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David G. Bogard - One of the best experts on this subject based on the ideXlab platform.

  • in the showerhead region of a gas turbine vane part i stagnation region and near Pressure Side
    2015
    Co-Authors: Marc D Polanka, Virginia C Witteveld, David G. Bogard
    Abstract:

    An experimental program to study film cooling adiabatic effectiveness in the showerhead region of a turbine vane was completed. A typical first stage turbine stator vane was modeled in a nine-times scale, low speed test facility. The leading edge showerhead consisted of a row of holes at the stagnation point of the vane as well as two rows on the Pressure surface and three rows on the suction surface. Film cooling performance on the Pressure Side of the showerhead is described in this paper. Experiments were conducted for a number of blowing ratios, from 0.3 to 2.9, and at a density ratio of nominally 1.8. The impact of high mainstream turbulence was studied by comparing performance at turbulence levels of 0.5% and 22%. The sensitivity to the stagnation line position was also investigated. The approaching stagnation line was positioned either directly on the stagnation row of holes or just to the suction Side of that row. With the stagnation line directly on the stagnation row of holes at low turbulence, the direction of the coolant jets was observed to vary along the span of the airfoil. At high turbulence, this same effect did not occur. Moreover, at the high mainstream turbulence level, the coolant from the entire stagnation row of holes seemed to be spread more evenly throughout the stagnation region. At high blowing ratios, this resulted in higher effectiveness values in the stagnation region. Farther downstream, the high mainstream turbulence reduced the effectiveness at all blowing ratios.

  • Measurements of Adiabatic Film and Overall Cooling Effectiveness on a Turbine Vane Pressure Side With a Trench
    Journal of Turbomachinery, 2013
    Co-Authors: Jason E. Albert, David G. Bogard
    Abstract:

    Film cooling performance is typically quantified by separating the external convective heat transfer from the other components of the conjugate heat transfer that occurs in turbine airfoils. However, it is also valuable to assess the conjugate heat transfer in terms of the overall cooling effectiveness, which is a parameter of importance to airfoil designers. In the current study, adiabatic film effectiveness and overall cooling effectiveness values were measured for the Pressure Side of a simplified turbine vane model with three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the Pressure Side. This was done by utilizing two geometrically identical models made from different materials. Adiabatic film effectiveness was measured using a very low thermal conductivity material, and the overall cooling effectiveness was measured using a material with a higher thermal conductivity selected such that the Biot number of the model matched that of a turbine vane at engine conditions. The theoretical basis for this matched-Biot number modeling technique is discussed in some detail. Additionally, two designs of Pressure Side body film cooling holes were conSidered in this study: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. This study was performed using engine representative flow conditions, including a coolant-to-mainstream density ratio of DR = 1.4 and a mainstream turbulence intensity of Tu = 20%. The results of this study show that adiabatic film and overall cooling effectiveness increase with blowing ratio for the showerhead and Pressure Side trenched holes. Performance decreases with blowing ratio for the standard holes due to coolant jet separation from the surface. Both body film designs have similar performance at a lower blowing ratio when the standard hole coolant jets remain attached. Far downstream of the cooling holes both designs perform similarly because film effectiveness decays more rapidly for the trenched holes.

  • Experimental Simulation of Contaminant Deposition on a Film-Cooled Turbine Vane Pressure Side With a Trench
    Journal of Turbomachinery, 2013
    Co-Authors: Jason E. Albert, David G. Bogard
    Abstract:

    An important issue in the use of coal- or biomass-derived synthetic gaseous (syngas) fuels is the deposition of contaminants on film-cooled turbine surfaces, which alter cooling and aerodynamic performance and increase material degradation. The current study applied a new experimental technique that simulated the key physical aspects of contaminant deposition on a film-cooled turbine vane. The depositing contaminants were modeled in a wind tunnel facility with a spray of molten wax droplets of a size range that matched the Stokes number of the contaminant particles in engine conditions. Most experiments were performed using a vane model with a thermal conductivity selected such that the model had the same Biot number of an actual engine airfoil, resulting in a cooler surface temperature. Some experiments were performed using an approximately adiabatic model for comparison. The film cooling design consisted of three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the Pressure Side. Two designs of Pressure Side body film cooling holes were conSidered: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. The results showed thin deposits formed in the trench, with the thickest deposits on its downstream wall between coolant jets. Adiabatic film effectiveness levels were essentially unchanged by the presence of deposits for either film configuration. Deposit formation was strongly influenced by the model surface temperature with cooler surfaces inhibiting deposition. There was evidence of a threshold surface temperature above which deposits became significantly thicker.

