• Title/Summary/Keyword: Rocket combustor

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Flow Analysis of the Oxidizer Manifold for a Liquid Rocket Combustor using OpenFOAM (OpenFOAM을 이용한 액체 로켓 연소기의 산화제 매니폴드 내 유동 해석)

  • Joh, Mi-Ok;Han, Sang-Hoon;Kim, Seong-Ku;Choi, Hwan-Seok
    • Journal of the Korean Society for Aeronautical & Space Sciences
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    • v.40 no.9
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    • pp.781-788
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    • 2012
  • Flow in the oxidizer manifold of a liquid rocket combustor has been analysed using an open source CFD toolbox, OpenFOAM. The applicability of OpenFOAM to the problems with complex geometries involving porous media zones for simulating the pressure drop induced by the injectors has been evaluated by performing turbulent, incompressible steady-state flow analysis. The usefulness and applicable area of the OpenFOAM as a design evaluation and analysis tool will be confirmed and enlarged by further evaluation with various computational cases representing major physical phenomena in rocket combustion devices.

System Analysis of a Gas Generator Cycle Rocket Engine

  • Cho, Won Kook;Kim, Chun IL
    • International Journal of Aerospace System Engineering
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    • v.6 no.2
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    • pp.11-16
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    • 2019
  • A system analysis program has been developed for a gas generator cycle liquid rocket engine of 30 ton class. Numerical models have been proposed for a combustor, a turbopump, a gas generator and pressure drop through a regenerative cooling system. Numerical algorithm has been validated by comparing with the published data of MC-1. The major source of error is not the numerical algorithm but the imperfect performance models of subsystems. So the precision of the program can be improved by revising the performance models using experimental data. The sea level specific impulse and vacuum specific impulse have been demonstrated for a 30 ton class gas generator engine. The optimal condition of combustor pressure and mixture ratio for specific impulse which is a typical characteristic of a gas generator cycle engine has been illustrated.

Rocket Engine Test Facility Improvement for Hot Firing Test of 75 ton-f Class Gas Generator and Cold Flow Test (75톤급 가스발생기 연소시험을 위한 시험장 개선 및 수류시험)

  • Kang, Dong-Hyuk;Lim, Byoung-Jik;Ahn, Kyu-Bok;Seo, Seong-Hyeon;Han, Yeoung-Min;Choi, Hwan-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2009.11a
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    • pp.29-33
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    • 2009
  • On the basis of the development experience of a gas generator for the 30 ton-f thrust liquid rocket engine combustor a Subscale Ground Firing Test Facility was designed and fabricated for a gas generator for the 75 ton-f thrust liquid rocket engine combustor. The Subscale Ground Firing Test Facility developed is going to be used to develop 75 ton-f class gas generator. Acquired data and test technique from this facility will be used to develope the high performance liquid rocket engine combustor and the Ground Firing Test Facility. This report describes the improved Subscale Ground Firing Test Facility for 75 ton-f class gas generator and results of the cold flow test.

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Combustion Stability Evaluation of 30 ton-f Class Liquid Rocket Engine Combustor (30톤급 엑체로켓엔진 연소기의 연소안정성 평가)

  • Lim, Byoung-Jik;Lee, Kwang-Jin;Kim, Mun-Ki;Kang, Dong-Hyuk;Yang, Seung-Ho;Seo, Seong-Hyeon;Han, Yeoung-Min;Choi, Hwan-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2008.05a
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    • pp.163-167
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    • 2008
  • This paper presents pressure fluctuation characteristics of a 30 ton-f class liquid rocket engine combustor. Combustion stability of the combustor was evaluated using the results 46 firing tests performed with a varying O/F ratio and chamber pressure. The RMS value of pressure fluctuation during the steady state combustion was less than 2.6% of the static chamber pressure, demonstrating static stability of the combustion phenomenon. The decay time of pressure fluctuation caused by forced disturbance of a pulse gun was found to be less than 3.5 msec verifying dynamic stability of the combustor.

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Preliminary Design of Test Facility for 75-tonf-Class Liquid Rocket Engine Combustor (75톤급 액체로켓엔진 연소기 시험설비 기본설계)

  • Lim, Byoung-Jik;Seo, Seong-Hyeon;Kim, Mun-Ki;Kang, Dong-Hyuk;Han, Yeong-Min;Choi, Hwan-Seok
    • Journal of the Korean Society of Propulsion Engineers
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    • v.14 no.5
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    • pp.84-91
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    • 2010
  • For the successful development of 75-tonf-class liquid rocket engine, a plenty of tests on each engine component have to be performed and this is equally true for a combustor. However the test facility which is in operation at Korea Aerospace Research Institute lacks its capacity to perform fire tests of a 75 tonf class combustor at its nominal thrust. Since the test facility has to be ready prior to the start of development tests, it is very urgent to establish the test facility. The preliminary design of a test facility for a 75 tonf class combustor which was performed according to such a necessity is described in the paper.

