• Journal of computational fluids engineering
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• v.11 no.2 s.33
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• pp.25-31
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• 2006

A two-dimensional backward facing step flow is simulated by using URANS and Detached Eddy Simulations(DES) approaches. Turbulence models adopted for URANS and DES simulations are Spalart-Allmaras(S-A) model and Shear Stress Transport(SST) model. The target flow with ER=1.125, $Re_H=37,500$ is experimentally studied by Driver & Seegmiller. Various versions of DES have been tested in this paper. Results of the simulations are compared with the experimental data available to evaluate the merits and demerits of URANS and several versions of DES. URANS simulation converges to a steady state and hence unsteady characteristics are not featured. DES simulations in general successfully mimic large scale structures and oscillation characteristics of the flow.

• Proceedings of the Korea Water Resources Association Conference
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• 2017.05a
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• pp.127-127
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• 2017

침수식생이 식재된 개수로에서 식생밀도에 따라 유동 및 난류의 특성이 변화된다. 이러한 특성은 식생에서의 유사, 영양물질, 용존 산소 등에 영향을 미치며 수중 생물의 서식에 변화를 준다. 따라서 침수식생이 식재된 개수로 흐름을 이해하는 것은 중요하게 여겨지고 있으며 많은 선행연구자들에 의해 실험 및 수치모의를 통해 활발히 연구되고 있다. 하지만 대부분의 RANS(Reynolds-Averaged Navier-Stokes)를 기반으로 한 선행연구에서는 침수식생의 흐름 특성을 반영하지 않은 모형을 이용하여 정확한 모의 결과를 도출하지 못 하였다. 이에 정확한 침수식생 흐름을 수치모의하기 위해서는 침수식생 흐름의 특성을 반영한 지배방정식을 이용해야 한다. 본 연구의 목적은 침수식생 흐름의 특성을 RANS 모형 중의 하나인 SA (Spalart-Allmaras) model에 반영하고, 식생밀도에 따른 유동 및 난류량을 실측치와 비교하는 것이다. RANS 방정식을 이용하여 난류모델링을 하였으며, 난류폐합문제를 해결하기 위해서 modified SA model을 이용하였다. 침수식생에서의 흐름을 해석하기 위해 운동량방정식에 식생항력을 추가하였다. 선행연구자의 식생수로 실험을 바탕으로 모형검증을 하였으며, 식생밀도에 따라 평균유속 및 난류구조를 확인하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.761-770
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• 2008

The report presents turbulent Navier Stokes computations on twin engine afterbody model with jet exhaust. The computations are carried out for free-stream Mach number of 0.8 to 1.20 and jet pressure ratio of 3.4 to 7.8. The Spalart-Allmaras turbulence model is used in the computations. Comparison is made with experimental data and Cp distribution around the afterbody is found to agree well with experiments. Flow features of the exhaust jet like under expansion, over expansion, Mach discs, etc are well captured. The effect of nozzle pressure ratio and flight Mach number are studied in detail. These computations serve as validation of the in-house code for twin jet afterbody.

• Journal of computational fluids engineering
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• v.13 no.4
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• pp.45-49
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• 2008

The present paper describes the results of performance analysis for UH-60A rotor blade in hover. For the numerical simulations, commercial CFD software, FLUENT was used with Spalart-Allmaras turbulence model. The flow solver was based on node based scheme and second order spatial accuracy options was used for simulations. For the enhancement of wake capturing capability, high resolution grid was used around tip vortex region. Granting that somewhat over-prediction of thrust was observed near blade tip region, performance was well correlated with experimental data within 3% accuracy in the operating region. Finally it was shown that the present flow solver can be used as a preliminary performance analysis tool for hovering helicopter rotor blades.

• Journal of computational fluids engineering
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• v.12 no.1
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• pp.53-59
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• 2007

An axisymmetric supersonic screeching jet is numerically simulated to examine the length scales of screech frequency as well as screech tone generation mechanism. The axisymmetric Reynolds-averaged Navier-Stokes equations in conjuction with a modified Spalart-Allmaras turbulence model are employed. It is demonstrated that the axisymmetric jet screech tones can be simulated correctly and the numerical results are in good agreement with the experimental data. Instability waves, shock-cell structures and the phenomena of shock motion are investigated in detail to identify the screech tone generation mechanism. Shock spacings and standing wave length are analyzed to determine the dominent length scale crucial to the screech frequency formulation.

