Film Cooling on a Modern HP Turbine Blade: Part I — Experimental and Computational Methodology and Validation

2002 ◽  
Author(s):  
Frederick A. Buck ◽  
D. Keith Walters ◽  
Jeffrey D. Ferguson ◽  
E. Lee McGrath ◽  
James H. Leylek
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Large Strain ◽  
Density Ratio ◽  
Gas Analysis ◽  
Computational Techniques ◽  
Mean Flow ◽  
Suction Surface ◽  
Computational Methodology ◽  
Compound Angle

State-of-the-art experimental and computational techniques are used to study film cooling on the suction and pressure surfaces of a modern turbine blade under realistic engine conditions. Measured data and predicted results are compared for coolant jets injected through a row of three fundamentally different configurations: (1) Compound-angle round (CAR) holes; (2) Axial shaped holes (ASH); and (3) Compound-angle shaped holes (CASH). Experiments employ a single-passage cascade for validation-quality adiabatic film effectiveness measurements using a gas analysis technique. Computations use a novel combination of geometry and grid generation techniques, discretization scheme, turbulence modeling, and numerical solvers to evaluate a “best practice” standard for use in the gas turbine industry. The gridding procedure uses a super-block, multi-topology, unstructured/adaptive, non-conformal, near-wall resolved mesh to accurately capture all of the mean flow features of the 3-D jet-in-crossflow interaction. The effects of blowing ratio (M) are examined, with M = 1.0, 1.5, and 2.0 on the suction surface and M = 1.5, 3.0, and 4.5 on the pressure surface. All simulations are run with a density ratio of 1.52. The simulations model the three-way coupling between a transonic blade passage flow, subsonic film-hole flow, and creeping plenum flow; high pressure gradients; high rates of curvature; and large strain-rates found in actual engines. Computed results are compared to experimental data in terms of aerodynamic loading and spanwise-averaged adiabatic effectiveness on the blade surfaces in order to validate the computational methodology for this class of problems and to explain the mechanisms responsible for the performance of CAR, ASH, and CASH configurations.

Film Cooling on a Modern HP Turbine Blade: Part II — Compound-Angle Round Holes

2002 ◽  
Author(s):  
D. Keith Walters ◽  
James H. Leylek ◽  
Frederick A. Buck
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Mean Flow ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Pressure Surface ◽  
Nominal Density ◽  
High Turbulence ◽  
The Mean ◽  
Compound Angle

A well-tested computational methodology and a companion experimental study are used to analyze the physics of compound-angle, cylindrical-hole film cooling on the pressure and suction surfaces of a modern high-pressure turbine airfoil. A single-passage cascade (SPC) is used to model the blade passage flow experimentally and computationally. Realistic engine conditions, including transonic flow, high turbulence levels, and a nominal density ratio of 1.52, are used to examine blowing ratios of 1.0, 1.5, and 2.0 on the suction surface (SS) and 1.5, 3.0, and 4.5 on the pressure surface (PS). The predicted results agree with experimental trends, and differences are explained in terms of known deficiencies in the turbulence treatment. The mean-flow physics downstream of coolant injection are influenced primarily by a single dominant vortex that entrains coolant and mainstream fluid, and by the effect of convex (SS) or concave (PS) curvature on the coolant jet.

Film Cooling on a Modern HP Turbine Blade: Part III — Axial Shaped Holes

2002 ◽  
Author(s):  
Jeffrey D. Ferguson ◽  
James H. Leylek ◽  
Frederick A. Buck
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Flow Characteristics ◽  
Quality Data ◽  
Suction Surface ◽  
Predictive Tool ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Hole Shaping

A well-tested computational methodology and high-quality data from a companion experimental study are used to analyze the physics of axial-injected, shaped-hole film cooling on the pressure and suction surfaces of a modern high-pressure turbine blade. Realistic engine conditions, including transonic flow, high turbulence levels, and a nominal density ratio of 1.52, are used to examine blowing ratios of 1.0, 1.5, and 2.0 on the suction surface (SS) and 1.5, 3.0, and 4.5 on the pressure surface (PS). SS results show excellent film-cooling performance with the hole shaping, but massive hot crossflow ingestion is found using similar hole shaping on the PS. Primary mechanisms governing the near and far-field cooling effectiveness and crossflow ingestion are identified, including: (1) the nature of the coolant entry into the film hole; (2) location of hole shaping relative to major coolant flow characteristics; and (3) susceptibility of low-momentum fluid to pressure gradients. Changes in blowing ratio, while not introducing new physical mechanisms, significantly alter the extent to which the mechanisms already present affect the flow. These effects are highly non-linear for both SS and PS geometries, highlighting the inadequacy of one-dimensional design practices and the potential usefulness of CFD as a predictive tool.

