• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.5
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• pp.403-413
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• 2010

Transition location of boundary layers over airfoils is predicted by using PSE(Parabolized Stability Equations) and $e^N$-method. Growth rates of disturbances are obtained from the PSE analysis and the N-factor curves are calculated by integrating the growth rates. The computational code developed in the present study is validated by comparing the computed results with the well known data for the cases of flat plate boundary layers and airfoils. Predictions of transition location are made for the boundary layers over NACA0012, NLF(1)-0414F, and NLF(1)-0416 airfoil. Predicted transition locations are found to be in good agreement with the experimental data.

• 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.

• Proceedings of the KSME Conference
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• 2003.11a
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• pp.804-809
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• 2003

In this study, 2D computations of the Aeolian tones for some obstacles (circular cylinder, square cylinder and NACA0012 airfoil) are simulated. First of all, we calculate the flow noise generated by a uniform flow around a two-dimensional circular cylinder at Re=150 are simulated by applying the finite difference lattice Boltzmann method (FDLBM). The third-order-accurate up-wind scheme (UTOPIA) is used for the spatial derivatives, and the second-order-accurate Runge-Kutta scheme is applied for the time marching. The results show that we successively capture very small acoustic pressure fluctuation with the same frequency of the Karman vortex street compared with the pressure fluctuation around a circular cylinder. The propagation velocity of the acoustic waves shows that the points of peak pressure are biased upstream due to the Doppler effect in the uniform flow. For the downstream, on the other hand, it is faster. To investigate the effect of the lattice dependence, furthermore, simulations of the Aeolian tones at the low Reynolds number radiated by a square cylinder and a NACA0012 airfoil with a blunt trailing edge at high incidence are also investigated.

• Proceedings of the KSME Conference
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• 2000.11b
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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.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.20 no.7
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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.

• Journal of computational fluids engineering
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• v.3 no.1
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• pp.54-62
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• 1998

본 논문은 2차원 해면효과의 수치계산 결과를 정리하였다. 지면으로부터의 높이변화에 따른 점성유동장을 계산하기 위하여 지배방정식으로는 비압축성 RANS 방정식을, 시간에 대하여서는 음해법으로 프로그램을 구성하였다. 압력항은 가상압축성과 4차 수치확산항을 추가하는 것에 의해 계산하였으며, 높은 레이놀즈 수에서의 효과적인 계산을 위해 Baldwin- Lomax 난류모델을 도입하였다. 해면효과가 없는 무한유중에서의 NACA-0012 단면 계산결과를 실험 데이터와 비교하는 것에 의해 프로그램의 타당성을 확인하였다. NACA-6409와 두께 비 4.6%의 날개에 대하여 해면효과를 고려한 계산을 수행하였다. 계산결과, 높이의 변화에 따라 계산된 무차원계수, 압력 및 속도분포는 해면효과의 특성을 잘 보여주고 있다.