introduced a disturbance to this model, and obtained the dynamic response in the time domain. And by applying the well-known Levy method to this temporal response, we could deduce the transfer function of that system that can give us various information of engine and can be manipulated to design the control system.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.93-100
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• 2020

The turbopump of the liquid rocket engine adapts an inducer to minimize the cavitation due to the variations of the propellants supply condition. However, the inducer introduces cavitation instabilities which are well-known problems in the engine development. In this paper, operational characteristics by the cavitation instabilities are analyzed and the reliability of the engine is checked when the first stage engine of the KSLV-II is tested at the low inlet pressure conditions. The characteristic frequencies representing the cavitation instabilities of the LOx pump are clearly found in various high frequency sensor signals around the entire engine in addition to the LOx and fuel pump.

• 한국연소학회:학술대회논문집
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• 2014.11a
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• pp.223-224
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• 2014

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.4
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• pp.334-340
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• 2015

Engine derating is often considered for reliability benefits because lower power operation reduces its failure probability. To be derated during operation, however, the engine must be initially overdesigned. The engine overdesign is cost effective only if reliability increased from derating is enough to offset the initial increase in the development cost caused from the overdesign. The purpose of this paper is to provide an analytical model to consider a trade-off between the engine overdesign and derating. We use a logistic regression model to explain reliability growth in the number of hot firing tests for a fixed power level. Using the Transcost model with the reliability growth model, we show that 10% overdesign of Titan rocket engine decreases its development cost by about 9% and 23% depending on the reliability requirement. We also point out that such a cost reduction depends on the fuel type a rocket uses.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.1
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• pp.56-64
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• 2010

A system design has been conducted of the liquid rocket engine for Korean launch vehicle (KSLV-II, Korea Space Launch Vehicle II). The present turbopump-fed liquid rocket engine of vacuum thrust 76 ton and vacuum specific impulse 297 sec adopts gas generator cycle. The combustion pressure of the regeneratively cooled combustor is 60 bar. The propellant is LOx/kerosene. The engine is started by pyrostarter and the combustor is ignited by TEA (TriEthylAluminium). The engine system performance and the subsystems performance requirements are given through energy balance analysis. The combustion pressure, specific impulse and the engine mass are analyzed to be reasonable comparing with the published data. The startup analysis method which will be used in the future has been validated against the turbopump-gas generator coupled test. The tuning method for performance variation of the engine which is not actively controled has been prepared by mode analysis and performance deviation analysis.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.60-71
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• 2007

The startup characteristic of liquid propellant rocket engine should be focused on the stable ignition of combustion chamber and gas generator. Also, to lessen the propellants consumption during this period which doesn't contribute to the flight thrust, the engine has to be transferred to the nominal mode quickly. Because of the risk of test, it is impossible to develop all the startup cyclogram or the specifications of engine by test, so the precedent numerical approach is quite necessary. In this study we developed a mathematical model for the startup phenomena in a liquid rocket engine driven by gas generator-turbopump system based on the commercial 1-D flow system analysis program, Flowmaster. Using this program we proposed a methodology to obtain the specifications of turbine starter and the opening time of shutoff valves for the stable startup of the engine. To verify this methodology we qualitatively compared the analysis results to the typical startup curve of the published engine, then found it is quite well matched.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.6
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• pp.71-75
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the lab-scale dump-cooled liquid rocket engine using LOX and kerosene as propellants. The nozzle of the rocket engine was film cooled with water as coolant. A special film cooling adapter was fabricated to introduce the film-coolant into the thrust chamber. The flow rates of film coolant was approximately 15~19 percent of the total propellant. The nozzle heat flux was determined from the measured temperature rise and flow rate of the coolant(water). Large reductions in the nozzle heat flux was resulted when film cooling adapter located directly upstream of the nozzle.