• EDISON SW 활용 경진대회 논문집
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• 제5회(2016년)
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• pp.647-652
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• 2016

To reduce drag force at supersonic speeds, sharp leading edge is hugely efficient. It is, however, incompatible with leading edge shape to have fine aerodynamic characteristics at subsonic and transonic speeds. It is critical to reduce drag force for enhanced cruise performance and higher efficiency. An air superiority fighter, however, required to have high maneuverability for survivability, and sharp leading edge is not proper. Consequently, variable leading edge is demanded to reduce drag force significantly at supersonic speeds for cruise performance. Leading edge altering system is constructed with rigid material to improve possibility of realization, and minimized movement of its components in altering for reduce effects on flight. It is compared with bi-convex airfoil and NACA 65-006 airfoil, which have comparable maximum thickness. At Mach number 1.7 and zero angle of attack, supersonic mode of designed airfoil indicates approximately 17% higher drag coefficient than the bi-convex airfoil indicates, it is, however, 23% lower than the NACA 65-006 indicates. Also, subsonic mode of the designed airfoil shows fine aerodynamic characteristics in comparison with NACA 65-006 airfoil in subsonic and transonic speed range. In this regard, design of the airfoil achieved the object of this study satisfactorily.

• International Journal of Naval Architecture and Ocean Engineering
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• 제1권1호
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• pp.1-12
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• 2009

Wing-in-Ground vehicles and aerodynamically assisted boats take advantage of increased lift and reduced drag of wing sections in the ground proximity. At relatively low speeds or heavy payloads of these craft, a flap at the wing trailing-edge can be applied to boost the aerodynamic lift. The influence of a flap on the two-dimensional NACA 4412 airfoil in viscous ground-effect flow is numerically investigated in this study. The computational method consists of a steady-state, incompressible, finite volume method utilizing the Spalart-Allmaras turbulence model. Grid generation and solution of the Navier-Stokes equations are completed using computer program Fluent. The code is validated against published experimental and numerical results of unbounded flow with a flap, as well as ground-effect motion without a flap. Aerodynamic forces are calculated, and the effects of angle of attack, Reynolds number, ground height, and flap deflection are presented for a split and plain flap. Changes in the flow introduced with the flap addition are also discussed. Overall, the use of a flap on wings with small attack angles is found to be beneficial for small flap deflections up to 5% of the chord, where the contribution of lift augmentation exceeds the drag increase, yielding an augmented lift-to-drag ratio.

• 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.

• 한국항공우주학회지
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• 제31권8호
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• pp.8-18
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• 2003

진동하는 에어포일의 근접후류에서의 레이놀즈수 효과를 조사하기 위한 실험적 연구가 수행되었다. NACA 4412에어포일은 1/4 시위 지점을 중심으로 조화적으로 피칭운동을 하고, 순간받음각은 +6$^{\circ}$에서 -6$^{\circ}$까지 진동되도록 하였다. 진동하는 에어포일의 근접후류에서의 난류강도를 측정하기 위하여 열선풍속계를 사용하였다. 본 연구에서 자유류의 속도는 3.4, 12.4, 26.2 m/s이다. 이러한 자유류 속도에 따른 시위 레이놀즈수는 $R_N=5.3{\times}10^4$, $1.9{\times}10^5$, $4.1{\times}10^5$이고, 무차원 진동수는 K=0.1이다. 레이놀즈수가 진동하는 에어포일의 근접후류에 미치는 영향을 나타내는 축방향 난류강도 분포를 제시하였다. 본 측정에서 모든 경우에 난류 강도는 $R_N=5.3{\times}10^4$인 경우에 아주 크고, $R_N=1.9{\times}10^5$과 $4.1{\times}10^5$인 경우에는 작다는 것을 관찰할 수 있었다. 진동하는 에어포일의 근접후류에서 레이놀즈수의 임계값은 층류분리인 경우, 분리가 발생하지 않거나 난류분리인 경우로 구분되며, 그 값은 $R_N=5.3{\times}10^4\;{\sim}\;1.9{\times}10^5$사이에 존재한다.

