Fully Cooled Single Stage HP Transonic Turbine—Part I: Influence of Cooling Mass Flow Variations and Inlet Temperature Profiles on Blade Internal and External Aerodynamics

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
C. W. Haldeman ◽  
M. G. Dunn ◽  
R. M. Mathison
Keyword(s):  
Heat Transfer ◽  
Inlet Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Pressure Loading ◽  
Turbine Stage ◽  
Suction Surface ◽  
Reference Case ◽  
Cooling Flows ◽  
Mass Flows

A fully cooled turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and high-pressure turbine (HPT) vane inlet temperature profile on the aerodynamics and heat transfer of the turbine stage operating at the proper design corrected conditions. Part I of this paper describes the overall experimental matrix, the influence of the cooling mass flows, and temperature profiles from an aerodynamic perspective. The measurements include internal and external pressures for the blade airfoil. Part II of this paper focuses on the influence of these parameters on the heat transfer to the blade airfoil and the stationary blade shroud. The major results show that cooling levels do not significantly affect the external pressure distributions over the majority of the blade and vane. However, aerodynamic effects of cooling levels and temperature profiles are seen for the vane and blade pressure loading on the suction surfaces. The magnitude of these effects ranges from 5% to 10% of the local measurement for the reference case, which is the uniform inlet profile with nominal cooling for this study. Inlet temperature profiles and cooling levels have comparable impacts on pressure loading, but their relative influence changes with location, and Reynolds number and corrected speed variations have the lowest impact on pressure loading changes, with changes below 5% of the local measurements. Another important result is that unlike uncooled experiments, the proper normalizing variable for pressures aft of the vane is not the inlet pressure but a “rotor reference pressure,” which adjusts the total inlet pressure by the increase in pressure resulting from the additional cooling mass flows. For the rotor, this consists primarily of the vane trailing edge cooling flows. This simplified model accounts for the effects of the vane cooling, and isolates the changes due to blade cooling. The spread of the cooling flows through the stage is important to the surface heat-flux, and has an impact on pressure loadings on the suction surface. The data establishes important guidelines for modelers of cooling flows. The changes observed on the suction side of the airfoils are real, but quite small from an engineering design perspective. Thus the pressure levels are stable and relatively independent of cooling levels, which is critical for good heat-transfer predictions.

Fully-Cooled Single Stage HP Transonic Turbine: Part I—Influence of Cooling Mass Flow Variations and Inlet Temperature Profiles on Blade Internal and External Aerodynamics

2010 ◽  
Author(s):  
C. W. Haldeman ◽  
M. G. Dunn ◽  
R. M. Mathison
Keyword(s):  
Heat Transfer ◽  
Inlet Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Pressure Loading ◽  
Turbine Stage ◽  
Suction Surface ◽  
Reference Case ◽  
Cooling Flows ◽  
Mass Flows

A fully cooled turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and HPT vane inlet temperature profile on the aerodynamics and heat transfer of the turbine stage operating at the proper design corrected conditions. Part I of this paper describes the overall experimental matrix, the influence of the cooling mass flows, and temperature profiles from an aerodynamic perspective. The measurements include internal and external pressures for the blade airfoil. Part II of this paper focuses on the influence of these parameters on the heat transfer to the blade airfoil and the stationary blade shroud. The major results show that cooling levels do not significantly affect the external pressure distributions over the majority of the blade and vane. However, aerodynamic effects of cooling levels and temperature profiles are seen for the vane and blade pressure loading on the suction surfaces. The magnitude of these effects range from 5 to 10% of the local measurement for the reference case, which is the uniform inlet profile with nominal cooling for this study. Inlet temperature profiles and cooling levels have comparable impacts on pressure loading, but their relative influence changes with location, Reynolds number and corrected speed variations have the lowest impact on pressure loading changes, with changes below 5% of the local measurements. Another important result is that unlike un-cooled experiments, the proper normalizing variable for pressures aft of the vane is not the inlet pressure but a “Rotor Reference Pressure”, which adjusts the total inlet pressure by the increase in pressure resulting from the additional cooling mass flows. For the rotor, this consists primarily of the vane trailing edge cooling flows. This simplified model accounts for the effects of the vane cooling, and isolates the changes due to blade cooling. The spread of the cooling flows through the stage is important to the surface heat flux, and has an impact on pressure loadings on the suction surface. The data establishes important guidelines for modelers of cooling flows. The changes observed on the suction side of the airfoils are real, but quite small from an engineering design perspective. Thus the pressure levels are stable and relatively independent of cooling levels, which is critical for good heat-transfer predictions.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part I: Experimental Configuration and Data Review With Inlet Temperature Profile Effects

