• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2011년 춘계학술대회논문집
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• pp.414-419
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• 2011

Missile am fighter aircraft have been challenged by low restoring nose-down pitching moment at high angle of attach. The consequence of weak nose-down pitching moment can be resulting in a deep stall condition. Especially, the pressure oscillation has a huge effect on noise generation, structure damage, aerodynamic performance and safety, because the flow has strong unsteadiness at high angle of attack. In this paper, the unsteady aerodynamics coefficients were analyzed at high angle of attack up to 60 degrees around two dimensional NACA0012 airfoil. The two dimensional unsteady compressible Navier-Stokes equation with a LES turbulent model was calculated by OHOC (Optimized High-Order Compact) scheme. The flow conditions are Mach number of 0.3 and Reynolds number of $10^5$. The lift, drag, pressure distribution, etc. are analyzed according to the angle of attack. The results at a low angle of attack are compared with other results before a stall condition. From a certain high angle of attack, the strong vortex formed by the leading edge are flowing downstream as like Karman vortex around a circular cylinder. Unsteady velocity field, periodic vortex shedding, the unsteady pressure distribution on the airfoil surface, and the acoustic fields are analyzed. The effects of these unsteady characteristics in the aerodynamic coefficients are analyzed.

• 한국전산유체공학회지
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• 제15권3호
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• pp.1-8
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• 2010

Laminar separation bubble and transitional flow over the NACA0012 are investigated at a moderate range of Reynolds numbers. A Reynolds-Averaged Navier-Stokes code is coupled with an empirical transition model that can predict transition onset points and the length of transition region. Without solving the boundary layer equations, approximated e-N method is directly applied to the RANS code and iteratively solved together. The computational results are compared with the experimental data for the NACA0012 airfoil. Results of transition onset point and the length are compared well with experimental data and Xfoil prediction. The present RANS results show at high angles of attack better agreement with experimental data than Xfoil results using the boundary layer equations.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2000년도 추계학술대회논문집B
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• pp.568-572
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• 2000

A numerical investigation was performed to determine the effect of airfoil thickness on the optimum Gurney flap height using NACA 00XX series airfoils. Seven airfoils which have 3% chord thickness difference were used. These were NACA 0006, 0009, 0012, 0015, 0018, 0021, and 0024. A Navier-Stokes code, FLUENT, was used to calculate the flow field about airfoil. The fully turbulent results were obtained using the standard $k-{\varepsilon}$ two-equation turbulence model. To provide a check case fur our computational method, numerical studies for NACA 4412 airfoil were made and compared with already existing experimental data for this airfoil by Wadcock. For every NACA 00XX airfoil, Gurney flap heights ranging from 0.5% to 2.0% chord were changed by 0.5% chord interval and their effects were studied. With the numerical solutions, the relationship between $(L/D)_{max}$ and airfoil thickness as a function of flap height and the relationship between $(L/D)_{max}$ and flap height as a function of airfoil thickness were investigated. The same relationship for $(C_l)_{max}$ also were shown. From these results, the optimum flap size for each airfoil thickness can be determined and vice versa.

• 한국마린엔지니어링학회:학술대회논문집
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• 한국마린엔지니어링학회 2001년도 춘계학술대회 논문집
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• pp.58-62
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• 2001

A numerical investigation was performed to determine the effect of airfoil on the optimum flap. height using NACA 0006, 0009, 0012, 0015, 0018, 0021 and 0024 airfoils. The six flaps which have 0.5% chord height difference were used. A Navier-Stokes code, FLUENT, was used to calculate the flow field of the airfoil. The code was first tested as a benchmark by modelling flow around a NACA 4412 airfoil. Predictions of local pressure coefficients are found to be in good agreement with the result of the experimental result. For every NACA 00XX airfoil, flap heights ranging from 0.0% to 2.5% chord were changed by 0.5% chord interval and their effects were also studied. Representative results from each case are presented graphically and discussed. It is concluded that this initial approach gives a promise for the future development of wind turbine optimum design.

• EDISON SW 활용 경진대회 논문집
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• 제3회(2014년)
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• pp.562-567
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• 2014

본 연구에서는 교육 및 연구를 위한 CFD 해석 프로그램인 EDISON_CFD를 이용하여 Symmetric airfoil(NACA0012)과 Cambered airfoil(NACA4412)의 뒷전 두께에 따른 공력 특성을 분석해보았다. Chord 길이의 0%, 1%, 2%, 3%, 4%에 해당하는 뒷전 두께를 가지는 Blunt trailing edge airfoil의 받음각에 따른 공력 특성을 비교 및 분석하고, 어떠한 장단점을 가지는지 확인하였다. 그 결과 Chord 길이의 1% 뒷전 두께를 가질 때를 제외하면 뒷전 두께가 두꺼워질수록 최대양력계수는 증가하였고, 양항비와 실속각은 감소하였다. 또한, 뒷전 두께가 두꺼워질수록 Symmetric airfoil에서는 받음각 $0^{\circ}$를 기준으로 양력곡선기울기가 증가하였고, Cambered airfoil에서는 전체적으로 양력계수가 증가함과 동시에 양력곡선기울기 또한 증가하는 것을 확인할 수 있었다.

