• Title/Summary/Keyword: Rocket combustor

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A Trade-off Analysis between Combustion and Cooling Performance of a Liquid Rocket Combustor with Fuel Film Cooling Scheme (연료 막냉각을 적용한 액체로켓 연소기의 연소/냉각 성능 간 trade-off 해석)

  • Joh, Mi-Ok;Kim, Seong-Ku;Choi, Hwan-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2012.05a
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    • pp.35-41
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    • 2012
  • Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films. Also is analyzed the effect of varying mixture ratios for the peripheral injectors on combustion performance enhancement. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

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Thrust and Propellant Mixture Ratio Control of Open Type Liquid Propellant Rocket Engine (개방형 액체추진제로켓엔진의 추력 및 혼합비 제어)

  • Jung, Young-Suk;Lee, Jung-Ho;Oh, Seung-Hyub
    • Proceedings of the KSME Conference
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    • 2007.05a
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    • pp.1143-1148
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    • 2007
  • LRE(Liquid propellant Rocket Engine) is one of the important parts to control the motion of rocket. For operation of rocket in error boundary of the set-up trajectory, it is necessarily to control the thrust of LRE according to the required thrust profile and control the mixture ratio of propellants fed into combustor for the constant mixture ratio. It is not easy to control thrust and mixture ratio of propellants since there are co-interferences among the components of LRE. In this study, the dynamic model of LRE was constructed and the dynamic characteristics were analyzed with control system as PID control and PID+Q-ILC(Iterative Learning Control with Quadratic Criterion) control. From the analysis, it could be observed that PID+Q-ILC control logic is more useful than standard PID control system for control of LRE.

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Study on Auto Ignition of Hybrid Rocket Using $N_2O$ Catalytic Decomposition ($N_2O$ 촉매 분해를 이용한 하이브리드 로켓 자연 점화 연구)

  • Yong, Sung-Ju;Kim, Tae-Gyu
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2010.05a
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    • pp.202-205
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    • 2010
  • Auto ignition of hybrid rocket using $N_2O$ catalytic decomposition was studied in the present study. The hybrid rocket consists of catalytic igniter, solid fuel, combustor, and nozzle. The Ru/$Al_2O_3$ catalyst for $N_2O$ decomposition was synthesized by an impregnation method, and $N_2O$ conversion as reaction temperatures was measured. The temperature change of the catalytic ignitor was measured at the operating condition, and the possibility for the auto ignition of hybrid rocket was validated.

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Effect of Flows on the Evolution of Sprays and Combustion in Ramjet Combustor (I) : Ram Air Flows in Combustion Chamber (램제트 연소기 내 유동조건에 따른 분무 및 연소천이 (I) : 연소실 램공기 유동)

  • 함희철;이진호;윤웅섭
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2002.04a
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    • pp.50-54
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    • 2002
  • With a view to estimating the effect of flows on evolving sprays and combustion in ramjet combustor and corresponding extent of combustion, ram air flows in combustion chamber is numerically experimented. Preconditioned three dimensional Navier-Stokes system of equations per transient, compressible, turbulent flows in IRR(Integral Rocket Ramjet) combustor is numerically integrated. Flow properties in the side-dump ramjet combustor, rectangular duct with two 60-deg curved inlets located radially at an angle of 180-deg, are addressed in terms of mixing quality and extent of combustion efficiency.

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램제트 연소기의 보염기 장착에 따른 연소기 특성 변화에 대한 수치적 연구

  • Kim, Seong-Don;Jeong, In-Seok
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2008.11a
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    • pp.455-456
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    • 2008
  • A numerical study was conducted on the effect of flame holder which could be added to the inlet duct of Integral Rocket-Ramjet. Two different types of flame holder installations, flame holder without sudden expansion region and flame holder with small sudden expansion region, were compared and showed different flame shapes and pressure rise in the combustor.

