• 한국전산유체공학회:학술대회논문집
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• 2001.05a
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• pp.51-59
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• 2001

An evaluation of one zero-equation and two one-equation eddy viscosity-transport turbulence closure models as implemented CFDS(Characteristic Flux Difference Splitting ) code is presented herein. Comparisons of Baldwin-Lomax model as zero-equation and Baldwin-Barth and Spalart-Allmaras model as one-equation are presented for three test cases, first inlvolving the 3 dimensional supersonic flow at M=1.98 over tangent ogive cylinder, second involving the 2 dimensional transonic flow at M=0.79 over RAE 2822 airfoil, third involving the 3 dimensional transonic flow at M=0.84 over ONERA M6 wing. The numerical results of CFDS code will also examined through direct comparison with experimental data.

• 한국전산유체공학회:학술대회논문집
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• 2007.10a
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• pp.88-91
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• 2007

In the shape design optimization of an airfoil, the shape function has been used to find the optimal airfoil shape for given conditions. The parameters determining the airfoil shape are used in the shape design optimization as design variables. However, they usually don't have physical meaning. The PARSEC (Parametric Shapes) function is a recently proposed shape function and its parameters have the physical meaning. In this study the usefulness of the PARSEC is tested for the RAE2822 airfoil in the transonic flow region to reduce the shock strength and the result is compared with Hicks-Henne function. The optimized airfoils reduce the shock strength and they show similar result.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2001.05a
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• pp.675-680
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• 2001

In this study, a fluid/structure coupled analysis system using computational fluid dynamics and computational structural dynamics has been developed. The unsteady flow fields are predicted using unstructured Euler code. Coupled time-integration method (CTIM) was applied to computer simulation of the flow-induced vibration phenomena. To investigate the interaction effect of shock motions, 2-DOF airfoil systems have been studied in the subsonic and transonic flow region. Also, aeroelastic analyses for the airfoil with an arbitrary object are performed to show the analysis capability and interference effects for the complex geometries. The present results show the flutter stabilities and characteristics of aeroelastic responses with moving shock effects.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.8
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• pp.21-29
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• 2002

In the present study, a passive control method, using a porous wall and cavity system, is applied to the shock wave/boundary layer interactions in transonic moist air flow. The two-dimensional, unsteady, compressible, Navier-Stokes equations, which are fully coupled with a droplet growth equation, are solved by the third-order MUSCL type TVD finite difference scheme. Baldwin-Lomax model is employed to close the governing equations. In order to investigate the effectiveness of the present control method, the total pressure loss of the flow and the time-dependent behaviour of shock motions are analyzed in detail. The computed results show that the present passive control method considerably reduces the total pressure losses due to the shock wave/boundary layer interaction in transonic moist air flow and suppresses the unsteady shock wave motions over the airfoil as well. It is also found that the location of the porous ventilation significantly affects the control effectiveness.

• Journal of the Korea Institute of Military Science and Technology
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• v.1 no.1
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• pp.145-153
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• 1998

In the direct numerical optimization method, the aerodynamic coefficients of the airfoil designed by one-point design can be deteriorated at other operating points. Therefore, the capacity of the multi-point design is indispensable for actual airfoil design. In this paper, the two-point design of transonic airfoils is studied based on the Navier-Stokes equations flow solver and the feasible direction optimization algorithm, and the effects of weighting parameter were analyzed and compared. The results show that the airfoils designed by two-point design satisfy the performances at the peripheral regions of two operating points concurrently and have the favorable aerodynamic characteristics at the point which has larger weighting parameter than the other point.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.601-613
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• 2017

An adjoint-based error estimation and mesh adaptation study is conducted for two-dimensional viscous flows on unstructured hybrid meshes. The error in an integral output functional of interest is estimated by a dot product of the residual vector and adjoint variable vector. Regions for the mesh to be adapted are selected based on the amount of local error at each nodal point. Triangular cells in the adaptive regions are refined by regular refinement, and quadrangular cells near viscous walls are bisected accordingly. The present procedure is applied to single-element airfoils such as the RAE2822 at a transonic regime and a diamond-shaped airfoil at a supersonic regime. Then the 30P30N multi-element airfoil at a low subsonic regime with a high incidence angle (${\alpha}=21deg.$) is analyzed. The same level of prediction accuracy for lift and drag is achieved with much less mesh points than the uniform mesh refinement approach. The detailed procedure of the adjoint-based mesh refinement for the multi-element airfoil case show that the basic flow features around the airfoil should be resolved so that the adjoint method can accurately estimate an output error.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.6
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• pp.23-29
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• 2003

