• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.31-34
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• 2011

Design and fabrication of Technology Demonstration Model(TDM) of 75 tonf regenerative cooling thrust chamber were described. It has design chamber pressure of 60 bar, propellant mass flow rate of 243.6 kg/s, and nozzle expansion ratio of 12. It has a single welded structure of the mixing head and the chamber. Design and fabrication technologies established through this TDM can be used to development of flight model.

• 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.

• 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.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.133-137
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• 2008

Results of combustion tests performed for a regenerative cooling combustor of a 30 tonf-class liquid rocket engine were described. The combustion chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. The combustion chamber is composed of mixing head, baffle injector, and regenerative cooling chamber. The hot firing tests were performed at design and off-design points. The test results show that the combustion characteristic velocity is in the range of 1738${\sim}$1751 m/sec and the specific impulse of the combustion chamber is in the range of 253${\sim}$270 sec. The peak of combustion characteristic velocity and specific impulse for this combustor is shown at mixture ratio of 2.35 and 2.5, respectively.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.21-21
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• 2000

통상적으로 액체로켓의 노즐은 재생냉각에 의해 고온의 연소가스로부터 보호된다. 그러나 재생냉각의 경우, 시스템에 상당한 투자가 요구되며, 잦은 엔진 결함의 원인을 제공하기도 한다. 최근 들어 액체로켓에 재생냉각을 사용하지 않고, 연소실과 노즐 보호를 위해 삭마재료가 사용되고 있다. 노즐재료에 대한 삭마량과 삭마형상 연구를 위해 500회 이상의 연소실험이 수행되었다. 그러나 연소실험을 통한 삭마특성은 전혀 예측할 수 없는 방향으로 진행되고 있으며, 실험에 사용된 액체로켓의 작동범위가 실제 로켓과 거의 유사하다는 것을 감안한다면, 삭마재질을 로켓에 적용하기 위해서는 상당한 주의가 필요할 것으로 판단된다. 실험변수는 추진제의 공급 순서, 인젝터의 형상, 점화기의 위치, 그리고 액체산소의 공급온도이다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.211-217
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• 2003

A liquid rocket engine(LRE) Using LO$_2$/LNG(Liquefied Natural Gas) propellants was experimentally evaluated. The purpose of this study was to investigate the performance of the LO$_2$/LNG rocket combustor that is composed of three sect ions(igniter spacer, cylinder and nozzle section), especially focused on the influence of regenerative cool ing effect in association with the phase of regenerative coolant Series of tests were conducted under the conditions of water cool ing and regenerative cool ing with LNG in the cylinder section and independent cool ing with water in the igniter spacer and nozzle sections. Parametric studies on the variation of a chamber pressure and mixture ratio were undertaken. In addition, effect of propellant(LNG) composition and its enthalpy on the performance is examined.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.2
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• pp.54-61
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• 2003

This paper describes a general design procedure of regenerative cooling system for liquid rocket engine(LRE). From this design logic, cooling channels are designed and fabricated. The measured heat flux from firing test is similar to the heat flux predicted by design logic. Therefore, proposed design procedure of cooling channel can be applied to real LRE system. Also the result of firing test indicates that soot from combustion products have strong influence on the cooling characteristics of LRE.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.139-143
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• 2007

The overall results of combustion tests performed for a 30 tonf-class full-scale combustion chambers of a liquid rocket engine were described. The combustion chambers have chamber pressure of 53${\sim}$60 bar and propellant mass flow rate of 89 kg/so The combustion chamber is composed of mixing head, SUS baffle, baffle injector, ablative chamber, channel cooling chamber and regenerative cooling chamber. The test results show that the combustion characteristic velocity is in the range of 1673${\sim}$1730 m/sec and the specific impulse of the combustion chamber is in the range of 254${\sim}$263 sec. As the recess number of the injectors increases, the combustion characteristic velocity increases. And as the combustion characteristic velocity increases, the specific impulse of the combustion chamber also increases.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.67-72
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• 2004

A regenerative cooling system has been designed through empirical 1-D analysis for a liquid rocket engine of 30-ton-level thrust. The hot-gas-side wall temperature from 1-D analysis shows 100K difference compared to 3D CFD analysis. Two variations of design with same cooling performance are suggested for different maximum channel widths i.e., 4mm and 2mm. The coolant pressure drop of the latter design is higher by 20%. The maximum liner temperature is about 700K when TBC and the thermal resistance of carbon deposit are considered. So film cooling is recommended to increase the cooling capacity as the present cooling capacity is insufficient