• 한국연소학회:학술대회논문집
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• 2001.06a
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• pp.127-132
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• 2001

A numerical study was carried out to investigate the combustion phenomena in a model Scramjet engine, which had been experimentally studied in the University of Tokyo using a high-enthalpy supersonic wind tunnel. The main airflow was 2.0 in Mach number and the total temperature of hot flow was 1800K. Equivalence ratio was set to be rather higher value of 0.26 than that of experiment to investigate the effect of strong precombustion shock. The results showed that self-ignition occurred at the rear bottom wall of the combustor and combined with the shear layer flame between fuel jet and main airflow. Then, precombustion shock was generated at the step location and reversely enhanced the mixing and combustion process behind the shock. Due to the high equivalence ratio, the precombustion shock moved upstream of the step compared with that of experiment.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.11a
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• pp.17-17
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• 1997

극초음속 여객기와 군사용 항공기에 대한 수요가 증가함에 따라서 새로운 개념의 다양한 추진기관이 연구가 진행되고 개발되어 왔다. 초음속 항공기의 속도 영역은 마하 10-20 정도가 되는데 이 속도 한계를 극복하기 위하여 초음속 연소 램제트 엔진(SCRamjet; Supersonic Combustion Ramjet)이 제안되었다. 스크램 제트를 개발하기 위해서는 연료와 산화제의 혼합 효율 문제, 화염의 안정화 문제, 벽면의 냉각에 관한 문제 등 몇 가지 기본적인 문제들을 해결해야 한다. Univ of Michigan에서 실험한 연소기를 모델로 본 연구에서는 연료와 공기의 혼합에 관한 수치 연구를 수행하였다. 다원 혼합기체에 관한 축대칭 Navier-Stokes 방정식을 지배 방정식을 이용하였고 비평형 화학반응식을 고려하였다. 공간 차분에는 유한 체적법을 이용하였다. 대류 플럭스 항은 Roe의 Upwind FDS 기법을 사용하여 차분하였고 점성항에는 중심 차분법을 이용하였다. 시간 적분법으로는 근사 자코비안과 LU분할 기법을 이용한 완전 내재적 방법이 쓰였다. 난류 모델로는 Mentor에 의해 제안된 2 방정식 k-$\varepsilon$/k-$\omega$ 혼합모델을 사용하였다. 유동장이 실험에서의 찍은 사진과 유사한 모습의 충격파 간섭을 수치 모사하였고 수소가 확산되는 모습과 함께 노즐 lip 주위의 재순환 영역에 대해서 살펴볼 수 있었다.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.1
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• pp.57-63
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• 2008

In combustor of SCRamjet of air-breathing engine type, the flow duration time is very short because of the supersonic air flow. In this short duration, the whole process of combustion should be done, so it is very important to study supersonic combustion technologies. In this study, we focus fuel-air mixing enhancement method using cavity and conducted 3-dimensional Navier-Stokes computational analysis. Cavity height is fixed by 10mm, length is changed from 0 to 40mm. There is a supersonic jet injection downstream of the cavity and the hole size is 1mm. As a result, the higher ratio of cavity length/height is, the higher value of vorticity gets. The increased area of vorticity expands to upper and sidewise combustor. However, the stagnation pressure loss which generates thrust loss becomes higher when the vorticity is higher. Considering these result, we can conclude that optimized design which considers the highest mixing performance and the least stagnation pressure loss is needed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.321-324
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• 2011

A computational analysis of nozzle flow characteristics and plume structure using Reynolds-averaged Navier-Stokes equations with $k-{\omega}$ SST turbulence model was conducted to examine performance of the supersonic nozzle employed in a small liquid-rocket engine for ground firing test. Computed results and experimental outcome of 2-D converging-diverging nozzle flow were compared for verifying the computational capability as well as the turbulence model validity. Numerical computations of 2-D axisymmetric nozzle flow was carried out with the selected model. As a result, flow separation with backflow appeared around the nozzle exit. This investigation was reported as a background data for the optimal nozzle design of small liquid-propellant rocket engine for ground test.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.4
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• pp.9-15
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• 2009