  • Measurements of adiabatic film and overall cooling effectiveness on a turbine vane Pressure Side with a trench
    Volume 5: Heat Transfer Parts A and B, 2011
    Co-Authors: Jason E. Albert, David G. Bogard
    Abstract:

    Film cooling performance is typically quantified by separating the external convective heat transfer from the other components of the conjugate heat transfer that occurs in turbine airfoils. However, it is also valuable to assess the conjugate heat transfer in terms of the overall cooling effectiveness, which is a parameter of importance to airfoil designers. In the current study, adiabatic film effectiveness and overall cooling effectiveness values were measured for the Pressure Side of a simplified turbine vane model with three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the Pressure Side. This was done by utilizing two geometrically identical models made from different materials. Adiabatic film effectiveness was measured using a very low thermal conductivity material, and the overall cooling effectiveness was measured using a material with a higher thermal conductivity selected such that the Biot number of the model matched that of a turbine vane at engine conditions. The theoretical basis for this matched-Biot number modeling technique is discussed in some detail. Additionally, two designs of Pressure Side body film cooling holes were conSidered in this study: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. This study was performed using engine representative flow conditions, including a coolant-to-mainstream density ratio of DR = 1.4 and a mainstream turbulence intensity of Tu = 20%. The results of this study show that adiabatic film and overall cooling effectiveness increase with blowing ratio for the showerhead and Pressure Side trenched holes. Performance decreases with blowing ratio for the standard holes due to coolant jet separation from the surface. Both body film designs have similar performance at a lower blowing ratio when the standard hole coolant jets remain attached. Far downstream of the cooling holes both designs perform similarly because film effectiveness decays more rapidly for the trenched holes.Copyright © 2011 by ASME

  • Experimental Simulation of Contaminant Deposition on a Film Cooled Turbine Vane Pressure Side With a Trench
    Volume 5: Heat Transfer Parts A and B, 2011
    Co-Authors: Jason E. Albert, David G. Bogard
    Abstract:

    An important issue in the use of coal- or biomass-derived synthetic gaseous (syngas) fuels is the deposition of contaminants on film cooled turbine surfaces, which alter cooling and aerodynamic performance and increase material degradation. The current study applied a new experimental technique that simulated the key physical aspects of contaminant deposition on a film cooled turbine vane. The depositing contaminants were modeled in a wind tunnel facility with a spray of molten wax droplets of a size range that matched the Stokes number of the contaminant particles in engine conditions. Most experiments were performed using a vane model with a thermal conductivity selected such that the model had the same Biot number of an actual engine airfoil, resulting in a cooler surface temperature. Some experiments were performed using an approximately adiabatic model for comparison. The film cooling design consisted of three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the Pressure Side. Two designs of Pressure Side body film cooling holes were conSidered: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. The results showed thin deposits formed in the trench, with the thickest deposits on its downstream wall between coolant jets. Adiabatic film effectiveness levels were essentially unchanged by the presence of deposits for either film configuration. Deposit formation was strongly influenced by the model surface temperature with cooler surfaces inhibiting deposition. There was evidence of a threshold surface temperature above which deposits became significantly thicker.Copyright © 2011 by ASME

Karen A. Thole - One of the best experts on this subject based on the ideXlab platform.

  • Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side
    Journal of Turbomachinery, 2008
    Co-Authors: N. Sundaram, M. D. Barringer, Karen A. Thole
    Abstract:

    Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the Pressure Side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the Pressure Side of the airfoil. Deposits having a semielliptical shape were placed along the Pressure Side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the Pressure Side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

  • Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side
    Volume 4: Turbo Expo 2007 Parts A and B, 2007
    Co-Authors: N. Sundaram, M. D. Barringer, Karen A. Thole
    Abstract:

    Film-cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film-cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film-cooling along the Pressure Side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the Pressure Side of the airfoil. Deposits having a semi-elliptical shape were placed along the Pressure Side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets towards the vane endwall junction on the Pressure Side. Results also indicated that there was a steady decrease in adiabatic effectiveness with a sequential increase in the number of rows with the deposits.Copyright © 2007 by ASME

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes—Part I: Adiabatic Effectiveness Measurements
    Journal of Turbomachinery, 2005
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    Durability of turbine blade tips has been and continues to be challenging, particularly since increasing turbine inlet temperatures is the driver for improving turbine engine performance. As a result, cooling methods along the blade tip are crucial. Film-cooling is one typically used cooling method whereby coolant is supplied through holes placed along the Pressure Side of a blade. The subject of this paper is to evaluate the adiabatic effectiveness levels that occur on the blade tip through blowing coolant from holes placed near the tip of a blade along the Pressure Side. A range of blowing ratios was studied whereby coolant was injected from holes placed along the Pressure Side tip of a large-scale blade model. Also present were dirt purge holes on the blade tip, which is part of a commonly used blade design to expel any large particles present in the coolant stream. Experiments were conducted in a linear cascade with a scaled-up turbine blade whereby the Reynolds number of the engine was matched. This paper, which is Part I of a two part series, compares adiabatic effectiveness levels measured along a blade tip, while Part 2 combines measured heat transfer coefficients with the adiabatic effectiveness levels to assess the overall cooling benefit of Pressure Side blowing near a blade tip. The results show much better cooling can be achieved for a small tip gap compared with a large tip gap with different flow phenomena occurring for each tip gap setting.

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes—Part II: Heat Transfer Measurements
    Journal of Turbomachinery, 2005
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    The clearance gap between a turbine blade tip and its associated shroud allows leakage flow across the tip gap from the Pressure Side to the suction Side of the blade. Understanding how this leakage flow affects heat transfer is critical in extending blade tip durability in terms of oxidation, erosion, clearance, and overall turbine performance. This paper is the second of a two part series that discusses the augmentation of tip heat transfer as a result of blowing from the Pressure Side of the tip as well as dirt purge holes placed on the tip. For the experimental investigation, three scaled-up blades were used to form a two-passage linear cascade in a low speed wind tunnel. The rig was designed to simulate different tip gap sizes and coolant flow rates. Heat transfer coefficients were quantified by measuring the total power supplied to a constant heat flux surface placed on the tip of the blade and measuring the tip temperatures. Results indicate that increased blowing leads to increased augmentations in tip heat transfer, particularly at the entrance region to the gap. When combined with adiabatic effectiveness measurements, the coolant from the Pressure Side holes provides an overall net heat flux reduction to the blade tip but is nearly independent of coolant flow levels.

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes: Part 2 — Heat Transfer Measurements
    Volume 3: Turbo Expo 2004, 2004
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    The clearance gap between a turbine blade tip and its associated shroud allows leakage flow across the tip gap from the Pressure Side to the suction Side of the blade. Understanding how this leakage flow affects heat transfer is critical in extending blade tip durability in terms of oxidation, erosion, clearance, and overall turbine performance. This paper is the second of a two part series that discusses the augmentation of tip heat transfer as a result of blowing from the Pressure Side of the tip as well as dirt purge holes placed on the tip. For the experimental investigation, three scaled-up blades were used to form a two-passage linear cascade in a low speed wind tunnel. The rig was designed to simulate different tip gap sizes and coolant flow rates. Heat transfer coefficients were quantified by measuring the total power supplied to a constant heat flux surface placed on the tip of the blade and measuring the tip temperatures. Results indicate that increased blowing leads to increased augmentations in tip heat transfer, particularly at the entrance region to the gap. When combined with adiabatic effectiveness measurements, the coolant from the Pressure Side holes provides an overall net heat flux reduction to the blade tip but is nearly independent of coolant flow levels.© 2004 ASME

Sumanta Acharya - One of the best experts on this subject based on the ideXlab platform.