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Combustion Performance Characteristics of a High Pressure Sub-scale Liquid Rocket Combustor (고압 축소형 연소기의 연소 성능 특성에 관한 연구)

  • Kim, Jong-Gyu;Lee, Kwang-Jin;Seo, Seong-Hyeon;Han, Yeoung-Min;Choi, Hwan-Seok
    • Journal of the Korean Society of Propulsion Engineers
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    • v.11 no.5
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    • pp.31-36
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    • 2007
  • Combustion performance characteristics of subscale high-pressure combustor were investigated at 70 bar combustion pressure. All tests were successfully performed without any damage on the combustor. The mixing characteristics and distribution pattern of the injectors were found to have considerable influence on the combustion performance. The characteristic velocity of the combustor was higher in the injector with internal mixing than that of external mixing and in the injector with smaller mass flowrate. The pressure fluctuations at the propellant manifolds and the combustion chamber were measured to be less than 3% of the mean combustion pressure to meet the combustion stability criterion and to prove stable combustion characteristics of the combustor.

Numerical Study of CH4/LOx Combustion of Shear-coaxial Injector in High Pressure Combustion Chamber of Liquid Rocket (액체로켓 동축인젝터(CH4/LOx)의 고압 연소실 내 연소 유동장에 대한 수치적 연구)

  • Kim, Jung Eun;Jeung, In-Seuck
    • 한국연소학회:학술대회논문집
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    • 2014.11a
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    • pp.311-313
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    • 2014
  • High pressure combustion with multiphase--liquid, gas, and supercritical phase--mixtures are widely used technology in the high efficiency liquid propellent rocket engine. This is the typical characteristics differentiate from the combustor of conventional air-breathing engines. Therefore, successful research of high pressure combustion at supercritical condition is essential to develope a high efficiency liquid rocket engine. Numerical studies have been carried out to explore capabilities of numerical method for LOx-CH4 non-premixed flames at high pressure. In this paper, corresponding numerical results are presented and compared with experimental result of MASCOTTE facility.

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A Study on Structural Safety of the Solid Fuel Grain by Hot Flow inside a Hybrid Rocket Combustor (Multi-port 하이브리드 로켓 연소기에서 고온 산화제 유동에 의한 고체연료의 구조적 안전성에 대한 연구)

  • Do, Gyu-Sung;Yoon, Chang-Jin;Kim, Jin-Kon;Moon, Hee-Jang
    • Journal of the Korean Society for Aviation and Aeronautics
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    • v.15 no.4
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    • pp.38-44
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    • 2007
  • This paper describes the structural safety of solid fuel in the Hybrid Rocket Motor (HRM). Hybrid rocket combustion has the distinct regression characteristics which include the process of thermal pyrolysis and fuel vaporization. Most of all, this regression characteristics would structurally affect the strength of the fuel having a multi-port configuration, and even may cause the breaking from the fuel grain. This problem would probably influence the performance and operating safety of HRM. Therefore, for the safe operation of HRM, the critical port radius which determines the structurally safe region was discussed from the heat analysis of the solid fuel.

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Modeling of 2D Axisymmetric Reacting Flow in Solid Rocket Motor with Preconditioning

  • Lee, S.N.;Baek, S.W.
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2008.03a
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    • pp.260-265
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    • 2008
  • A numerical scheme for solid propellant rocket has been studied using preconditioning method to research unsteady combustion processes for the double-base propellant with a converging-diverging nozzle. The Navier-Stokes equation is solved by dualtime stepping method with finite volume method. The turbulence model uses a shear stress transport modeling. The species equation follows up the method of Xinping WI, Mridul Kumar and Kenneth K. Kuo. A preconditioned algorithm is applied to solve incompressible regime inside the combustor and compressible flow at nozzle. Mass flux was evaluated using modified advective upwind splitting method. The simulated result the comparison a fully coupled implicit method and a semi implicit method in terms of accuracy and efficiency. This report shows the result of solid rocket propellant combustion.

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