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.

Performance Prediction of a Novel Design for Turbine Vane Film Cooling With Low Aerodynamic Loss and High Effectiveness

2004 ◽  
Author(s):  
William D. York ◽  
James H. Leylek
Keyword(s):  
Film Cooling ◽  
Stokes Equations ◽  
Density Ratio ◽  
Leading Edge ◽  
Cost Effective ◽  
Design Environment ◽  
Suction Surface ◽  
Reference Case ◽  
Aerodynamic Loss ◽  
Turbine Vane

A new film-cooling scheme for the suction surface of a gas turbine vane in a transonic cascade is studied numerically. The concept of the present design is to inject a substantial amount of coolant at a very small angle, approaching a “wall-jet,” through a single row of relatively few, large holes near the vane leading edge. The near-match of the coolant stream and mainstream momentums, coupled with the low coolant trajectory, theoretically results in low aerodynamic losses due to mixing. A minimal effect of the film cooling on the vane loading is also important to realize, as well as good coolant coverage and high adiabatic effectiveness. A systematic computational methodology, developed in the Advanced Computational Research Laboratory (ACRL) and tested numerous times on film-cooling applications, is applied in the present work. For validation purposes, predictions from two previous turbine airfoil film-cooling studies, both employing this same numerical method, are presented and compared to experimental data. Simulations of the new film-cooling configuration are performed for two blowing ratios, M=0.90 and M=1.04, and the density ratio of the coolant to the mainstream flow is unity in both cases. A solid vane with no film cooling is also studied as a reference case in the evaluation of losses. The unstructured numerical mesh contains about 5.5 million finite-volumes, after solution-based adaption. Grid resolution is such that the full boundary layer and all passage shocks are resolved. The Renormalization Group (RNG) k-ε turbulence model is used to close the Reynolds-averaged Navier-Stokes equations. Predictions indicate that the new film-cooling scheme meets design intent and has negligible impact on the total pressure losses through the vane cascade. Additionally, excellent coolant coverage is observed all the way to the trailing edge, resulting in high far-field effectiveness. Keeping the design environment in mind, this work represents the power of validated computational methods to provide a rapid and reasonably cost-effective analysis of innovative turbine airfoil cooling.

Scaling of Performance for Varying Density Ratio Coolants on an Airfoil With Strong Curvature and Pressure Gradient Effects

2000 ◽  
Author(s):  
Marcia I. Ethridge ◽  
J. Michael Cutbirth ◽  
David G. Bogard
Keyword(s):  
Pressure Gradient ◽  
Film Cooling ◽  
Density Ratio ◽  
Flux Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Cooling Performance ◽  
Injection Angle ◽  
Adiabatic Effectiveness ◽  
Strong Curvature

An experimental study was conducted to investigate the film cooling performance on the suction side of a first stage turbine vane. Tests were conducted on a nine times scale vane model at density ratios of DR = 1.1 and 1.6 over a range of blowing conditions, 0.2 ≤ M ≤ 1.5 and 0.05 ≤ I ≤ 1.2. Two different mainstream turbulence intensity levels, Tu∞ = 0.5% and 20%, were also investigated. The row of coolant holes studied was located in a position of both strong curvature and strong favorable pressure gradient. In addition, its performance was isolated by blocking the leading edge showerhead coolant holes. Adiabatic effectiveness measurements were made using an infrared camera to map the surface temperature distribution. The results indicate that film cooling performance was greatly enhanced over holes with a similar 50° injection angle on a flat plate. Overall, adiabatic effectiveness scaled with mass flux ratio for low blowing conditions and with momentum flux ratio for high blowing conditions. However, for M < 0.5 there was a higher rate of decay for the low density ratio data. High mainstream turbulence had little effect at low blowing ratios, but degraded performance at higher blowing ratios.