  • Large Eddy Simulation of the Laminar Heat Transfer Augmentation on the Pressure Side of a Turbine Vane Under Freestream Turbulence
    Journal of Turbomachinery, 2019
    Co-Authors: Yousef Kanani, Sumanta Acharya, Forrest E. Ames
    Abstract:

    Vane Pressure Side heat transfer is studied numerically using large eddy simulation (LES) on an aft-loaded vane with a large leading edge over a range of turbulence conditions. Numerical simulations are performed in a linear cascade at exit chord Reynolds number of Re = 5.1 × 105 at low (Tu ≈ 0.7%), moderate (Tu ≈ 7.9%), and high (Tu ≈ 12.4%) freestream turbulence with varying length scales as prescribed by the experimental measurements of Varty and Ames (2016, “Experimental Heat Transfer Distributions Over an Aft Loaded Vane With a Large Leading Edge at Very High Turbulence Levels,” ASME Paper No. IMECE2016-67029). Heat transfer predictions on the vane Pressure Side are in a very good agreement with the experimental measurements and the heat transfer augmentation due to the freestream turbulence is well captured. At Tu ≈ 12.4%, freestream turbulence enhances the Stanton number on the Pressure surface without boundary layer transition to turbulence by a maximum of about 50% relative to the low freestream turbulence case. Higher freestream turbulence generates elongated structures and high-velocity streaks wrapped around the leading edge that contain significant energy. Amplification of the velocity streaks is observed further downstream with max rms of 0.3 near the trailing edge but no transition to turbulence or formation of turbulence spots is observed on the Pressure Side. The heat transfer augmentation at the higher freestream turbulence is primarily due to the initial amplification of the low-frequency velocity perturbations inSide the boundary layer that persist along the entire chord of the airfoil. Stanton numbers appear to scale with the streamwise velocity fluctuations inSide the boundary layer.

  • LES Study of the Laminar Heat Transfer Augmentation on the Pressure Side of a Turbine Vane Under Freestream Turbulence
    Volume 5C: Heat Transfer, 2018
    Co-Authors: Yousef Kanani, Sumanta Acharya, Forrest E. Ames
    Abstract:

    Vane Pressure Side heat transfer is studied numerically using Large Eddy Simulation (LES) on an aft loaded vane with a large leading edge over a range of turbulence conditions. Numerical simulations are performed in a linear cascade at exit chord Reynolds number of Re = 5.1 × 105 at low (Tu≈0.7%), moderate (Tu≈7.9%) and high (Tu≈12.4%) freestream turbulence with varying length scales as prescribed by the experimental measurements of Varty and Ames (2016). Heat transfer predictions (i.e. Stanton number based on exit condition) on the vane Pressure Side are in a very good agreement with the experimental measurements and the heat transfer augmentation due to the freestream turbulence is well captured. At Tu≈12.4%, freestream turbulence enhances the Stanton number on the Pressure surface without boundary layer transition to turbulence by a maximum of about 50% relative to the low freestream turbulence case (Tu≈0.7%). Higher freestream turbulence generates elongated structures and high-velocity streaks wrapped around the leading edge that contain significant energy. Amplification of the velocity streaks is observed further downstream with max r.m.s of 0.3 near the trailing edge but no transition to turbulence or formation of turbulence spots is observed on the Pressure Side. The heat transfer augmentation at the higher freestream turbulence is primarily due to the initial amplification of the low-frequency velocity perturbations inSide the boundary layer that persist along the entire chord of the airfoil. Stanton numbers appear to scale with the streamwise velocity fluctuations inSide the boundary layer. Görtler vortices are not observed for this airfoil geometry.

  • Turbine Blade Tip Film Cooling With Blade Rotation—Part I: Tip and Pressure Side Coolant Injection
    Journal of Turbomachinery, 2016
    Co-Authors: Onieluan Tamunobere, Sumanta Acharya
    Abstract:

    This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and Pressure Side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped Pressure Side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer result for the no coolant injection shows a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction Side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction Side as opposed to the Pressure Side in stationary studies can primarily be attributed to the effects of the blade relative motion.