• Journal of Mechanical Science and Technology
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• v.15 no.10
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• pp.1434-1441
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• 2001

A computational study has been performed to determine the effects of divergent trailing edge (DTE) modification to a supercritical airfoil in transonic flow field. For this, the computational result with the original DLBA 186 supercritical airfoil was compared to that of the modified DLBA 283. A wavier-Stokes code, Fluent 5. 1, was used with Spalart-Allmaras's one-equation turbulence model. Results in this study showed that the reduction in drag due to the DTE modification is associated with weakened shock and delayed shock appearance. The decrease in drag due to the DTE modification is greater than the increase in base drag. The effect of the recirculating flow region on lift increase was also observed. An airfoil with DTE modification achieved the same lift coefficient at a lower angle of attack while giving a lower drag coefficient. Thus, the lift-to-drag ratio increases in transonic flow conditions compared to the original airfoil. The lift coefficient increases considerably whereas the lift slope increases just a little due to DTE modification.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2014.04a
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• pp.558-563
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• 2014

To reduce the transmission of large gun-generated noise from the firing test range to the community, we have tested a silencer to be used with howitzer(155mm KM114A2) after simulation. Numerical analysis was conducted by using a commercial CFD code, FLUENT. To analyze complex blast flow fields, the Spalart-Allmaras model was applied under 2 dimensional and axisymmetric conditions. Firing tests were also performed with the KM114A2 howitzer while the silencer was installed. This paper describes a result of comparison between results of computer analysis and test outcomes which were gotten by firing 155mm projectiles at the testing range. This paper will also be informative to the muffler design which will be adapted to 155mm large gun in the future.

• International Journal of Aeronautical and Space Sciences
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• v.8 no.2
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• pp.44-53
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• 2007

A three-dimensional viscous flow solver has been developed for the prediction of the aerodynamic performance of hovering helicopter rotor blades using unstructured hybrid meshes. The flow solver utilized a vertex-centered finite-volume scheme that is based on the Roe's flux-difference splitting with an implicit Jacobi/Gauss-Seidel time integration. The eddy viscosity are estimated by the Spalart- Allmaras one-equation turbulence model. Calculations were performed at three operating conditions with varying tip Mach number and collective pitch setting for the Caradonna-Tung rotor in hover. Additional computations are made for the UH-60A rotor in hover. Reasonable agreements were obtained between the present results and the experiment in both blade loading and overall rotor performance. It was demonstrated that the present vertex-centered flow solver is an efficient and accurate tool for the assessment of rotor performance in hover.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.289-292
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• 2008

Highly over-expanded nozzle of the rocket engines will be excited by non-axial forces due to flow separation at sea level operations. Since rocket engines are designed to produce axial thrust to power the vehicle, non-axial static and/or dynamic forces are not desirable. Several engine failures were attributed to the side loads. Present work investigate the unsteady flow in an over-expanded rocket nozzle in order to estimate side load during a shutdown/starting. Numerical computations has been carried out with density based solver on multi-block structured grid. Present solver is explicit in time and unsteady time step is calculated using dual time step approach. AUSMDV is considered as a numerical scheme for the flux calculations. One equation Spalart-Allmaras turbulence model is selected. Results presented here is for two nozzle pressure ratio i.e. 100 and 20. At 100 NPR, restricted shock separation (RSS) pattern is observed while, 20 NPR shows free shock separation (FSS) pattern. Side load is observed during the transition of separation pattern at different NPR.

• 한국전산유체공학회:학술대회논문집
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• 2007.10a
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• pp.73-76
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• 2007

The present paper describes the results of performance analysis for UH-60A rotor blade in hover. For the numerical simulations, commercial CFD software, FLUENT was used with Spalart-Allmaras turbulence model. The flow solver was based on node based scheme and second order spatial accuracy options was used for simulations. For the enhancement of wake capturing capability, high resolution grid was used around tip vortex region. Granting that somewhat over prediction of thrust was observed near blade tip region, performance was well correlated with experimental data within 3% accuracy in the operating region. Finally it was shown that the present flow solver can be used for preliminary performance analysis tool for hovering helicopter rotor blades.