Aerothermal Impact of Stator-Rim Purge Flow and Rotor-Platform Film Cooling on a Transonic Turbine Stage

2010 ◽  
Vol 132 (2) ◽  
Author(s):  
M. Pau ◽  
G. Paniagua ◽  
D. Delhaye ◽  
A. de la Loma ◽  
P. Ginibre
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Fluctuations ◽  
Pressure Sensors ◽  
Suction Side ◽  
Fast Response ◽  
Purge Flow ◽  
Rotor Disk ◽  
Layer Thin Film ◽  
The Impact

The sealing of the stator-rotor gap and rotor-platform cooling are vital to the safe operation of the high-pressure turbine. Contrary to the experience in subsonic turbines, this paper demonstrates the potential to improve the efficiency in transonic turbines at certain rim seal rates. Two types of cooling techniques were investigated: purge gas ejected out of the cavity between the stator rim and the rotor disk, and cooling at the rotor-platform. The tests were carried out in a full annular stage fed by a compression tube at M2is=1.1, Re=1.1×106, and at temperature ratios reproducing engine conditions. The stator outlet was instrumented to allow the aerothermal characterization of the purge flow. The rotor blade was heavily instrumented with fast-response pressure sensors and double-layer thin film gauges. The tests are coupled with numerical calculations performed using the ONERA’s code ELSA. The results indicate that the stator-rotor interaction is significantly affected by the stator-rim seal, both in terms of heat transfer and pressure fluctuations. The flow exchange between the rotor disk cavity and the mainstream passage is mainly governed by the vane trailing edge shock patterns. The purge flow leads to the appearance of a large coherent vortex structure on the suction side of the blade, which enhances the overall heat transfer coefficient due to the blockage effect created. The impact of the platform cooling is observed to be restricted to the platform, with negligible effects on the blade suction side. The platform cooling results in a clear attenuation of pressure pulsations at some specific locations. The experimental and computational fluid dynamics results show an increase in the turbine performance compared with the no rim seal case. A detailed loss breakdown analysis helped to identify the shock loss as the major loss source. The presented results should help designers improve the protection of the rotor platform while minimizing the amount of coolant used.

Aerothermal Impact of Stator-Rim Purge Flow and Rotor-Platform Film Cooling on a Transonic Turbine Stage

2008 ◽  
Author(s):  
M. Pau ◽  
G. Paniagua ◽  
D. Delhaye ◽  
A. de la Loma ◽  
P. Ginibre
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Fluctuations ◽  
Pressure Sensors ◽  
Suction Side ◽  
Turbine Stage ◽  
Purge Flow ◽  
Rotor Disk ◽  
Transonic Turbine ◽  
The Impact

This paper describes the effects on the mainstream flow of two types of cooling techniques in a transonic turbine stage: purge gas ejected out of the cavity between the stator rim and the rotor disk, as well as film cooling gas discharged from the rotor-platform. The tests were carried out in a full annular stage fed by a compression tube, at M2is = 1.1, Re = 1.1×106, and at temperature ratios reproducing engine conditions. The stator outlet was instrumented to allow the aerothermal characterization of the purge flow. The rotor blade was heavily instrumented with fast-response pressure sensors and double-layer thin film gauges. The tests are coupled with numerical calculations performed using the ONERA’s code elsA. The stator-rotor interaction is seen to be significantly affected by the stator-rim seal, both in terms of heat transfer and pressure fluctuations. The flow exchange between the rotor disk cavity and the mainstream passage is mainly governed by the vane shock patterns. The purge flow leads to the appearance of a large coherent vortex structure on the suction side of the blade which enhances the overall heat transfer coefficient due to the blockage effect created. Secondly, the impact of the platform cooling is observed to be restricted to the platform, with negligible effects on the blade suction side. The platform cooling results in a clear attenuation of pressure pulsations at some specific locations. Finally the turbine performance was analyzed, comparing measured and CFD results. A detailed loss breakdown analysis has been done using correlations, in order to isolate the different loss component contributions. The presented results should help designers improve the protection of the rotor platform and minimize the amount of coolant used.

Comparison of Gas Turbine Vane Pressure Side and Suction Side Film Cooling Performance and the Applicability of Superposition

2012 ◽  
Author(s):  
Mats Kinell ◽  
Esa Utriainen ◽  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Botond Barabas
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Gas Turbines ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
The Impact