The flow of combustor in scramjet engine has supersonic speed so that the residence time and mixing ratio are very important factors for the efficient combustion. This study used open cavity(L/D=4.8) as a fuel/air mixing model. Laser schlieren visualization and pressure measurement were carried out to observe the flow characteristics around a jet orifice and a cavity at the time of fuel injection. As a result of 10ns laser schlieren, unsteady flow which was around the cavity could be observed effectively. Pressure was measured that momentum flux ratio(J) was changed. And the change of critical ignition point could be observed by the momentum flux ratio changed.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.65-71
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• 1999

An experimental study was carried out in order to investigate the effect of shock waves on the supersonic hydrogen-air jet flames stabilized in the Mach 2.5 model scramjet combustor. This experiment was the first reacting flow experiment interacting with shock waves. Two identical $10^{\cire}$ wedges were mounted on the diverging sidewalls of the combustor in order to produce oblique shock waves that interacted with the flame. Schlieren visualization pictures, wall static pressures, and combustion efficiency at two different air stagnation temperatures were measured and compared to corresponding flames without shock wave-flame interaction. It was observed that shock waves significantly altered the shape of supersonic jet flames, but had different effects on combustion efficiency depending on air temperatures. At the higher air stagnation temperature and higher fuel flow rates, combustion of efficiency showed a better result.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.232-235
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• 2017

Scramjet combustor materials are exposed at ultra high temperature over 2000K and severe erosion environment. Inconel alloys are usually applied for combustor material however its mechanical properties are decreased beyond temperature of 1000K so that is impossible for long term operation and reuse. In this study, fiber reinforced ceramic material was used as scramjet combustor material and its feasibility studied. To increase combustion efficiency, regenerative combustor system developed and channel fabrication in composite material also studied.

• 한국연소학회:학술대회논문집
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• 2006.10a
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• pp.209-219
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• 2006

The supersonic combustion experiments are carried out using T3 free-piston shock tunnel. Different shock tube fill pressures have various inflow conditions. $15^{\circ}$ inclined hydrogen fuel injection is located before the cavity. Oblique shock is generated at the trailing edge of the cavity and reflects off the top and bottom wall. For non-reacting flow, static pressures in low equivalence ratio are similar to those in no fuel injection. As equivalence ratio is increased, static pressures are increased in the duct. In the similar equivalence ratio, static pressures are increased when total enthalpy is decreased. For reacting flow, the flame is occurred near the cavity. The combustion is weak locally in the middle of the duct. The up and down pressure distribution in the duct means that the supersonic combustion is generated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.6
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• pp.61-68
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• 2002

Combustion characteristics of dual transverse injection of hydrogen in supersonic air flows were studied using computational methods. Three-dimensional Navier-Stokes with a non-equilibrium chemical reaction model and the k-$\omega$ SST turbulence model were used. A parametric study was conducted with the variation of the distance between two injectors. Combustion characteristics of dual injection are very different from those of single injection. The combustion characteristics of two injection flows are very different from each other, and the ignition and combustion characteristics of the rear injection flow are strongly influenced by those of the front injection flow. The increase of the distance between two injectors up to a specific distance results in the increase of burning rate. However, the increase of the distance over the specific distance gives no increase of burning rate but makes more losses of stagnation pressure. From the results it can be stated that there exists a distance between two injectors for optimum combustion characteristics.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.484-488
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• 2012

Korea Aerospace Research Institute (KARI) performed the conceptual design of rocket engine test facility for the development and qualification of the 3rd stage liquid rocket engine for KSLV-II. The 3rd stage rocket engine test facility, which are to be constructed at Naro Space Center, will supply propellants and high-pressure gases to engine for firing test at ground and altitude conditions. The altitude test condition is obtained using a supersonic diffuser operated by the self-ejecting jet from the liquid rocket engine.