Stephen P. Lynch

Heat Transfer for a Turbine Blade With Non-Axisymmetric Endwall Contouring

Complex vortical secondary flows that are present near the endwall of an axial gas turbine blade are responsible for high heat transfer rates and high aerodynamic losses. The application of non-axisymmetric, three-dimensional contouring to the endwall surface has been shown to reduce the strength of the vortical flows and decrease total pressure losses when compared to a flat endwall. The reduction of secondary flow strength with non-axisymmetric contouring might also be expected to reduce endwall heat transfer. In this study, measurements of endwall heat transfer were taken for a low-pressure turbine blade geometry with both flat and three-dimensional contoured endwalls. Endwall oil flow visualization indicated a reduction in the passage vortex strength for the contoured endwall geometry. Heat transfer levels were reduced by 20 percent in regions of high heat transfer with the contoured endwall, as compared to the flat endwall. The heat transfer benefit of the endwall contour was not affected by changes in the cascade Reynolds number.Copyright

Heat Transfer for a Turbine Blade With Nonaxisymmetric Endwall Contouring

turbine blade are responsible for high heat transfer rates and high aerodynamic losses. The application of nonaxisymmetric, three-dimensional contouring to the endwall surface has been shown to reduce the strength of the vortical flows and decrease total pressure losses when compared with a flat endwall. The reduction in secondary flow strength with nonaxisymmetric contouring might also be expected to reduce endwall heat transfer. In this study, measurements of endwall heat transfer were taken for a low-pressure turbine blade geometry with both flat and three-dimensional contoured endwalls. Endwall oil flow visualization indicated a reduction in the passage vortex strength for the contoured endwall geometry. Heat transfer levels were reduced by 20% in regions of high heat transfer with the contoured endwall, as compared with the flat endwall. The heat transfer benefit of the endwall contour was not affected by changes in the cascade Reynolds number. DOI: 10.1115/1.4000542 A fundamental issue in the design of gas turbine engines is the detrimental effects of vortical flows that are present near the junction of the airfoil and its endwall. These flows, generally termed secondary flows, increase aerodynamic losses through the turbine leading to lower overall engine efficiency. Furthermore, the vortical flows result in high heat transfer rates due to mean flow distortion and high turbulence levels. For highly loaded parts such as the low-pressure turbine blade in this study, part durability can be significantly impacted by high heat transfer. It is obvious that reducing or eliminating the effects of the secondary flows would result in appreciable benefits in an engine, and thus, much research has been conducted to understand these flows and control them. Past research has shown that three-dimensional modifications to the endwall surface, generally referred to as nonaxisymmetric contouring, are an effective method of reducing or eliminating secondary flows. The peaks and valleys in the endwall locally decrease or increase the static pressure, respectively, which can reduce the pressure gradient between adjacent airfoils that strengthens secondary flows. Weaker secondary flows result in lower aerodynamic losses, as has been documented by many researchers. However, few studies have considered the effect of nonaxisymmetric contouring on endwall heat transfer, and thus, the overall benefit of contouring in an engine is not completely understood as of yet. This paper discusses the effect of nonaxisymmetric contouring on the endwall heat transfer for a low-pressure turbine blade. The effect of varying the cascade Reynolds number for both the flat and contoured endwalls is also considered.

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

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 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-e 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 nonaxisymmetric contouring limits the spread of film-cooling flow over the endwall depending upon the interaction of the film with the contour geometry.

The Effect of Combustor-Turbine Interface Gap Leakage on the Endwall Heat Transfer for a Nozzle Guide Vane

To enable turbine components to withstand high combustion temperatures, they are cooled by air routed from the compressor, which can leak through gaps between components. These gaps vary in size from thermal expansions that take place. The leakage flow between the interface of the combustor and the turbine, in particular, interacts with the flowfield along the endwall. This study presents measurements of adiabatic cooling effectiveness and heat transfer coefficients on the endwall of a first vane, with the presence of leakage flow through a flush slot upstream of the vane. The effect of axial contraction of the slot width due to thermal expansion of the engine was tested for two blowing rates. Contracting the slot width, while maintaining the slot mass flow, resulted in a larger coolant coverage area and higher effectiveness values, as well as slightly lower heat transfer coefficients. Matching the momentum flux ratio of the leakage flow from the nominal and contracted slot widths lowered both cooling effectiveness and heat transfer coefficients for the contracted slot flow. Comparison of the coolant coverage pattern to the measured endwall shear stress topology indicated that the trajectory of the slot coolant was dictated by the complex endwall flow.

The Effect of the Combustor-Turbine Slot and Midpassage Gap on Vane Endwall Heat Transfer

Turbine vanes are generally manufactured as singleor double-airfoil sections that are assembled into a full turbine disk. The gaps between the individual sections, as well as a gap between the turbine disk and the combustor upstream, provide leakage paths for relatively higher-pressure coolant flows. This leakage is intended to prevent ingestion of the hot combustion flow in the primary gas path. At the vane endwall, this leakage flow can interfere with the complex vortical flow present there and thus affect the heat transfer to that surface. To determine the effect of leakage flow through the gaps, heat transfer coefficients were measured along a first-stage vane endwall and inside the midpassage gap for a large-scale cascade with a simulated combustor-turbine interface slot and a midpassage gap. For increasing combustor-turbine leakage flows, endwall surface heat transfer coefficients showed a slight increase in heat transfer. The presence of the midpassage gap, however, resulted in high heat transfer near the passage throat where flow is ejected from that gap. Computational simulations indicated that a small vortex created at the gap flow ejection location contributed to the high heat transfer. The measured differences in heat transfer for the various midpassage gap flowrates tested did not appear to have a significant effect. DOI: 10.1115/1.4002950

