scholarly journals Heat Transfer on a Film-Cooled Rotating Blade Using a Two-Equation Turbulence Model

1998 ◽  
Vol 4 (3) ◽  
pp. 201-216 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Turbulence Model ◽  
Turbulence Models ◽  
Leading Edge ◽  
Equation Model ◽  
Tip Clearance ◽  
Suction Surface ◽  
Pressure Surface ◽  
Rotating Blade

A three-dimensional Navier–Stokes code has been used to compare the heat transfer coefficient on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only filmcooled rotating blade over which experimental data is available for comparison. Over 2.278 million grid points are used to compute the flow over the blade including the tip clearance region, using Coakley'sq-ωturbulence model. Results are also compared with those obtained by Garg and Abhari (1997) using the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the experimental data is obtained on the suction surface for both the turbulence models. At the leading edge, the B-L model yields a better comparison than theq-ωmodel. On the pressure surface, however, the comparison between the experimental data and the prediction from either turbulence model is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects due to stator-rotor interaction in the experiments which are not modeled in the present computations. Prediction using the two-equation model is in general poorer than that using the zero-equation model, while the former requires at least 40% more computational resources.

scholarly journals Comparison of Predicted and Experimental Heat Transfer on a Film-Cooled Rotating Blade Using a Two-Equation Turbulence Model

1997 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Turbulence Model ◽  
Turbulence Models ◽  
Leading Edge ◽  
Equation Model ◽  
Tip Clearance ◽  
Suction Surface ◽  
Pressure Surface ◽  
Rotating Blade

A three-dimensional Navier-Stokes code has been used to compare the heat transfer coefficient on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only film-cooled rotating blade over which experimental data is available for comparison. Over 2.278 million grid points are used to compute the flow over the blade including the tip clearance region, using Coakley’s q-ω turbulence model. Results are also compared with those obtained by Garg and Abhari (1996) using the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the experimental data is obtained on the suction surface for both the turbulence models. At the leading edge, the B-L model yields a better comparison than the q-ω model. On the pressure surface, however, the comparison between the experimental data and the prediction from either turbulence model is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects due to stator-rotor interaction in the experiments which are not modeled in the present computations. Prediction using the two-equation model is in general poorer than that using the zero-equation model, while the former requires at least 40% more computational resources.

scholarly journals Prediction of Turbine Blade Passage Heat Transfer Using a Zero and a Two-Equation Turbulence Model

1994 ◽  
Author(s):  
Ali A. Ameri ◽  
Andrea Arnone
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Turbulence Model ◽  
Turbulence Models ◽  
Algebraic Model ◽  
Equation Model ◽  
Accurate Solution ◽  
Bypass Transition ◽  
Transfer Rates ◽  
Blade Passage

Predictions of the heat transfer rates on the hot surfaces of a turbine cascade blade passage as influenced by the turbulence models was examined. A zero equation turbulence model supplemented by a bypass transition model and a two equation low Reynolds number model were chosen for this study. The experimental data of Graziani et. al. were used for comparison. The comparisons suggest that at least for the experimental data considered in this work the use of a two-equation model does not provide an overall more accurate solution than the zero equation model. This conclusion is strengthened if one takes into account the relative economy of computations with the algebraic model.

scholarly journals Navier-Stokes Predictions of Transition, Loss and Heat Transfer in a Turbine Cascade

1987 ◽  
Author(s):  
N. T. Birch
Keyword(s):  
Heat Transfer ◽  
Turbulence Model ◽  
Leading Edge ◽  
Solution Procedure ◽  
Navier Stokes ◽  
Turbine Cascade ◽  
Suction Surface ◽  
Pressure Surface ◽  
Transitional Boundary Layer ◽  
Data Correlations

The loss and heat transfer of a turbine cascade are strongly influenced by the location and extent of the transitional boundary layer. In this paper, two approaches are adopted to predict the onset and extent of transition within a 2-D explicit Navier-Stokes solution procedure. In the first, transition is predicted by coupling transition data correlations with an algebraic turbulence model. In the second, a low Reynolds Number one-equation turbulence model is used. Comparison is made with the turbine cascade data of Nicholson et al. (1982). This indicates that the first model gives good predictions of suction surface behaviour but poor predictions on the pressure surface. The model is also difficult to apply in a N-S method. The second model gives good predictions of pressure surface behaviour but consistently predicts transition near the leading edge on the suction surface. The latter is attributed to a Mach Number over-speed and leading edge effects.