scholarly journals Effects of Geometry on Slot–Jet Film Cooling Performance

1996 ◽  
Author(s):  
Daniel G. Hyams ◽  
Kevin T. McGovern ◽  
James H. Leylek
Keyword(s):  
Film Cooling ◽  
Thermal Field ◽  
Density Ratio ◽  
Width Ratio ◽  
Surface Phenomena ◽  
Reference Case ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Slot Length ◽  
Computational Methodology

The physics of the film cooling process for shaped, inclined slot–jets with realistic slot–length–to–width ratios (L/s) is studied for a range of blowing ratio (M) and density ratio (DR) parameters typical of gas turbine operations. The effect of inlet and exit shaping of the slot–jet on both flow and thermal field characteristics is isolated, and the dominant mechanisms responsible for differences in these characteristics are documented. A previously documented computational methodology was applied for the study of four distinct configurations: (1) slot with straight edges and sharp comers (reference case); (2) slot with shaped inlet region; (3) slot with shaped exit region; and (4) slot with both shaped inlet and exit regions. Detailed field results as well as surface phenomena involving adiabatic film effectiveness (η) and heat transfer coefficient (h) are presented. It is demonstrated that both η and h results are vital in the proper assessment of film cooling performance. The key parameters M and DR were varied from 1.0 to 2.0 and 1.5 to 2.0, respectively, to show their influence. Simulations were repeated for slot length–to–width ratio (L/s) of 3.0 and 4.5 in order to explain the effects of this important parameter. The computational simulations showed exceptionally strong internal consistency. Moreover, the ability of using a state–of–the–art computational methodology to sort the relative performance of different slot–jet film cooling configurations was clearly established.

A Systematic Computational Methodology Applied to a Three-Dimensional Film-Cooling Flowfield

1997 ◽  
Vol 119 (4) ◽  
pp. 777-785 ◽  
Author(s):  
D. K. Walters ◽  
J. H. Leylek
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Turbulence Modeling ◽  
Density Ratio ◽  
Three Dimensional ◽  
Elliptic Solution ◽  
Injection Angle ◽  
Free Stream Turbulence ◽  
Acceleration Techniques ◽  
Computational Methodology

Numerical results are presented for a three-dimensional discrete-jet in crossflow problem typical of a realistic film-cooling application in gas turbines. Key aspects of the study include: (1) application of a systematic computational methodology that stresses accurate computational model of the physical problem, including simultaneous, fully elliptic solution of the crossflow, film-hole, and plenum regions; high-quality three-dimensional unstructured grid generation techniques, which have yet to be documented for this class of problems; the use of a high-order discretization scheme to reduce numerical errors significantly; and effective turbulence modeling; (2) a three-way comparison of results to both code validation quality experimental data and a previously documented structured grid simulation; and (3) identification of sources of discrepancy between predicted and measured results, as well as recommendations to alleviate these discrepancies. Solutions were obtained with a multiblock, unstructured/adaptive grid, fully explicit, time-marching, Reynolds-averaged Navier–Stokes code with multigrid, local time stepping, and residual smoothing type acceleration techniques. The computational methodology was applied to the validation test case of a row of discrete jets on a flat plate with a streamwise injection angle of 35 deg, and two film-hole length-to-diameter ratios of 3.5 and 1.75. The density ratio for all cases was 2.0, blowing ratio was varied from 0.5 to 2.0, and free-stream turbulence intensity was 2 percent. The results demonstrate that the prescribed computational methodology yields consistently more accurate solutions for this class of problems than previous attempts published in the open literature. Sources of disagreement between measured and computed results have been identified, and recommendations made for future prediction of film-cooling problems.

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.

