• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.86-94
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• 1999

Measured shock wave effects were investigated by changing shock strength and position with particular emphasis on the stability limits of hydrogen-air jet flames. For this purpose, a supersonic nonpremixed, jet-like flame was stabilized along the axis of a Mach 2.5 wind tunnel, and wedges were mounted on the sidewall in order to interact oblique shock waves with the flame. This experiment was the first reacting flow experiment interacting with shock waves. Schilieren visualization pictures, wall static pressures, and flame stability limits were measured and compared to corresponding flames without shock-flame interaction. Substantial improvements in the flame stability limits were achieved by properly interacting the shock waves with the flameholding recirculation zone. The reason for the significant improvement in flame stability limits is believed to be the adverse pressure gradient caused by the shock, which can elongate the recirculation zone.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.133-137
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• 2009

Numerical simulations of flows in an axisymmetric supersonic inlet with bleed regions were performed. For the simulations, the existing code which solves the RANS(Reynolds Averaged Navier-Stokes) equations and 2-equation turbulence model equations was transformed to axisymmetric form and bleed boundary condition was applied to the code. In this paper, the modified code was validated by comparing the results against an experimental data and other computational results for flow on a bump and over an oblique shock with bleed region. Using the code, numerical simulations were performed for the flow in the inlet with multiple bleed regions.

• Proceeding of EDISON Challenge
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• 2013.04a
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• pp.365-370
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• 2013

현대에 이르러 초음속 운영 영역에서의 항공기에 대한 많은 연구가 진행되고 있으나, 음속 폭음 현상과 충격파 현상에서 야기된 높은 항력 및 연료 효율성 저하로 인하여 그 한계에 부딪치고 있다. Busemann 복엽 익형은 이와 같은 문제를 해결하기 위한 형상이며, 상하 형상에 의한 파동 상쇄효과 및 파동 감소효과를 통해 충격파의 강도와 음속 폭음 효과를 감소시키는 형상이다. 하지만 본 익형은 탈설계 조건에서 항력 계수가 급격하게 증가하는 등의 문제점을 가지고 있다. 본 연구에서는 EDISON_CFD를 이용하여 Busemann 복엽 익형의 주변 유동 특성에 대하여 면밀한 고찰을 수행하였다. 우선 Busemann 복엽 익형의 초기 형상에 대한 유동 조건별 해석을 통하여 탈설계 조건에서의 항력 성능 저하 문제에 대한 고찰을 하였다. 이후 3개의 형상 변수에 대한 매개 변수 연구를 통하여 익형의 각 형상 변수가 탈설계 영역 및 해당 영역에서의 최대 항력 계수에 미치는 영향에 대한 고찰을 수행하였으며, 이를 통하여 기존 형상보다 좁은 탈설계 구간을 가지고, 최대 항력 계수가 약 34.8% 감소한 부스만 복엽 익형을 설계하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.61-61
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• 2002

스크램 제트 엔진의 설계에서 초음속으로 유입된 공기의 짧은 잔류시간으로 인한 연료-공기의 혼합은 가장 중요하며 해결하기 힘든 문제이다. 전헝적인 비행 조건에서 흡입 공기가 극초음속 비행기 엔진 내에서 잔류하는 시간의 단위는 1 ms 정도이어서 짧은 시간 동안 연료와 공기는 효율적으로 혼합되어야 하며, 최대의 추진력을 얻기 위하여 과도한 공력저항없이 연소 가능한 연료-공기 혼합기를 생성시킬 수 있는 효율적인 연료-공기의 혼합 방법이 요구된다. 현재까지 가장 많이 연구되어 온 혼합 방법은 엔진 입구로 들어오는 공기 유동에 수직 방향으로 연료를 분사하는 것으로 이 방법은 연료 유동 방향과 공기 유동 방향이 수직이기 때문에 추력 손실이 생기는 단점을 갖고 있지만, 초음속으로 유입되는 공기에 수직으로 연료를 분사하게되면 분사 위치 앞에 궁형 충격파가 생겨서 감속되어 유동이 회전하는 재순환영역이 생기고 연료의 혼합이 잘 이루어지는 장점이 있다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.331-336
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• 2006

