• Title/Summary/Keyword: Hybrid Rocket Engine

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Flamelet Modeling for Combustion Processes of Hybrid Rocket Engine (화염편 모델을 이용한 하이브리드 로켓의 연소과정 해석)

  • Lim, Jae-Bum;Kang, Sung-Mo;Kim, Yong-Mo;Yoon, Myung-Won
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2006.11a
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    • pp.237-240
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    • 2006
  • Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. Accordingly, the recent research efforts are focused on the improvement of engine efficiency and regressionrate in the hybrid rocket engine. The present study has numerically investigated the combustion processes and the flame structure in the hybrid rocket engine. The turbulent combustion is represented by the flamelet model and Low Reynolds number $k-{\varepsilon}$turbulent model is employed to reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect. Numerical results suggest that the present approach is capable of realistically simulating the combustion characteristics of the hybrid rocket engines.

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Flamelet Modeling for Combustion Processes of Hybrid Rocket Engine (화염편 모델을 이용한 하이브리드 로켓의 연소과정 해석)

  • Lim, Jae-Bum;Kim, Yong-Mo;Yoon, Myung-Won
    • 유체기계공업학회:학술대회논문집
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    • 2006.08a
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    • pp.245-248
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    • 2006
  • Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. Accordingly, the recent research efforts are focused on the improvement of engine efficiency and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the flamelet model and Low Reynolds number $k-{\varepsilon}$ turbulent model is employed to reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect. Based on numerical results, the detailed discussions have been made for the effects of oxygen injection methods and oxygen injection flow rate on flame structure and regression rate in the vortex hybrid rocket engines

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하이브리드 로켓의 연소기술동향 분석

  • 김용모;윤명원;김윤곤
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2001.11a
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    • pp.55-60
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    • 2001
  • Recently, there have been many research efforts to improve the fuel regression rate in the hybrid rocket engines. In the present study, the ongoing research and development of the next general ion hybrid rocket engine are systematically reviewed. The detailed discussions have been made for the innovative combust ion technologies including the vortex hybrid rocket engines , cryogenic sol id propellant hybrid rocket engine, and the gas generator hybrid rocket engines.

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Analysis for Combustion Characteristics of Hybrid Rocket Motor (하이브리드 로켓의 연소특성 해석)

  • 김후중;김용모;윤명원
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2001.11a
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    • pp.61-67
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    • 2001
  • Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. The recent research efforts are focused on the improvement of volume limitation and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the eddy breakup model and Hiroyasu and Nagle and Strickland-Constable model are used for soot formation and soot oxidation. Radiative heat transfer is modeled by finite volume method. To reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect, the Low Reynolds number k-$\varepsilon$ turbulent model is employed. Based on numerical results, the detailed discussion has been made for the turbulent combustion processes in the vortex hybrid rocket engine.

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Development of a 1500N-thrust Swirling-Oxidizer-Flow-Type Hybrid Rocket Engine

  • Sakurazawa, Toshiaki;Kitagawa, Koki;Hira, Ryuji;Matsuo, Yuji;Sakurai, Takashi;Yuasa, Saburo
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2008.03a
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    • pp.849-854
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    • 2008
  • We have been developing a 1500N-thrust Swirling-Oxidizer-Flow-Type hybrid rocket engine. In order to put the engine into practical use, we conducted long duration burning experiments up to 25s to examine the influence of configuration change of fuel grain on the engine performance and designed an LOX vaporization nozzle to supply GOX for the 1500N-thrust engine. The experiment with a small hybrid rocket engine showed that combustion was stable and the engine performance was approximately constant during combustion. There was no essential problem to with increasing combustion time. The LOX vaporization nozzle designed had 30 rectangular channels with a depth of 0.5mm. During passing through the nozzle, the LOX increased in temperature and vaporized sufficiently.

