Bump and Trench Modifications to Film-Cooling Holes at the Vane-Endwall Junction

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first-stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases toward it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely, trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

Bump and Trench Modifications to Film-Cooling Holes at the Vane Endwall Junction

2007 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases towards it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

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.

Film-Cooling Flowfields With Trenched Holes on an Endwall

2008 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
High Heat ◽  
Transfer Rates ◽  
Baseline Study ◽  
High Heat Transfer ◽  
Stator Vane ◽  
Horseshoe Vortices ◽  
Film Cooling Holes

The leading edge region along the endwall of a stator vane experiences high heat transfer rates resulting from the formation of horseshoe vortices. Typical gas turbine endwall designs include a leakage slot at the combustor-turbine interface as well as film-cooling holes. Past studies have documented the formation of a horseshoe vortex at the leading edge of a vane, but few studies have documented the flowfield in the presence of an interface slot and film-cooling jets. In this paper, a series of flowfield measurements are evaluated at the leading edge with configurations including: a baseline with neither film-cooling holes nor an upstream slot, a row of film-cooling holes and an interface slot, and a row of film-cooling holes in a trench and an interface slot. The results indicated the formation of a second vortex present for the case with film-cooling holes and a slot relative to the baseline study. In addition, turbulence intensity levels as high as 50% were measured at the leading edge with film-cooling holes and a slot compared to the 30% measured for the baseline study. A trench was shown to provide improved overall cooling relative to the no trench configuration as more of the coolant stayed attached to the endwall surface with the trench.

Heat Transfer and Effectiveness on the Film Cooled Tip and Inner Rim Surfaces of a Turbine Blade

2010 ◽  
Author(s):  
Jun Su Park ◽  
Dong Hyun Lee ◽  
Hyung Hee Cho ◽  
Dong-Ho Rhee ◽  
Shin-Hyung Kang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
High Heat ◽  
Chord Length ◽  
Tip Clearance ◽  
Film Cooling Effectiveness ◽  
High Heat Transfer ◽  
Film Cooling Holes

Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a rotor blade with a squealer rim. The blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, the axial chord length (Cx) of which was 237 mm with a turning angle of 126°. The mainstream Reynolds number based on the axial chord was 1.5×105. The turbulence intensity level at the cascade inlet was approximately 12%. Measurements were made at three different rim heights (H) of about 3%, 6%, and 9% of the axial chord length. The tip clearance (C) ranges were 1–3% of the axial chord length. Also, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and the suction side rim. In particular, a coolant was injected at an incline of 45° from near the suction side film cooling holes. The film cooling experiments were done with a fixed tip clearance and rim height at 1% and 6% of the axial chord length. The blowing rate was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface in some cases, due to the reattachment of tip leakage flow. The peak values moved toward the suction-side edge, and the magnitude and area of high heat transfer increased near the leading edge as the tip clearance increased. The heat transfer decreased on the tip surface with increases in the rim height. In the film-cooling cases, the high heat transfer and film-cooling effectiveness region appeared near the film-cooling holes.

Film-Cooling Flowfields With Trenched Holes on an Endwall

2009 ◽  
Vol 131 (4) ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
High Heat ◽  
Transfer Rates ◽  
Baseline Study ◽  
High Heat Transfer ◽  
Stator Vane ◽  
Horseshoe Vortices ◽  
Film Cooling Holes

The leading edge region along the endwall of a stator vane experiences high heat transfer rates resulting from the formation of horseshoe vortices. Typical gas turbine endwall designs include a leakage slot at the combustor-turbine interface as well as film-cooling holes. Past studies have documented the formation of a horseshoe vortex at the leading edge of a vane, but few studies have documented the flowfield in the presence of an interface slot and film-cooling jets. In this paper, a series of flowfield measurements is evaluated at the leading edge with configurations including a baseline with neither film-cooling holes nor an upstream slot, a row of film-cooling holes and an interface slot, and a row of film-cooling holes in a trench and an interface slot. The results indicated the formation of a second vortex present for the case with film-cooling holes and a slot relative to the baseline study. In addition, turbulence intensity levels as high as 50% were measured at the leading edge with film-cooling holes and a slot compared with the 30% measured for the baseline study. A trench was shown to provide improved overall cooling relative to the no trench configuration as more of the coolant stayed attached to the endwall surface with the trench.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2007 ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

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.

