• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.27-35
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• 2004

In this paper a method for finding solution of acoustic, vortex and entropy wave-equations in a cylindrical tube with variable section area was suggested under the consideration of that the high frequency instability in a rocket engine combustion chamber is an acoustic phenomena, which is coupled with combustion reaction, and that a combustion chamber and exhaust nozzle are usually shaped cylindrically. As a consequence of that some method, which enable the quantitative analysis of the influence of entropy and vortex waves to acoustic wave, was suggested.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.04a
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• pp.4-4
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• 1999

최근 충남대에서는 다양한 추진제와 연소 조건으로 액체로켓 연소 실험이 진행되고 있으며 그에 비례하여 많은 사고들이 발생하고 있다. 한 예로 GO$_2$/kerosene을 추진제로 하는 노즐삭마 시험용 로켓엔진(추력 1001bf, 연소실 압력 600psia) 개발 중에 화염이 인젝터면에 형성되어 인젝터면을 손상시키는 사고를 여러 번 경험하게 되었다. 본 연구는 인젝터 손상의 원인을 규명하여 안정적인 인젝터 설계에 도움을 주기 위한 목적으로 실험용 동축 인젝터를 제작하여 화염 부상 특성을 실험적으로 연구하였다. 사용된 인젝터는 연료인 Kerosene을 접선형 선회기로부터 90$^{\circ}$의 원뿔 각을 가지고 분무되도록 설계하였으며, 그 주위로 산하제인 GO$_2$가 연소실의 축방향에 수평하게 분무되도록 설계하였다. 2-유체 동축 인젝터의 난류 확산 화염에서 연료와 산화제의 혼합은 화염 특성을 결정하는 주요 변수이므로 인젝터로부터 분무되는 추진제간의 유량을 변화시켜 화염 부상 특성을 연구하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.04a
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• pp.1-1
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• 1999

액체로켓엔진은 인젝터, 추력실, 노즐 등으로 구성되어 있으며, 이들 구성요소 중에서 연소성능과 연소안정성에 가장 지배적인 요소는 인젝터이다. 본 연구에서는 액체로켓엔진에 사용될 인젝터의 F-O-O-F 충돌형 Uni-element 인젝터를 설계하였으며 특히, F-O-O-F 충돌형 인젝터는 두 단계의 충돌과정으로 구성되는데, 첫단계 충돌은 산화제와 연료 오리피스의 분무로 충돌이 이루어져 Liquid Sheet가 형성되며 다음 단계로는 이러한 양쪽의 Liquid Sheet가 서로 충돌하여 미립화가 진행되는 두단계 충돌과정으로 이루어지게 된다. 그러므로 인젝터의 가공 정밀도가 매우 높아야 하며 인젝터 오리피스 크기나 각도 등의 어느 하나라도 정밀도가 낮게 가공된다면 연소성능과 연소안정성에 심각한 영향을 주게 된다.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.2
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• pp.25-33
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• 2012

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.4
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• pp.76-81
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• 2003

To investigate the convective heat transfer phenomena inside the Lox/Kerosene liquid rocket combustion chamber, hot fire tests were performed by using a water-cooled calorimetric chamber. The calorimetric chamber consists of one cylindrical section and nozzle section with independent cooling passage. To measure the heat flux, thermocouples were installed inlet and outlet of cooling passage of each section. The investigated range of combustion chamber pressure is from 100 psi to 300psi at fixed O/F ratio of 2.0 and radiation heat transfer from the hot gas to the surface is not considered. The measured heat flux was almost linearly depended on the chamber pressure.

• 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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• 2011.11a
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• pp.996-999
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• 2011

The flow and combustion dynamics in the ignition and ramjet sustainer phase of an integrated rocket-ramjet(IRR) engine are investigated. The physical model includes the entire engine flowpath, from the freestream in front of the inlet to the exit of the exhaust nozzle. The flowfield obtained from a rocket booster study is used as the initial condition for the present analysis, so that the complete operation history of the engine can be obtained. The analysis for the primary factor governing flame propagation during the ignition and the key mechanisms for driving and sustaining the flow oscillations are performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.127-134
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• 2011

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.16-22
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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 and mixture ratios for the peripheral injectors. 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.1-8
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• 2005

An axisymmetric compressible flow solver accounting for chemical equilibrium has been developed as an analysis tool exclusively suitable for performance prediction and optimum contour design of LRE thrust chamber. By virtue of several features focusing on user-friendliness and effectiveness including automatical grid generation and iterative calculations with changes in design parameters prescribed through only one keyword-type input file, a design engineer can evaluate very fast and easily the influences of various design inputs such as geometrical parameters and operating conditions on propulsive performance. Validations have been carried out for various aspects by detailed comparisons with the result of CEA code, experimental data of JPL nozzle, actual data for two historical engines, and ReTF data for KSR-III.