• 한국전산유체공학회:학술대회논문집
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• 2003.10a
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• pp.75-76
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• 2003

Direct numerical simulation results of aeolian tones generated by a two-dimensional obstacle (circular cylinder, square cylinder, NACA0012 airfoil) in a uniform flow are presented and the generation and propagation mechanisms of the sound are discussed. The unsteady compressible Navier-Stokes equations are solved by a highly-accurate finite difference scheme over the entire region from near to far fields. The direct numerical simulation results are also compared with the results obtained by Curle's acoustic analogy.

• Journal of the Society of Naval Architects of Korea
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• v.44 no.1 s.151
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• pp.16-25
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• 2007

Studies of unsteady-airfoil flows have been motivated mostly by efforts to avoid. or reduce such undesirable effects as flutter, noise and vibrations, dynamic stall. In this paper, we carry out a computational study of viscous flows around a two-dimensional oscillating airfoil to investigate unsteady effects in these important and challenging flows. A fully implicit incompressible RANS solver has been used for calculating unsteady viscous flows around an airfoil. The cell-centered End order finite volume method is utilized to discretize governing equations. in order to ease the flow computation for fluid region changing in time, improve the qualify of solution and simplify the grid generation for an oscillating airfoil flow, the computational method adopts a moving and deforming grid generation technique based on the multi-block grid topology. The numerical method is applied for calculating viscous flows of an oscillating NACA 0012 in uniform flow. The computational results are compared with available experimental data. Computed results are compared with experimental data and flow characteristics of the experiment are reproduced well In the computed results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.4
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• pp.18-27
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• 2002

Aerodynamic shape optimization of two-dimensional airfoils in inviscid compressible flows is performed using a continuous adjoint formulation on unstructured meshes. Accurate evaluation of the gradient is achieved by using a reconstruction scheme based on the Laplacian averaging. A least-square method with extended stencil is used for flow gradient calculations. Proper convergence criterion is studied on Euler and adjoint equations for efficient design. The present method has been applied to RAE2822 and NACA0012 airfoils such that wave drag can be minimized by removing the shock wave. An inverse design is also performed to recover the shock wave on the designed RAE2822 airfoil.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.1
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• pp.12-22
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• 2012

Performance variation of autorotating rotor was investigated. The shaft angle of the rotor is reduced while the flight velocity is increased. The BO-105 helicopter rotor blade was replaced by untwisted NACA 0012 airfoil and the rotor was simulated by using Transient Simulation Method(TSM) to judge the autorotation region for the variables. To simulate the compressibility effect at high speed flight, two-dimensional aerodynamic data was analyzed by compressible Navier-Stokes solver and Pitt/Peters inflow theory was adopted to simulate the induced velocity field. Thrust and lift coefficients, lift to drag ratio variations were investigated, also the lift and power were compared to those of BO-105 helicopter. Sharing lift and power between the autorotating rotor and wing was considered when the compound aircraft concept is introduced.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.565-574
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• 2007

The flow control effect of blending Gurney flap with jet flap for flow around an NACA 0012 airfoil was numerically investigated through parameter variation of each flow control mechanism on unstructured meshes. The aerodynamic force and moment variations due to flow control were examined, and the results were compared between the blending control and each individual flow control. The results showed that the blending control required less energy input to achieve the same level of lift increment than that of the jet flap, and at the same time alleviated drag increment caused by introducing the Gurney flap.

• Journal of Aerospace System Engineering
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• v.10 no.4
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• pp.11-16
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• 2016

In this paper, an experimental study on the aerodynamic characteristics of a variable span wing flying inside a channel guideway is accomplished using wind tunnel testing. A variable span wing with a NACA 0012 airfoil section was fabricated and actuated using a linear servo motor. The aerodynamic effects of 1) wing aspect ratio, 2) ground effect, and 3) the gap between the wingtip and the wing fence were investigated. It was found that both ground effect and wing fence gap increased lift. Also, the wing fence gap does not significantly affect drag. Therefore, it was found that a variable span mechanism can be used as an effective high lift device when flap use is limited.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.12
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• pp.1153-1161
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• 2010

An efficient solution method for harmonic balance techniques with Fourier transform is presented for periodic unsteady flow problems. The present partially-implicit harmonic balance treats the flux terms implicitly and the harmonic source term is solved explicitly. The convergence of the partially Implicit method is much faster than the explicit Runge-Kutta harmonic balance method. The method does not need to compute the additional flux Jacobian matrix from the implicit harmonic source term. Compared with fully implicit harmonic balance method, this partial approach turns out to have good convergence property. Oscillating flows over NACA0012 airfoil are considered to verify the method and to compare with results of explicit R-K(Runge-Kutta) and dual time stepping methods.

