• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.45-54
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• 2011

A program aiming at predicting dynamic characteristics of a Liquid Rocket Engine(LRE) was developed and examined to trace entire LRE operation. In the startup period, transient characteristics of the propellant flows were predicted and validated with hydraulic tests data. An arrangement of each component for the pipelines was based on an operating circuit of open cycle LRE. The flow rate ratio for the gas generator and the main chamber was determined to mimic that of real open cycle LRE. Individual component modeling at its transient was completed and was integrated into the system prediction program. Essential parameters of the component dynamic characteristics were examined in an integrated fashion.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.91-99
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• 2020

A liquid rocket engine goes through many tests to prove its performance before liftoff. It means the tests for setting ignition and start-up conditions or a test on design condition, which verifies the design performance. However, the development process requires verification of performance under off-design conditions through tests involving different operating conditions, which affects the duration of engine development. The off-design performance test is performed by altering the conditions of the propellant supplied to the engine in conjunction with the engine performance test that varies the opening of the control valves in the engine. This paper is based on the results of the engine tests performed at the KSLV-II engine test facilities in the Naro Space Center and describes the operations of the test facility for off-design condition test that changes the inlet conditions of the turbo-pump due to changes in the pressure and temperature of the propellant supplied to the test engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.34-37
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• 2010

One dimensional transient analysis was studied for turbopump-fed liquid propellant rocket engine(LRE) system in starting using AMESim. The effects of timing of gas generator fuel valve opening and gas generator ignition to start-up stability were researched for open cycle type system using LOX/RP-1 to propellants. Result show that the parameters and sequence on start-up should be considered to design optimized turbopump-fed LRE system.

• Aerospace Engineering and Technology
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• v.9 no.1
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• pp.93-97
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• 2010

The analysis of specific impulse of the liquid rocket engine employing gas generator cycle with LOx/kerosene as propellant has been performed. The relative error of performance of 300 ton level engine is 0.1%s for specific impulse and 12% for optimal combustion pressure comparing with the published data. The difference of the performance model and the material property models of gas generator product gas are the presumed major reason of discrepancy. The optimal condition of 30 ton level engine is combustion pressure of 68 bar and mixture ratio of 2.2 for maximum specific impulse. This optimal condition can be changed by performance models.

• 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.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.2
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• pp.54-61
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• 2003

This paper describes a general design procedure of regenerative cooling system for liquid rocket engine(LRE). From this design logic, cooling channels are designed and fabricated. The measured heat flux from firing test is similar to the heat flux predicted by design logic. Therefore, proposed design procedure of cooling channel can be applied to real LRE system. Also the result of firing test indicates that soot from combustion products have strong influence on the cooling characteristics of LRE.

• 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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• 2007.11a
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• pp.174-177
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• 2007

In this work, we researched static and dynamic characteristics concern to hydraulic regulator control process and parameter setting, which is used on liquid rocket engine regulation. The hydraulic regulator of 8K14 "SCUD" 9D21 engine is analyzed and on the basis of the developed mathematical modeling the analysis of response time and certification on automatic control accuracy is carried out. In this process, we find out specific design configuration of needle valve flow section that effects on pressure regulation performance.

• 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.