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Harika S. Kahveci ◽  
Charles W. Haldeman ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Turbine Stage ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Pressure Surface ◽  
Cooling Flow ◽  
The Impact

This paper investigates the vane airfoil and inner endwall heat transfer for a full-scale turbine stage operating at design corrected conditions under the influence of different vane inlet temperature profiles and vane cooling flow rates. The turbine stage is a modern 3D design consisting of a cooled high-pressure vane, an un-cooled high-pressure rotor, and a low-pressure vane. Inlet temperature profiles (uniform, radial, and hot streaks) are created by a passive heat exchanger and can be made circumferentially uniform to within ±5% of the bulk average inlet temperature when desired. The high-pressure vane has full cooling coverage on both the airfoil surface and the inner and outer endwalls. Two circuits supply coolant to the vane, and a third circuit supplies coolant to the rotor purge cavity. All of the cooling circuits are independently controlled. Measurements are performed using double-sided heat-flux gauges located at four spans of the vane airfoil surface and throughout the inner endwall region. Analysis of the heat transfer measured for the uncooled downstream blade row has been reported previously. Part I of this paper describes the operating conditions and data reduction techniques utilized in this analysis, including a novel application of a traditional statistical method to assign confidence limits to measurements in the absence of repeat runs. The impact of Stanton number definition is discussed while analyzing inlet temperature profile shape effects. Comparison of the present data (Build 2) to the data obtained for an uncooled vane (Build 1) clearly illustrates the impact of the cooling flow and its relative effects on both the endwall and airfoils. Measurements obtained for the cooled hardware without cooling applied agree well with the solid airfoil for the airfoil pressure surface but not for the suction surface. Differences on the suction surface are due to flow being ingested on the pressure surface and reinjected on the suction surface when coolant is not supplied for Build 2. Part II of the paper continues this discussion by describing the influence of overall cooling level variation and the influence of the vane trailing edge cooling on the vane heat transfer measurements.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part I: Influence of Cooling Variation

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Flows ◽  
Cooling Flow

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design-corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film-cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the uncooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton number distribution on the blade.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine: Part I—Influence of Cooling Variation

2010 ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Flows ◽  
Cooling Flow

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the un-cooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton Number distribution on the blade.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine: Part I—Experimental Configuration and Data Review With Inlet Temperature Profile Effects

2011 ◽  
Author(s):  
Harika S. Kahveci ◽  
Charles W. Haldeman ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Turbine Stage ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Pressure Surface ◽  
Cooling Flow ◽  
The Impact

This paper investigates the vane airfoil and inner endwall heat transfer for a full-scale turbine stage operating at design corrected conditions under the influence of different vane inlet temperature profiles and vane cooling flow rates. The turbine stage is a modern 3-D design consisting of a cooled high-pressure vane, an un-cooled high-pressure rotor, and a low-pressure vane. Inlet temperature profiles (uniform, radial and hot streaks) are created by a passive heat exchanger and can be made circumferentially uniform to within ±5% of the bulk average inlet temperature when desired. The high-pressure vane has full cooling coverage on both the airfoil surface and the inner and outer endwalls. Two circuits supply coolant to the vane, and a third circuit supplies coolant to the rotor purge cavity. All of the cooling circuits are independently controlled. Measurements are performed using double-sided heat-flux gauges located at four spans of the vane airfoil surface and throughout the inner endwall region. Analysis of the heat transfer measured for the uncooled downstream blade row has been reported previously. Part I of this paper describes the operating conditions and data reduction techniques utilized in this analysis, including a novel application of a traditional statistical method to assign confidence limits to measurements in the absence of repeat runs. The impact of Stanton Number definition is discussed while analyzing inlet temperature profile shape effects. Comparison of the present data (Build 2) to the data obtained for an un-cooled vane (Build 1) clearly illustrates the impact of the cooling flow and its relative effects on both the endwall and airfoils. Measurements obtained for the cooled hardware without cooling applied agree well with the solid air-foil for the airfoil pressure surface but not for the suction surface. Differences on the suction surface are due to flow being ingested on the pressure surface and re-injected on the suction surface when coolant is not supplied for Build 2. Part II of the paper continues this discussion by describing the influence of overall cooling level variation and the influence of the vane trailing edge cooling on the vane heat transfer measurements.