• 한국항공우주학회지
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• 제34권2호
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• pp.17-25
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• 2006

본 연구에서는 2차원 비정렬 중첩격자계를 이용한 서로 겹쳐진 물체간의 상대운동 해석기법을 개발하였다. 물체들이 교차되는 교차점에서 가장 가까운 거리에 있는 각 격자계의 물체 경계면에 위치한 격자점들을 교차점으로 이동시키고, 이 교차점을 기준으로 계산영역에 포함되는 격자와 계산에서 제외되는 격자가 구분되도록 하였다. 비정상 유동의 계산을 위해 계산에서 제외되는 비활성 격자점에 적절한 유동값을 부여하여, 상대운동이 진행됨에 따라 새롭게 활성 격자점으로 분류될 때 이전 시간에서의 유동값으로 이용될 수 있도록 하였다. 해석기법의 검증을 위해 단순플랩의 진동에 따른 에어포일 주위의 유동을 해석하여 타 연구자의 해석결과와 비교하였고, NACA0012 에어포일의 내부에서 사출되는 발사체의 운동에 대하여 해석을 수행하였다.

• International Journal of Aeronautical and Space Sciences
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• 제18권1호
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• pp.8-16
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• 2017

Most of small air vehicles with moving wing fly at low Reynolds number condition and the reduced frequency of the moving wing ranges from 0.0 to 1.0. The physical phenomena over the wing dramatically vary with the reduced frequency. This study examines experimentally the effect of the reduced frequency at low Reynolds number. The NACA0012 airfoil performs sinusoidal pitching motion with respect to the quarter chord with the four reduced frequencies of 0.1, 0.2, 0.4 and 0.76 at the Reynolds number $2.3{\times}10^4$. Smoke-wire flow visualization, unsteady surface pressure measurement, and unsteady force calculation are conducted. At the reduced frequency of 0.1 and 0.2, various boundary layer events such as reverse flow, discrete vortices, separation and reattachment change the amplitude and the rotation direction of the unsteady force hysteresis. However, the boundary layer events abruptly disappear at the reduced frequency of 0.4 and 0.76. Especially at the reduced frequency of 0.76, the local variation of the unsteady force with respect to the angle of attack completely vanishes. These results lead us to the conclusion that the unsteady aerodynamic characteristics of the reduced frequency of 0.2 and 0.4 are clearly distinguishable and the unsteady aerodynamic characteristics below the reduced frequency of 0.2 are governed by the boundary layer events.

• 한국항공우주학회지
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• 제34권1호
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• pp.18-23
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• 2006

본 연구에서는 Gurney 플랩이 달린 NACA 0012 익형에 대한 플러터 해석을 시간 영역에서 수행하였다. 2차원 비정상 압축성 Navier-Stokes 방정식과 Lagrange 방정식으로 부터 유도한 2계 자유도 plunge & pitch 모델을 지배방정식으로 하여 연성 결합 기법을 통해 플러터 해석을 수행하였다. 계산 결과 Gurney 플랩을 장착할 경우 NACA 0012에 비해 플러터가 발생하는 속도가 낮아졌고, 마하수가 0.85보다 작은 영역에서는 Gurney 플랩의 플러터 경계 곡선은 안전 여유를 상회하는 영역에 위치하였다. 그러나, 마하수가 0.85에서 0.9사이일 경우에는 안전 여유에 근접하게 되므로 이러한 운용 영역에서는 Gurney 플랩의 사용에 주의를 요한다.

• Wind and Structures
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• 제32권3호
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• pp.239-248
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• 2021

In this study, three airfoil families, NACA, FX and S, in each case three from each series with different shapes were investigated at different angles of attack using Computational Fluid Dynamics (CFD) method. To verify the CFD model, simulation results of the NACA 0012 airfoil was compared against the available experimental data and k-ω SST was used as the turbulence model. Lift coefficients, lift to drag ratios and pressure distributions around airfoils were obtained from the CFD simulations and compared each other. The simulations were performed at three Reynolds numbers, Re=2×105, 1×106and 2×106, and angle of attack was varied between -6 and 12 degrees. According to the results, similar lift coefficient values were obtained for symmetric airfoils reaching their maximum values at similar angles of attack. Maximum lift coefficients were obtained for FX 60-157 and S 4110 airfoils having lift coefficient values around 1.5 at Re=1×106 and 12 degrees of angle of attack. Flow separation occurred close to the leading edge of some airfoils at higher angles of attack, while some other airfoils were more successful in keeping the flow attached on the surface.

• 대한기계학회논문집B
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• 제20권7호
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• pp.2347-2355
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• 1996

Purpose of the present study is to develop a computational design program for shape optimization, combining the numerical optimization technique with the flow analysis code. The present methodology is then validated in three cases of aerodynamic shape optimization. In the numerical optimization, a feasible direction optimization algorithm and shape functions are considered. In the flow analysis, the Navier-Stokes equations are discretized by a cell-centered finite volume method, and Roe's flux difference splitting TVD scheme and ADI method are used. The developed design code is applied to a transonic channel flow over a bump, and an external flow over a NACA0012 airfoil to minimize the wave drag induced by shock waves. Also a separated subsonic flow over a NACA0024 airfoil is considered to determine a maximum allowable thickness of the airfoil without separation.