Computational Predictions of Heat Transfer and Film-Cooling for a Turbine Blade With Nonaxisymmetric Endwall Contouring

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
Stephen P. Lynch ◽  
Karen A. Thole ◽  
Atul Kohli ◽  
Christopher Lehane
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Turbine Engine ◽  
Cooling Flow ◽  
Endwall Contouring ◽  
Dynamics Simulations ◽  
Endwall Film Cooling

Three-dimensional contouring of the compressor and turbine endwalls in a gas turbine engine has been shown to be an effective method of reducing aerodynamic losses by mitigating the strength of the complex vortical structures generated at the endwall. Reductions in endwall heat transfer in the turbine have been also previously measured and reported in literature. In this study, computational fluid dynamics simulations of a turbine blade with and without nonaxisymmetric endwall contouring were compared to experimental measurements of the exit flowfield, endwall heat transfer, and endwall film-cooling. Secondary kinetic energy at the cascade exit was closely predicted with a simulation using the SST k-ω turbulence model. Endwall heat transfer was overpredicted in the passage for both the SST k-ω and realizable k-ε turbulence models, but heat transfer augmentation for a nonaxisymmetric contour relative to a flat endwall showed fair agreement to the experiment. Measured and predicted film-cooling results indicated that the nonaxisymmetric contouring limits the spread of film-cooling flow over the endwall depending on the interaction of the film with the contour geometry.

Computational Predictions of Heat Transfer and Film-Cooling for a Turbine Blade With Non-Axisymmetric Endwall Contouring

2010 ◽  
Author(s):  
Stephen P. Lynch ◽  
Karen A. Thole ◽  
Atul Kohli ◽  
Christopher Lehane
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Turbine Engine ◽  
Cooling Flow ◽  
Endwall Contouring ◽  
Dynamics Simulations ◽  
Endwall Film Cooling

Three-dimensional contouring of the compressor and turbine endwalls in a gas turbine engine has been shown to be an effective method of reducing aerodynamic losses by mitigating the strength of the complex vortical structures generated at the endwall. Reductions in endwall heat transfer in the turbine have been also previously measured and reported in the literature. In this study, computational fluid dynamics simulations of a turbine blade with and without non-axisymmetric endwall contouring were compared to experimental measurements of the exit flowfield, endwall heat transfer and endwall film-cooling. Secondary kinetic energy at the cascade exit was closely predicted with a simulation using the SST k-ω turbulence model. Endwall heat transfer was overpredicted in the passage for both the SST k-ω and realizable k-ε turbulence models, but heat transfer augmentation for a non-axisymmetric contour relative to a flat endwall showed fair agreement to the experiment. Measured and predicted film-cooling results indicated that the non-axisymmetric contouring limits the spread of film-cooling flow over the endwall depending upon the interaction of the film with the contour geometry.

scholarly journals Comparison of Two-Equation Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades

1997 ◽  
Author(s):  
Vijay K. Garg ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Computer Time ◽  
Suction Surface ◽  
The Heat Transfer Coefficient

A three-dimensional Navier-Stokes code has been used to compute the heat transfer coefficient on two film-cooled turbine blades, namely the VKI rotor with six rows of cooling holes including three rows on the shower head, and the C3X vane with nine rows of holes including five rows on the shower head. Predictions of heat transfer coefficient at the blade surface using three two-equation turbulence models, specifically, Coakley’s q-ω model, Chien’s k-ε model and Wilcox’s k-ω model with Menter’s modifications, have been compared with the experimental data of Camci and Arts (1990) for the VKI rotor, and of Hylton et al. (1988) for the C3X vane along with predictions using the Baldwin-Lomax (B-L) model taken from Garg and Gaugler (1995). It is found that for the cases considered here the two-equation models predict the blade heat transfer somewhat better than the B-L model except immediately downstream of the film-cooling holes on the suction surface of the VKI rotor, and over most of the suction surface of the C3X vane. However, all two-equation models require 40% more computer core than the B-L model for solution, and while the q-ω and k-ε models need 40% more computer time than the B-L model, the k-ω model requires at least 65% more time due to slower rate of convergence. It is found that the heat transfer coefficient exhibits a strong spanwise as well as streamwise variation for both blades and all turbulence models.

