• 한국소음진동공학회:학술대회논문집
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• 한국소음진동공학회 2002년도 춘계학술대회논문집
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• pp.140-146
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• 2002

The fluid induced vibration (FIV) phenomena of a 2-D.O.F airfoil system have been investigated in low Reynolds number incompressible flow region. Unsteady flows with viscosity are computed using two-dimensional incompressible Navier-stokes code. To validate developed Navier-Stokes code, steady and unsteady flow fields around airfoil are analyzed. The present fluid/structure interaction analysis is based on the most accurate computational approach with computational fluid dynamics (CSD) and computational structural dynamics (CSD) techniques. The highly nonlinear fluid/structure interaction phenomena due to severe flow separations have been analyzed fur the low Reynolds region (R$_{N}$ =500~5000) that has a dominancy of flow viscosity. The effect of R$_{N}$ on the fluid/structure coupled vibration instability of 2-DOF airfoil system is presented and the effect of initial angle of attack on the dynamic instability are also shown.own.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2008년 영문 학술대회
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• pp.671-679
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• 2008

The effect of Pre-cooled Turbojet Engine installation and nozzle exhaust jet on Hypersonic Turbojet EXperimental aircraft(HYTEX aircraft) were investigated by three-dimensional numerical analyses to obtain aerodynamic characteristics of the aircraft during its in-flight condition. First, simulations of wind tunnel experiment using small scale model of the aircraft with and without the rectangular duct reproducing engine was performed at M=5.1 condition in order to validate the calculation code. Here, good agreements with experimental data were obtained regarding centerline wall pressures on the aircraft and aerodynamic coefficients of forces and moments acting on the aircraft. Next, full scale integrated analysis of the aircraft and the engine were conducted for flight Mach numbers of M=5.0, 4.0, 3.5, 3.0, and 2.0. Increasing the angle of attack $\alpha$ of the aircraft in M=5.0 flight increased the mass flow rate of the air captured at the intake due to pre-compression effect of the nose shockwave, also increasing the thrust obtained at the engine plug nozzle. Sufficient thrust for acceleration were obtained at $\alpha=3$ and 5 degrees. Increase of flight Mach number at $\alpha=0$ degrees resulted in decrease of mass flow rate captured at the engine intake, and thus decrease in thrust at the nozzle. The thrust was sufficient for acceleration at M=3.5 and lower cases. Lift force on the aircraft was increased by the integration of engine on the aircraft for all varying angles of attack or flight Mach numbers. However, the slope of lift increase when increasing flight Mach number showed decrease as flight Mach number reach to M=5.0, due to the separation shockwave at the upper surface of the aircraft. Pitch moment of the aircraft was not affected by the installation of the engines for all angles of attack at M=5.0 condition. In low Mach number cases at $\alpha=0$ degrees, installation of the engines increased the pitch moment compared to no engine configuration. Installation of the engines increased the frictional drag on the aircraft, and its percentage to the total drag ranged between 30-50% for varying angle of attack in M=5.0 flight.

• 한국전산유체공학회지
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• 제16권4호
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• pp.14-20
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• 2011

The initial unfolding motion simulation of a sub-munition with drag ribbon for precision guidance and reliable operation has been investigated by analyzing its unsteady aerodynamic load and fluid structure interaction. The effects of change in the ribbon configuration and flow angle are numerically studied using a commercial software "XFLOW" based on Lattice-Boltzmann Method. It is shown that the motion is affect adversely by the separation bubble formed posterior part of the fuselage. The rolling moment for arming of the sub-munition is increased with angle of attack and rotational movement.

• 한국군사과학기술학회지
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• 제15권1호
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• pp.101-110
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• 2012

The side jet effects between jet flow and free-stream on a missile body were investigated by experimentally and numerically for modeling aerodynamic coefficients in pitch plane. K-factors for normal force and pitching moment were introduced to estimate the side jet effects. The main parameters of the jet interaction phenomena were angle of attack, jet pressure ratio, Mach number and jet bank angle. The K-factors for normal force coefficient and pitching moment coefficients in pitch plane were analysed.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2002년도 춘계 학술대회논문집
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• pp.88-93
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• 2002

In the present research, a parametric study of aerodynamic characteristics for multi-element airfoils is performed. The major geometric parameters of interest are the gap distance between airfoils and relative deflection angle of slat/flap. The present results are mainly obtained by using inviscid flow calculation, and the aerodynamic characteristics are focused on the surface pressure distribution and the lifts. The results of the present research may be used as not only qualitative data but also quantitative data for small angle of attack flows, where the viscous effect does not play major role in terms of surface pressure distribution and lifts. A further research in this subject including viscous calculation and more geometric parameters is to be performed in the future.

