• 한국전산유체공학회:학술대회논문집
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• 2004.03a
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• pp.161-165
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• 2004

In the vortex particle method based on the vorticity-velocity formulation for solving the Wavier-Stokes equations, the unsteady, incompressible, viscous laminar flow over a NACA 0012 foil is simulated. By applying an operator-splitting method, the 'convection' and 'diffusion' equations are solved sequentially at each time step. The convection equation is solved using the vortex particle method, and the diffusion equation using the particle strength exchange(PSE) scheme which is modified to avoid a spurious vorticity flux. The scheme is improved for variety body shape using one image layer scheme. For a validation of the present method, we illustrate the early development of the viscous flow about an impulsively started NACA 0012 foil for Reynolds number 550.

• Proceeding of EDISON Challenge
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• 2012.04a
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• pp.53-56
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• 2012

항공기에 작용하는 공기역학적 힘인 양력과 항력은 항공기 날개 설계에서 성능을 좌우하는 성능지수로 주로 이용된다. 본 연구에서는 NACA0012 airfoil 모델의 공력특성을 EDISON 열유체 시뮬레이션 프로그램(이하 EDISON)을 이용해 분석하고 검증해 보았다. 아음속 유동의 특정 조건에서 받음각과 유동형태에 따른 공력특성 분석을 수행하여 받음각에 따라 변하는 양력계수, 항력계수, 양항비, 실속각과 천음속 유동 조건에 맞추어진 마하수 0.5~1.22 영역에서 변하는 항력계수를 기존 데이터와 비교 검증했다.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.20 no.5
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• pp.1689-1699
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• 1996

A study on the transitional boundary layer with arbitrary pressure gradient under various upstream conditions is very important for engineering applications like the performance predictions of the turbomachineries under various and strong disturbances. Experimental data on the transitional boundary layer for real cascades of the turbomachinery are rare because of difficulties in boundary layer experiments. Flow on NACA0012 airfoil is more similar to the real case than that on the flat plate with which many researches are done. The data of the transitional flow on the airfoil could be used to verify or to develop a turbulence model for numerical simulations. The experiment was performed with two cases of Reynolds number at a=0$^{0}$ and one case of Reynolds number at a=5$^{0}$ . The measured data are the transition length and the wall shear stresses. These two characteristic values are measured within 25%~90% of the airfoil chord by Computation Preston tube Method(CPM) proposed by Nitsche et al.(1983). At a=0$^{0}$ , transition occured at 70% and 55% of chord length when R $e_{c}$=6*10$^{5}$ and 8* 10$^{5}$ , respectively. It started when R {\theta}=500 regardless of R $e_{c}$, and ended when R {\theta}=750, and 850 respectively. The transition length was 15~20% of the chord length. At a=5$^{0}$ (R $e_{c}$=6*10$^{5}$ ), boundary layer on the pressure side does not undergo transition, but on the suction side transition occured at .chi.$_{c}$/c=0.16 and ended at .chi.$_{c}$/c=0.22.c//c=0.22./c=0.22.c//c=0.22.

• 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.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.

• Journal of the Society of Naval Architects of Korea
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• v.28 no.2
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• pp.146-158
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• 1991

The flow around a two-dimensional foil section Is measured by a LDV(Laser Doppler Velocimetry) system which is capable of measuring the datailed flow field without interfering the original flow field. A 2-color 3-beam LDV system, which is capable of mea,;tiring 2 velocity components simultaneously and uses 2W Ar-Ion laser source, is used to measure the flow field around an NACA0012 foil section. The measured flow velocities are analysed iii order to study the boundary layer characteristics, flow separation and the detail structure of the flow near the trailing edge of the foil. The boundary layer characteristics are compared with the results by the head's momentum integral method. For the case of small angle of attack at relatively higher Reynolds number, both results show good agreements. The measured data of the velocity field around an NACA0012 foil section would be valuable data to validate the CFD(Computational Fluid Dynamic) calculation results. The developed experimental technique to evaluate the characteristics of two-dimensional foil sections is essential tool to develope new blade sections which have good lift characteristics and better cavitation performances.

• 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 Society of Naval Architects of Korea
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• v.31 no.1
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• pp.84-93
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• 1994

The purpose of present study is to investigate the surface roughness on tip-vortex cavitation and its induced noise, emanating from an elliptic wing of NACA 0012 section. Roughness elements of $200{\mu}m$ are applied to the 10% portion of wing tip, and then, the wing tip as well as the leading edge. It is shown from cavitation observation that the cavitation inception is first visible at about half chord downstream of wing tip for most experimental conditions, and developed into the tip-vortex cavitation and finally the fully developed cavitation as cavitation number is decreased. Acoustic noise generated by a tip-vortex cavitation has its frequency range of 3 kHz to 50 kHz, while the fully-developed cavitation at lower cavitation number induces a broad band spectrum. It is also shown that, when the roughness elements are applied to the wing tip and the leading edge, the cavitation characteristics and its induced noise are improved. Moreover, it is appeared that the condition at which the rough surface is at pressure side gives a better result. although its lift-drag ratio is reduced.

• Journal of Ocean Engineering and Technology
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• v.13 no.3 s.33
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• pp.114-122
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• 1999

준쇄파(쇄파발생정 리플형태의 불안정파)를 수치적으로 해석하고 예측하기 위하여 수중날개를 대상으로 점성유동장의 계산을 수행하였다. Navier-Stokes 방정식을 사용하여 자유수면을 계산하였고 정도향상을 위하여 Euler 형태의 자유표면조건에 고차의 유한차문법을 적용하였다. 이산화 과정에서 자유표면 격자에 3차 풍상미분항을 적용시켜 수치계산을 수행하였고, 계산결과를 사용하여 준쇄파의 생성조건을 규명하였다.

• Journal of Ocean Engineering and Technology
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• v.11 no.1
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• pp.55-64
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• 1997

A wing operating in close proximity to the ground exhibits a reduction in induced drag, which increase the lift/drag ratio. The poert-augmented ram (RAR) phenomenon involves directiing the efflux from forward mounted propulsion ststem under the wings, with the efflux nearly stagnated under the wings. In the present paper, 3 dimentional PAR was numerically studied by solving the Navier-Stokes equations. Pressure distribution and velocity vectors are calculated around the wing surface and the ground. Through the numerical simulation, Cp values and lift/drag ratio are carefully reviewed by changing the height/chord; 0.05, 0.1, 0.3 and 0.8. The shape of model is NACA 0012 with a span/chord ratio of 3.0. According to the numerical results, the relationship between lift/drag and height/chord is fairly reasonable.