• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.73-80
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• 2018

At the early stages of development of high-speed propulsion systems, associated uncertainties cannot be easily modeled into probabilistic distributions, owing to the lack of test data, cost, and difficulty of simulating real-flight environments on the ground. To tackle this issue, in this research, the combustion efficiencies of dual-combustion ramjet engines are assumed to have been provided by experts and quantified by evidence theory. Using quantified uncertainty, the inlet area and combustor exit are optimized while satisfying reliability margins of thrust and thermal choking. The result shows a reasonable design of the engine under uncertain circumstances.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.804-809
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• 2008

In this study, the cooling performance of a C/C composite material structure with metallic cooling tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2900 K and heat flux to the combustion chamber wall was about 9 $MW/m^2$. No thermal damage was observed on the stainless steel tubes which were in contact with the C/C composite materials. Results of the heating test showed that such a metallic-tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure(also as a heat exchanger), as well as indicating the possibility of reducing the amount of the coolant even if the thermal load to the engine is high. Thus, application of the metallic-tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined cycle engine is expected.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.14-14
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• 1998

램제트 엔진은 초음속에서 공기가 충격파를 통해 아음속으로 속도가 낮아지고 압력이 증가하는 램 압축 현상을 이용하되 압축기를 사용하지 않고 아음속 상태에서 연소하는 구조로 되어 있다. 따라서 각 부품의 성능은 독립적이지 않으며 전체적인 성능을 규명하기 위해서는 공기 흡입구와 연소실, 연료 분사체계 등의 상호작용을 고려하여 유동의 특성과 그에 따른 연소현상의 특성을 함께 고려해야만 한다. 본 연구에서는 속도 범위 Mach 5이내, 고도 30km이내의 운항조건을 갖는 유도 무기에 장착될 램제트 엔진 개발을 위한 연구의 첫 번째 단계로써, 램제트 연소실내의 유동장해석을 실험과 수치해석 두 분야로 나누어 수행하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.383-389
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• 2010

The vapor concentration of hexane in a liquid spray jet in crossflow was qualitatively measured on the basis of the infrared (IR) extinction techniques. The objectives of the present study are to understand the whole evaporation process from droplet breakup to vapor and to confirm the usefulness of IR emission method in a lab-scale ramjet combustor. From the experimental results, we concluded that hexane vapor mole fraction increased with temperature rise and kept nearly constant during the variation of fuel to air momentum ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.4
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• pp.25-31
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• 2010

The vapor concentration of hexane in a liquid spray jet in crossflow was qualitatively measured on the basis of the infrared (IR) extinction techniques. The objectives of the present study are to understand the whole evaporation process from droplet breakup to vapor and to confirm the usefulness of IR emission method in a lab-scale ramjet combustor. From the experimental results, we concluded that hexane vapor mole fraction increased with temperature rise and kept nearly constant during the variation of fuel to air momentum ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.713-721
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• 2004

Variable air intake and variable exhaust nozzle of hypersonic engines are designed and tested in this study. Dimensions for variable geometry air intake, ram combustor and variable geometry exhaust nozzle are defined based on the requirements of a pre-cooled turbojet engine. Hypersonic Ramjet Engine is designed as a scaled test bed for each component. Actuation forces of moving parts for variable intake and variable nozzle are reduced by balancing the other force in the opposite direction. A demonstrator engine which includes variable intake and variable nozzle is designed and the components are fabricated. Composite material with silicone carbide is applied for high temperature parts under oxidation environment such as leading edge of the variable intake and combustor liner. Internal cooling structure is adopted for both moving and static parts of the variable nozzle. Pressure recovery and mass capture ratio of the variable intake at Mach 5 is obtained by a hypersonic wind tunnel test. Flow characteristics of the variable nozzle are obtained by a low temperature flow test. Wall temperature and heat flux of the nozzle at Mach 3 is obtained by a firing test. As results, the intake and the nozzle are proved to be used at designed pressure and temperature environment.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.579-583
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• 2011

The component based propulsion modeling and simulation of an air-breathing engine such as ramjet and scramjet is studied. The simulation model has been realized considering the characteristics of the air-breathing engine which is composed of air intake, combustor and nozzle including engine controller and fuel supply system. To estimate the engine performance and to verify the engine controller, real time based Hardware in the Loop System simulating actual environment is constructed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.43-46
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• 2006

Storage air heater(SAH) is a general purpose facility that is used to simulate the high altitude condition of supersonic ground test facility, thurst compensation test of rocket engine nozzle and gas turbine engine combustor test. SAH in KARI is built to simulate the total temperature of the supersonic ground test facility which has a wide flight envelope from altitude 0km, Mach 2 to altitude 25km, Mach 5 and operates up to 1300K, 3.5MPa. In this paper, we introduces the SAH in JAXA which is model of SAH in KARI and summarizes the design process and manufacture of ours.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.1-1
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• 2005

Japan Aerospace Exploration Agency(JAXA) has been conducting research and development of the Scramjet engines and their derivative combined cycle engines as hypersonic propulsion system for space access. Its history will be introduced first, and its recent advances, focusing on the engine performance progress, will follow. Finally, future plans for a flight test of scramjet and ground test of combined cycle engine will be introduced. Two types of test facilities for testing those hypersonic engines. namely, the 'Ramjet Engine Test Facility (RJTF)' and the 'High Enthalpy Shock Tunnel (HIEST)' were designed and fabricated during 1988 through 1996. These facilities can test engines under simulated flight Mach numbers up to 8 for the former, whereas beyond 8 for the latter, respectively. Several types of hydrogen-fueled scramjet engines have been designed, fabricated and tested under flight conditions of Mach 4, 6 and 8 in the RJTF since 1996. Initial test results showed that the thrust was insufficient because of occurrence of flow separation caused by combustion in the engines. These difficulty was later eliminated by boundary-layer bleeding and staged fuel injection. Their results were compared with theory to quantify achieved engine performances. The performances with regards to combustion, net thrust are discussed. We have reached the stage where positive net thrust can be attained for all the test coditions. Results of these engine tests will be discussed. We are also intensively attempting the improvement of thrust performance at high speed condition of Mach 8 to 15 in High Enthalpy Shock Tunnel (HIEST). Critical issues for this purposemay be air/fuel mixing enhancement, and temperature control of combustion gas to avoid thermal dissociation. To overcome these issues we developed the Hypermixier engine which applies stream-wise vortices for mixing enhancement, and the M12-engines which optimizes combustor entrance temperature. Moreover, we are going to conduct the flight experiment of the Hypermixer engine by utilizing flight test infrastructure (HyShot) provided by the University of Queensland in fall of 2005 for comparison with the HIEST result. The plan of the flight experiment is also presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.432-437
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• 2010

A new device injecting secondary fuel behind flameholder was invented and tested in order to reduce low frequency combustion instability of combustor using V-gutter flameholder. Specially designed combustion device could make large combustion instability up to 180 dB successfully, and newly invented device made a success to reduce 110~120Hz low frequency pressure pulsation up to 84%. It was found that the fuel flow rate of secondary fuel supplying behind flameholder was the only parameter which dominates reduction of instability. It is considered that stabilized flame with sufficient secondary fuel can lead to break the connection between combustion system and acoustic system due to independence of flame from fluctuation of main fuel resulted from synchronization with acoustic wave.