• 한국전산유체공학회:학술대회논문집
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• 2002.05a
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• pp.109-114
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• 2002

The effect of nose radius on aerodynamic heating are investigated by using the Wavier-Stokes code extended to thermochemical nonequilibrium airflow. A spherical blunt body, whose radius varies from 0.003048 m to 0.6096 m, flying at Mach 25 at an altitude of 53.34 km is considered. Comparison of heat flux at stagnation point with the solution of Viscous Shock Layer and Fay-Riddell are made. Obtained result reveals that the flow chemistry for very small radius is nearly frozen, and therefore the contribution of heat flux due to chemical diffusion is smaller than that of translational energy. As the radius becomes larger, the portion of diffusion heat flux becomes greater than translational heat flux and approaches to a constant value.

• 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.5 no.1
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• pp.10-17
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• 2001

In this paper are presented main power and gasdynamic characteristics of hypersonic test cell of Research Test Center (RTC) of Central Institute of Aviation Motors (CIAM). The distributions of temperature and Mach number at the exit of the aerodynamic nozzle of test cell are received at simulation conditions of flight at Mf=6. Values of available pressure difference and throttling characteristics for various operational modes of test cell, including the loading of working section by Scramjet model without the heating of air at entrance to the aerodynamic nozzle and with the heating of air, are received too.

• Proceeding of EDISON Challenge
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• 2016.11a
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• pp.49-54
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• 2016

본 연구에서는 유/무인 우주발사체의 재진입모듈에 적용되는 Heat-Shield(열차단막)의 형상에 따른 유동특성을 수치해석으로 분석했다. 재진입모듈이 지구의 대기권을 다시 진입하는 환경(고도 70km)을 기준으로 해석을 진행했다. 열차단막의 형상은 평판, 곡률이 다른 타원으로 나누었으며, 각 형상별 유동특성을 확인했다. 결론적으로 재진입 모듈의 형상에 따라 압력과 속도분포를 기준으로 열적인 분포를 예상했으며, 계산된 항력계수를 비교했다. 단순한 유동으로 열적 분포를 예상한 것에 한계가 있지만 대기권 재진입 모듈의 2차원 설계에 도움이 될 형상기준을 제시했다. 수치해석은 모두 Edison_CFD에서 제공하는 툴을 이용해 수행했으며, 전처리에는 e-mega (structured)을 후처리에는 e-dava를 이용했다. 해석 solver는 '정렬격자기반 2차원 압축성 유동 범용해석 소프트웨어'를 이용했다.

• Journal of ILASS-Korea
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• v.26 no.4
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• pp.171-181
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• 2021

This study presents a prediction methodology of thermodynamic properties by using RK-PR Equation of State in a wide range of temperature and pressure conditions including both sub-critical and super-critical regions, in order to obtain thermophysical properties for hydrocarbon aviation fuels and their products resulting from endothermic reactions. The density and the constant pressure specific heat are predicted in the temperature range from 300 to 1000 K and the pressure from 0.1 to 5.0 MPa, which includes all of the liquid and gas phases and the super-critical region of three representative hydrocarbon fuels, and then compared with those data obtained from the NIST database. Results show that the averaged relative deviations of both predicted density and constant pressure specific heat are below 5% in the specified temperature and pressure conditions, and the major sources of the errors are observed near the saturation line and the critical point of each fuel.

• Journal of Industrial Technology
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• v.43 no.1
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• pp.33-41
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• 2023

This paper provides a general overview of surface catalytic recombination in hypersonic flow. The surface catalytic recombination phenomena is elaborated in terms of its general overview and numerical modeling associated with it. The general overview of the surface catalytic recombination phenomena describes the elementary surface reactions for the surface catalytic and the role of the surface catalytic recombination efficiency in the heat transfer determination. In the numerical modeling, the surface catalytic recombination is described based on the stagnation-point boundary layer analysis, and finite-rate surface reaction modeling. Throughout this overview manuscript, a general understanding of this phenomena is obtained and can be used as foundation for deeper application with the numerical computational fluid dynamics (CFD) flow solver to estimate the surface heat transfer in the hypersonic vehicles.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.6
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• pp.83-90
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• 2016

Ramjet mode combustion test was performed for a dual-mode ramjet engine model. The engine model consists of an air intake, a combustor and a nozzle. The combustor in the model has a large backward-facing step, designed to be used as a part of a rocket-based combined cycle engine. The test was performed at the flight speed of Mach 5 and the altitude of 24 km. Strong combustion was established only when the fuel was injected from both of the bottom-side and cowl-side wall. When the total fuel stoichiometric ratio was 1.0, distributed as 0.5 on the cowl side and 0.5 on the bottom side, the flow became subsonic at some portion in the combustor by thermal choking, i.e., ramjet mode was established for this condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.330-335
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• 2017

Recently many studies have been made for the development of propulsion system with wide range flight from supersonic to hypersonic. Dual Mode scramjet engine as a hybrid cycle with advantage of ramjet and scramjet has one combustor. It works under the ramjet mode (subsonic combustion) and scramjet mode (supersonic combustion) respectively. In this study, Experimental results of hot firing tests of dual scramjet engine designed on the condition of Mach 3.5~6 as a flight Mach number are discussed. The tests were carried out on a ground test bench under free stream condition of Mach 6 at 27.6km altitude. In the tests, the adopted design and technological solutions were verified and efficient operation of the dual mode ramjet engine with Kerosene combustion during 5 seconds was demonstrated.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.33-41
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• 2015

Scaled model of a dual-mode ramjet engine with large backward-facing step, as a component of the rocket-based combined cycle engine, was designed. Design parameters were derived for this engine with the consideration of application for the rocket-based combined cycle engine. Design methodology was established for these design parameters. The design was partially verified through numerical study. Flow characteristics of the dual-mode ramjet engine with large backward-facing step was investigated experimentally. The design methodology for relevant design parameters established in this study was verified as feasible.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.394-398
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• 2005

Ramjet engine have been usually operated on Mach $1.5\sim3$ as the vehicle of supersonic cruising engine and studied about the higher performance above Mach 4. The research of Duel mode Scramjet engine which have duel operating mode of ramjet/Scramjet are in progress actively nowadays. This paper suggests the effect the flow characteristics and the effects of back pressure, angle of attack, angle of yow on the supersonic air intake on mach 4 through the Schlieren/Oil flow visualization, and pressure measurement on experimental model.