Endwall Film Cooling Effects on Secondary Flows in a Contoured Endwall Nozzle Vane

2010 ◽  
Vol 132 (4) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi ◽  
Marco Quattrore
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Secondary Flows ◽  
Intensity Level ◽  
Planar Case ◽  
Nozzle Vane ◽  
Cooling Flow ◽  
Cooling Effects ◽  
Endwall Contouring ◽  
Gas Turbine Nozzle

The present paper investigates the effects of endwall injection of cooling flow on the aerodynamic performance of a nozzle vane cascade with endwall contouring. Tests have been performed on a seven vane cascade with a geometry typical of a real gas turbine nozzle vane. The cooling scheme consists of four rows of cylindrical holes. Tests have been carried out at low speed (Ma2is=0.2) with a low inlet turbulence intensity level (1.0%) and with a coolant to mainstream mass flow ratio varied in the range from 0% (solid endwall) to 2.5%. Energy loss coefficient, secondary vorticity, and outlet angle distributions were computed from five-hole probe measured data. Contoured endwall results, with and without film cooling, were compared with planar endwall data. Endwall contouring was responsible for a significant overall loss decrease, as a result of the reduction in both profile and planar side secondary flows losses; a loss increase on the contoured side was instead observed. Like as for the planar endwall, even for the contoured endwall, coolant injection modifies secondary flows, reducing their intensity, but the relevance of the changes is reduced. Nevertheless, for all the tested injection conditions, secondary losses on the contoured side are always higher than in the planar case, while contoured cascade overall losses are lower. A unique minimum overall loss injection condition was found for both tested geometries, which corresponds to an injected mass flow rate of about 1.0%.

Endwall Film Cooling Effects on Secondary Flows in a Contoured Endwall Nozzle Vane

2008 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi ◽  
Marco Quattrore
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Secondary Flows ◽  
Intensity Level ◽  
Planar Case ◽  
Nozzle Vane ◽  
Cooling Flow ◽  
Cooling Effects ◽  
Endwall Contouring ◽  
Gas Turbine Nozzle

The present paper investigates the effects of endwall injection of cooling flow on the aerodynamic performance of a nozzle vane cascade with endwall contouring. Tests have been performed on a 7 vane cascade with a geometry typical of a real gas turbine nozzle vane. The cooling scheme consists of four rows of cylindrical holes. The same cooling scheme, applied to a flat endwall, was already investigated by the authors. Tests have been carried out at low speed (M2is = 0.2) with a low inlet turbulence intensity level (1.0%) and with a coolant to mainstream mass flow ratio varied in the range from zero (solid endwall) to 2.5%. Energy loss coefficient, secondary vorticity and outlet angle distributions were computed from 5-hole probe measured data. Contoured endwall results, with and without film cooling, were compared to planar endwall data. Endwall contouring was responsible for a significant overall loss decrease, thanks to the reduction of both profile and planar side secondary flows losses; a loss increase on the contoured side was instead observed. Like as for the planar endwall, even for contoured endwall coolant injection modifies secondary flows, reducing their intensity, but the relevance of the changes is reduced. Nevertheless for all the tested injection conditions, secondary losses on the contoured side are always higher than in the planar case, while contoured cascade overall losses are lower. A unique minimum overall loss injection condition was found for both tested geometries, corresponding to an injected mass flow rate of about 1.0%.

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

2015 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Intensity Level ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.

Turbine Nozzle Endwall Inlet Film Cooling: The Effect of a Back-Facing Step

2003 ◽  
Author(s):  
Luzeng Zhang ◽  
Hee Koo Moon
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Mass Flow ◽  
Film Cooling ◽  
Secondary Flows ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Wall Flow ◽  
Cooling Air ◽  
Near Wall

Film cooling effectiveness was measured on a contoured endwall surface using the pressure sensitive paint (PSP) technique. A double staggered row of holes was adopted to supply cooling air in front of the nozzle leading edges. To simulate realistic engine configuration, a back-facing step was built, which was located upstream from the film injection. Nitrogen gas was used to simulate film cooling flow as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to be from 0.5% to 3.0% of the mainstream mass flow. Film effectiveness distributions were measured on the endwall surface for both smooth (baseline) and back-facing step inlet configurations. For the smooth inlet case, film effectiveness increased nonlinearly with mass flow rate, indicating a strong interference between the cooling jets and the secondary flows. At lower mass flow ratios, the secondary flow dominated the near wall flow field, resulting in a low film effectiveness value. At higher mass flow ratios, the cooling jet momentum dominated the near wall flow field, resulting in a higher film effectiveness. For the back-facing step inlet configuration, the values of film effectiveness were reduced significantly, suggesting a stronger secondary flow interaction. In addition to the comparison between the smooth and back-facing step inlet configurations, comparison to previous data by the authors on a flat endwall was also made.

