• Proceedings of the Korean Nuclear Society Conference
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• 1997.05a
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• pp.114-118
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• 1997

현재 PWR에서는 반응도 변화를 보상하기 위해서 강한 중성자 흡수체인 수용성 붕소가 사용되고 있고 수용성 붕소내의 $B^{10}$은 연소하게 된다. 최근 대부분의 발전소가 장주기 운전전략을 채택하고 있고 발전소 이용률이 증가하고 있어 $B^{10}$ 연소의 중요성이 부각되고 있다. 본 연구에서는 국내의 웨스팅하우스형 3-Loop 발전소에 대해 $B^{10}$의 연소거동을 모사하는 프로그램을 개발하여 $B^{10}$ Percent를 연소도의 함수로 계산하였고 실제로 냉각재내의 $B^{10}$ Percent를 측정하여 검증하였다. 또한 $B^{10}$ Percent 변화에 따른 Boron Letdown Curve를 보정하여 측정값과 비교, 검증하였다. 검증결과 $B^{10}$ 예측 프로그램의 예상값과 측정값이 비교적 잘 일치하는 것으로 나타났다. 따라서 발전소에서 나타나고 있는 임계붕산농도의 설계값과 측정값 간의 일반적인 차이에 대한 원인이 규명되었고, $B^{10}$ 연소에 의한 상대적인 흡수체의 감소는 정지여유붕소량 계산시 고려되어야할 필요성이 있는 것으로 나타났다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.474-478
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• 2010

Hybrid rocket combustion experiments using liquefying diaphragm made by blended liquefying fuel with 10 wt% of LDPE were performed. Results of experiments were compared to the those of pure paraffin. In case of using liquefying diaphragm, regression rate of rear fuel grain, characteristic velocity and specific impulse highly increased due to the induced turbulent intensity and heat transfer. The serious combustion instability was not observed in analysis of combustion instability. These results can imply that the liquefying diaphragm is efficient to improve low combustion efficiency in hybrid rocket using liquefying fuel.

• Fire Science and Engineering
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• v.25 no.5
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• pp.69-75
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• 2011

It was performed to test the combustive properties of low density polyethylene and ethylene vinyl acetate (LDPE-EVA) composite by the addition of magnesium hydroxide. Flame retardant of natural magnesium hydroxide was added to the mixture of LDPE-EVA in 40 to 80 wt% concentration. The composite was compounded to prepare specimen for combustive analysis by cone calorimeter (ISO 5660-1). Comparing with virgin LDPE-EVA, the specimens including the magnesium hydroxide had lower flashover possibility. It is supposed that the combustive properties in the composites decreased due to the endothermic decomposition of magnesium hydroxide. The specimens with magnesium hydroxide showed both the lower total heat release rate (THR) and lower CO production rate than those of virgin polymer. As the magnesium hydroxide content increases, the total smoke release (THR) and smoke extinction area (SEA) decreased.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.61-66
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• 2004

The objective of the present study is to develop methodology for the assessment of combustion stability of liquid rocket injectors. To simulate actual combustion occurring inside of a thrust chamber, a fullscale injector has been employed in the study, which bums gaseous oxygen and mixture of methane and propane. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a fullscale thrust chamber. A single split triplet injector has been used with an open-end cylindrical combustion chamber. The characteristics revealed by excited dynamic pressures in gaseous combustion show degrees of relative acoustic damping depending on operating conditions. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.119-122
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• 2007

The present study employs experimental approach to identify combustion stability characteristics of fuel-rich gas generators. The gas generator of interest, fueled by LOx and Jet A-1, experienced combustion instability coupled to a longitudinal resonant mode of the combustion chamber at about 1200 Hz. The occurrence of instability is strongly associated with acoustic boundary condition at the exit .and axial location of maximum heat release. As a result, stretching heat release zone in the axial direction by increase of the fuel nozzle diameter has dramatically stabilized combustion.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.69-73
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• 2009

Combustion test results of 1/2.5-scale thrust chamber for 75tonf-class liquid rocket engine were described. The thrust chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion ratio of 12. The combustion tests were conducted to verify the combustion performance, the regenerative cooling performance and the durability of thrust chamber at design point condition, and then were performed to confirm the operation and the combustion performance at low combustion pressure condition. All the tests had been successfully executed without the damage of the hardware. These test results present a possibility of hot firing test at low combustion pressure condition, and can be used as fundamental data to predict the combustion performance at design point condition for 75 tonf thrust chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.8-14
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• 2010

For the development of a liquid rocket engine, hot-firing tests of a regeneratively cooled thrust chamber were performed at chamber pressures of approximately 30 and 60 bars. In the paper, pressure fluctuation data, which were obtained from the dynamic pressure transducers installed in propellant manifolds and combustion chamber, were analyzed. Compared to the data at chamber pressure of 60 bar, the results at chamber pressure of 30 bar showed low-frequency oscillations around 150 Hz in the combustion chamber. The low-frequency waves in the combustion chamber were coupled with those in the manifolds. However, the RMS values of the chamber pressure fluctuations at chamber pressure of 30 bar were only 0.8% of the chamber pressures. Thus, it can be inferred that the thrust chamber operates in the stability boundary even at low chamber pressure.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.3
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• pp.89-100
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• 2021

This paper deals with the construction of a combustion test facility and preliminary tests for hot-firing tests of a methane engine. First, the combustion test facility for a 1 kN-class thrust chamber using liquid oxygen/gas methane as propellants was designed and built. Before hot-firing tests, the cold-flow tests of each propellant line and the ignition tests of torch igniter/afterburner were performed to verify propellant supply stability of the combustion test facility, operation of the control and measurement system, and successful ignition. Finally, a preliminary hot-firing test was conducted to measure the combustion efficiency, heat flux, and combustion stability of a thrust chamber prototype. The constructed combustion test facility will be helpfully used for basic research and development of methane engine thrust chambers.

• Journal of the Korean Society for Nondestructive Testing
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• v.23 no.5
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• pp.488-496
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• 2003

To establish nondestructive test method for rocket motor assembly with rubber and aerospace composite materials, practicable quality control acoustic emission test method is presented. Structural analysis for motor assembly is performed by ABAQUS code and analysis output result is confirmed by strain gage and AE data. Various specimens were tested and analyzed using strain gage and acoustic emission data. The hit rate of acoustic emission was closely related with case/rubber debonding. This report also describes practicable acoustic emission nondestructive method for evaluating motor case assembly quality assurance in the industrial field.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.258-262
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• 2011

The hybrid combustion experiments using non-combustible diaphragm were performed for characteristic of regression rate and combustion efficiency. Results of experiments using diaphragm were showed that the regression rate and efficiency were increased. In addition, the larger difference between fuel grain port and diaphragm port increase the regression rate and efficiency. The modified regression rate equation was proposed with the port area ratio of fuel and diaphragm.