A New Unsteady-Based Turbulence Model to Predict Shear Layer Rollup and Breakdown

2004 ◽  
Author(s):  
D. Scott Holloway ◽  
James H. Leylek
Keyword(s):  
Shear Layer ◽  
Reynolds Stress ◽  
Turbulence Model ◽  
Turbine Blades ◽  
Turbulence Models ◽  
Separated Flow ◽  
Leading Edge ◽  
Tip Clearance ◽  
Aircraft Carrier ◽  
Flow Through

This paper documents the computational investigation of the unsteady rollup and breakdown of a turbulent separated shear layer. This complex phenomenon plays a key role in many applications, such as separated flow at the leading edge of an airfoil at off-design conditions; flow through the tip clearance of a rotor in a gas turbine; flow over the front of an automobile or aircraft carrier; and flow through turbulated passages that are used to cool turbine blades. Computationally, this problem poses a significant challenge in the use of traditional RANS-based turbulence models for the prediction of unsteady flows. To demonstrate this point, a series of 2-D and 3-D unsteady simulations have been performed using a variety of well-known turbulence models, including the “realizable” k-ε model, a differential Reynolds stress model, and a new model developed by the present authors that contains physics that account for the effects of local unsteadiness on turbulence. All simulations are fully converged and grid independent in the unsteady framework. A proven computational methodology is used that takes care of several important aspects, including high-quality meshes (2.5 million finite volumes for 3-D simulations) and a discretization scheme that will minimize the effects of numerical diffusion. To isolate the shear layer breakdown phenomenon, the well-studied flow over a blunt leading edge (Reynolds number based on plate half-thickness of 26,000) is used for validation. Surprisingly, none of the traditional eddy-viscosity or Reynolds stress models are able to predict an unsteady behavior even with modifications in the near-wall treatment, repeated adaption of the mesh, or by adding small random perturbations to the flow field. The newly developed unsteady-based turbulence model is shown to predict some important features of the shear layer rollup and breakdown.

Design and Numerical Investigation of Rudder With Leading-Edge Protuberances

2018 ◽  
Author(s):  
Wencai Zhu ◽  
Hongtao Gao
Keyword(s):  
Numerical Investigation ◽  
Turbulence Model ◽  
Lift Coefficient ◽  
Pressure Coefficient ◽  
Leading Edge ◽  
Numerical Results ◽  
Spanwise Direction ◽  
Suction Surface ◽  
Pressure Surface ◽  
Edge Profile

The marine rudder with leading-edge protuberances is numerically investigated by SST k-ω turbulence model in present investigations. The newly designed rudder has a sinusoidal leading-edge profile along the spanwise direction. The numerical results show that the newly designed rudder helps to improve the lift coefficient of the rudder. The efficiency of the rudder is improved by adopting the leading-edge protuberances. The results are analyzed by means of streamlines and pressure coefficient. The leading-edge protuberances can delay or overcome the stall. The effect of leading-edge protuberances on the pressure coefficient of pressure surface is very small. However, the pressure coefficient of the suction surface is changed in the vicinity of leading-edge.

Prediction of a Laminar Separation Bubble Over a Controlled-Diffusion Compressor Blade

2000 ◽  
Author(s):  
G. V. Hobson ◽  
S. Weber
Keyword(s):  
Turbulence Model ◽  
Separation Bubble ◽  
Turbulence Models ◽  
Leading Edge ◽  
Equation Model ◽  
Implicit Time ◽  
Controlled Diffusion ◽  
Time Marching ◽  
The One ◽  
Edge Separation

The paper describes the comparison of the prediction of the flow through a cascade of controlled-diffusion compressor blades with two Navier-Stokes solvers. Both codes solved the thin-layer N-S equations, however; one code performed implicit time marching whereas the other performed explicit time marching. Flow predictions were accomplished with the implicit code using the algebraic turbulence model of Baldwin and Lomax and the one-equation model of Spalart and Allmaras, while predictions were made with the explicit code using the two-equation model by Wilcox. Predictions were made of the detailed laser-anemometry measurements of the flow field taken previously in a low-speed cascade wind tunnel. Comparisons were also made with the experimentally measured blade surface pressures and flow visualization of the extent of the laminar leading edge separation bubble. The one-equation turbulence model was combined with an intermittency based transition-length model for comparisons with fully turbulent calculations. Both codes predicted the leading-edge separation bubble satisfactorily when using higher order turbulence models.

scholarly journals Predictions for the Effects of Freestream Turbulence on Turbine Blade Heat Transfer

2004 ◽  
Author(s):  
R. J. Boyle ◽  
Forrest E. Ames ◽  
P. W. Giel
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Steady State ◽  
Reynolds Numbers ◽  
Navier Stokes ◽  
Two Dimensional ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Pressure Surface ◽  
Wide Range