Turbine Nozzle Endwall Inlet Film Cooling: The Effect of a Back-Facing Step and Velocity Ratio

2004 ◽  
Author(s):  
Luzeng Zhang ◽  
Hee Koo Moon
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Velocity Ratio ◽  
Secondary Flows ◽  
Tracer Gas ◽  
Nitrogen Gas ◽  
Cooling Flow ◽  
Pressure Sensitive ◽  
Jet Velocity ◽  
Cooling Air

Endwall inlet film cooling serves two purposes: to suppress the secondary flows and to provide effective cooling. To optimize endwall inlet film cooling, the combined effects of a back facing step and jet velocity ratio were studied in a warm cascade simulating realistic engine conditions. Film effectiveness distribution was measured on a nozzle endwall surface using the pressure sensitive paint (PSP) technique. A double staggered row of holes was used to supply cooling air in front of the nozzle leading edges. Changing the diameter of the film injection hole varied the velocity ratio and the back-facing step was designed to simulate the discontinuity of the nozzle inlet to the combustor exit cone. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to be from 0.5% to 3.0% of the mainstream mass flow. The film effectiveness distribution was locally measured for each of the cooling mass flows. It was demonstrated that by optimizing the jet velocity ratio the adverse effect of the back-facing step could be reduced, particularly for the range of mass flow practical in design. The pattern of the film effectiveness distribution suggested the opposite effect of the film injection and the back-facing step on the secondary flows, while one suppresses and the other enhances it.

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.

Platform Film Cooling Investigation on an HP Nozzle Vane Cascade With Discrete Shaped Holes and Slot Film Cooling

2020 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
L. Abba ◽  
L. Pestelli
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Heat Transfer Performance ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Intensity Level ◽  
Transfer Performance ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

Abstract The present paper reports on an experimental investigation on the aerodynamic and heat transfer performance of different platform cooling schemes: two based on cylindrical and shaped holes and one featuring a slot located upstream of the leading edge plane simulating the combustor to stator interface gap. Tests were run on a 6-vane cascade operated at an isentropic cascade exit Mach number of 0.4 and a significant inlet turbulence intensity level of about 9%. The cooling schemes were first tested to quantify their impact on secondary flows and related losses for variable injection conditions. Heat transfer performance was then assessed through adiabatic film cooling effectiveness and heat transfer coefficient measurements. The Net Heat Flux Reduction parameter was then computed to critically assess the cooling schemes. When compared with the cylindrical hole scheme, shaped holes outperform for all tested injection rates, while the slot alone is able to thermally protect only the front of the passage. Discrete holes are required to cool the platform region along the whole pressure side and the suction side leading edge region.

Cross Flow Effects on Endwall Heat Transfer in Film-Cooled Turbine Nozzle Guide Vanes

2013 ◽  
Author(s):  
Rebecca Reviol ◽  
Roman Franze ◽  
Martin Böhle ◽  
Kenichiro Takeishi ◽  
Alexander Wiedermann
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cross Flow ◽  
Turbulence Models ◽  
Secondary Flows ◽  
Stagnation Region ◽  
Naphthalene Sublimation ◽  
Sublimation Method ◽  
Cooling Effects ◽  
Nozzle Guide

Film cooling effects on endwalls in the stagnation point region are of special interest since the heat transfer is influenced drastically by secondary flows. Additionally, a complex vortex structure exists along the stagnation streamline which influences heat transfer on the endwall. The flow phenomenon is described and discussed in the open literature but it is still difficult to predict the heat transfer on the endwall and the turbine profile by CFD methods with sufficient accuracy. In this paper it is examined how the flow field in the stagnation region should be simulated using CFD. The effect of meshes with various grid resolutions and turbulence models as k-ε-, k-ω-SST- and DES-turbulence models have been investigated. The CFD-data are compared with the experimental results obtained by Naphthalene Sublimation Method, Pressure Sensitive Paint, Laser Induced Fluorescence and Particle Image Velocimetry. Three cases, namely film cooling on a flat plate, the endwall flow near a symmetrical airfoil and the symmetrical airfoil with endwall film cooling, are examined in detail.

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.

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.

Flow Migration and Mixing in a High Pressure Turbine Rotor Passage With Axisymmetrically Contoured Endwall and Leakage Flow

2015 ◽  
Author(s):  
Reema Saxena ◽  
Arya Ayaskanta ◽  
Terrence W. Simon ◽  
Hee-Koo Moon ◽  
Luzeng J. Zhang
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Film Cooling ◽  
Thermal Loading ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Flow Rates ◽  
Flow Mixing ◽  
Cooling Effects

The flow field in the passage of a high pressure gas turbine is quite complex, involving strong secondary flows, transverse pressure gradients and strong streamwise acceleration. This complexity may have an adverse effect on cooling of the hub endwall, which is subjected to high thermal loading due to the flat combustor exit temperature profile of modern low-NOx systems. Therefore, given material limitations, better cooling management techniques that can be included with certainty in new gas turbine designs are needed. In the present study, film cooling has been investigated experimentally in a stationary linear cascade. The flow is representative of a high pressure gas turbine rotor with combustor liner coolant introduced to the approach flow. Focus is on the endwall axisymmetric contouring and the cooling effect of leakage flow bled from the compressor through the stator-rotor disc cavity. Two endwall contours, ‘shark nose’ (gradual slope over a larger distance) and ‘dolphin nose’ (steep slope over a shorter distance), are considered and comparison is made under conditions of three mass flow rates (MFR) of leakage, 0.5%, 1.0% and 1.5% of the approach flow rate. The performance of both endwall contours is compared at different streamwise locations in terms of adiabatic effectiveness values over the endwall. This study gives enhanced insight into the physics of coolant flow mixing, migration and subsequent coverage over the endwall. The results show the cooling effects of the contoured shapes over a range of leakage flow rates in the strong secondary flow environment. It is found that the leakage flow plays a crucial role in enhancing coolant coverage over the endwall. To add to our knowledge of mixing effects, detailed thermal field data are taken in the leakage flow discharge region. Doing so helps explain the behavior of the flow as it is ejected into the passage and interacts with the mainstream flow.

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