In order to protect a solid surface exposed to high temperature gaseous flows, e.g. gas turbines and rocket engines, a second gas at lower temperature may be introduced into the hot boundary layer, i.e. one obtains a three temperature problem. The impact of the film cooling on a prototype vane due to variation in blowing ratio, the shape of the hole-outlet and position has been experimentally investigated. The semi-infinite and low conductive test object, initially at a uniform temperature, was exposed to a sudden step change in main flow temperature and a time-resolved surface temperature was measured using an IR camera. By assuming constant values of the heat transfer coefficient and the film cooling effectiveness over time, the heat equation was solved using least squares. The prototype vane was tested for different film cooling row positions on the pressure and suction side. Both cylindrical as well as fan shaped holes were investigated with and without showerhead cooling. The resulting heat transfer coefficient and film cooling effectiveness on the pressure side is compared to flat plate studies and to the results from the suction side. Also, the applicability of using superposition on showerhead cooling and on single/double rows is investigated. Furthermore, the results are compared to other published airfoil film cooling experiments and to CFD analysis for which conclusions are drawn on quantitative and qualitative capabilities of this tool.

scholarly journals An Experimental Study on the Aerodynamics and the Heat Transfer of a Suction Side Film Cooled Large Scale Turbine Cascade

1999 ◽  
Author(s):  
Wolfgang Ganzert ◽  
Leonhard Fottner
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Film Cooling ◽  
Total Pressure ◽  
Large Scale ◽  
Suction Side ◽  
Surface Heat ◽  
Two Dimensional ◽  
Turbine Cascade ◽  
Surface Heat Transfer

As a part of a more complex research program systematic isothermal investigations on the aerodynamics and heat transfer of a large scale turbine cascade with suction side film cooling were carried out. The film cooling through a row of holes at forty percent chord length on the suction side was supplied by a large plenum chamber. Two injection geometries were hitherto tested and evaluated: cylindrical holes with thirty respectively fifty degrees axial inclination angle and no lateral inclination. Typical engine conditions for the Mach and Reynolds number as well as the inlet turbulence level were maintained. The aerodynamic studies are based on steady state pressure measurements. The static profile pressure distribution together with oil-and-dye flow visualisation gives information on the surface flow conditions and boundary layer development especially in the near hole region. The measured data also comprise local and integral total pressure loss coefficients obtained by pressure probe traversing at mid span downstream of the cascade. The heat transfer examination set-up is based on the steady state liquid crystal technique using a compound of a thermochromic sheet combined with an electrical surface heating layer attached on an adiabatic blade corpus. Two dimensional pseudo colour plots are used for the documentation of the local surface heat transfer coefficient distribution and hot spot estimation. Laterally averaged and statistically analysed data of the surface heat transfer is applied in overall heat transfer examinations. All this data is used for a joint aerodynamic flow and surface heat transfer optimisation of a blowing configuration in suction side film cooled turbine cascade. The most important conclusions can be summarised as follows: Aiming at an optimised design of cylindrical film cooling configurations the axial inclination of the holes should be kept low thus diminishing the suction peak value at the cooling position in the profile pressure distribution and decreasing the mainstream deceleration area upstream of the jets. This also leads to reduced total pressure losses. Through the high influence of the blowing on the aerodynamics the flow in the near hole mixing region is highly three-dimensional. This shows significant effects in the two-dimensional surface distribution and the laterally averaged heat transfer coefficient. Oil-and-dye pictures confirm the observations qualitatively.

Clocking Effects of Inlet Non-Uniformities in a Fully Cooled High-Pressure Vane: A Conjugate Heat Transfer Analysis

2015 ◽  
Author(s):  
Duccio Griffini ◽  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Cooling System ◽  
Hot Spot ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Heat Transfer Analysis

A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations have been performed. A preliminary grid sensitivity analysis with uniform inlet flow has been used to quantify the effect of spatial discretization. Turbulence model has been assessed in comparison with available experimental data. The effects of the relative alignment between combustion chamber and high-pressure vanes are then investigated considering realistic inflow conditions in terms of hot spot and swirl. The inlet profiles used are derived from the EU-funded project TATEF2. Two different clocking positions are considered: the first one where hot spot and swirl core are aligned with passage and the second one where they are aligned with the leading edge. Comparisons between metal temperature distributions obtained from conjugate heat transfer simulations are performed evidencing the role of swirl in determining both the hot streak trajectory within the passage and the coolant redistribution. The leading edge aligned configuration is resulted to be the most problematic in terms of thermal load, leading to increased average and local vane temperature peaks on both suction side and pressure side with respect to the passage aligned case. A strong sensitivity of both injected coolant mass flow and heat removed by heat sink effect has also been highlighted for the showerhead cooling system.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

The Influence of Purge Flow Parameters on Heat Transfer and Film Cooling in Turbine Center Frames

2021 ◽  
Author(s):  
Patrick R. Jagerhofer ◽  
Marios Patinios ◽  
Tobias Glasenapp ◽  
Emil Göttlich ◽  
Federica Farisco
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Convective Heat Transfer Coefficient ◽  
Density Ratio ◽  
Constant Heat Flux ◽  
High Pressure Turbine ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow

Abstract Due to stringent environmental legislation and increasing fuel costs, the efficiencies of modern turbofan engines have to be further improved. Commonly, this is facilitated by increasing the turbine inlet temperatures in excess of the melting point of the turbine components. This trend has reached a point where not only the high-pressure turbine has to be adequately cooled, but also components further downstream in the engine. Such a component is the turbine center frame (TCF), having a complex aerodynamic flow field that is also highly influenced by purge-mainstream interactions. The purge air, being injected through the wheelspace cavities of the upstream high-pressure turbine, bears a significant cooling potential for the TCF. Despite this, fundamental knowledge of the influencing parameters on heat transfer and film cooling in the TCF is still missing. This paper examines the influence of purge-to-mainstream blowing ratio, purge-to-mainstream density ratio and purge flow swirl angle on the convective heat transfer coefficient and the film cooling effectiveness in the TCF. The experiments are conducted in a sector-cascade test rig specifically designed for such heat transfer studies using infrared thermography and tailor-made flexible heating foils with constant heat flux. The inlet flow is characterized by radially traversing a five-hole-probe. Three purge-to-mainstream blowing ratios and an additional no purge case are investigated. The purge flow is injected without swirl and also with engine-similar swirl angles. The purge swirl and blowing ratio significantly impact the magnitude and the spread of film cooling in the TCF. Increasing blowing ratios lead to an intensification of heat transfer. By cooling the purge flow, a moderate variation in purge-to-mainstream density ratio is investigated, and the influence is found to be negligible.

THE INFLUENCE OF PURGE FLOW PARAMETERS ON HEAT TRANSFER AND FILM COOLING IN TURBINE CENTER FRAMES

2021 ◽  
pp. 1-26
Author(s):  
Patrick René Jagerhofer ◽  
Marios Patinios ◽  
Tobias Glasenapp ◽  
Emil Goettlich ◽  
Federica Farisco
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Density Ratio ◽  
Constant Heat Flux ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Turbofan Engines ◽  
Blowing Ratio ◽  
Purge Flow

Abstract The imperative improvement in the efficiency of turbofan engines is commonly facilitated by increasing the turbine inlet temperature. This development has reached a point where also components downstream of the high-pressure turbine have to be adequately cooled. Such a component is the turbine center frame (TCF), known for a complex aerodynamic flow highly influenced by purge-mainstream interactions. The purge air, being injected through the wheelspace cavities of the upstream high-pressure turbine, bears a significant cooling potential for the TCF. Despite this, fundamental knowledge of the influencing parameters on heat transfer and film cooling in the TCF is still missing. This paper examines the influence of purge-to-mainstream blowing ratio, density ratio and purge swirl angle on heat transfer and film cooling in the TCF. The experiments are conducted in a sector-cascade test rig specifically designed for such heat transfer studies using infrared thermography and tailor-made flexible heating foils with constant heat flux. Three purge-to-mainstream blowing ratios and an additional no purge case are investigated. The purge flow is injected without swirl and also with engine-similar swirl angles. The purge swirl and blowing ratio significantly impact the magnitude and the spread of film cooling in the TCF. Increasing blowing ratios lead to an intensification of heat transfer. By cooling the purge flow, a moderate variation in purge-to-mainstream density ratio is investigated, and the influence is found to be negligible.

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.

Overall Effectiveness of a Blade Endwall With Jet Impingement and Film Cooling

2013 ◽  
Vol 136 (3) ◽  
Author(s):  
Amy Mensch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Jet Impingement ◽  
Heat Transfer Analysis ◽  
Transfer Model ◽  
Convective Cooling ◽  
Impingement Cooling ◽  
Internal Impingement ◽  
The Impact ◽  
Overall Effectiveness

Ever-increasing thermal loads on gas turbine components require improved cooling schemes to extend component life. Engine designers often rely on multiple thermal protection techniques, including internal cooling and external film cooling. A conjugate heat transfer model for the endwall of a seven-blade cascade was developed to examine the impact of both convective cooling and solid conduction through the endwall. Appropriate parameters were scaled to ensure engine-relevant temperatures were reported. External film cooling and internal jet impingement cooling were tested separately and together for their combined effects. Experiments with only film cooling showed high effectiveness around film-cooling holes due to convective cooling within the holes. Internal impingement cooling provided more uniform effectiveness than film cooling, and impingement effectiveness improved markedly with increasing blowing ratio. Combining internal impingement and external film cooling produced overall effectiveness values as high as 0.4. A simplified, one-dimensional heat transfer analysis was used to develop a prediction of the combined overall effectiveness using results from impingement only and film cooling only cases. The analysis resulted in relatively good predictions, which served to reinforce the consistency of the experimental data.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1—Experimental Technique and Results

1992 ◽  
Vol 114 (4) ◽  
pp. 734-740 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine nondimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio, and gas-to-coolant density ratio. The results show that the secondary flow and horseshoe vortex act on the coolant, which is convected toward the suction side of the NG V endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50 to 75 percent.

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