Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine: Part II—Effect of Cooling Variation on the Vane Airfoil and Inner Endwall

2011 ◽  
Author(s):  
Harika S. Kahveci ◽  
Charles W. Haldeman ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
External Heat ◽  
Stanton Number ◽  
Inlet Temperature ◽  
Trailing Edge ◽  
Flow Rates ◽  
External Heat Transfer ◽  
Pressure Surface ◽  
The Impact

The impact of film cooling on heat transfer is investigated for the high-pressure vane of a one-and-one-half stage high-pressure turbine operating at design corrected conditions. Cooling is supplied through three independently controllable circuits to holes in the inner and outer endwall, vane leading edge showerhead, and the pressure and suction surfaces of the airfoil in addition to vane trailing edge slots. Four different overall cooling flow rates are investigated and one cooling circuit is varied independently. All results reported in this part of the paper are for a radial inlet temperature profile, one of the four profiles reported in Part I of this paper. Part I describes the experimental setup, data quality, influence of inlet temperature profile, and influence of cooling when compared to a solid vane. This part of the paper shows that the addition of coolant reduces airfoil Stanton Number by up to 60%. The largest reductions due to cooling are observed close to the inner endwall because the coolant to the majority of the vane is supplied by a plenum at the inside diameter. While the introduction of cooling has a significant impact on Stanton Number, the impact of changing coolant flow rates is only observed for gauges near 5% span and on the inner endwall. This indicates that very little of the increased coolant mass flow reaches all the way to 90% span and the majority of the additional mass flow is injected into the core flow near the plenum. Turning off the vane outer cooling circuit that supplies coolant to the outer endwall holes, vane trailing edge slots, and three rows of holes on the pressure surface of the airfoil, has a local impact on Stanton Number. Changes downstream of the holes on the airfoil pressure surface indicate that internal heat transfer from the coolant flowing inside the vane is important to the external heat transfer, suggesting that a conjugate heat-transfer solution may be required to achieve good external heat-transfer predictions in this area. Measurements on the inner endwall show that temperature reduction in the vane wake due to the trailing edge cooling is important to many points downstream of the vane.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part II: Effect of Cooling Variation on the Vane Airfoil and Inner End Wall

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Harika S. Kahveci ◽  
Charles W. Haldeman ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
External Heat ◽  
Stanton Number ◽  
Inlet Temperature ◽  
Trailing Edge ◽  
Flow Rates ◽  
Pressure Surface ◽  
End Wall ◽  
The Impact

The impact of film cooling on heat transfer is investigated for the high-pressure vane of a 1-1/2 stage high-pressure turbine operating at design corrected conditions. Cooling is supplied through three independently controllable circuits to holes in the inner and outer end wall, vane leading edge showerhead, and the pressure and suction surfaces of the airfoil, in addition to vane trailing edge slots. Four different overall cooling flow rates are investigated and one cooling circuit is varied independently. All results reported in this part of the paper are for a radial inlet temperature profile, one of the four profiles reported in part I of this paper. Part I describes the experimental setup, data quality, influence of inlet temperature profile, and influence of cooling when compared to a solid vane. This part of the paper shows that the addition of coolant reduces airfoil Stanton number by up to 60%. The largest reductions due to cooling are observed close to the inner end wall because the coolant to the majority of the vane is supplied by a plenum at the inside diameter. While the introduction of cooling has a significant impact on Stanton number, the impact of changing coolant flow rates is only observed for gauges near 5% span and on the inner end wall. This indicates that very little of the increased coolant mass flow reaches all the way to 90% span and the majority of the additional mass flow is injected into the core flow near the plenum. Turning off the vane outer cooling circuit that supplies coolant to the outer end wall holes, vane trailing edge slots, and three rows of holes on the pressure surface of the airfoil, has a local impact on Stanton number. Changes downstream of the holes on the airfoil pressure surface indicate that internal heat transfer from the coolant flowing inside the vane is important to the external heat transfer, suggesting that a conjugate heat-transfer solution may be required to achieve good external heat-transfer predictions in this area. Measurements on the inner end wall show that temperature reduction in the vane wake due to the trailing edge cooling is important to many points downstream of the vane.

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 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.