This paper studies the effects of the nozzle-rotor axial clearance of a supersonic turbine on turbine performance. The nozzle-rotor axial clearance of the supersonic turbine developed to drive a turbopump for 30 ton class liquid rocket engines was varied and a numerical analysis of the turbines having the different nozzle-rotor axial clearances was conducted. It has been found that turbine performance degrades with an increasing axial clearance due to the increased stagnation pressure loss in the axial clearance region.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.332-335
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• 2008

The purpose of this study is analysis of flow field where is around of injector of supersonic combustor which is bluff-body stabilized flame and hyper-mixer type of supersonic combustor injector by using hydrogen or hydrocarbon fuel. Various schemes are evaluated to supersonic backward step flow filed with massive separation region in validation step. Compounded scheme of 5th-order TVD-MUSCL, Roe FDS, S-A DES/DDES has a good performance in base and base-bleed flow.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.67-74
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• 1999

Problems created by supersonic jet impinging on solid objects or ground arise in a variety of situations. For example multi-stage rocket separation, deep-space docking, V/STOL aircraft, jet-engine exhaust, gas-turbine blade, terrestrial rocket launch, and so on. These impinging jet flows generally contain a complex structures. (mixed subsonic and supersonic regions, interacting shocks and expansion waves, regions of turbulent shear layer) This paper describes experimental works on the phenomena (surface pressure distribution, flow visualization) when underexpanded supersonic jets impinge on the perpendicular, inclined plate using a supersonic cold-(low system. The used supersonic nozzle is convergent-divergent type, exit Mach number 2, The maximum on the plate when it was inclined was much larger than perpendicular plate, owing to high pressure recoveries through multiple shocks. Surface pressure distribution as to underexpanded ratio showed similar patterns together.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.1
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• pp.27-32
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• 2007

this paper the acoustic signature from a supersonic nozzle is measured and compared to the result of a program developed for a gas turbine noise prediction. In order to measure the jet Mach Number, the pressure and temperature at the settling chamber was measured along with pressures from a pitot-tube placed near the exit. The results are also compared to the ones obtained with a shadow graph technique. Jet noise produced by an imperfectly expanded jet contains shock associated noise, which consist of broadband noise and screech tone noise. For subsonic condition, the directivity is dominant to the downstream direction due to turbulence mixing noise. For supersonic conditions, however, the directivity is dominant toward upstream direction due to shock associated noise. The comparison with a jet exhaust noise prediction code shows good agreement at supersonic conditions but needs to be improved at subsonic speeds.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.29-38
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• 2015

In this study, 3-dimensional URANS simulation was performed to analyze the effect of the nozzle-rotor axial gap spacing of a supersonic impulse turbine on turbine performance. The computations were conducted for four different axial gap cases corresponding to about 6%, 10%, 20% and 30% of the blade height, respectively. The results show a good agreement with previous studies and the turbine efficiency decreases drastically in certain range. It is examined that the turbine performance characteristics could change depending on the influence of leading edge shock to the nozzle outlet. It is also found that the entropy rise distributions along the span differ from each other.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.2
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• pp.33-39
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• 2008

To improve the mixing efficiency and to reduce the pressure loss, it is needed to develope a new mixing device for supersonic combustion. The vent mixer is introduced as the new supersonic fuel-air mixer. Computational analyses, that include pressure profile, density contour, and streamline tracing, have been carried out. The expansion wave generates at the end of the extended upper wall of the mixer. And it reduces the shock wave from the hole. Incoming air flow through the hole makes several recirculation regions which increase the mixing efficiency, and the separation region at the downward wall expends the boundary layer which reduces the pressure loss.