Film Cooling on a Modern HP Turbine Blade: Part IV — Compound-Angle Shaped Holes

2002 ◽  
Author(s):  
E. Lee McGrath ◽  
James H. Leylek ◽  
Frederick A. Buck
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Thermal Protection ◽  
Density Ratio ◽  
Lower Percentage ◽  
Suction Surface ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Flow Features ◽  
Compound Angle

The performance and physics of film cooling with compound-angle shaped holes on a modern high-pressure turbine airfoil is studied in detail using state-of-the-art computational simulations. Computations model high-speed single-airfoil-passage cascade experiments, and computational results show good agreement with experimental data. Evaluation of physics includes examination of flow features and adiabatic effectiveness. The blowing ratios (M) simulated on the pressure surface (PS) of the blade are 1.5, 3.0, and 4.5, with a single density ratio of 1.52. On the pressure surface the dominant mechanism affecting coolant behavior is vorticity, which increasingly tucks hot crossflow under the coolant as the blowing ratio increases. Thus at high blowing ratios, a lower percentage of the coolant provides thermal protection for the blade until the vortices dissipate far downstream. Also, the vortex structures cause large lateral temperature gradients despite the lateral motion of the flow induced by the compound-angle injection. The dominance of vorticity can be attributed to poor diffusion of the coolant inside the diffuser of the film hole. On the suction surface (SS), the simulated blowing ratios are 1.0, 1.5, and 2.0, with a single density ratio of 1.52. Pressure gradients normal to the SS result in the flow pushing the coolant onto the blade. Also, vorticity is less dominant since diffusion of coolant inside the film hole is better due to low blowing ratios and due to a hole metering section that is almost 3 times longer than that of the PS hole. Hot crossflow ingestion into the film hole is observed at M = 2.0. Ingested crossflow causes heating of the surface inside the hole that extends down to the end of the hole metering section, where the surface temperatures are approximately equal to an average of the crossflow and coolant temperatures. These results demonstrate the inadequacy of 1-D, empirical design tools and demonstrate the need for a validated CFD-based film cooling methodology.

scholarly journals A Detailed Analysis of Film Cooling Physics: Part IV — Compound–Angle Injection With Shaped Holes

1997 ◽  
Author(s):  
Robert A. Brittingham ◽  
James H. Leylek
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Diameter Ratio ◽  
Valuable Insight ◽  
Combined Effects ◽  
Hole Shaping ◽  
Cylindrical Holes ◽  
Shaped Hole ◽  
Computational Methodology ◽  
Compound Angle

The flow physics of film cooling with compound–angle shaped holes is documented for realistic gas turbine parameters. For the first time in the open literature, the combined effects of compound–angle injection and hole shaping are isolated and the dominant mechanisms are examined. Results provide valuable insight into the flowfield of this class of film–cooling jets. Computational and experimental results are presented for a row of holes injected at 35° on a flat plate with three distinct geometric configurations: (1) streamwise injected cylindrical holes (reference case); (2) 15° forward–diffused holes injected at a 60° compound angle; and (3) 12° laterally–diffused holes injected at a 45° compound angle. Detailed field and surface data, including adiabatic effectiveness (η) and heat transfer coefficient (h), of the two compound–angle shaped holes are provided and compared to: (i) the reference streamwise cylindrical case; (ii) results from Part II detailing the compound–angle flowfield for cylindrical holes; (in) results of Part III detailing the streamwise injected shaped–hole flowfield; and (iv) experimental data. The 60° compound–angle forward–diffused holes provided excellent lateral coolant distribution, but suffered from crossflow ingestion at the film–hole exit plane. The 45° compound–angle lateral–diffused hole had much steeper lateral effectiveness variations. A previously documented and validated computational methodology was utilized. Computations were performed using a multi–block, unstructured–adaptive grid, fully–implicit pressure–correction Navier–Stokes code with multi–grid and under–relaxation type convergence accelerators. All simulations had fixed length–to–diameter ratio of 4.0, pitch–to–diameter ratio of 3.0, nominal density ratio of 1.55 and film–hole Reynolds number of 17350, which allowed isolation of the combined effects of compound–angle injection and hole shaping for nominal blowing ratios of 1.25 and 1.88. The results demonstrate the ability of the prescribed computational methodology to accurately predict the complex flowfield associated with compound–angle shaped–hole film–cooling jets.