Fully-Cooled Single Stage HP Transonic Turbine: Part II—Influence of Cooling Mass Flow Changes and Inlet Temperature Profiles on Blade and Shroud Heat-Transfer

2010 ◽  
Author(s):  
C. W. Haldeman ◽  
M. G. Dunn ◽  
R. M. Mathison
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Temperature Profile ◽  
Mass Flow ◽  
Stanton Number ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Turbine Stage ◽  
Flow Rates ◽  
Cooling Effects

A fully cooled transonic HP turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and vane inlet temperature profile on heat transfer to the blades with the stage operating at the proper design corrected conditions. For this series of experiments, both the vane row and the blade row were fully cooled. The matrix of experimental conditions included varying the cooling flow rates and the vane inlet temperature profiles to observe the overall effect on airfoil heat-transfer. The data presented in Part I focused on the aerodynamics of the fully cooled turbine for a subset of the cases investigating two vane inlet temperature profiles (uniform and radial), and three different cooling levels (none, nominal and high) for the high Reynolds number condition. This part of the paper focuses on the time-average heat-flux measurements on the blade and shroud region for the same cooling mass flow rates and vane inlet temperature profiles. The cooling effects are shown to be small and are centered primarily on the suction side of the airfoil. This relatively small influence is due to the ratio of the cooling gas to metal temperature being closer to 1 than the design value would dictate. The vane inlet temperature profile effects are more dominant, and using a Net Stanton Number Reduction Factor to compare the cases, an effect on the order of about 0.25 is demonstrated. This effect is due primarily to the change in the reference temperature used for the Stanton number calculation. The differences due to profile effects are small, but observable towards the trailing edge of both the blade and rotor shroud. This data set forms an excellent baseline for heat-flux calculations, as the variation in the main input conditions are well documented and do not produce large changes in the heat-flux. It provides insight into the flow physics of an actual engine and guidelines about proper normalization of variables for a cooled turbine stage, supporting further development of computational heat-flux modeling techniques.

Fully Cooled Single Stage HP Transonic Turbine—Part II: Influence of Cooling Mass Flow Changes and Inlet Temperature Profiles on Blade and Shroud Heat-Transfer

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
C. W. Haldeman ◽  
M. G. Dunn ◽  
R. M. Mathison
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Temperature Profile ◽  
Mass Flow ◽  
Stanton Number ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Turbine Stage ◽  
Flow Rates ◽  
Cooling Effects

A fully cooled transonic high-pressure turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and vane inlet temperature profile on heat transfer to the blades with the stage operating at the proper design corrected conditions. For this series of experiments, both the vane row and the blade row were fully cooled. The matrix of experimental conditions included varying the cooling flow rates and the vane inlet temperature profiles to observe the overall effect on airfoil heat-transfer. The data presented in Part I focused on the aerodynamics of the fully cooled turbine for a subset of the cases investigating two vane inlet temperature profiles (uniform and radial) and three different cooling levels (none, nominal, and high) for the high Reynolds number condition. This part of the paper focuses on the time-average heat-flux measurements on the blade and shroud region for the same cooling mass flow rates and vane inlet temperature profiles. The cooling effects are shown to be small and are centered primarily on the suction side of the airfoil. This relatively small influence is due to the ratio of the cooling gas to metal temperature being closer to 1 than the design value would dictate. The vane inlet temperature profile effects are more dominant, and using a net Stanton number reduction factor to compare the cases, an effect on the order of about 0.25 is demonstrated. This effect is due primarily to the change in the reference temperature used for the Stanton number calculation. The differences due to profile effects are small but observable toward the trailing edge of both the blade and rotor shroud. This data set forms an excellent baseline for heat-flux calculations, as the variation in the main input conditions are well documented and do not produce large changes in the heat-flux. It provides insight into the flow physics of an actual engine and guidelines about proper normalization of variables for a cooled turbine stage, supporting further development of computational heat-flux modeling techniques.

