• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.1-14
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• 2015

In case of a gas turbine engine for supersonic operation, the engine have a wide range of operating inlet mass flow rate and required high performance such as thrust and fuel consumption. Therefore, variable system and its optimal control logic are essentially needed. In this work, a method for performance prediction of a gas turbine engine with variable system compressor and its control scheme were developed. Conceptual design of compact acuation system for the operation of the variable system was also conducted. The performance of a low-bypass ratio mixed flow turbofan engine was analyzed, and it was observed that the surge margin of the engine is improved at off-design condition by applying the control scheme.

• 한국연소학회:학술대회논문집
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• 2004.11a
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• pp.141-145
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess length have been tested for the comparative study of CH flame structure and dynamic characteristics. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a full-scale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of ILASS-Korea
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• v.25 no.3
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• pp.126-131
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• 2020

Throttleable rocket engines are in high demand due to the diversification of space missions. Pintle injector is known to be suitable for throttleable rocket engines, because of its high efficiency in overall thrust zone. In this study, the relationship between spray angle of a throttleable pintle injector and total momentum ratio based on hot fire test conditions was investigated. As a result, the spray angle in 100% and 60% throttling level is higher than the spray angle obtained by the case which considers only propellant mass flow rate, owing to higher total momentum ratio (TMR). The results of this study may be useful for predicting spray angle in hot fire test.

• Journal of the Korean Society of Visualization
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• v.14 no.3
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• pp.22-27
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• 2016

A sweeping jet actuator (SJA) is an instrument generating pulsing jets with no moving elements. Because of its simple design and high durability to shock and vibration, SJA has recently drawn increasing attention for the application to flow control such as aerodynamic control of a wing and thrust vectoring of a jet engine. However, experimental and numerical studies on SJA have been limited to internal flow structure of SJA. In this study, we investigated the flow structure and its variation in the outlet of SJA. We carried out the experiment to understand the flow structures using PIV (Particle Image Velocimetry). The flow structure varies with a degree of the outlet and volume flow rate. There is leaking process during half jetting cycle. The process of the main jet can occur because the jet moving time increased from one side to the other side.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.19 no.2
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• pp.58-63
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• 2011

A Continuous Descent Approach(CDA) is defined as a simple, cost-effective, noise and emission abatement technique for any powered aircraft on approach. CDA also can be optimised within energy, speed and safety constraints by avoiding unnecessary flap, air brake and engine thrust. This study includes comparison on fuel consumption between Continuous Descent type and Step Down type by using flight data. In particular, we investigated fuel flow per hour, calibrated airspeed and pressure altitude for all flight time. During descent flight, the fuel consumption of Continuous Descent type was less than the fuel consumption of Step Down type.

• Journal of the Korea Institute of Military Science and Technology
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• v.6 no.4
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• pp.113-123
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• 2003

Both numerical analysis and experiment of cold and hot tests were performed to obtain basic design data for the swirl coaxial type Injector and to predict the combustion performance. Mass distribution, mixing distribution, mixing efficiency, characteristic velocity efficiency were measured by the cold tests and numerical analysis using the commercial thermo-hydraulic program. Test and analysis variables were recess, pressure drop, velocity ratio, mixing spray, mixture ratio. Hot tests were performed for the Uni-element injector to compare the performance with the cold test results, and, hot tests for Multi-element injector were performed to compare the performance with Uni-element injector. Designed thrust of the Uni-element injector liquid rocket was 35kgf at sea level and combustion chamber pressure, 20bar. Kerosene and Lox were used as a propellant.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.32 no.7
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• pp.541-548
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• 2008

To improve the operational stability of the 100 Watts class Micro Gas Turbine, the air foil bearing with additional damping material has been investigated. The key of structure is that a viscoelastic material is coated under the top foil. The compliant foil journal bearing and thrust bearing are designed to withstand high load of vibrations at the operational speed 870,000 rpm. Test is executed in room temperature. Rotor has stably operated above 480,000 rpm. It is over 55% of the designed speed 870,000 rpm. Synchronous and subsynchronous vibrations are both well controlled. Vibration amplitude diminished over 50%. With the help of increased damping resulting from the viscoelasticity, the rotor stability of Micro turbocharger has been improved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1996.11a
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• pp.209-212
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• 1996

접촉발화성 액체추진제를 사용하는 소형 추진기관의 성능을 확인하기 위해 핵심 부품의 상온 성능 시험 및 추진기관 조립품의 연소성능시험을 실시하여 추력 400 파운드급의 로켓엔진의 설계성능을 확인하였다. 로켓엔진의 성능측정은 요소분사기의 미립화, 혼합성능의 측정을 위한 상온수류시험과 추진기관 조립체의 지상연소시험으로 실시하였으며, 연소성능은 산화제로 질산을 연료로는 아민계 혼합물로 구성된 접촉발화형 이원 액체추진제를 사용하여 가압방식에 의해 연소실에 분사하는 방법으로 수행하였다. 성능시험결과는 설계성능 및 이상성능과 비교함으로서 이들의 상관관계를 파악하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.289-292
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• 2008

Highly over-expanded nozzle of the rocket engines will be excited by non-axial forces due to flow separation at sea level operations. Since rocket engines are designed to produce axial thrust to power the vehicle, non-axial static and/or dynamic forces are not desirable. Several engine failures were attributed to the side loads. Present work investigate the unsteady flow in an over-expanded rocket nozzle in order to estimate side load during a shutdown/starting. Numerical computations has been carried out with density based solver on multi-block structured grid. Present solver is explicit in time and unsteady time step is calculated using dual time step approach. AUSMDV is considered as a numerical scheme for the flux calculations. One equation Spalart-Allmaras turbulence model is selected. Results presented here is for two nozzle pressure ratio i.e. 100 and 20. At 100 NPR, restricted shock separation (RSS) pattern is observed while, 20 NPR shows free shock separation (FSS) pattern. Side load is observed during the transition of separation pattern at different NPR.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.111-114
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• 2009

A catalyst ignitor of small thrust engine using hydrogen peroxide and kerosene was designed and fabricated, which confirmed mass flow rate for design pressure through the water cold-flow test in this study. In order to investigate ignition performance, it was changed that mixture ratio for kerosene mass flow rate in a position which heat of hydrogen peroxide decomposition comes to a steady state. And we confirmed stable ignition property in a wide range of mixture ratio.