scholarly journals The Design of an Improved Endwall Film-Cooling Configuration

1998 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
High Static Pressure ◽  
Endwall Film Cooling ◽  
Aerodynamic Losses

The endwall film-cooling cooling configuration investigated by Friedrichs et al. (1996, 1997) had in principle sufficient cooling flow for the endwall, but in practice, the redistribution of this coolant by secondary flows left large endwall areas uncooled. This paper describes the attempt to improve upon this datum cooling configuration by redistributing the available coolant to provide a better coolant coverage on the endwall surface, whilst keeping the associated aerodynamic losses small. The design of the new, improved cooling configuration was based on the understanding of endwall film-cooling described by Friedrichs et al. (1996, 1997). Computational fluid dynamics were used to predict the basic flow and pressure field without coolant ejection. Using this as a basis, the above described understanding was used to place cooling holes so that they would provide the necessary cooling coverage at minimal aerodynamic penalty. The simple analytical modelling developed in Friedrichs et al. (1997) was then used to check that the coolant consumption and the increase in aerodynamic loss lay within the limits of the design goal. The improved cooling configuration was tested experimentally in a large scale, low speed linear cascade. An analysis of the results shows that the redesign of the cooling configuration has been successful in achieving an improved coolant coverage with lower aerodynamic losses, whilst using the same amount of coolant as in the datum cooling configuration. The improved cooling configuration has reconfirmed conclusions from Friedrichs et al. (1996, 1997); firstly, coolant ejection downstream of the three-dimensional separation lines on the endwall does not change the secondary flow structures; secondly, placement of holes in regions of high static pressure helps reduce the aerodynamic penalties of platform coolant ejection; finally, taking account of secondary flow can improve the design of endwall film-cooling configurations.

Overall Effectiveness and Flowfield Measurements for an Endwall With Nonaxisymmetric Contouring

2015 ◽  
Vol 138 (3) ◽  
Author(s):  
Amy Mensch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Transfer Model ◽  
Internal Cooling ◽  
Blowing Ratio ◽  
Passage Vortex ◽  
Endwall Contouring ◽  
Overall Effectiveness

Endwall contouring is a technique used to reduce the strength and development of three-dimensional secondary flows in a turbine vane or blade passage in a gas turbine. The secondary flows locally affect the external heat transfer, particularly on the endwall surface. The combination of external and internal convective heat transfer, along with solid conduction, determines component temperatures, which affect the service life of turbine components. A conjugate heat transfer model is used to measure the nondimensional external surface temperature, known as overall effectiveness, of an endwall with nonaxisymmetric contouring. The endwall cooling methods include internal impingement cooling and external film cooling. Measured values of overall effectiveness show that endwall contouring reduces the effectiveness of impingement alone, but increases the effectiveness of film cooling alone. Given the combined case of both impingement and film cooling, the laterally averaged overall effectiveness is not significantly changed between the flat and the contoured endwalls. Flowfield measurements indicate that the size and location of the passage vortex changes as film cooling is added and as the blowing ratio increases. Because endwall contouring can produce local effects on internal cooling and film cooling performance, the implications for heat transfer should be considered in endwall contour designs.