Effect of Temperature Nonuniformity on Heat Transfer in an Unshrouded Transonic HP Turbine: An Experimental and Computational Investigation

2010 ◽  
Author(s):  
Imran Qureshi ◽  
Andy D. Smith ◽  
Kam S. Chana ◽  
Thomas Povey
Keyword(s):  
Heat Transfer ◽  
Three Dimensional ◽  
Aerodynamic Characteristics ◽  
Test Facility ◽  
Effect Of Temperature ◽  
Inlet Temperature ◽  
Experimental Measurements ◽  
Turbine Stage ◽  
Cfd Simulations ◽  
Turbine Inlet

Detailed experimental measurements have been performed to understand the effects of turbine inlet temperature distortion (hot-streaks) on the heat transfer and aerodynamic characteristics of a full-scale unshrouded high pressure turbine stage at flow conditions that are representative of those found in a modern gas turbine engine. To investigate hot-streak migration, the experimental measurements are complemented by three-dimensional steady and unsteady CFD simulations of the turbine stage. This paper presents the time-averaged measurements and computational predictions of rotor blade surface and rotor casing heat transfer. Experimental measurements obtained with and without inlet temperature distortion are compared. Time-mean experimental measurements of rotor casing static pressure are also presented. CFD simulations have been conducted using the Rolls-Royce code Hydra, and are compared to the experimental results. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough UK. This is a short duration transonic facility, which simulates engine representative M, Re, Tu, N/T and Tg /Tw at the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation temperature distortion generator, capable of simulating well-defined, aggressive temperature distortion both in the radial and circumferential directions, at the turbine inlet.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

2005 ◽  
Author(s):  
T. Povey ◽  
K. S. Chana ◽  
T. V. Jones ◽  
J. Hurrion
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Cooling System ◽  
Numerical Study ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Wall Heat Transfer ◽  
Hot Streak ◽  
End Wall

Pronounced non-uniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of end-wall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the affects of these temperature non-uniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimise the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and end-wall heat transfer rate for an HP nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two non-uniform temperature profiles. The temperature profiles were non-dimensionally similar to profiles measured in an engine. A difference of one half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and end-wall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualisation study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short duration engine-size turbine facility. Mach number, Reynolds number and gas-to-wall temperature ratios were correctly modelled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

The Effects of Spraying Parameters on Spray Cooling Heat Transfer Performance in the Non-Boiling Regime

2017 ◽  
Author(s):  
Azzam S. Salman ◽  
Jamil A. Khan
Keyword(s):  
Heat Transfer ◽  
Cooling System ◽  
Target Surface ◽  
Spray Cooling ◽  
Operating Conditions ◽  
Droplet Velocity ◽  
Inlet Pressure ◽  
Working Fluid ◽  
Inlet Temperature ◽  
Spraying Parameters

Experiments were conducted in a closed loop spray cooling system working with deionized water as a working fluid. This study was performed to investigate the effect of the spraying parameters, such as Sauter mean diameter (SMD), the droplet velocity, and the residual velocity on the spray cooling heat transfer in the non-boiling region. Thermal effects on plain and modified surfaces with circular grooves were examined under different operating conditions. The inlet pressure of the working fluid was varied from 78.6 kPa to 183.515kPa, and the inlet temperature was kept between 21–22 °C. The distance between the nozzle and the target surface 10 mm. The results showed that increasing the coolant inlet pressure increases the droplet velocity and the number of droplets produced while decreasing the droplet size. As a consequence of these changes, increasing inlet pressure improved the heat transfer characteristics of both surfaces.

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