Computational Analysis of Conjugate Heat Transfer and Particulate Deposition on a High Pressure Turbine Vane

2009 ◽  
Author(s):  
Weiguo Ai ◽  
Thomas H. Fletcher
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Particle Method ◽  
Experimental Results ◽  
Turbine Vane ◽  
Particulate Deposition ◽  
Cooling Hole ◽  
Lagrangian Particle Method ◽  
Film Cooling Holes

Numerical computations were conducted to simulate flyash deposition experiments on gas turbine disk samples with internal impingement and film cooling using a CFD code (FLUENT). The standard k-ω turbulence model and RANS were employed to compute the flow field and heat transfer. The boundary conditions were specified to be in agreement with the conditions measured in experiments performed in the BYU Turbine Accelerated Deposition Facility (TADF). A Lagrangian particle method was utilized to predict the ash particulate deposition. User-defined subroutines were linked with FLUENT to build the deposition model. The model includes particle sticking/rebounding and particle detachment, which are applied to the interaction of particles with the impinged wall surface to describe the particle behavior. Conjugate heat transfer calculations were performed to determine the temperature distribution and heat transfer coefficient in the region close to the film-cooling hole and in the regions further downstream of a row of film-cooling holes. Computational and experimental results were compared to understand the effect of film hole spacing, hole size and TBC on surface heat transfer. Calculated capture efficiencies compare well with experimental results.

scholarly journals A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

1999 ◽  
Author(s):  
James D. Heidmann ◽  
David L. Rigby ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Film Cooling ◽  
Three Dimensional ◽  
High Heat ◽  
Operating Conditions ◽  
High Heat Flux ◽  
Nasa Lewis Research ◽  
Turbine Vane ◽  
Film Cooling Holes

A three-dimensional Navier-Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Lewis Research Center experiment and has both circular cross-section and shaped film cooling holes. This complex geometry is modeled using a multi-block grid which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat flux in the showerhead region due to low film effectiveness and high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges.

Effects on Film Cooling Performance in the Showerhead From Geometric Parameterization of Shaped Hole Designs

2021 ◽  
Author(s):  
Jacob D. Moore ◽  
Christopher C. Easterby ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Thermal Protection ◽  
Leading Edge ◽  
Field Measurements ◽  
High Heat ◽  
Area Ratio ◽  
Cooling Performance ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Marginal Improvement

Abstract The high heat loads at the leading-edge regions of turbine vanes and blades necessitate the most robust thermal protection, typically accomplished via a dense array of film cooling holes, nicknamed the “showerhead.” Although research has shown that film cooling using shaped holes provides more reliable thermal protection than that using cylindrical holes, the effects on cooling performance from varying the geometric details of the shaped hole design are not well characterized. In this study, adiabatic effectiveness and off-the-wall thermal field measurements were conducted for two shaped hole geometries designed as successors to a baseline hole geometry presented in a previous study. One geometry with a 40% increase in area ratio exhibited only a marginal improvement in adiabatic effectiveness (∼10%). A second design with a 12° forward and lateral expansion angle with a breakout area 40% larger performed marginally worse than its matched area ratio counterpart (∼15% lower), suggesting a negative sensitivity to breakout area. Such changes in performance for different shaped hole designs were small compared to the boost in performance gained by switching from a cylindrical hole to a shaped hole, which suggests cooling performance is insensitive to specific shaped hole details provided the exterior coolant flow is well-attached.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2014 ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the under-prediction by superposition analysis are presented.

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