• International Journal of Fluid Machinery and Systems
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• v.1 no.1
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• pp.1-9
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• 2008

In the first part of the paper, Computational Fluid Dynamics analysis of the combusting flow within a high-swirl lean premixed gas turbine combustor and over the $1^{st}$ row nozzle guide vanes is presented. In this analysis, the focus of the investigation is the fluid dynamics at the combustor/turbine interface and its impact on the turbine. The predictions show the existence of a highly-rotating vortex core in the combustor, which is in strong interaction with the turbine nozzle guide vanes. This has been observed to be in agreement with the temperature indicated by thermal paint observations. The results suggest that swirling flow vortex core transition phenomena play a very important role in gas turbine combustors with modern lean-premixed dry low emissions technology. As the predictability of vortex core transition phenomena has not yet been investigated sufficiently, a fundamental validation study has been initiated, with the aim of validating the predictive capability of currently-available modelling procedures for turbulent swirling flows near the sub/supercritical vortex core transition. In the second part of the paper, results are presented which analyse such transitional turbulent swirling flows in two different laboratory water test rigs. It has been observed that turbulent swirling flows of interest are dominated by low-frequency transient motion of coherent structures, which cannot be adequately simulated within the framework of steady-state RANS turbulence modelling approaches. It has been found that useful results can be obtained only by modelling strategies which resolve the three-dimensional, transient motion of coherent structures, and do not assume a scalar turbulent viscosity at all scales. These models include RSM based URANS procedures as well as LES and DES approaches.

• 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.

• 한국전산유체공학회:학술대회논문집
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• 2010.05a
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• pp.530-533
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• 2010

This paper presents numerical analysis and design optimization of various turbine blade cooling techniques with three-dimensional Reynolds-averaged Navier-Stokes(RANS) analysis. The fluid flow and heat transfer have been performed using ANSYS-CFX 11.0. A fan-shaped hole for film-cooling has been carried out to improve film-cooling effectiveness with the radial basis neural network method. The injection angle of hole, lateral expansion angle of hole and ratio of length-to-diameter of the hole are chosen as design variables and spatially averaged film-cooling effectiveness is considered as an objective function which is to be maximized. The impingement jet cooling has been performed to investigate heat transfer characteristic with geometry variables. Distance between jet nozzle exit and impingement plate, inclination of nozzle and aspect ratio of nozzle hole are considered as geometry variables. The area averaged Nusselt number is evaluated each geometry variables. A rotating rectangular channel with staggered array pin-fins has been investigated to increase heat transfer performance ad to decrease friction loss using KRG modeling. Two non-dimensional variables, the ratio of the eight diameter of the pin-fins and ratio of the spacing between the pin-fins to diameter of the pin-fins selected as design variables. A rotating rectangular channel with staggered dimples on opposite walls are formulated numerically to enhance heat transfer performance. The ratio of the dimple depth and dimple diameter are selected as geometry variables.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.40 no.2
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• pp.75-83
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• 2016

Since a turbomachine has complex flow characteristics, which are caused by adverse pressure gradient and high speed motion, an elaborate turbulence model is needed to accurately predict the flow. Some turbulence models such as an algebraic or a two-equation eddy viscosity model have been used for in-house RANS-code, but it is difficult to obtain good result for several complex flows. In this study, Durbin's V2-F turbulence model, which has been known for better prediction for severe flow separation, is applied to T-Flow. It was validated for simple cases such as channel and compressor cascade, and its applicability to turbomachinery was shown by analyzing internal flow of a single rotor. As a result, the V2-F turbulence model shows better blade surface pressure distribution than the one-and-two equation turbulence model.

• Journal of computational fluids engineering
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• v.18 no.2
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• pp.66-71
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• 2013

In this paper, the supersonic flows around space launch vehicles have been numerically simulated by using a 3-D RANS flow solver. The focus of the study was made for investigating plume-induced flow separation(PIFS). For this purpose, a vertex-centered finite-volume method was utilized in conjunction with 2nd-order Roe's FDS to discretize the inviscid fluxes. The viscous fluxes were computed based on central differencing. The Spalart-Allmaras model was employed for the closure of turbulence. The Gauss-Seidel iteration was used for time integration. To validate the flow solver, calculation was made for the 0.04 scale model of the Saturn-5 launch vehicle at the supersonic flow condition without exhaust plume, and the predicted results were compared with the experimental data. Good agreements were obtained between the present results and the experiment for the surface pressure coefficient and the Mach number distribution inside the boundary layer. Additional calculations were made for the real scale of the Saturn-5 configuration with exhaust plume. The flow characteristics were analyzed, and the PIFS distances were validated by comparing with the flight data. The KSLV-1 is also simulated at the several altitude conditions. In case of the KSLV-1, PIFS was not observed at all conditions, and it is expected that PIFS is affected by the nozzle position.

• Wind and Structures
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• v.27 no.2
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• pp.101-109
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• 2018

Wall jet flow exists widely in engineering applications, including the simulation of thunderstorm downburst outflows, and has been investigated extensively by both experimental and numerical methods. Most previous studies focused on the scaling laws and self-similarity, while the effect of lip thickness and external stream height on mean velocity has not been examined in detail. The present work is a numerical study, using steady Reynolds-Averaged Navier Stokes (RANS) simulations at a Reynolds number of $3.5{\times}10^4$, of a turbulent plane wall jet with an external stream to investigate the influence of the wall jet domain on downstream development of the flow. The comparisons of flow characteristics simulated by the Reynolds stress turbulence model closure (Stress-omega, SWRSM) and experimental results indicate that this model may be considered reasonable for simulating the wall jet. The confined wall jet is further analyzed in a parametric study, with the results compared to the experimental data. The results indicate that the height and the width of the wind tunnel and the lip thickness of the jet nozzle have a great effect on the wall jet development. The top plate of the tunnel does not confine the development of the wall jet within 200b of the nozzle when the height of the tunnel is more than 40b (b is the height of jet nozzle). The features of the centerline flow in the mid plane of the 3D numerical model are close to those of the 2D simulated plane wall jet when the width of the tunnel is more than 20b.