Effect of Various Injection Hole Shapes on Film Cooling of Turbine Blade

2003 ◽  
Author(s):  
S.-M. Kim ◽  
Youn J. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Three Dimensional ◽  
Cylindrical Body ◽  
Navier Stokes ◽  
Cylinder Surface ◽  
Injection Hole ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Cooling Flow

Dispersion of coolant jets in a film cooling flow field is the result of a highly complex interaction between the film cooling jets and the mainstream. In order to investigate the effects of injection hole shapes and injection angle on the film cooling of turbine blade, four models having cylindrical and laterally-diffused holes were used. Three-dimensional Navier-Stokes code with k – ε model was used to compute the film cooling coefficient on the turbine blade. A multi-block grid system was generated that was nearly orthogonal to the various surfaces. Mainstream Reynolds number based on the cylinder diameter was 7.1 × 104. The turbulence intensity kept at 5.0% for all inlets. The effect of coolant flow rates was studied for blowing ratios of 0.9, 1.3 and 1.6, respectively. The temperature distribution of the cylindrical body surface is visualized by infrared thermography (IRT) and compared with computational results. Results show that the effects of injection hole shape and injection angle increase as the blowing ratio increases. As lateral injection angle increases, the adiabatic film cooling effectiveness is more broadly distributed and the area protected by coolant increases. The mass flow rate of the coolant through the first-row holes is less than that through the second-row holes due to the pressure distribution around the cylinder surface.

Influence of Film Cooling Unsteadiness on Turbine Blade Leading Edge Heat Flux

2012 ◽  
Vol 134 (7) ◽  
Author(s):  
James L. Rutledge ◽  
Paul I. King ◽  
Richard B. Rivir
Keyword(s):  
Heat Flux ◽  
Turbine Blade ◽  
Film Cooling ◽  
Flow Behavior ◽  
Leading Edge ◽  
Coolant Flow ◽  
Adiabatic Wall ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Cooling Flow

Film cooling in the hot gas path of a gas turbine engine can protect components from the high temperature main flow, but it generally increases the heat transfer coefficient h partially offsetting the benefits in reduced adiabatic wall temperature. We are thus interested in adiabatic effectiveness η and h which are combined in a formulation called net heat flux reduction (NHFR). Unsteadiness in coolant flow may arise due to inherent unsteadiness in the external flow or be intentionally introduced for flow control. In previous work it has been suggested that pulsed cooling flow may, in fact, offer benefits over steady blowing in either improving NHFR or reducing the mass flow requirements for matched NHFR. In this paper we examine this hypothesis for a range of steady and pulsed blowing conditions. We use a new experimental technique to analyze unsteady film cooling on a semicircular cylinder simulating the leading edge of a turbine blade. The average NHFR with pulsed and steady film cooling is measured and compared for a single coolant hole located 21.5° downstream from the leading edge stagnation line, angled 20° to the surface and 90° to the streamwise direction. We show that for moderate blowing ratios at blade passing frequencies, steady film flow yields better NHFR. At higher coolant flow rates beyond the optimum steady blowing ratio, however, pulsed film cooling can be advantageous. We present and demonstrate a prediction technique for unsteady blowing at frequencies similar to the blade passing frequency that only requires the knowledge of steady flow behavior. With this important result, it is possible to predict when pulsing would be beneficial or detrimental.

scholarly journals Two-Equation Turbulence Models for Prediction of Heat Transfer on a Transonic Turbine Blade

2001 ◽  
Author(s):  
Vijay K. Garg ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Suction Side ◽  
Navier Stokes ◽  
Production Term ◽  
Passage Vortex ◽  
Sst Model ◽  
Transonic Turbine

Two versions of the two-equation k-ω model and a shear stress transport (SST) model are used in a three-dimensional, multi-block, Navier-Stokes code to compare the detailed heat transfer measurements on a transonic turbine blade. It is found that the SST model resolves the passage vortex better on the suction side of the blade, thus yielding a better comparison with the experimental data than either of the k-ω models. However, the comparison is still deficient on the suction side of the blade. Use of the SST model does require the computation of distance from a wall, which for a multi-block grid, such as in the present case, can be complicated. However, a relatively easy fix for this problem was devised. Also addressed are issues such as (1) computation of the production term in the turbulence equations for aerodynamic applications, and (2) the relation between the computational and experimental values for the turbulence length scale, and its influence on the passage vortex on the suction side of the turbine blade.

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.

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