• 한국항공우주학회지
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• 제34권8호
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• pp.8-15
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• 2006

고받음각의 NACA23012익형에 대하여 synthetic jet을 이용하여 박리 제어를 수행하였다. 단일 synthetic jet을 이용하여 익형에 발생하는 앞전 박리를 효과적으로 지연시킬 수 있고, 또한 실속 특성을 개선 할 수 있음을 확인하였다. 그때 발생하는 비정상 유동 특성을 파악하였다. 또한, 현실적으로 구현 가능한 jet 속도를 얻기 위하여 multi-array synthetic jet의 특성을 파악하였다. 그리고, 단일 위치에 장착된 synthetic jet을 이용하여 박리를 제어 하였을 경우 익형 윗면에 발생하는 작은 와동을 제거하기 위하여, multi-location synthetic jet을 이용하였다. 작은 와동을 제거하고 안정적인 유동을 확보하기 위하여, 높은 진동수의 synthetic jet을 이용하여 국부적으로 효과적인 박리 제어를 통한 익형 주변의 유동의 전체적 특성을 안정화 시킬 수 있음을 확인하였다. Multi-location synthetic jet의 phase 변화를 이용하여 multi-array/multi-location synthetic jet의 성능 및 특성을 향상 시킬 수 있음을 확인하였다.

• 대한기계학회논문집B
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• 제32권4호
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• pp.300-306
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• 2008

In the present study, the second order response surface method (RSM) is carried out to get optimum thermal design for enhancing heat transfer in a square channel with bleed flow. The RSM is used as an optimization technique. To calculate the heat transfer, RNG k-epsilon model and enhanced wall function are used. To design optimum rib turbulators, two design variables such as attack angle of rib $({\alpha})$ and rib pitch-to-rib height ratio (p/e) are optimized. In these analyses, the channel inlet Reynolds number was fixed at 10,000 in both non-bleeding and bleeding cases. The response surfaces of two design variables are constructed in cases with and without bleed flow. As a result, the optimum (or highest) heat transfer values are almost the same in ranges of two cases with and without bleed flow. However, the friction losses in the case with bleed flow are lower than those without bleed flow.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2006년도 PARALLEL CFD 2006
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• pp.94-99
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• 2006

As part of a Department of Defense Grand Challenge Project, advanced high performance computing (HPC) time-accurate computational fluid dynamics (CFD) techniques have been developed and applied to a new area of aerodynamic research on microjets for control of small and medium caliber projectiles. This paper describes a computational study undertaken to determine the aerodynamic effect of flow control in the afterbody regions of spin-stabilyzed projectiles at subsonic and low transonic speeds using an advanced scalable unstructured flow solver in various parallel computers such as the IBM SP4 and Linux Cluster. High efficiency is achieved for both steady and time-accurate unsteady flow field simulations using advanced scalable Navier-Stokes computational techniques. Results relating to the code's portability and its performance on the Linux clusters are also addressed. Numerical simulations with the unsteady microjets show the jets to substantially alter the flow field both near the jet and the base region of the projectile that in turn affects the forces and moments even at zero degree angle of attack. The results have shown the potential of HPC CFD simulations on parallel machines to provide to provide insight into the jet interaction flow fields leading to improve designs.

• Journal of Advanced Marine Engineering and Technology
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• 제23권4호
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• pp.462-472
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• 1999

De-NOx facility using Selective Catalytic Reduction method is the most widely applied one that removes NOx from flue gas emitted from combustion facility such as boiler for power generation engine incinerator etc. Reductant $NH_3\;or\;NH_4OH$ is sprayed into flue gas to convert NOx into $H_2O$ and $N_2.$ Good mixing between flue gas and $NH_3$ is the most important factor to increase reduction in catalytic layer and to reduce unreacted NH3 slip. Therefore the development of mixer device for mixing effect is one of the important part for SCR facility. Objectives of this study are to investigate the relation between flow and concentration field by observation at the wake of delta-wing type mixer. At the first stage qualitative measurement of flow field is conducted by flow visualization using laser light sheet in lab. scale wind tunnel. Also we have conducted the quantitative analysis by comparing flow field measurement using LDV with numerical simulation. On the basis of qualitative and quantitative analysis we investigate the dis-tribution of flow and concentration in flow model facility. The results of an experimental and compu-tational examination of the vortex structures shed from delta wing type vortex generator having $40^{\circ}$ angle of attack are presented, The effects of vortex structure on the gas mixing is discussed, too.

• 한국가시화정보학회지
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• 제18권3호
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• pp.69-76
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• 2020

In the present study, we investigate the flow separation characteristics of a paraglider canopy model by tuft visualization. The experiment is conducted at Re = 3.3×105 in a wind tunnel large enough to contain the three-dimensional paraglider canopy model, where Re is Reynolds number based on the mean chord length and the free-stream velocity. The flow separation characteristics of the canopy model near the wing root are similar to those of a two-dimensional airfoil with a cross-section similar to the model. On the other hand, near the wingtip region, the flow separation is suppressed by the downwash induced by the wingtip vortex. As a result, as the angle of attack increases, the flow separation occurs from the wing root region of the canopy model and develops toward the wingtip.