• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.31-35
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• 2008

This test cell is developed to provide the fundamentals of operational mechanism and structural configuration, and further to verify thermodynamic calculation with this test data to the institutes or laboratories research and study gas turbine engine for academic purpose. The test cell is installed to monitor and collect real-time data as to temperature, pressure, thrust, fuel flow, and air flow etc. using by NI DAQ(Data acquisition)device and LabVIEW program based on 30lbf-micro turbojet engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.46-51
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• 2012

This study aims to develop the performance analysis program on the staged combustion cycle of the liquid rocket engine using liquid oxygen(LOx) as oxidizer, liquid hydrogen(LH2) and RP-1 as fuel. The developed analysis program can obtain the propellant mass flow rate, the specific impulse, and representative design values of engine components for the required thrust satisfying pressure, mass flow, and energy balance conditions. The analysis results show that the the specific impulses (Isp) compared to those of the real engines have been less than 1%. With additional constraints, the program will be improved for the system optimization.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1995.11a
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• pp.59-66
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• 1995

인공위성의 궤도진입, 궤도수정 및 자세제어를 담당하는 추진체계는 위성의 용도와 궤도위치, 자세/궤도제어 방식 등을 고려하여 설계하여야 한다. 현재까지도 널리 사용되고 있는 인공위성 추진체계는 AKM(Apogee Kick Motor)과 단일 추진제 추력기로 구성된 "통상 추진체계"이나 최근에는 AKE(Apogee Kick Engine)과 이원 추진제추력기로 구성된 "통합 추진체계" 그리고 이와 유사한 "완전통합 추진체계", "이중 추진체계" 등이 기술적 선택방안으로 제안되어 일부 적용되고 있는 실정이며 이러한 추진체계의 효과적 실용화를 위해서 단일 추진제(하이드라진) 추력기의 성능향상 및 이원 추진제 엔진과 이원 추진제 추력기의 기술개선 연구가 이루어지고 있다.원 추진제 엔진과 이원 추진제 추력기의 기술개선 연구가 이루어지고 있다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.55-62
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• 2010

In addition to the previous study for development of a 1,200 N-class bipropellant rocket engine with concentrated hydrogen peroxide, the effect of characteristic length and thrust measurement were experimentally evaluated. Tests with characteristic lengths of 0.95, 1.07, and 1.20 m were performed and $C^*$ and Isp efficiencies were increased as increasing characteristic length. The maximum $C^*$ and Isp efficiencies were 98.4% and 93.1% respectively. Based on the evaluation of the designed engine, the optimized characteristic length was proposed in using the engine adapted decomposed hydrogen peroxide and the engine performance at vacuum-level was evaluated using thrust and Isp efficiency at the designed equivalence ratio. As a result, 218.4 s at sea-level, 253.3 s at vacuum-level, and vacuum thrust of 1035.3 N can be estimated.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.3
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• pp.1-8
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• 2011

In addition to the previous study for development of a 1,200 N-class bipropellant rocket engine with concentrated hydrogen peroxide, the effect of characteristic length and thrust measurement were experimentally evaluated. Tests with characteristic lengths of 0.95, 1.07, and 1.20 m were performed and $C^*$ and Isp efficiencies were increased as increasing characteristic length. The maximum $C^*$ and Isp efficiencies were 98.4% and 93.1% respectively. Based on the evaluation of the designed engine, the optimized characteristic length was proposed in using the engine adapted decomposed hydrogen peroxide and the engine performance at vacuum-level was evaluated using thrust and Isp efficiency at the designed equivalence ratio. As a result, 218.4 s at sea-level, 253.3 s at vacuum-level, and vacuum thrust of 1035.3 N can be estimated.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.2
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• pp.23-28
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• 2011

The engine ground and altitude tests were carried out to investigate the effect of jet fuel alteration on the performance of a small turbojet engine. JP-S was supplied 8% higher than JP-8 by fuel metering system at the same command. The employment of JP-S showed the similar starting characteristic to that of JP-8, however, difference in the ignition time and acceleration rate of engine speed due to the difference of fuel flow rate by fuel metering system was observed. In spite of jet fuel alteration, the test results yield the similar steady-state engine performance in net thrust, air flow, exhaust gas temperature, etc. On the other hand, the fuel consumption of JP-S increased by 5 % compared with that of JP-8. In point of specific fuel consumption (SFC), SFC of JP-S was approximately 1.1~2.6 %, 5 % higher than that of JP-8 in ground and altitude tests respectively at the same thrust.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.410-415
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• 2010

The engine ground and altitude tests were carried out to investigate the effect of jet fuel alteration on the performance of a small turbojet engine. JP-S was supplied 8% higher than JP-8 by fuel metering system at the same command. The employment of JP-S showed the similar starting characteristic to that of JP-8, however, difference in the ignition time and acceleration rate of engine speed due to the difference of fuel flow rate by fuel metering system was observed. In spite of jet fuel alteration, the test results yield the similar Steady-State engine performance in Net thrust, Air flow, Exhaust Gas Temperature, etc. On the other hand, the Fuel consumption of JP-S increased by 5 % compared with that of JP-8. In point of Specific Fuel Consumption (SFC), SFC of JP-S was approximately 1.1~2.6 %, 5 % higher than that of JP-8 in ground and altitude tests respectively at the same thrust.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.1-8
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• 2013

A liquid-monopropellant hydrazine thruster has several outstanding advantages such as relatively-simple structure, long/stable propellant storability, clean exhaust products, and so on. Therefore hydrazine thruster has such a wide application as orbit and attitude control system (ACS) for space vehicles. A hydrazine thruster with the medium-level thrust to be used in the ACS of space launch vehicles (SLV) has been developed, and its ground firing test result is presented in terms of thrust, impulse bit, temperature, and chamber pressure. It is verified through the performance test that the response and repeatability of thrust are very excellent, and the thrust efficiencies compared to its ideal requirement are larger than 93%.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.4
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• pp.334-340
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• 2015

Engine derating is often considered for reliability benefits because lower power operation reduces its failure probability. To be derated during operation, however, the engine must be initially overdesigned. The engine overdesign is cost effective only if reliability increased from derating is enough to offset the initial increase in the development cost caused from the overdesign. The purpose of this paper is to provide an analytical model to consider a trade-off between the engine overdesign and derating. We use a logistic regression model to explain reliability growth in the number of hot firing tests for a fixed power level. Using the Transcost model with the reliability growth model, we show that 10% overdesign of Titan rocket engine decreases its development cost by about 9% and 23% depending on the reliability requirement. We also point out that such a cost reduction depends on the fuel type a rocket uses.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.85-90
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• 2003

To control spacecraft including satellite, we should understand precisely the performance of propulsion system and the program logic with appropriate format for satellite operations. In this study, the thruster performance functions was generated by using the best curve fitting for performance data from bi-propellant thrusters. Detailed thruster performance data are, in general, company proprietary information, therefore real firing tests were performed to understand the basic characteristics of the performance curve. Experimental rocket motor utilize liquid oxygen and kerosine as propellant and designed average thrust was 100 pound.