• Title/Summary/Keyword: Liquid Rocket Engine Thrust Chamber

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The Combustion Characteristics of a Subscale Engine of KSRIII(I) (KSR-III 엔진 축소형 모델 연소 특성(I))

  • Kim, Young-Han;Kim, Yong-Wook;Ko, Young-Sung;Lee, Soo-Yong;Ryu, Chul-Song;Seol, Woo-Seok
    • Proceedings of the KSME Conference
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    • 2001.06d
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    • pp.846-851
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    • 2001
  • For the successful development of the main engine of KSR(Korea Sounding Rocket)-III, Korea Aerospace Research Institute(KARI) carried out the experimental study on the subscale model engines. Several types of engines were tested on the Small Liquid Rocket Engine Test Facility. One of the typical test results of a Sub. engine(Sub. Mod.3) is presented here. It uses the Jet A-1 as fuel, liquid oxygen as oxidizer, and Tri-Ethyl Aluminium(TEA1) as ignition agent. The gas pressure feed system is adopted as a feeding mechanism and the design chamber pressure is 200psia. The physical phenomena are described in three regimes(ignition, transient, and steady state) with the pressure, thrust, flowrate and image data. And the pressure oscillation is analyzed in Fourier domain (<500Hz). Then we conclude that in this experiment, the engine shows the characteristic low frequency of 80Hz and it is stable for that frequency of pressure oscillation.

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Ignition Experiments of a High Pressure Liquid Propellant Thrust Chamber (실물형 연소기의 점화시험)

  • Moon Ilyoon;Kim SeungHan;Kim Jonggyu;Lim Byoungjik;Lee Kwangjin;Kim Intae
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • v.y2005m4
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    • pp.265-268
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    • 2005
  • A series of ignition tests had been conducted for a thrust chamber propelled by Jet A-1 and liquid oxygen with a chamber pressure of 52.5 bara and a thrust of 30 tonf. The chamber ignited by a hypergolic fluid, TEAL, keeps its first constant pressure low at $63\%$ of the design value by $61\%$ of a liquid oxygen mass flow rate and $67\%$ of fuel for 0.25 sec. The operating O/F ratio of the chamber was kept at lower values than that of the design operating condition throughout the whole ignition procedure. Surge of the chamber pressure is below $6\%$ of the design value.

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Study on the Performance of Small Size Liquid Rocket Engine (축소형 엔진의 성능에 관한 연구)

  • Namkoung Hyuck-Joon;Han Poong-Gyoo;Kim Dong-Hwan;Lee Kyoung-Hoon;Kim Young-Soo
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • v.y2005m4
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    • pp.139-143
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    • 2005
  • Combustion Test Facility for Liquid Rocket Engine using kerosene and liquid oxygen has been developed for the purpose of cooling and performance study. Test engine of thrust 0.5 KN(design thrust) is tested to confirm the normal operation. Therefore, water-cooled firing tests using kerosene engine with injectors of fuel-centered coaxial type are conducted. With the viewpoint of characteristic velocity, and specific impulse at sea level, and chamber pressure on OF mixture ratio are analyzed.

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Experimental Study on the Combustion Stability of Full Scale Rocket Combustor (실물형 액체로켓 연소기의 연소안정성에 대한 시험적인 고찰)

  • Lee Kwang-Jin;Seo Seong-Hyeon;Kang Dong-Hyeuk;Song Ju-Young;Lim Byoung-Jik;Han Yeoung-Min
    • 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.

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Experimental Study of Film Cooling in Liquid Rocket Engine(I) (액체로켓엔진의 막냉각에 관한 실험적 연구(I))

  • Choi, Young-Hwan;Jeong, Hae-Seung;Kim, Yoo;Kim, Sun-Jin
    • 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.

Experimental Investigation of a Regression rate On Hybrid Rocket Engine

  • Park, J. W.;S. Krishnan;Lee, C. W.;M. W. Yoon
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2004.03a
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    • pp.524-527
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    • 2004
  • Hybrid rocket had many advantage with compared to solid and liquid rockets. However, the engines have not yet been used in practical rocket systems, due mainly to the disadvantage of hybrid combustion, such as low fuel regression rate. In this study, lab-scale hybrid motor was designed and manufactured. And the methods of regression rate improvement were considered. Test firings with thrusts up to 300 N were conducted with GOX and transparent PMMA. Thrust was calculated with the pressure of the combustion chamber and the regression rate was measured in with variation of oxidizer flow rate. The regression rates showed a strong dependency on GOX mass flux. The frequency analysis technique of the bulk-mode oscillation of motor was applied to a hybrid rocket motor and was based on the principle that this frequency was inversely proportional to the square root of the chamber volume. Several problems and solutions of operating hybrid rocket were presented.

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Predictions of Fatigue Life of Copper Alloy for Regenerative Cooling Channel of Thrust Chamber (연소기 재생냉각 채널용 구리합금의 피로수명예측)

  • Lee, Keumoh;Ryu, Chulsung;Heo, Seongchan;Choi, Hwanseok
    • Journal of the Korean Society of Propulsion Engineers
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    • v.21 no.6
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    • pp.73-82
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    • 2017
  • Low-cycle thermal fatigue problem resulting from multiple use of a liquid rocket engine has to be considered for the development of a reusable launch vehicle. In this study, life prediction equations suggested by previous researchers were compared as applied to various copper alloy cases to predict fatigue lives from tensile test data. The present study has revealed that among the presently considered life prediction methods, universal slopes method provides the best life prediction result for the copper alloys, and the modified Mitchell's method provides the best life prediction result for oxygen free high conductivity (OFHC) copper.

Optimal Design and Test of Fuel-Rich Gas Generator

  • Lee, Changjin;Kwon, Sun-Tak
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2004.03a
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    • pp.560-564
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    • 2004
  • The optimal design and combustion analysis of the gas generator for Liquid Rocket Engine (LRE) were performed. A fuel-rich gas generator in open cycle turbopump system was designed for 10ton$_{f}$ in thrust with RP-1/Lox propellant. The optimal design was done for maximizing specific impulse of main combustion chamber with constraints of combustion temperature and power matching required by turbopump system. Design variables were selected as total mass flow rate to gas generator, O/F ratio in gas generator, turbine injection angle, partial admission ratio, and turbine rotational speed. Results of optimal design show the dimension of length, diameter, and contraction ratio of gas generator. Also, the combustion test was conducted to evaluate the performance of injector and combustion chamber. And the effect of the turbulence ring was investigated on the mixing enhancement in the chamber.r.

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Preliminary Design Plan for Determining Combustor Configuration of Regenerative-cooled Liquid Rocket Engine (재생냉각식 액체로켓엔진의 연소기 형상 결정을 위한 예비 설계 방안)

  • Son, Min;Seo, Min-Kyo;Koo, Ja-Ye;Cho, Won-Kook;Seol, Woo-Seok
    • Journal of the Korean Society of Propulsion Engineers
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    • v.15 no.1
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    • pp.83-89
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    • 2011
  • A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted by one-dimensional and experimental correlations. Finally, determinable plan for the contour of combustor were presented through Rao nozzle design method.

Preliminary Design Plan for Determining Combustor Configuration of Regenerative-cooled Liquid Rocket Engine (재생냉각식 액체로켓엔진의 연소기 형상 결정을 위한 예비 설계 방안)

  • Son, Min;Seo, Min-Kyo;Koo, Ja-Ye;Cho, Won-Kook;Seol, Woo-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2010.11a
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    • pp.37-42
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    • 2010
  • A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted using one-dimensional and experimental equations. Finally, determinable plan for contour of combustor were presented through Rao nozzle design method.

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