Combustion Performance Tests of High Pressure Subscale Liquid Rocket Combustors (고압 축소형 연소기의 연소 성능 시험)

  • Kim, Jong-Gyu;Lee, Kwang-Jin;Seo, Seong-Hyeon;Lim, Byoung-Jik;Ahn, Kyu-Bok;Han, Yeoung-Min;Choi, Hwan-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2007.04a
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    • pp.128-134
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    • 2007
  • Combustion performance and characteristics of high-pressure subscale liquid rocket combustors were studied experimentally. Four different models of combustor were considered in this paper. The high-pressure subscale combustor is composed of the mixing head, the water cooling cylinder and the nozzle. One model of the combustors employed regenerative cooling combustor in that the kerosene used for the chamber cooling is burned. This combustor was damaged due to a high frequency combustion instability occurred during a firing test. The results of the firing tests, comparison of performance, and characteristics of static and dynamic pressures of the combustors are described.

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Comparison Study on System Design Parameters of Gas Generator Cycle Liquid Rocket Engine (가스발생기 사이클 액체로켓엔진의 시스템 설계 인자 비교)

  • Nam Chang-Ho;Park Soon-Young;Moon Yoon-Wan
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2005.11a
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    • pp.220-223
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    • 2005
  • System design parameters of gas generator cycle liquid rocket engines were investigated and compared in the present study. Characteristic velocity of combustor, pressure drop of combustor injector, exit pressure of pump, pump efficiency and specific power of turbine were considered as a system design parameter. The result shows the characteristic velocity is in the range of 1700-1770 m/s, pressure drop of combustor injector, 4-10 bar, pump exit pressure ratio to combustion pressure, 120-230%, pump efficiency, 60-80%, specific power of turbine, $0.28-0.58MW{\cdot}s/kg$.

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Characteristics of Unielement Injector Combustion with Flow rates and Chamber Pressures (유량 및 연소압에 따른 액체로켓 단위분사기 연소특성 변화)

  • Moon Il-Yoon;Kim Jong-Gyu;Han Yeoung-Min;Yoo Jin;Lee Yang-Seok;Ko Young-Sung
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2005.11a
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    • pp.247-250
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    • 2005
  • In the case of appling a unielement injector developed for a full scale liquid rocket combustor, a operating condition or configuration of the injector is changed by combustion pressure, arrangement and injector quantity of a full scale liquid rocket combustor. In order to verify application, swirl coaxial injectors propelled by jet-A1 and liquid oxygen are tested at different conditions of a combustion pressure, a flowrate and an injector length. As a test result, the application of the present swirl coaxial injectors is excellent because an efficiency of a characteristic velocity is increased at the each test condition beyond that variation of dynamic pressure intensity is small.

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Film Cooling Modeling for Combustion and Heat Transfer within a Regeneratively Cooled Rocket Combustor (막냉각 모델을 이용한 재생냉각 연소기 성능/냉각 해석)

  • Kim, Seong-Ku;Joh, Mi-Ok;Choi, Hwan-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2011.11a
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    • pp.636-640
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    • 2011
  • Film cooling technique has been applied to effectively reduce thermal load on liquid rocket combustion chambers by direct injection of a portion of propellant, which flows through the regeneratively cooling channels, into the chamber wall. This study developed a comprehensive model to quantitatively predict the effects of kerosene film cooling on propulsive performance and wall cooling at supercritical pressure conditions, and assessed the predictive capability against hot-firing tests of an actual combustor. The present model is expected to be utilized as a design and analysis tool to meet the conflicting requirements in terms of performance, cooling, pressure loss and weight.

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Thermal Barrier Efficiency and Endurance of Ni-Cr Coating in Liquid Rocket Engine Combustor (액체로켓엔진 연소기에 적용된 니켈-크롬 코팅의 열차폐 효율과 내구성)

  • Lee, Kwang-Jin;Lim, Byoung-Jik;Kim, Jong-Gyu;Han, Yeoung-Min;Choi, Hwan-Seok
    • Aerospace Engineering and Technology
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    • v.8 no.1
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    • pp.138-143
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    • 2009
  • Thermal barrier efficiency and endurance of coatings in liquid rocket engine combustor were evaluated for air plasma spray coating and electro/electroless plating. The result of firing tests has revealed occasional occurrence of local delamination of $ZrO_2$, NiCrAlY coating obtained by the method of air plasma spray in the region of supersonic flow and it necessitated a new coating method as a substitution. It was found that Ni-Cr coating by means of electro/electroless plating can substitute $ZrO_2$, NiCrAlY coatings of air plasma spray in terms of thermal barrier efficiency and endurance.

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