• Journal of Advanced Marine Engineering and Technology
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• v.25 no.5
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• pp.1091-1097
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• 2001

A numerical investigation was performed to determine the effect of airfoil on the optimum flap height using NACA 00XX and 44XX airfoils. The six flaps which have 0.5% chord height difference were selected . 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 results. For every NACA 00XX and 44XX 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 conclued that this initial approach gives an idea for the future development of the wind turbine optimum design.

• Journal of Mechanical Science and Technology
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• v.14 no.9
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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.

• Journal of computational fluids engineering
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• v.3 no.1
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• pp.46-53
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• 1998

An inverse method for the subsonic and transonic airfoil design was developed using the Euler equations. Two testcases were performed. One was a verification of the method using the supercritical airfoil of the Korean mid-sized (100 passengers class) transport aircraft. The other was the design of an airfoil showing a good cruising performance (L/D ratio) in the high subsonic flow regime. These testcases demonstrated the efficiency and the robustness of the design method in the present study.

• Journal of Industrial Technology
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• v.20 no.B
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• pp.71-76
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• 2000

A numerical investigation was performed to determine the effect of the Gurney flap on NACA 0015 airfoil. A Navier-Stokes code. FLUENT, was used to calculate the flow field about the airfoil. The fully-turbulent results were obtained using the standard ${\kappa}-{\varepsilon}$ two-equation turbulence model. The numerical solutions showed the Gurney flap increased both lift and drag. These results suggested that the Gurney flap served to increase the effective camber of the airfoil. Gurney flap provided a significant increase in lift-to-drag ratio relatively at low angle of attack and for high lift coefficient. It turned out that 0.75% chord size of flap was best. The numerical results exhibited detailed flow structures at the trailing edge and provided a possible explanation for the increased aerodynamic performance.

• Proceedings of the Korean Society of Marine Engineers Conference
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• 2001.05a
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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.

• Proceedings of the KSME Conference
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• 2004.04a
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• pp.1619-1624
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• 2004

Experimental study of separated flow over a NACA0012 airfoil is conducted at $Re=2{\times}10^5$ when periodic wakes pass over the airfoil. The wakes are periodically generated by circular cylinders upstream of the airfoil. The measurement of surface pressure and surface visualization at various angles of attack are carried out without and with passing wakes. Without passing wakes, a separation bubble at the leading edge of the suction surface is formed at an angle of attack, found from a local plateau in the streamwise pressure distribution and two distinct lines in the surface flow visualization. With passing wakes, however, the bubble disappears. Owing to passing wakes, the lift increases at high angle of attack and the angle of stall also increases.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.9
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• pp.729-738
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• 2021

Feedback flow control using an artificial neural network was numerically investigated for NACA0015 Airfoil to suppress flow separation on an airfoil. In order to achieve goal of flow control which is aimed to reduce the size of separation on the airfoil, Blowing&Suction actuator was implemented near the separation point. In the system modeling step, the proper orthogonal decomposition was applied to the pressure field. Then, some POD modes that are necessary for flow control are extracted to analyze the unsteady characteristics. NARX neural network based on decomposed modes are trained to represent the flow dynamics and finally operated in the feedback control loop. Predicted control signal was numerically applied on CFD simulation so that control effect was analyzed through comparing the characteristic of aerodynamic force and spatial modes depending on the presence of the control. The feedback control showed effectiveness in pressure drag reduction up to 29%. Numerical results confirm that the effect is due to dramatic pressure recovery around the trailing edge of the airfoil.

• Journal of computational fluids engineering
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• v.4 no.1
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• pp.53-61
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• 1999

Two-dimensional, unsteady, incompressible and compressible Navier-Stokes codes are developed for the computation of the viscous turbulent flow over high-lift airfoils. The compressible code involves a conventional upwind-differenced scheme for the convective terms and LU-SGS scheme for temporal integration. The incompressible code with pseudo-compressibility method also adopts the same schemes as the compressible code. Three two-equation turbulence models are evaluated by computing the flow over single and multi-element airfoils. The compressible and incompressible codes are validated by predicting the flow around the RAE 2822 transonic airfoil and the NACA 4412 airfoil, respectively. In addition, both the incompressible and compressible code are used to compute the flow over the NLR 7301 airfoil with flap to study the compressible effect near the high-loaded leading edge. The grid systems are efficiently generated using Chimera overlapping grid scheme. Overall, the κ-ω SST model shows closer agreement with experiment results, especially in the prediction of adverse pressure gradient region on the suction surfaces of high-lift airfoils.

• International Journal of Aeronautical and Space Sciences
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• v.14 no.3
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• pp.201-209
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• 2013

Much numerical and experimental research has been done for the flow around an oscillating airfoil. The main research topics are vortex shedding, dynamic stall phenomenon, MAV's lift and thrust generation. Until now, researches mainly have been concentrated on analyzing the wake flow for the variation of frequency and amplitude at a low angle of attack. In this study, wake structures and acoustic wave propagation characteristics were studied for a plunging airfoil at high angle of attack. The governing equations are the Navier-Stokes equation with LES turbulence model. OHOC (Optimized High-Order Compact) scheme and 4th order Runge-Kutta method were used. The Mach number is 0.3, the Reynolds number is, and the angle of attack is from $20^{\circ}$ to $50^{\circ}$. The plunging frequency and the amplitude are from 0.05 to 0.15, and from 0.1 to 0.2, respectively. Due to the high resolution numerical method, wake vortex shedding and pressure wave propagation process, as well as the propagation characteristics of acoustic waves can be simulated. The results of frequency analysis show that the flow has the mixed characteristics of the forced plunging frequency and the vortex shedding frequency at high angle of attack.

• 유체기계공업학회:학술대회논문집
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• 2002.12a
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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.