An approach to predicting the effects of freestream turbulence on turbine vane and blade heat transfer is described. Four models for predicting the effects of freestream turbulence were incorporated into a Navier-Stokes CFD analysis. Predictions were compared with experimental data in order to identify an appropriate model for use across a wide range of flow conditions. The analyses were compared with data from five vane geometries and from four rotor geometries. Each of these nine geometries had data for different Reynolds numbers. Comparisons were made for twenty four cases. Steady state calculations were done because all experimental data were obtained in steady state tests. High turbulence levels often result in suction surface transition upstream of the throat, while at low to moderate Reynolds numbers the pressure surface remains laminar. A two-dimensional analysis was used because the flow is predominantly two-dimensional in the regions where freestream turbulence significantly augments surface heat transfer. Because the evaluation of models for predicting turbulence effects can be affected by other factors, the paper discusses modeling for transition, relaminarization, and near wall damping. Quantitative comparisons are given between the predictions and data.

Investigation of Film Cooling Effectiveness on Squealer Tip of a Gas Turbine Blade

2009 ◽  
Author(s):  
Dianliang Yang ◽  
Zhenping Feng ◽  
Xiaobing Yu
Keyword(s):  
Film Cooling ◽  
Turbulence Models ◽  
Tip Clearance ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Height ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Rotating Blade ◽  
Division Line

The effect of the film cooling holes arrangements and the blowing ratio on the tip film cooling effectiveness in a rotating blade with the squealer tip was investigated by using numerical methods in this paper. The first stage rotor blade with squealer tip of GE-E3 engine high pressure turbine was adopted to perform this study. The tip clearance was specified as 1% of the blade height, and the groove depth was specified as 2% of the blade height. The different turbulence models were checked by Kim’s experiment data [1] in 1995, and the standard k-ε turbulence model was chosen to predict the film cooling effectiveness on the blade tip. The film holes were arranged at the tip camber line, the tip division line, the tip pressure side and the pressure surface near tip, respectively. The effect of the holes position on the tip film cooling effectiveness in the rotating blade was studied. The effect of the blowing ratio was analyzed for the cases that the film holes were placed at the tip division line and the pressure surface near tip. The results show that the area-averaged tip film cooling effectiveness reaches the highest when the film holes are placed along the tip division line, and the tip leakage mass flow rate can be reduced by placing the film holes on the pressure surface near tip.

scholarly journals Comparison of Predicted and Experimental Nusselt Number for a Film-Cooled Rotating Blade

1996 ◽  
Author(s):  
Vijay K. Garg ◽  
Reza S. Abhari
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Turbine Blade ◽  
Film Cooling ◽  
Three Dimensional ◽  
Tip Clearance ◽  
Computational Domain ◽  
Pressure Surface ◽  
Rotating Blade ◽  
Film Cooling Holes

The predictions from a three-dimensional Navier-Stokes code have been compared to the Nusselt number data obtained on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only film-cooled rotating blade over which experimental heat transfer data is available for the present comparison. Over 2.25 million grid points are used to compute the flow over the blade. Usually in a film cooling computation on a stationary blade, the computational domain is just one spanwise pitch of the film-cooling holes, with periodic boundary conditions in the span direction. However, for a rotating blade, the computational domain consists of the entire blade span from hub to tip, as well as the tip clearance region. As far as the authors are aware of, the present work offers the first comparison of the prediction of surface heat transfer using a three dimensional CFD code with film injection and the measured heat flux on a fully film-cooled rotating transonic turbine blade. In a detailed comparison with the measured data on the suction surface, a reasonably good comparison is obtained, particularly near the hub section. On the pressure surface, however, the comparison between the data and the prediction is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects due to stator-rotor interaction in the experiments which are not modeled in the present numerical computations.

scholarly journals Origins and Structure of Spike-Type Rotating Stall

2015 ◽  
Vol 137 (5) ◽  
Author(s):  
G. Pullan ◽  
A. M. Young ◽  
I. J. Day ◽  
E. M. Greitzer ◽  
Z. S. Spakovszky
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Suction Surface ◽  
Tip Leakage ◽  
Pressure Surface ◽  
Sequence Of Events ◽  
High Incidence ◽  
Experimental Findings

In this paper, we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a separation at the leading edge due to high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The computed structure of the spike is shown to be consistent with rotor leading edge pressure measurements from the casing of several compressors: the centre of the vortex is responsible for a pressure drop and the partially blocked passages associated with leading edge separations produce a pressure rise. The simulations show leading edge separation and shed vortices over a range of tip clearances including zero. The implication, in accord with recent experimental findings, is that they are not part of the tip clearance vortex. Although the computations always show high incidence to be the cause of the spike, the conditions that give rise to this incidence (e.g., blockage from a corner separation or the tip leakage jet from the adjacent blade) do depend on the details of the compressor.

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