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Optimal battery selection for hybrid rocket engine

  • Filippo, Masseni
    • Advances in aircraft and spacecraft science
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    • v.9 no.5
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    • pp.401-414
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    • 2022
  • In the present paper, the optimal selection of batteries for an electric pump-fed hybrid rocket engine is analyzed. A two-stage Mars Ascent Vehicle, suitable for the Mars Sample Return Mission, is considered as test case. A single engine is employed in the second stage, whereas the first stage uses a cluster of two engines. The initial mass of the launcher is equal to 500 kg and the same hybrid rocket engine is considered for both stages. Ragone plot-based correlations are embedded in the optimization process in order to chose the optimal values of specific energy and specific power, which minimize the battery mass ad hoc for the optimized engine design and ascent trajectory. Results show that a payload close to 100 kg is achievable considering the current commercial battery technology.

Analysis for Combustion Characteristics of Hybrid Rocket Motor (하이브리드 로켓의 연소특성 해석)

  • 김후중;김용모;윤명원
    • Journal of the Korean Society of Propulsion Engineers
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    • v.6 no.1
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    • pp.21-29
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    • 2002
  • Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. The recent research efforts are focused on the improvement of volume limitation and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the eddy breakup model and Hiroyasu and Nagle and Strickland-Constable model are used for soot formation and soot oxidation. Radiative heat transfer is modeled by finite volume method. To reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect, the Low Reynolds number $\kappa-\varepsilon$ turbulent model is employed. Based on numerical results, the detailed discussion has been made for the turbulent combustion processes in the vortex hybrid rocket engine.

Fuel-rich Combustion with AP added Propellant in a Staged Hybrid Rocket Engine (다단 하이브리드 로켓에서 AP 첨가 추진제의 연료과농 연소)

  • Lee, Dongeun;Lee, Changjin
    • Journal of the Korean Society for Aeronautical & Space Sciences
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    • v.44 no.7
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    • pp.576-584
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    • 2016
  • In this study, AP added propellant has been proposed as a method of enhancing the low specific impulse performance found for staged hybrid rocket engine. Experimental tests were carried out to analyze and evaluate the effect of AP added propellant on specific impulse performance as well as fuel-rich combustion characteristics in a staged hybrid rocket engine. Upper limit of AP content in propellant was set to be 15 wt% to maintain the hybrid rocket engine advantages. As a result, 15 wt% AP added propellant showed 3% higher specific impulse performance compared to 0 wt% AP added propellant. Moreover, AP addition proved to offer less injected oxidizer mass flow, less O/F variation, and less combustion pressure while producing fuel-rich gas of the same combustion temperature. Future studies will carry out more combustion tests with metal additives to further enhance specific impulse.

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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Manufacture & Launch of Small PE/$LN_2O$ Hybrid Rocket with 50 kgf Thrust Level (추력 50 kgf 급 PE/$LN_2O$ 소형 하이브리드 로켓 제작 및 시험발사)

  • Kim, Hyeon-Woo;Jeon, Min-Ho;Oh, Ji-Sung;Han, See-Hee;Kang, Min-Seok;Jang, Hyoung-Gui;Kim, Hee-Yong;Bae, Tae-Hyun;Lee, Sun-Jae;Kim, Jin-Kon
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2009.11a
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    • pp.507-510
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    • 2009
  • The small size of hybrid rocket using PE-$LN_2O$ was designed, constructed and launched for a development basic technology of Hybrid rocket vehicle. The hybrid engine ignition system was designed with valve system using external actuator and confirmed working without any fault. To design fuel grain an internal ballistics design was carried out, and to estimate rockets flight path an external ballistics analysis was carried out. So the rocket was designed and constructed, and the launch test proves that hybrid rocket's design was suitable. The hybrid rocket(weight : 9kg, diameter : 110 mm, height : 1.7 m) was launched successfully. But parachute was deployed on mid-flight and the mission could not finish its purposed flight. Some of problems were found in this activity but next launch vehicle will be improved.

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