• 한국태양에너지학회 논문집
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• 제26권2호
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• pp.1-7
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• 2006

A numerical investigation was performed to determine the effect of airfoil on the optimum flap height using NACA0015 Wells turbine. The five double flaps which have 0.5% difference were selected. A Navier-Stokes code, CFX-TASCflow, was used to calculate the flow field of the Wells turbine. The basic feature of the Wells turbine is that even though the cyclic airflow produces oscillating axial forces on the airfoil blades, the tangential force on the rotor is always in the same direction. Geometry used to define the three dimension numerical grid is based upon that of an experimental test rig. This paper tries to disign the double flap of Wells turbine with the numerical analysis.

• International Journal of Naval Architecture and Ocean Engineering
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• 제6권2호
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• pp.282-296
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• 2014

In this paper, an iterative boundary element method (IBEM) was proposed to solve for a wing-in-ground (WIG) effect. IBEM is a fast and accurate method used in many different fields of engineering and in this work; it is applied to a fluid flow problem assessing a wing in ground proximity. The theory and the developed code are validated first with other methods and the obtained results with the proposed method are found to be encouraging. Then, time consumptions of the direct and iterative methods were contrasted to evaluate the efficiency of IBEM. It is found out that IBEM dominates direct BEM in terms of time consumption in all trials. The iterative method seems very useful for quick assessment of a wing in ground proximity condition. After all, a NACA6409 wing section in ground vicinity is solved with IBEM to evaluate the WIG effect.

• 한국추진공학회지
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• 제8권1호
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• pp.16-26
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• 2004

본 연구의 목적은 예조건화 방법론을 난류모델을 포함한 Navier-Stokes 방정식에 적용하는 것이다. 가상 음속 개념을 적용하였다 공간차분을 위해 Roe의 FDS를 사용하였고, 시간 적분을 위해 LU-SGS 기법을 사용하였다. 알고리즘을 검증하기 위하여 NACA 날개 주위의 저속유동, 초음속 노즐 유동을 계산하였다. 본 연구에서 개발된 알고리즘은 저속유동 및 초음속 유동의 계산에서 충분한 계산 성능을 보이는 것으로 판단된다.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2009년 춘계학술대회논문집
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• pp.87-89
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• 2009

In the present paper, transition turbulence model is applied to the NACA64(3)618 and detailed flow features are studied. The turbulence model is sensitive to the boundary layer grid quality and y+ of the grid was limited to 1. The prediction of the transition region is dependent on the local flow condition. The pressure coefficient distribution of the transition turbulence model is compared with that of the fully turbulent mode and the drag distribution of the transition turbulence model was compared with that of the wind tunnel test.

• 유체기계공업학회:학술대회논문집
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• 유체기계공업학회 2002년도 유체기계 연구개발 발표회 논문집
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• pp.433-438
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• 2002

The flow field around a symmetrical airfoil in a uniform flow under the generation of noise was numerically studied and compared with experimental datum. The numerical simulation was carried out by LES which employs a deductive dynamic model as subgrid-scale model. The result of an attack angle of $6^{\circ}$ indicate that the discrete frequency noise is generated when the separated laminar flow reattaches near the trailing edge of the pressure side and the turbulent boundary layer is formed over the suction side of the airfoil near the trailing edge. The periodic behavior of vortex formation was observed around the trailing edge and it persists further downstream in the wake. The frequency of the vortex formation in the wake was consistent with that of the discrete frequency noise.

• Journal of Mechanical Science and Technology
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• 제14권9호
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• pp.1013-1019
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• 2000

A numerical investigation was performed to determine the effect of the Gurney flap on a NACA 23012 airfoil. A Navier-Stokes code, RAMPANT, was used to calculate the flow field about the airfoil. Fully-turbulent results were obtained using the standard ${\kappa}-{\varepsilon}$ two-equation turbulence model. The numerical solutions showed that the Gurney flap increased both lift and drag. These results suggested that the Gurney flap served to increase the effective camber of the airfoil. The Gurney flap provided a significant increase in the lift-to-drag ratio relatively at low angle of attack and for high lift coefficient. It turned out that 0.6% chord size of flap was the best. The numerical results exhibited detailed flow structures at the trailing edge and provided a possible explanation for the increased aerodynamic performance.