• Journal of computational fluids engineering
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• v.21 no.3
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• pp.15-23
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• 2016

In the present study, two phase flows around a projectile vertically launched from an underwater platform have been numerically investigated by using a three dimensional multi-phase RANS flow solver based on pseudo-compressibility and a homogeneous mixture model on unstructured meshes. The relative motion between the platform and projectile was described by six degrees of freedom equations of motion with Euler angles and a chimera technique. The propulsive power of the projectile was modeled as the fluid force acting on the lower surface of the body by the compressed air emitted from the underwater platform. Various flow conditions were considered to analyze the fluid-dynamics motion parameters of the projectile. The water level of platform and the current speed around the projectile were the main parametric variables. The numerical calculations were conducted up to 0.75sec in physical time scale. The dynamics tendency of the projectile was almost identical with respect to the water level variation due to the constant buoyancy term. The moving speed of the projectile along the vertical axis inside the platform decreased when the current speed increased. This is because the inflow from outside of the platform impeded development of the compressed air emitted from the floor surface of the launch platform. As a result, the fluid force acting on the lower surface of the projectile decreased, and injection time of the projectile from the platform was delayed.

• Journal of computational fluids engineering
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• v.19 no.2
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• pp.86-92
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• 2014

In this paper, prediction of separation trajectory for Two-stage-To-Orbit space launch vehicle has been numerically simulated by using an aerodynamic database based on steady state analysis. Aerodynamic database were obtained for matrix of longitudinal and vertical positions. The steady flow simulations around the launch vehicle have been made by using a 3-D RANS flow solver based on unstructured meshes. For this purpose, a vertex-centered finite-volume method was adopted to discretize inviscid and viscous fluxes. Roe's finite difference splitting was utilized to discretize the inviscid fluxes, and the viscous fluxes were computed based on central differencing. To validate this flow solver, calculations were made for the wind-tunnel experiment model of the LGBB TSTO vehicle configuration on steady state conditions. Aerodynamic database was constructed by using flow simulations based on test matrix from the wind-tunnel experiment. ANN(Artificial Neural Network) was applied to construct interpolation function among aerodynamic variables. Separation trajectory for TSTO launch vehicle was predicted from 6-DOF equation of motion based on the interpolated function. The result of present separation trajectory calculation was compared with the trajectory using experimental database. The predicted results for the separation trajectory shows fair agreement with reference[4] solution.

• 한국전산유체공학회:학술대회논문집
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• 2006.05a
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• pp.311-314
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• 2006

The supersonic flow around tandem cavities was investigated by three- dimensional numerical simulations using the Reynolds-Averaged Navier-Stokes(RANS) equation with the $\kappa-\omega$ thrbulence model. The flow around a cavity is characterized as unsteady flow because of the formation and dissipation of vortices due to the interaction between the freestream shear layer and cavity internal flow, the generation of shock and expansion waves, and the acoustic effect transmitted from wake flow to upstream. The upwind TVD scheme based on the flux vector split using van Leer's limiter was used as the numerical method. Numerical calculations were performed by the parallel processing with time discretizations carried out by the 4th-order Runge-Kutta method. The aspect ratio of cavities are 3 for the first cavity and 1 for the second cavity. The ratio of cavity interval to depth is 1. The ratio of cavity width to depth is 1 in the case of three dimensional flow. The Mach number and the Reynolds number were 1.5 and $4.5{\times}10^5$, respectively. The characteristics of the dominant frequency between two-dimensional and three-dimensional flows were compared, and the characteristics of the second cavity flow due to the fire cavity flow cavity flow was analyzed. Both two dimensional and three dimensional flow oscillations were in the 'shear layer mode', which is based on the feedback mechanism of Rossiter's formula. However, three dimensional flow was much less turbulent than two dimensional flow, depending on whether it could inflow and outflow laterally. The dominant frequencies of the two dimensional flow and three dimensional flows coincided with Rossiter's 2nd mode frequency. The another dominant frequency of the three dimensional flow corresponded to Rossiter's 1st mode frequency.

• Advances in aircraft and spacecraft science
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• v.7 no.6
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• pp.537-551
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• 2020

Investigation of leading edge impingement cooling for first stage rotor blades in an aero-engine turbine, its effect on rotor temperature and trailing edge wake loss have been undertaken in this study. The rotor is modeled with the nozzle for attaining a more accurate simulation. The rotor blade is hollowed in order for the coolant to move inside. Also, plenum with the 15 jet nozzles are placed in it. The plenum is fed by compressed fresh air at the rotor hub. Engine operational and real condition is exerted as boundary condition. Rotor is inspected in two states: in existence of cooling technique and non-cooling state. Three-dimensional compressible and steady solutions of RANS equations with SST K-ω turbulent model has been performed for this numerical simulation. The results show that leading edge is one of the most critical regions because of stagnation formation in those areas. Another high temperature region is rotor blade tip for existence of tip leakage in this area and jet impingement cooling can effectively cover these regions. The rotation impact of the jet velocity from hub to tip caused a tendency in coolant streamlines to move toward the rotor blade tip. In addition, by discharging used coolant air from the trailing edge and ejecting it to the turbines main flow by means of the slot in trailing edge, which could reduce the trailing edge wake loss and a total decrease in the blade cooling loss penalty.