  • Turbine Blade Tip Film Cooling With Blade Rotation: Part I — Tip and Pressure Side Coolant Injection
    Volume 5B: Heat Transfer, 2015
    Co-Authors: Onieluan Tamunobere, Sumanta Acharya
    Abstract:

    This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and Pressure Side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped Pressure Side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer results for the no coolant injection show a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction Side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction Side as opposed to the Pressure Side in stationary studies can primarily be attributed to the effects of the blade relative motion.Copyright © 2015 by ASME

  • Blade Tip Leakage Flow and Heat Transfer with Pressure-Side Winglet
    International Journal of Rotating Machinery, 2006
    Co-Authors: Arun K. Saha, Sumanta Acharya, Ron Bunker, Chander Prakash
    Abstract:

    A numerical study has been conducted to explore the effect of a Pressure-Side winglet on the flow and heat transfer over a blade tip. Calculations are performed for both a flat tip and a squealer tip. The winglet is in the form of a flat extension, and is shaped in the axial chord direction to have the maximum thickness at the chord location, where the Pressure difference is the largest between the Pressure and suction Sides. For the flat tip, the Pressure-Side winglet exhibits a significant reduction in the leakage flow strength. The low heat transfer coefficient “sweet-spot” region is larger with the Pressure-Side winglet, and lower heat transfer coefficients are also observed along the Pressure Side of the blade. For the flat tip, the winglet reduces the heat transfer coefficient locally by as much as 30%, while the average heat transfer coefficient is reduced by about 7%. In the presence of a squealer, the role of the winglet decreases significantly, and a 5% reduction in the Pressure loss coefficient is achieved with the winglet with virtually no reduction in the average heat transfer coefficient. On the other hand, the suction-Side squealer with constant width winglet shows lower heat transfer (reduction of 5.5%) and Pressure loss coefficient (reduction of 26%) than its baseline counterpart.

Frank J. Cunha - One of the best experts on this subject based on the ideXlab platform.

  • Heat Transfer in an Airfoil Trailing Edge Configuration With Shaped Pedestals Mounted Internal Cooling Channel and Pressure Side Cutback
    Volume 3: Heat Transfer Parts A and B, 2006
    Co-Authors: S. P. Chen, Frank J. Cunha, Minking K. Chyu, William Abdel-messeh
    Abstract:

    Described in this paper is an experimental study of heat transfer over a trailing edge configuration preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pedestals and oblong shaped blocks. Downstream to the pedestal array, the trailing edge features Pressure Side cutback partitioned by the oblong shaped blocks. The local heat transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined by using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat transfer coefficient on the endwall surface uncovered by the pedestals. The heat transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and endwalls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the downstream most section of the suction Side, the land, due to Pressure Side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat transfer coefficient on the land section are characterized by using the transient liquid crystal technique.Copyright © 2006 by ASME

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes—Part I: Adiabatic Effectiveness Measurements
    Journal of Turbomachinery, 2005
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    Durability of turbine blade tips has been and continues to be challenging, particularly since increasing turbine inlet temperatures is the driver for improving turbine engine performance. As a result, cooling methods along the blade tip are crucial. Film-cooling is one typically used cooling method whereby coolant is supplied through holes placed along the Pressure Side of a blade. The subject of this paper is to evaluate the adiabatic effectiveness levels that occur on the blade tip through blowing coolant from holes placed near the tip of a blade along the Pressure Side. A range of blowing ratios was studied whereby coolant was injected from holes placed along the Pressure Side tip of a large-scale blade model. Also present were dirt purge holes on the blade tip, which is part of a commonly used blade design to expel any large particles present in the coolant stream. Experiments were conducted in a linear cascade with a scaled-up turbine blade whereby the Reynolds number of the engine was matched. This paper, which is Part I of a two part series, compares adiabatic effectiveness levels measured along a blade tip, while Part 2 combines measured heat transfer coefficients with the adiabatic effectiveness levels to assess the overall cooling benefit of Pressure Side blowing near a blade tip. The results show much better cooling can be achieved for a small tip gap compared with a large tip gap with different flow phenomena occurring for each tip gap setting.

