• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.1-12
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• 2019

A fundamental study on a gas injection variable thrust method for thrust throttling in a liquid rocket engine was conducted. The gas injection variable thrust has the advantage of not only being able to control the thrust with a simple structure but it also increases the atomization performance through the injection pressure drop that increases in direct proportion to the density reduction. In this study, spray characteristics such as spray instability, spray pattern, spray angle, and breakup length based on changes in the liquid mass flow rate and amount of injected gas were investigated using effervescent swirl injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.66-70
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• 2010

The performance of 75 ton liquid rocket engine combustion chamber for a space launch vehicle was predicted through firing tests at low pressure. In low pressure tests of 75 ton LRE combustor chamber, the combustion characteristic velocity of 1750 m/sec and the specific impulse of 240 sec were obtained which are higher than the low pressure performance of 30ton combustion chamber. The combustion characteristic velocity of 1770 m/sec and the specific impulse of 278 sec at design point for 75 ton LRE combustion chamber were predicted by using the low/high pressure performance correlation of 30ton LRE combustion chamber.

• Journal of the KSME
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• v.56 no.9
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• pp.38-43
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• 2016

초임계 난류유동이 활용되고 있는 가장 대표적인 시스템은 액체로켓연소기이다. 액체로켓엔진은 목적하는 임무에 따라 추력이 결정되고 엔진의 추력이 결정된다. 이러한 추력을 결정하는 핵심부품이 분사기(injector)이다. 실용 인공위성을 발사하기 위한 액체로켓엔진의 연소기는 수백 개의 분사기를 통해 연료와 산화제가 혼합되는 구조를 가지고 있다. 따라서 로켓연소기의 연소특성 및 성능은 분사기의 혼합특성에 좌우된다. 그러므로 단일 분사기의 연료/산화제 혼합특성에 대한 많은 실험과 해석연구가 진행되고 있다. 그런데 초임계압력에서는 액체의 표면장력이 사라지게 되어 독특한 혼합특성을 가지고 있기 때문에 성능을 높여주기 위하여 초임계 압력을 선택할 경우 분사기에 대한 연구가 선행되어야 한다. 이 글에서는 이러한 초임계 작동압력에서 분사기에 대한 연구들을 소개하고자 한다.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.4
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• pp.1-8
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• 2005

An experimental investigation of a microthruster that uses hydrogen peroxide as a monopropellant is described. The study comprises of preparation method of silver as a catalyst and performance evaluation of a mesoscale reactor. Reduction of silver in $H_2\;at\;500^{\circ}C$ resulted in the best reactivity of all the treatment method tested. A mesoscale reactor was built to find the optimum configuration for full decomposition of propellant. The catalyst bed was made of a glass wafer substrate sputtered with silver and had a length of 20 mm. We measured the conversion rate with varying feed rate of $H_2O_2$ and preheating temperature. With the feed rate of $H_2O_2$, the space time within the reactor varies as well. For the bed length of 20 mm, space time more than 480 s was required for full conversion.

• Aerospace Engineering and Technology
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• v.9 no.1
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• pp.93-97
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• 2010

The analysis of specific impulse of the liquid rocket engine employing gas generator cycle with LOx/kerosene as propellant has been performed. The relative error of performance of 300 ton level engine is 0.1%s for specific impulse and 12% for optimal combustion pressure comparing with the published data. The difference of the performance model and the material property models of gas generator product gas are the presumed major reason of discrepancy. The optimal condition of 30 ton level engine is combustion pressure of 68 bar and mixture ratio of 2.2 for maximum specific impulse. This optimal condition can be changed by performance models.

• Proceedings of the KIPE Conference
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• 2002.07a
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• pp.133-137
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• 2002

리니어 모터 중에서 위치 정밀도가 우수하고 개루프 제어가 가능한 리니어 펄스모터(LPM : Linear Pulse Motor)에 대해 추력 성능 및 위치제어 정밀도 향상을 위해서 자기등가회로 모델링법에 의해 추력을 주로 계산하고 있다. 퍼미언스(Permeance) 모델법, 수치해석법(FEM), 속도 기전력법 등을 이용하여 자속밀도(B)를 구하는데 있어 해석적인 추력편차가 존재한다. 정확한 추력을 구하기 위해서는 이들 각각의 방법을 이용한 로렌츠 포스법(Lorentz Force Method), 자기수반 에너지법, Maxwell 응력법 등으로 추력을 계산하여 비교함으로써 모터용량, 파라미터 등의 각종 성능 비교시 중요 요소가 될 수 있다. 이에 본 연구는 위의 방법으로 LPM의 추력을 수식화하여 값을 얻고, 실제 추력값을 측정하여 얻어진 값과 비교 분석하고자한다. 이러므로 보다 정확한 LPM 추력 계산 방법을 얻을 수 있을 것으로 기대된다.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.19-26
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• 2014

A ground hot-firing test (HFT) was carried out to make a close examination into the pressure instability for the 70 N-class hydrazine thruster under development. Monopropellant grade hydrazine was adopted as a propellant for the HFT, and catalyst-bed was filled with $Ir/Al_2O_3$ catalyst. In order to investigate the effects of thrust-chamber diameter on combustion stability, evaluation tests for the development models were performed on three kinds of lower thrust chambers having the length-to-diameter ratio (L/D) of 1.03, 1.13, and 1.26. As results, it was found that low frequency instability (~ 50 Hz) was inherent in the models, and in addition, increase of the L/D and decrease of the operating pressure led to an amplification of pressure oscillation in the test condition specified.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.28 no.10
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• pp.1264-1270
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• 2004

In this paper research on micro solid propellant thruster is reported. Micro solid propellant thruster has four basic components; micro combustion chamber, micro nozzle, solid propellant and micro igniter. In this research igniter, solid propellant and combustion chamber are focused. Micro igniter was fabricated through typical micromachining and the effect of geometry was evaluated. The characteristic of solid propellant was investigated to observe burning characteristic and to obtain burning velocity. Change of thrust force and the amount of energy loss following scale down at micro combustion chamber were estimated by numerical simulation based on empirical data and through the calculation normalized specific impulses were compared to figure out the efficiency of combustion chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.147-150
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• 2003

The purpose of this research was to develope components of micro solid propellant thruster. Micro solid propellant thruster had four basic components: combustion chamber, nozzle, solid propellant and micro heater for ignition. A performance of micro heater and characteristic of solid propellant was investigated. Micro heater was fabricated by conventional MEMS process and Platinum layer was used for heating element. Effect of geometry parameters on micro heater was tested. The temperature responses of heater with respect to each parameters was compared for a given electrical power. The characteristic of solid propellant(HTPB/AP) was investigated to obtain burning velocity in small chamber. Additionally, a capacity of filling propellant with high viscosity in small chamber was checked to guarantee for the micro fabrication.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.942-948
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• 2011

Solid propulsion systems have simple structures compared to other propulsion systems and are suitable for long-term storage. However the systems generally have limits on control of thrust levels. In this paper we introduce controllers for combustion chamber pressure using on-off control techniques which have been known for relatively easy implementation and energy efficiency. For this, we use a simple pressure change model by considering only mass conservation within the combustion chamber and we design a classical controllers and on-off controllers with are Pulse Width Modulation(PWM) and Pulse Width Pulse Frequency Modulation (PWPFM). Then we compare the performance results of the controllers through numerical simulations.