• Ocean Systems Engineering
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• v.6 no.3
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• pp.245-264
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• 2016

The iterative boundary element method (IBEM) developed originally before for cavitating two-dimensional (2-D) and three-dimensional (3-D) hydrofoils moving under free surface is modified and applied to the case of 2-D (two-dimensional) airfoils and 3-D (three-dimensional) wings over water. The calculation of the steady-state flow characteristics of an inviscid, incompressible fluid past 2-D airfoils and 3-D wings above free water surface is of practical importance for air-assisted marine vehicles such as some racing boats including catamarans with hydrofoils and WIG (Wing-In-Ground) effect crafts. In the present paper, the effects of free surface both on 2-D airfoils and 3-D wings moving steadily over free water surface are investigated in detail. The iterative numerical method (IBEM) based on the Green's theorem allows separating the airfoil or wing problems and the free surface problem. Both the 2-D airfoil surface (or 3-D wing surface) and the free surface are modeled with constant strength dipole and constant strength source panels. While the kinematic boundary condition is applied on the airfoil surface or on the wing surface, the linearized kinematic-dynamic combined condition is applied on the free surface. The source strengths on the free surface are expressed in terms of perturbation potential by applying the linearized free surface conditions. No radiation condition is enforced for downstream boundary in 2-D airfoil and 3-D wing cases and transverse boundaries in only 3-D wing case. The method is first applied to 2-D NACA0004 airfoil with angle of attack of four degrees to validate the method. The effects of height of 2-D airfoil from free surface and Froude number on lift and drag coefficients are investigated. The method is also applied to NACA0015 airfoil for another validation with experiments in case of ground effect. The lift coefficient with different clearance values are compared with those of experiments. The numerical method is then applied to NACA0012 airfoil with the angle of attack of five degrees and the effects of Froude number and clearance on the lift and drag coefficients are discussed. The method is lastly applied to a rectangular 3-D wing and the effects of Froude number on wing performance have been investigated. The numerical results for wing moving under free surface have also been compared with those of the same wing moving above free surface. It has been found that the free surface can affect the wing performance significantly.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.4
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• pp.337-345
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• 2008

this study, three dimensional surface pressure distributions of SWIM whose main wing has NACA4412 airfoil with NACA0012 flaps were experimentally measured by pressure sensitive paint. Surface pressures on suction and pressure sides of the wing were measured by changing an angle of attack at a Reynolds number of 3.1x105 in KARI 1m subsonic wind tunnel. The experimental results showed that as an angle of attack increases minimum pressure region on a suction side moved from the wing root to the tip and low pressure region around trailing edge of the wing tip which causes wing tip vortex was observed. Although low pressure region at the tip still observed at an angle of attack 15 deg., other area on a suction side showed flat pressure distribution in a span-wise direction. It was also observed that the mean value of pressure coefficients was about 0.077 through a comparison between PSP and pressure taps at the same test conditions.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.29 no.10 s.241
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• pp.1163-1171
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• 2005

The shape optimization of an airfoil has been performed for an incompressible viscous flow. In this study, Pareto frontier sets, which are global and non-dominated solutions, can be obtained without various weighting factors by using the multi-objective genetic algorithm An NACA0012 airfoil is considered as a baseline model, and the profile of the airfoil is parameterized and rebuilt with four Bezier curves. Two curves, front leading to maximum thickness, are composed of five control points and the rest, from maximum thickness to tailing edge, are composed of four control points. There are eighteen design variables and two objective functions such as the lift and drag coefficients. A generation is made up of forty-five individuals. After fifteenth evolutions, the Pareto individuals of twenty can be achieved. One Pareto, which is the best of the . reduction of the drag furce, improves its drag to $13\%$ and lift-drag ratio to $2\%$. Another Pareto, however, which is focused on increasing the lift force, can improve its lift force to $61\%$, while sustaining its drag force, compared to those of the baseline model.