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes—Part II: Heat Transfer Measurements
    Journal of Turbomachinery, 2005
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    The clearance gap between a turbine blade tip and its associated shroud allows leakage flow across the tip gap from the Pressure Side to the suction Side of the blade. Understanding how this leakage flow affects heat transfer is critical in extending blade tip durability in terms of oxidation, erosion, clearance, and overall turbine performance. This paper is the second of a two part series that discusses the augmentation of tip heat transfer as a result of blowing from the Pressure Side of the tip as well as dirt purge holes placed on the tip. For the experimental investigation, three scaled-up blades were used to form a two-passage linear cascade in a low speed wind tunnel. The rig was designed to simulate different tip gap sizes and coolant flow rates. Heat transfer coefficients were quantified by measuring the total power supplied to a constant heat flux surface placed on the tip of the blade and measuring the tip temperatures. Results indicate that increased blowing leads to increased augmentations in tip heat transfer, particularly at the entrance region to the gap. When combined with adiabatic effectiveness measurements, the coolant from the Pressure Side holes provides an overall net heat flux reduction to the blade tip but is nearly independent of coolant flow levels.

  • Cooling the Tip of a Turbine Blade Using Pressure Side Holes: Part 2 — Heat Transfer Measurements
    Volume 3: Turbo Expo 2004, 2004
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    The clearance gap between a turbine blade tip and its associated shroud allows leakage flow across the tip gap from the Pressure Side to the suction Side of the blade. Understanding how this leakage flow affects heat transfer is critical in extending blade tip durability in terms of oxidation, erosion, clearance, and overall turbine performance. This paper is the second of a two part series that discusses the augmentation of tip heat transfer as a result of blowing from the Pressure Side of the tip as well as dirt purge holes placed on the tip. For the experimental investigation, three scaled-up blades were used to form a two-passage linear cascade in a low speed wind tunnel. The rig was designed to simulate different tip gap sizes and coolant flow rates. Heat transfer coefficients were quantified by measuring the total power supplied to a constant heat flux surface placed on the tip of the blade and measuring the tip temperatures. Results indicate that increased blowing leads to increased augmentations in tip heat transfer, particularly at the entrance region to the gap. When combined with adiabatic effectiveness measurements, the coolant from the Pressure Side holes provides an overall net heat flux reduction to the blade tip but is nearly independent of coolant flow levels.© 2004 ASME

  • Cooling the tip of a turbine blade using Pressure Side holes - Part 1: Adiabatic effectiveness measurements
    Volume 3: Turbo Expo 2004, 2004
    Co-Authors: J. R. Christophel, Karen A. Thole, Frank J. Cunha
    Abstract:

    Sealing and durability for turbine blade tips have been challenging problems since the development of gas turbine engines. Blade tip designs are extremely important in terms of sealing and engine performance. In general, overall engine performance can be improved by increasing turbine inlet temperatures. As a result, cooling methods along the blade tip need to be devised and applied effectively. Film-cooling is typically used as a blade tip cooling method, whereby coolant is supplied through holes placed along the Pressure Side of a blade. Experiments were conducted in a linear cascade with a scaled-up turbine blade whereby the Reynolds number of the engine was matched. A range of blowing ratios was studied whereby coolant was injected from holes placed along the Pressure Side tip of the blade as well as from dirt purge holes placed on the blade tip. This paper, which is Part 1 of a two part series, compares adiabatic effectiveness levels measured along a blade tip, while Part 2 combines measured heat transfer coefficients with the adiabatic effectiveness levels to assess the overall cooling benefit of Pressure Side blowing near a blade tip. The results show better cooling can be achieved for a small tip gap compared with a large tip gap with different flow phenomena occurring for each tip gap setting.Copyright © 2004 by ASME

N. Sundaram - One of the best experts on this subject based on the ideXlab platform.

  • Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side
    Journal of Turbomachinery, 2008
    Co-Authors: N. Sundaram, M. D. Barringer, Karen A. Thole
    Abstract:

    Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the Pressure Side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the Pressure Side of the airfoil. Deposits having a semielliptical shape were placed along the Pressure Side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the Pressure Side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

  • Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side
    Volume 4: Turbo Expo 2007 Parts A and B, 2007
    Co-Authors: N. Sundaram, M. D. Barringer, Karen A. Thole
    Abstract:

    Film-cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film-cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film-cooling along the Pressure Side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the Pressure Side of the airfoil. Deposits having a semi-elliptical shape were placed along the Pressure Side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets towards the vane endwall junction on the Pressure Side. Results also indicated that there was a steady decrease in adiabatic effectiveness with a sequential increase in the number of rows with the deposits.Copyright © 2007 by ASME