Transonic two-phase flow of Saturated humid air, in which relative humidity is 100%, with various condensation processes around thin airfoils is investigated. The study uses an extended transonic small-disturbance(TSD) model of Rusak and Lee [11, 12] which includes effects of heat addition to the flow due to condensation. Two possible limit types of condensation processes are considered. In the nonequilibrium and homogeneous process, the condensate mass fraction is calculated according to classical nucleation and droplet growth rate models. In the equilibrium process, the condensate mass fraction is calculated by assuming an isentropic process. The flow and condensation equations are solved numerical1y by iterative computations. Results under same upstream conditions describe the flow structure, field of condensate, and pressure distribution on airfoil's surfaces. It is found that flow characteristics, such as position and strength of shock waves and airfoil’s pressure distribution, are different for the two condensation processes. Yet, in each case, heat addition as a result of condensation causes significant changes in flow behavior and affects the aerodynamic performance of airfoils.

• Journal of computational fluids engineering
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• v.17 no.1
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• pp.54-60
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• 2012

Transition prediction results are validated with experimental data obtained from a transonic wind tunnel for the KU109C airfoil. A Reynolds-Averaged Navier-Stokes code is simultaneously coupled with the transition transport model of Langtry and Menter and applied to the numerical prediction of aerodynamic performance of the KU109C airfoil. Drag coefficients from the experiment are better correlated to the numerical prediction results using a transition transport model rather than the fully turbulent simulation results. Maximum lift coefficient and drag divergence at the zero-lift condition with Mach number are investigated. Through the present validation procedure, the accuracy and usefulness of both the experiment and the numerical prediction are assessed.

• 한국전산유체공학회:학술대회논문집
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• 1999.11a
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• pp.17-22
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• 1999

Convergence characteristics and efficiency of three implicit approximate factorization schemes(ADI, DDADI and MAF) are examined using 2-Dimensional compressible upwind Navier-Stokes code. Second-order CSCM(Conservative Supra Characteristic Method) upwind flux difference splitting method with Fromm scheme is used for the right-hand side residual evaluation, while generally first-order upwind differencing is used for the implicit operator on the left-hand side. Convergence studies are performed using an example of the flow past a NACA0012 airfoil at steady transonic flow condition, i. e. Mach number 0.8 at $1.25^{\circ}$ angle of attack. The results were compared with other computational results in order to validate the current numerical analysis. The results from the implicit AF algorithms were compared well in low surface with the other computational results; however, not well in upper surface. It might be due to lack of the grid around the shock position. Because the algorithm minimizes the errors of the approximate decomposition, the improved convergence rate with MAF were observed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.3
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• pp.233-240
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• 2017

The flowfield around transonic wing-body configuration was simulated using in-house CFD code and compared with the experimental data to understand the influence of several features of CFD(Computational Fluid Dynamics) ; grid dependency, turbulence models, spatial discretization, and viscosity. The wing-body configuration consists of a simple planform RAE Wing 'A' with an RAE 101 airfoil section and an axisymmetric body. The in-house CFD code is a compressible Euler/Navier-Stokes solver based on unstructured grid. For the turbulence model, the $k-{\omega}$ model, the Spalart-Allmaras model, and the $k-{\omega}$ SST model were applied. For the spatial discretization method, the central differencing scheme with Jameson's artificial viscosity and Roe's upwind differencing scheme were applied. The results calculated were generally in good agreement with experimental data. However, it was shown that the pressure distribution and shock-wave position were slightly affected by the turbulence models and the spatial discretization methods. It was known that the turbulent viscous effect should be considered in order to predict the accurate shock wave position.