• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.163-168
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• 2006

The combustion performance tests of a 30 tonf-class full-scale combustion chamber performed with kerosene as a coolant were described. The combustion chamber has chamber pressure of 53bara and propellant flow mass rate of 90kg/s. Since it was first firing test for 30tonf-class combustion chamber using kerosene cooling, kerosene coolant mass flow rate of 32kg/s which correspond to 120% of design mass flow rate were performed. Then, the firing test with kerosene mass flow rate of 25kg/s were successfully performed. The test results are described and the results showed that the kerosene cooling performance of this combustion chamber is sufficient and the firing test with regenerative cooling is feasible.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.1-9
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• 2020

Using kerosene and liquid oxygen, 1.5-tonf class liquid-liquid pintle injector with rectangular two-row orifice was designed and manufactured. The combustion test of the pintle injector was carried out to verify the combustion performance and combustion stability under a supercritical condition which is the actual operation condition of the liquid rocket engine. The combustion test result showed that the pintle tip was damaged by the high temperature combustion gas in the high-mixed ratio recirculation zone of the combustion chamber. To solve this problem, the insert nozzle was installed in the pintle injector to increase cooling performance at the pintle tip. As a result of the hot firing test, installation of the insert nozzle, AR and BF had a great effect on pintle tip cooling performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.147-150
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• 2007

In this study, cooling characteristics of combustion gas were investigated by injecting liquid nitrogen into liquid rocket combustion chamber. A injection ring of liquid nitrogen was installed between a combustion chamber and a mixing chamber which was designed for mixing of combustion gas and nitrogen. At first, a ignition test of liquid rocket engine was conducted to verify a stable combustion process and 10 second combustion tests were successfully conducted. The results showed that combustion gas of LRE could be cooled by using liquid nitrogen.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.36-42
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• 2004

The Properties of material, C18200 which is used for development of high performance liquid rocket engine combustor chamber were obtained by tension tests. The specimen for regenerative combustor was designed by structural analysis using that Properties. After the designed specimen was manufactured by the same manufacturing process of regenerative combustor. the yielding stress and yielding strain were obtained by strength tests. The properties of C18200 was degraded very much after brazing. The estimation of yielding pressure by structural analysis was almost same as that of strength test. The collector Part was yielded and failed previously than that of cooling channel part during strength test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.636-640
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• 2011

Film cooling technique has been applied to effectively reduce thermal load on liquid rocket combustion chambers by direct injection of a portion of propellant, which flows through the regeneratively cooling channels, into the chamber wall. This study developed a comprehensive model to quantitatively predict the effects of kerosene film cooling on propulsive performance and wall cooling at supercritical pressure conditions, and assessed the predictive capability against hot-firing tests of an actual combustor. The present model is expected to be utilized as a design and analysis tool to meet the conflicting requirements in terms of performance, cooling, pressure loss and weight.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.30-30
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• 2000

액체로켓엔진의 단일추진제 가스발생기는 연료공급 시스템의 터보펌프를 구동시키기 위한 작동가스 생성을 목적으로 사용된다. 고체추진제 가스발생기와 비교할 경우 작동시간이 보다 길고 연소생성물에 의한 터빈 블레이드의 삭마가 없으며 제어가 용이하므로 초기 액체로켓엔진 개발시부터 사용되어 왔다. 80년대 이후 개발된 액체로켓엔진은 이원추진제 가스발생기 또는 연소가스 FEEDBACK 시스템을 채용하고 있지만 단일추진제 가스발생기는 과산화수소수 또는 하이드라진과 같은 별도의 추진제 공급 시스템을 필요로 하는 단점에도 불구하고 상대적으로 낮은 온도의 무연 작동 가스를 발생하므로 가스발생기 자체를 위한 냉각시스템을 제거 또는 최소화 시켜 간단한 구조로 전체 시스템 설계를 가능하게 하므로 중소형 액체로켓엔진에 사용되고 있다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.10
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• pp.966-978
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• 2011

Transpiration cooling is the most effective cooling technique for the high-performance liquid rockets and air-breathing engines operating in aggressive environments with higher pressures and temperatures. When applying transpiration cooling, combustor liners and turbine blades/vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. Practical implementation of the cooling technique has been hampered by the limitations of available porous materials. But advances in metal-joining techniques have led to the development of multi-laminate porous structures such as Lamilloy$^{(R)}$ fabricated from several diffusion-bonded, etched metal thin sheets. And also with the availability of lightweight, ceramic matrix composites(CMC), transpiration cooling now seems to be a promising technique for high-performance engine cooling. This paper reviews recent research activities of transpiration cooling and its applications to gas turbines, liquid rockets, and the engines for hypersonic vehicles.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.69-77
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• 2020

A quasi one-dimensional multi-phase reaction flow analysis code is developed for the performance analysis of liquid pintle thrusters. Unsteady flow field, droplet evaporation, finite reaction and film cooling models are composed as the major models of the performance analysis. The droplet vaporization takes account of Abramzon's vaporization model, and the combustion employs a flamelet model based on detail chemical reactions. Shine's model is applied for the film cooling calculation. To verify each model, the Sod shock tube, single droplet vaporization, kerosene droplets combustion, and film length are evaluated.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.3
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• pp.38-44
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• 2000

In general liquid rocket nozzles are protected from hot combustion gas by regenerative cooling techniques. But due to the complexity of the cooling system, it causes increase of system cost and frequently source of the system malfunction. Recently, instead of regenerative cooing, ablative material are used to protect combustion chamber wall and nozzle. To determine the nozzle material erosion rate and erosion shape, more than 500 hot fire test were performed by using 100 lb thrust experimental liquid rocket. Test variable were propellant feed sequence, injector, position of igniter and liquid oxygen supply temperature.