Influence of Coolant Density on Turbine Blade Film-Cooling With Axial and Compound Shaped Holes

2014 ◽  
Vol 136 (4) ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Axial Angle ◽  
Lift Off ◽  
Compound Angle

Adiabatic film-cooling effectiveness is examined systematically on a typical high pressure turbine blade by varying three critical flow parameters: coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three coolant density ratios 1.0, 1.5, and 2.0 are chosen for this study. The average blowing ration and the turbulence intensity are 1.5% and 10.5%, respectively. Conduction-free pressure sensitive paint (PSP) technique is used to measure film-cooling effectiveness. Foreign gases are used to study the effect of coolant density. Two test blades feature axial angle and 45 deg compound-angle shaped holes on the suction side and pressure side. Both designs have 3 rows of 30 deg radial-angle cylindrical holes around the leading edge region. The inlet and the exit Mach number are 0.27 and 0.44, respectively. Reynolds number based on the exit velocity and blade axial chord length is 750,000. Overall, the compound angle design performs better film coverage that axial angle. Greater coolant-to-mainstream density ratio results in lower coolant-to-mainstream momentum and prevents coolant to lift-off.

scholarly journals Numerical Prediction of Mainstream Pressure Gradient Effects in Film Cooling

1999 ◽  
Author(s):  
William D. York ◽  
James H. Leylek
Keyword(s):  
Pressure Gradient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Applied Pressure ◽  
Density Ratio ◽  
Pressure Gradients ◽  
Suction Surface ◽  
Reference Case ◽  
Injection Angle ◽  
Computational Methodology

A documented, systematic, computational methodology is applied to singularly investigate the effects of mainstream pressure gradients on film cooling over a flat surface for realistic gas turbine parameters. Key aspects of the study include: (1) validation of the ability of computational fluid dynamics to simulate film cooling in regions of mainstream pressure gradients, accomplished through the isolation of this parameter and the careful modeling of a published experimental study; (2) documentation of the effects of the applied pressure gradient on film cooling adiabatic effectiveness, as compared to the zero-pressure gradient case; and (3) detailed discussion of the pertinent physical mechanisms involved, with appropriate flowfield results. The imposed pressure gradient is typical of the suction surface of a gas turbine airfoil, with a strong favorable pressure gradient (the acceleration parameter was K = 1.5×10−6 at injection) transitioning to a mild adverse pressure gradient region beyond 30 diameters downstream. A single row of cylindrical film-cooling holes had an injection angle of 35°, with hole length-to-diameter ratio of 4.0 and a lateral spacing of 3.0 diameters. The simulated mass flux ratios were M = 0.6, 1.0, and 1.5, and the density ratio was held constant at 1.6. Solutions were obtained using a multi-block, multi-topology grid and a pressure-correction based, fully-implicit Navier-Stokes solver. A “realizeable” k-ε turbulence model, which eliminates the documented unrealistic turbulence production of the standard k-ε model in regions of large flow strain, was employed to obtain practical results economically. The applied pressure gradient resulted in a small advantage in center-line effectiveness, while laterally averaged effectiveness was slightly lower as compared to the zero-pressure gradient reference case. The results of this study demonstrate the ability of the applied computational methodology to accurately model film cooling in the presence of mainstream pressure gradients and resolve one of the key fundamental issues in turbine airfoil film cooling.

