• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.4
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• pp.415-421
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• 2013

This study describes the aerodynamic design procedure for the axial turbines of a small heavy-duty gas turbine engine being developed by Doosan Heavy Industries. The design procedure mainly consists of three parts: namely, flowpath design, airfoil design, and 3D performance calculation. To design the optimized flowpath, through-flow calculations as well as the loss estimation are widely used to evaluate the effect of geometric variables, for example, shape of meridional plane, mean radius, blades axial gap, and hade angle. During the airfoil design procedure, the optimum number of blades is calculated by empirical correlations based on the in/outlet flow angles, and then 2D airfoil planar sections are designed carefully, followed by 2D B2B NS calculations. The designed planar sections are stacked along the spanwise direction, leading to a 3D surfaced airfoil shape. To consider the 3D effect on turbine performance, 3D multistage Euler calculation, single row, and multistage NS calculations are performed.

• 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 the Korean Society for Aeronautical & Space Sciences
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• v.45 no.3
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• pp.259-267
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• 2017

Flow fields around a KARI-11-180 airfoil, SDM and transonic body are numerically simulated by using an unstructured meshes based compressible flow solver developed at KAIST. RANS equations are solved to analyse the flow fields and Roe's FDS method is adopted to evaluate convective fluxes. Turbulence effect of the flow fields is modeled by a SA model, SST model and ${\gamma}-{\widetilde{Re}}_{{\theta}t}$ model. It is found that smaller drag coefficients are predicted for the KARI-11-180 airfoil when a transition phenomenon is considered and small deviations exist between CFD and EFD results. For the SDM, flow separation is observed at a leading edge and calculated aerodynamic properties show similar tendencies to experimental results. A shock wave on main wings of the transonic body is successfully captured by the present flow solver at a Mach number 0.9. Estimated pressure profiles by means of the present CFD method also agree well with those of wind tunnel results.

• 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.

• 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.

• 한국전산유체공학회:학술대회논문집
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• 2006.10a
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• pp.193-198
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• 2006

It takes a lot of time and effort to generate grids for numerical analysis of problems with ground effect because the relative attitude and height of airfoil should be maintained to the ground as well as the inflow. A low Mach number preconditioned turbulent flow solver using the overlap grid technique has been developed and applied to the ground effect simulation. It has been validated that the present method using the multi-block grid gives us highly accurate solutions comparing with the experimental data of the RAE 101 airfoil in an unbounded condition. Present numerical method has been extended to simulate ground effect problems by using the overlapped grid system to avoid tedious work in generating multi-block grid system. An extended method using the overlapped grid has been verified and validated by comparing with results of multi-block method and experimental data as well. Consequently, the overlapped grid method can provide not only sufficiently accurate solutions but also the efficiency to simulate ground effect problems. It is shown that the pressure and aerodynamic centers move backward by the ground effect as the airfoil approaches to the ground.

• Advances in aircraft and spacecraft science
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• v.2 no.4
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• pp.427-443
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• 2015

The present paper is the follow-on of a former work in which the influence of the gas-surface interaction models was evaluated on the aerodynamic coefficients of an aero-space-plane and on a section of its wing. The models by Maxwell and by Cercignani-Lampis-Lord were compared by means of Direct Simulation Monte Carlo (DSMC) codes. In that paper the diffusive, fully accommodated, semi-specular and specular accommodation coefficients were considered. The results pointed out that the influence of the interaction models, considering the above mentioned accommodation coefficients, is pretty strong while the Cercignani-Lampis-Lord and the Maxwell models are practically equivalent. In the present paper, the comparison of the same models is carried out considering the dependence of the accommodation coefficients on the angle of incidence (or partial accommodation coefficients). More specifically, the normal and the tangential momentum partial accommodation coefficients, obtained experimentally by Knetchel and Pitts, have been implemented. Computer tests on a NACA-0012 airfoil have been carried out by the DSMC code DS2V-64 bits. The airfoil, of 2 m chord, has been tested both in clean and flapped configurations. The simulated conditions were those at an altitude of 100 km where the airfoil is in transitional regime. The results confirmed that the two interaction models are practically equivalent and verified that the use of the Knetchel and Pitts coefficients involves results very close to those computed considering a diffusive, fully accommodated interaction both in clean and flapped configurations.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.27 no.3
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• pp.334-346
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• 2003

An experimental study was carried out to investigate near-wake characteristics of an elliptic airfoil oscillating in pitch. The airfoil was sinusoidally pitched about the half chord point between -5$^{\circ}$and +25$^{\circ}$angles of attack at the freestream velocities of 3.4 and 23.1 m/s. The corresponding Reynolds numbers based on the chord length were 3.3$\times$10$_{4}$ and 2.2$\times$10$^{5}$ , respectively. A hot-wire anemometer was used to measure the near-wake flow variables at the reduced frequency of 0.1. Ensemble-averaged velocity and turbulence intensity profiles were presented to examine the near-wake characteristics depending on the Reynolds number. The axial velocity deficit in the near-wake region tends to decrease with the increase in the Reynolds number as found in many stationary airfoil tests. Turbulence intensity in the near-wake region have a tendency to decrease with the -increase in the Reynolds number during the pitch-up motion, whereas it shows different feature during the pitch-down motion according to the separation characteristics.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.14 no.3
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• pp.7-15
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• 2006

Aerodynamic analysis of an airfoil in the transonic region was automated in order to enable parametric study by using the journal file of the commercial analysis code FLUENT, pre/post process Gambit and computational mathematics code MATLAB. The automated capability was illustrated via NACA 0012 and RAE 2822 airfoils. This analysis was carried out at Mach numbers ranged from 0.70 to 0.80, angles of attack; 1$^{\circ}$, 2$^{\circ}$ and 4$^{\circ}$, Reynolds numbers; 4.0${\times}$106, 6.5${\times}$106. The analysis results of a pressure coefficient were verified by comparing with the experimental data which were measured in terms of chord length because the pressure coefficient of an airfoil surface is a good estimator of flow characteristics. The results of two airfoils show that this analysis code is useful enough to be used in the design optimization of airfoil.

• 한국전산유체공학회:학술대회논문집
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• 2006.10a
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• pp.155-158
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• 2006

In the present paper, the scale effects of two-dimensional airfoil and three-dimensional wind turbine were investigated by using FLUENT software. For two dimensional analysis, flow around S809 airfoil with various Reynolds No. and Mach No. conditions were simulated. For three dimensional analysis, scaled NREL Phase VI wind turbine models from 6% to 1,600% were simulated under the same tip speed ratio condition. Finally, aerodynamic comparisons between two-dimensional flow and three dimensional wind turbine flow are made for the feasibility study of scale effect corrections. Currently, KARI(Korea Aerospace Research Institute) is preparing for the wind tunnel test of 12% NREL Phase VI wind turbine and the performance analysis of the scaled NREL wind turbine model will be validated by the wind tunnel test.