• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.239-242
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• 2004

In this paper, the general mathematical model of LRE(Liquid Rocket Engine) is presented. For the analysis about the trend of dynamics and the stability of open type LRE, it is transformed to linear model by Laplace transform and synthesized to the linear complex model of LRE with Matlab/Simulink.

• 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.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.2
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• pp.52-60
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• 2020

This paper deals with an optimal output tracking control for open-cycle liquid propellant rocket engine. For this purpose, we modeled simplified mathematical model of open-cycle liquid propellant rocket engine and designed optimal output feedback control system using combustion chamber pressure. For design the closed-loop system of open-cycle liquid propellant rocket engine, we designed optimal output feedback linear quadratic tracking control system using the linearized model and demonstrated the performance of the controller through numerical simulation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.240-246
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• 2005

A series of combustion tests of a 30-tonf-class full scale liquid rocket thrust chamber under development has been carried out to verify its design. The test results revealed decent performance in the aspects of efficiency. The combustion stability is one of the most important parameters of liquid rocket engine in addition to the efficiency. Assessment tests of combustion stability must be accomplished to confirm the possibility of combustion instability due to spontaneous or external disturbances. The combustion stability rating tests of the full scale thrust chamber with temporary baffles made of stainless steel were carried out utilizing pulse guns to estimate combustion stability characteristics. The tests results show highly stable combustion stability characteristics. The outcome acquired from the present experimental study will be used to design an actively cooled baffle that can survive for the life time operation of the thrust chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.120-130
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• 2013

A study was conducted for the performance characteristics of methane taking recently the limelight in the world as a next-generation propellant, with the survey for state of the art in the development of methane/oxygen rocket engine being accompanied. Liquid methane as a rocket fuel has the favorable characteristics such as non-toxic, low cost, regenerative cooling capability, and potential for in-situ resource utilization (ISRU). The combination of liquid methane and liquid oxygen also provides the excellent performance including high specific impulse and low system mass. For these reasons, many researches have been actively carried out on the methane/oxygen engine, nevertheless, its technology readiness level is not that high enough just yet. Therefore, it is judged that it is the time to mitigate the technical gap with the space technology of advanced countries through a swift onset of the development of methane rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.46-51
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• 2012

This study aims to develop the performance analysis program on the staged combustion cycle of the liquid rocket engine using liquid oxygen(LOx) as oxidizer, liquid hydrogen(LH2) and RP-1 as fuel. The developed analysis program can obtain the propellant mass flow rate, the specific impulse, and representative design values of engine components for the required thrust satisfying pressure, mass flow, and energy balance conditions. The analysis results show that the the specific impulses (Isp) compared to those of the real engines have been less than 1%. With additional constraints, the program will be improved for the system optimization.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.6
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• pp.104-110
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• 2018

Operating mode analysis of a liquid propellant rocket engine(LRE) is a crucial tool through the development of an engine. The operating mode analysis of an engine based on a collection of the acceptance tests of components shows discrepancies when compared to the test results. We propose a correction method for performance parameters to develop an engine analysis model for the gas generator cycle of an LRE. In order to simulate engine behavior, the performance parameters for the analysis model are tuned based on the test results of the 75tf engine of KSLV-II.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.213-219
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• 2005

Small liquid rocket engine (SLRE) with calorimeter were developed and tested to evaluate cooling characteristics in the liquid rocket engine. Therefore, cooling performance analysis was performed to predict the heat transfer coefficient on gas side wall in 10 calorimeter channel. A heat transfer empirical formula was determined by results of firing test and computational simulation.

• Aerospace Engineering and Technology
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• v.7 no.2
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• pp.157-161
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• 2008

This study introduce a newly developed program to calculate the combustion process of combustion chamber and gas generator of liquid rocket engine by use of Gibbs free energy minimization method based on chemical equilibrium. The simulation results of the new program and CEA code of NASA were compared and showed good agreement, thus proving the validity of the newly developed in-house program for combustion analysis.