Assessment of a Double Hole Film Cooling Geometry Using S-PIV and PSP

2013 ◽  
Author(s):  
Lesley M. Wright ◽  
Stephen T. McClain ◽  
Charles P. Brown ◽  
Weston V. Harmon
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Three Dimensional ◽  
Field Measurements ◽  
Secondary Flows ◽  
Cylindrical Hole ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Hole Geometry ◽  
Compound Angle

A novel, double hole film cooling configuration is investigated as an alternative to traditional cylindrical and fanshaped, laidback holes. This experimental investigation utilizes a Stereo-Particle Image Velocimetry (S-PIV) to quantitatively assess the ability of the proposed, double hole geometry to weaken or mitigate the counter-rotating vortices formed within the jet structure. The three-dimensional flow field measurements are combined with surface film cooling effectiveness measurements obtained using Pressure Sensitive Paint (PSP). The double hole geometry consists of two compound angle holes. The inclination of each hole is θ = 35°, and the compound angle of the holes is β = ± 45° (with the holes angled toward one another). The simple angle cylindrical and shaped holes both have an inclination angle of θ = 35°. The blowing ratio is varied from M = 0.5 to 1.5 for all three film cooling geometries while the density ratio is maintained at DR = 1.0. Time averaged velocity distributions are obtained for both the mainstream and coolant flows at five streamwise planes across the fluid domain (x/d = −4, 0, 1, 5, and 10). These transverse velocity distributions are combined with the detailed film cooling effectiveness distributions on the surface to evaluate the proposed double hole configuration (compared to the traditional hole designs). The fanshaped, laidback geometry effectively reduces the strength of the kidney-shaped vortices within the structure of the jet (over the entire range of blowing ratios considered). The three-dimensional velocity field measurements indicate the secondary flows formed from the double hole geometry strengthen in the plane perpendicular to the mainstream flow. At the exit of the double hole geometry, the streamwise momentum of the jets is reduced (compared to the single, cylindrical hole), and the geometry offers improved film cooling coverage. However, moving downstream in the steamwise direction, the two jets form a single jet, and the counter-rotating vortices are comparable to those formed within the jet from a single, cylindrical hole. These strong secondary flows lift the coolant off the surface, and the film cooling coverage offered by the double hole geometry is reduced.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator–Rotor Purge Flow and Compound-Angle Holes

2014 ◽  
Vol 6 (4) ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Fundamental Parameters ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Exit Mach Number ◽  
Compound Angle

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator–rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study—a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and was based on the exit velocity and chord length of the blade. The measurement technique adopted was the conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with an increase of blowing ratio. Results also indicate that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

Experimental Study of the Effects of an Oscillating Approach Flow on Overall Cooling Performance of a Simulated Turbine Blade Leading Edge

2009 ◽  
Author(s):  
Ross Johnson ◽  
Jonathan Maikell ◽  
David Bogard ◽  
Justin Piggush ◽  
Atul Kohli ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Oscillation Frequencies ◽  
Overall Effectiveness ◽  
Blade Leading Edge

When a turbine blade passes through wakes from upstream vanes it is subjected to an oscillation of the direction of the approach flow resulting in the oscillation of the position of the stagnation line on the leading edge of the blade. In this study an experimental facility was developed that induced a similar oscillation of the stagnation line position on a simulated turbine blade leading edge. The overall effectiveness was evaluated at various blowing ratios and stagnation line oscillation frequencies. The location of the stagnation line on the leading edge was oscillated to simulate a change in angle of attack between α = ± 5° at a range of frequencies from 2 to 20 Hz. These frequencies were chosen based on matching a range of Strouhal numbers typically seen in an engine due to oscillations caused by passing wakes. The blowing ratio was varied between M = 1, M = 2, and M = 3. These experiments were carried out at a density ratio of DR = 1.5 and mainstream turbulence levels of Tu ≈ 6%. The leading edge model was made of high conductivity epoxy in order to match the Biot number of an actual engine airfoil. Results of these tests showed that the film cooling performance with an oscillating stagnation line was degraded by as much as 25% compared to the performance of a steady flow with the stagnation line aligned with the row of holes at the leading edge.

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