HEAT TRANSFER AND FILM COOLING ON A CONTOURED BLADE ENDWALL WITH PLATFORM GAP LEAKAGE

Turbine blade components in an engine are typically designed with gaps between parts due to manufacturing, assembly, and operational considerations. Coolant is provided to these gaps to limit the ingestion of hot combustion gases. The interaction of the gaps, their leakage flows, and the complex vortical flow at the endwall of a turbine blade can significantly impact endwall heat transfer coefficients and the effectiveness of the leakage flow in providing localized cooling. In particular, a platform gap through the passage, representing the mating interface between adjacent blades in a wheel, has been shown to have a significant effect. Other important turbine blade features present in the engine environment are non-axisymmetric contouring of the endwall, and an upstream rim seal with a gaspath cavity, which can reduce and increase endwall vortical flow, respectively. To understand the platform gap leakage effect in this environment, measurements of endwall heat transfer and film cooling effectiveness were performed in a scaled blade cascade with a non-axisymmetric contour in the passage. A rim seal with a cavity, representing the overlap interface between a stator and rotor, was included upstream of the blades and a nominal purge flowrate of 0.75% of the mainstream was supplied to the rim seal. Results indicated that endwall heat transfer coefficients increased as platform gap net leakage increased from 0% to 0.6% of the mainstream flowrate, but net heat flux to the endwall was reduced due to high cooling effectiveness of the leakage flow.Copyright

Comparison of the Three-Dimensional Boundary Layer on Flat Versus Contoured Turbine Endwalls

The boundary layer on the endwall of an axial turbomachine passage is influenced by streamwise and cross-stream pressure gradients, as well as a large streamwise vortex, that develop in the passage. These influences distort the structure of the boundary layer and result in heat transfer and friction coefficients that differ significantly from simple two-dimensional boundary layers. Three-dimensional contouring of the endwall has been shown to reduce the strength of the large passage vortex and reduce endwall heat transfer, but the mechanisms of the reductions on the structure of the endwall boundary layer are not well understood.This study describes three-component measurements of mean and fluctuating velocities in the passage of a turbine blade obtained with a laser Doppler velocimeter. Friction coefficients obtained with the oil film interferometry method were compared to measured heat transfer coefficients. In the passage, the strength of the large passage vortex was reduced with contouring. Regions where heat transfer was increased by endwall contouring corresponded to elevated turbulence levels compared to the flat endwall, but the variation in boundary layer skew across the passage was reduced with contouring.Copyright

Simulations of Multiphase Particle Deposition on a Nonaxisymmetric Contoured Endwall With Film-Cooling

Designing turbine components for maximum aerodynamic performance with adequate cooling is a critical challenge for gas turbine engineers, particularly at the endwall of a turbine, due to complex secondary flows. To complicate matters, impurities from the fuel and intake air can deposit on film-cooled components downstream of the combustor. Deposition-induced roughness can reduce cooling effectiveness and aerodynamic performance dramatically. One method commonly used for reducing the effects of secondary flows on aerodynamic performance is endwall contouring. The current study evaluates deposition effects on endwall contouring given the change to the secondary flow pattern. For the current study, deposition was dynamically simulated in a turbine cascade to determine its effects on film-cooling with and without endwall contouring. Computationally predicted impactions were in qualitative agreement with experimental deposition simulations, showing that contouring reduced deposition around strategically placed film-cooling holes. Deposition reduced cooling effectiveness by 50% on a flat endwall and 40% on an identically cooled contoured endwall. Although 40% is still a dramatic reduction in effectiveness, the method of using the endwall contouring to alter deposition effects shows promise.

Time-resolved PIV Measurements of the Effect of Freestream Turbulence on Horseshoe Vortex Dynamics

The horseshoe vortex system is a common flow feature in many natural and industrial flows occurring near the junction of a blunt obstacle with the endwall surface. In industrial settings, such as in high temperature gas turbine engines, the dynamic behavior of the horseshoe vortex has been shown to contribute significantly to the pressure loading and heat transfer behavior on surfaces near the leading edge of the obstacle. Fundamental studies of the horseshoe vortex have characterized its time mean and dynamic behavior at low freestream turbulence conditions, and studies using industry relevant geometries, such as cylindrical pin fin arrays common to cooling applications, have captured dynamic behavior of the vortex at high freestream turbulence. The isolated effect of high freestream turbulence on the dynamic behavior of the vortex, independent of upstream wake effects found in pin fin arrays and other industry geometries, however, is not well understood. This study seeks use high-speed time resolved stereo particle image velocimetry (SPIV) measurements of the horseshoe vortex system taken at varied freestream turbulence levels in front of a single Rood wing obstacle to better understand the isolated effect of freestream turbulence on the vortex position and vortex breakdown dynamics.

Aerodynamic Loss for a Turbine Blade With Endwall Leakage Features and Contouring

Secondary flows near the endwall of a turbine blade contribute to a loss in aerodynamic performance in the engine. Further reductions in performance occur when the secondary flows interact with leakage flow from necessary clearance features, such as the clearance gap between the blade rotor and an upstream stator or gaps between adjacent blade platforms. Non-axisymmetric endwall contouring has been shown to reduce the strength of secondary flows near the endwall, but relatively little research has been done to test the sensitivity of the contouring to the endwall leakage features. This paper describes aerodynamic measurements taken downstream of a cascade with representative leakage features. In general, upstream leakage flow with swirl relative to the blade increased aerodynamic loss, relative to leakage that was matched to the blade wheelspeed. Non-axisymmetric contouring for an endwall without a platform gap reduced underturning of the flow but had no effect on overall loss, relative to a flat endwall without a gap. A contoured endwall with a platform gap had 12% higher mixed-out loss than a contoured endwall without a gap.Copyright