• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.72-77
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• 2007

The combustion test results of the sub-scale combustor having dual swirl injector with internal mixing for a liquid rocket engine are described. The sub-scale combustor uses liquid oxygen(LOx) and kerosene as propellants and has an injector head, an ablative material combustor wall and a water cooled nozzle. The injector head has LOx manifold, fuel manifold, fire face plate, one center swirl injector and 18 main swirl injectors of internal mixing. The combustion tests were successfully performed at design and off-design points without any damages on the injectors. Combustion characteristics velocity of 1756m/s was measured at design point. High frequency combustion instability was not observed but low frequency pulsations occurred at off-design conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.288-293
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• 2006

The combustion performances and characteristics of the subscale liquid rocket combustion chamber are discussed in this paper. Subscale combustion chamber is composed of mixing head, ablative cooling cylinder, and water cooling nozzle. The mixing head has eighteen coaxial swirl injectors and one center coaxial swirl injector for ignition. The mixing heads employing the injectors of low different recess length are considered in this paper. The results of the firing test, comparison of performance, and characteristics of static and dynamic pressures of the four different mixing heads are described. The characteristics of combustion at design and of design points are also discussed.

• 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, $0.28-0.58MW{\cdot}s/kg$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.117-120
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• 2007

This study investigated the spray modeling of the pressure-swirl atomizer installed in liquid rocket engine and the effect of drop distribution function especially. The $X^2$, originally implemented to KIVA, Rosin-Rammler and modified Rosin-Rammler distribution functions were investigated theoretically and compared to each other. Also, they were applied to pressure-swirl atomizer similar to the injector installed in liquid rocket engine to evaluate the feasibility for LRE injector. Among the distribution functions, original Rosin-Rammler distribution function was the most compatible with predicting the spray characteristics of pressure-swirl atomizer installed in liquid rocket engine.

• Proceedings of the KWS Conference
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• 2004.05a
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• pp.87-89
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• 2004

Brazing is an indispensable manufacturing technology for liquid rocket engine. In this study, for LRE injector, stainless steel 316L was used of base metal and Ni based MBF-20 of insert metal. The brazing and diffusion was carried out under various conditions. There are solid phase and. residual liquid phase in the brazed joint. With increment of holding time, the amount of solid phase increased and the elements of base metal and insert metal compositionally graded. Boron diffused from insert metal came into base metal and made boride with Cr and Mo at the brazed joint of base metal and insert metal.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.47-53
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• 2015

Design parameters required for Methane/oxygen bipropellant small-rocket-engine were derived through a theoretical performance analysis. The theoretical performance of the rocket engine was analyzed by using CEA and optimal propellant mixture ratio, characteristic length, and optimal expansion ratio were calculated by assuming chemical equilibrium. A coaxial-type swirl injector was chosen because of its outstanding atomization performance and high combustion efficiency compared to other types of injector and also a bell nozzle with 80% of its full length was designed. The rocket engine configuration with 1.72 MPa of chamber pressure, 0.18 kg/s in total propellant mass flow, and O/F ratio of 2.7 was proposed as a ground-firing test model.

• Journal of the Korea Institute of Military Science and Technology
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• v.9 no.4
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• pp.137-144
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• 2006

The basic data obtained in this research for single element performance were directly applied to the design of injector head(7 elements). Designed performance of the 7-element Swirl Coaxial injector was $245kg_f$ sea level thrust with 20bar combustion chamber pressure. Numerical analysis were performed to obtain the change of spray pattern for the design of injector head, and we confirmed the feasibility and application of those results. Hot tests were performed for the multi-element injector to compare with the performance of the single element injector and those can be applied to the design of scaled liquid rocket engine. The basic data obtained in this research can be directly applied to the real liquid rocket injector design.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.4
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• pp.12-19
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• 2001

The discharge characteristics of the oxidizer injector of liquid rocket engine were investigated. The discharge performance was obtained numerically which agreed quantitatively with the measured data. The discharge coefficient is proportional to the cavitation number for cavitating flow and constant for non-cavitating flow. The Reynolds number, however, affects little the discharge coefficient. The discharge coefficient decreased slightly as the Reynolds number increased because the friction loss decreased relatively at high Reynolds number flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.11a
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• pp.5-5
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• 2000

이 연구의 목적은 한국항공우주연구소가 개발 중인 액체추진제 로켓엔진의 축소형 엔진에 대하여 인젝터 배열의 변화가 성능 및 연소특성에 미치는 영향을 분석하는데 있다. 인젝터의 배열방식에 따라 방사형(radial) 및 직교형(H-type) 인젝터를 연구대상으로 하였으며 충돌각의 변화에 2차원 및 3차원 해석을 수행하였다. 로켓엔진에는 스월러 인젝터를 고려하지 않았기 때문에 인젝터의 배열 및 충돌 각은 엔진성능뿐만 아니라 연소특성에도 중요하게 영향을 미치는 인자가 된다.(중략)

• 한국연소학회:학술대회논문집
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• 2012.11a
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• pp.291-294
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• 2012

The research and development status of combustion chamber of liquid rocjet engine for Korea Space Launch Vehicle(KSLV-II) are briefly described. The cold and hot firing tests of uni-element injector, the performance/heat flux measurement/hot firing tests of subscale combustion chamber and the performance/stability rating/regenerative cooling/hot fire tests of 30ton-class combustion chamber were successfully performed. Based on these results, the research and development of combustion chamber for 75ton-class liquid rocket engine are underway.