Design and Optimization of the Internal Cooling Channels of a High Pressure Turbine Blade—Part I: Methodology

2010 ◽  
Vol 132 (2) ◽  
Author(s):  
Sergio Amaral ◽  
Tom Verstraete ◽  
René Van den Braembussche ◽  
Tony Arts
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Thermal Stress ◽  
Turbine Blade ◽  
Inlet Temperature ◽  
Analysis Tool ◽  
Internal Cooling ◽  
High Pressure Turbine ◽  
Cooling Channels ◽  
The Impact

This first paper describes the conjugate heat transfer (CHT) method and its application to the performance and lifetime prediction of a high pressure turbine blade operating at a very high inlet temperature. It is the analysis tool for the aerothermal optimization described in a second paper. The CHT method uses three separate solvers: a Navier–Stokes solver to predict the nonadiabatic external flow and heat flux, a finite element analysis (FEA) to compute the heat conduction and stress within the solid, and a 1D aerothermal model based on friction and heat transfer correlations for smooth and rib-roughened cooling channels. Special attention is given to the boundary conditions linking these solvers and to the stability of the complete CHT calculation procedure. The Larson–Miller parameter model is used to determine the creep-to-rupture failure lifetime of the blade. This model requires both the temperature and thermal stress inside the blade, calculated by the CHT and FEA. The CHT method is validated on two test cases: a gas turbine rotor blade without cooling and one with five cooling channels evenly distributed along the camber line. The metal temperature and thermal stress distribution in both blades are presented and the impact of the cooling channel geometry on lifetime is discussed.

Evaluation of the External Heat Transfer Coefficients in the High-Pressure Turbine of a Full-Scale Core Engine

1996 ◽  
Author(s):  
A. R. Narcus ◽  
H. R. Przirembel ◽  
F. O. Soechting
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Data Reduction ◽  
Turbine Blades ◽  
External Heat ◽  
Transfer Coefficients ◽  
High Pressure Turbine ◽  
External Heat Transfer ◽  
Heat Transfer Coefficients

The external heat transfer coefficients, necessary for efficient and accurate turbine blade design, have been quantified using three independent methods of data reduction for the high-pressure turbine blades tested in a core engine. Two of the methods utilized external and internal thermocouple data to determine the heat transfer coefficient levels while the third method required the applied heat-flux levels to determine the coefficients. The heat-flux was calculated from the measured potential difference between thermocouple pairs embedded in the external and internal walls of the turbine blades. The instrumented airfoils were calibrated in a laboratory prior to engine testing. The results of the experimental test showed external heat transfer coefficients could be obtained in an engine environment with a ±3.2% minimum absolute uncertainty. All three data reduction methods produced external heat transfer coefficients within a high degree of accuracy and precision for all data locations on the instrumented airfoils. The three data reduction approaches are presented as well as the data for a specific location on a turbine blade for each method of data reduction. In addition, pre-test calibration procedures and data are discussed along with supporting engine instrumentation used to verify the data acquired during the experimental evaluation.

Rotor Blade Heat Transfer Characteristics for High Pressure Turbine Stage Under Inlet Temperature and Velocity Traverses

2014 ◽  
Author(s):  
A. Rahim ◽  
L. He ◽  
E. Romero
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Load ◽  
Inlet Temperature ◽  
Turbine Stage ◽  
Heat Transfer Characteristics ◽  
Lean Burn ◽  
Transfer Characteristics ◽  
Hot Streak ◽  
The Impact

One of the key considerations in high pressure (HP) turbine design is the heat load experienced by rotor blades. The impact of turbine inlet non-uniformities on the blades in the form of combined temperature and velocity traverses, typical for a lean burn combustor exit, has rarely been studied. For general HP turbine aerothermal designs, it is also of interest to understand how the behavior of a lean burn combustor traverses (hot streak and swirl) might contrast with those for rich burn combustion (largely hot streak only). In the present work, a computational study has been carried out on the aerothermal performance of a HP turbine stage under non-uniform temperature and velocity inlet profiles. The analyses are primarily conducted for two combined hot streak and swirl inlets, with opposite swirl directions. In addition, comparisons are made against a hot streak only case and a uniform inlet. The effects of three NGV shape configurations are investigated; namely, straight, compound lean (CL) and reverse compound lean (RCL). The present results show that there is a qualitative change in the roles played by heat transfer coefficient (HTC) and fluid driving (‘adiabatic wall’) temperature, Taw. It has been shown that the blade heat load distribution for a uniform inlet is dominated by HTC, whilst for a hot streak only case it is wholly influenced by Taw. However, in contrast to the hot streak only case, the case with a combined hot streak and swirl shows a role reversal with the HTC being dominant in determining the heat load. Additionally, it is seen that the swirling flow radially redistributes the hot fluid within the NGV passage considerably, leading to a much ‘flatter’ rotor inlet temperature profile compared to its hot streak only counterpart. Further, the rotor heat transfer characteristics for the cases with the combined traverses are shown to be strongly dependent on the NGV shaping and the inlet swirl direction, indicating the potential for future design space exploration. The present findings underline the need to clearly define relevant combustor exit temperature and velocity profiles when designing and optimizing NGVs for HP turbine aerothermal performance.

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.

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