Scaling of Performance for Varying Density Ratio Coolants on an Airfoil With Strong Curvature and Pressure Gradient Effects

2000 ◽  
Vol 123 (2) ◽  
pp. 231-237 ◽  
Author(s):  
Marcia I. Ethridge ◽  
J. Michael Cutbirth ◽  
David G. Bogard
Keyword(s):  
Pressure Gradient ◽  
Film Cooling ◽  
Density Ratio ◽  
Flux Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Cooling Performance ◽  
Injection Angle ◽  
Adiabatic Effectiveness ◽  
Strong Curvature

An experimental study was conducted to investigate the film cooling performance on the suction side of a first-stage turbine vane. Tests were conducted on a nine times scale vane model at density ratios of DR=1.1 and 1.6 over a range of blowing conditions, 0.2⩽M⩽1.5 and 0.05⩽I⩽1.2. Two different mainstream turbulence intensity levels, Tu∞=0.5 and 20 percent, were also investigated. The row of coolant holes studied was located in a position of both strong curvature and strong favorable pressure gradient. In addition, its performance was isolated by blocking the leading edge showerhead coolant holes. Adiabatic effectiveness measurements were made using an infrared camera to map the surface temperature distribution. The results indicate that film cooling performance was greatly enhanced over holes with a similar 50 deg injection angle on a flat plate. Overall, adiabatic effectiveness scaled with mass flux ratio for low blowing conditions and with momentum flux ratio for high blowing conditions. However, for M<0.5, there was a higher rate of decay for the low density ratio data. High mainstream turbulence had little effect at low blowing ratios, but degraded performance at higher blowing ratios.

A Systematic Computational Methodology Applied to a Three–Dimensional Film–Cooling Flowfield

1996 ◽  
Author(s):  
Dibbon K. Walters ◽  
James H. Leylek
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Density Ratio ◽  
Three Dimensional ◽  
Test Case ◽  
Elliptic Solution ◽  
Injection Angle ◽  
Free Stream Turbulence ◽  
Acceleration Techniques ◽  
Computational Methodology

Numerical results are presented for a three–dimensional discrete–jet in crossflow problem typical of a realistic film–cooling application in gas turbines. Key aspects of the study include: (1) Application of a systematic computational methodology that stresses accurate computational model of the physical problem, including simultaneous, fully–elliptic solution of the crossflow, film–hole, and plenum regions; high quality 3–D unstructured grid generation techniques which have yet to be documented for this class of problems; the use of a high order discretization scheme to significantly reduce numerical errors; and effective turbulence modelling; (2) A three–way comparison of results to both code validation quality experimental data and a previously documented structured grid simulation; and (3) Identification of sources of discrepancy between predicted and measured results, as well as recommendations to alleviate these discrepancies. Solutions were obtained with a multi–block, unstructured/adaptive grid, fully explicit, time–marching, Reynolds averaged Navier–Stokes code with multi-grid, local time stepping, and residual smoothing type acceleration techniques. The computational methodology was applied to the validation test case of a row of discrete jets on a flat plate with a streamwise injection angle of 35°, and two film–hole length–to–diameter ratios of 3.5 and 1.75. The density ratio for all cases was 2.0, blowing ratio was varied from 0.5 to 2.0, and free–stream turbulence intensity was 2%. The results demonstrate that the prescribed computational methodology yields consistently more accurate solutions for this class of problems than previous attempts published in the open literature. Sources of disagreement between measured and computed results have been identified, and recommendations made for future prediction of this class of problems.

Numerical Investigation of Effects of Blockage, Inclination Angle, and Hole-Size on Film Cooling Effectiveness at Concave Surface

2021 ◽  
Vol 143 (2) ◽  
Author(s):  
Fu-qiang Wang ◽  
Jian Pu ◽  
Jian-hua Wang ◽  
Wei-dong Xia
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Numerical Study ◽  
Density Ratio ◽  
Concave Surface ◽  
Blockage Ratio ◽  
Coverage Area ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle

Abstract Film-hole can be often blocked by thermal-barrier coatings (TBCs) spraying, resulting in the variations of aerodynamic and thermal performances of film cooling. In this study, a numerical study of the blockage effect on the film cooling effectiveness of inclined cylindrical-holes was carried out on a concave surface to simulate the airfoil pressure side. Three typical blowing ratios (BRs) of 0.5, 1.0, and 1.5 were chosen at an engine-similar density ratio (DR) of 2.0. Two common inclination angles of 30 deg and 45 deg were designed. The blockage ratios were adjusted from 0 to 20%. The results indicated the blockage could enhance the penetration of film cooling flow to the mainstream. Thus, the averaged effectiveness and coolant coverage area were reduced. Moreover, the pressure loss inside of the hole was increased. With the increase of BR, the decrement of film cooling effectiveness caused by blockage rapidly increased. At BR = 1.5, the decrement could be acquired up to 70% for a blockage ratio of 20%. The decrement of film cooling effectiveness caused by blockage was nearly nonsensitive to the injection angle; however, the larger angle could generate the higher increment of pressure loss caused by blockage. A new design method for the couple scheme of film cooling and TBC was proposed, i.e., increasing the inlet diameter according to the blockage ratio before TBC spraying. In comparison with the original unblocked-hole, the enlarged blocked-hole not only kept the nearly same area-averaged effectiveness but also reduced slightly the pressure loss inside of the hole. Unfortunately, application of enlarged blocked-hole at large BR could lead to a more obvious reduction of effectiveness near hole-exit, in comparison with the original common-hole.

Numerical Investigation of Coolant-to-Mainstream Scaling Parameters With Film Cooling on Pressure and Suction Side of a Gas Turbine Blade

2017 ◽  
Author(s):  
Lingyu Zeng ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Momentum Flux ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Most experiments of blade film cooling are conducted with density ratio lower than that of turbine conditions. In order to accurately model the performance of film cooling under a high density ratio, choosing an appropriate coolant to mainstream scaling parameter is necessary. The effect of density ratio on film cooling effectiveness on the surface of a gas turbine twisted blade is investigated from a numerical point of view. One row of film holes are arranged in the pressure side and two rows in the suction side. All the film holes are cylindrical holes with a pitch to diameter ratio P/d = 8.4. The inclined angle is 30°on the pressure side and 34° on the suction side. The steady solutions are obtained by solving Reynolds-Averaged-Navier-Stokes equations with a finite volume method. The SST turbulence model coupled with γ-θ transition model is applied for the present simulations. A film cooling experiment of a turbine vane was done to validate the turbulence model. Four different density ratios (DR) from 0.97 to 2.5 are studied. To independently vary the blowing ratio (M), momentum flux ratio (I) and velocity ratio (VR) of the coolant to the mainstream, seven conditions (M varying from 0.25 to 1.6 on the pressure side and from 0.25 to 1.4 on the suction side) are simulated for each density ratio. The results indicate that the adiabatic effectiveness increases with the increase of density ratio for a certain blowing ratio or a certain momentum flux ratio. Both on the pressure side and suction side, none of the three parameters listed above can serve as a scaling parameter independent of density ratio in the full range. The velocity ratio provides a relative better collapse of the adiabatic effectiveness than M and I for larger VRs. A new parameter describing the performance of film cooling is introduced. The new parameter is found to be scaled with VR for nearly the whole range.

Turbine Platform Cooling and Blade Suction Surface Phantom Cooling From Simulated Swirl Purge Flow

2015 ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Jiyeon Lee ◽  
Je-Chin Han ◽  
Luzeng Zhang ◽  
Hee-Koo Moon
Keyword(s):  
Relative Motion ◽  
Film Cooling ◽  
Density Ratio ◽  
Flow Rate Ratio ◽  
Swirl Ratio ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Coolant Injection ◽  
Purge Flow ◽  
Cylindrical Holes

The paper presents the swirl purge flow on platform and a modeled land-based turbine rotor blade suction surface. Pressure sensitive paint (PSP) mass transfer technique provides detailed film cooling effectiveness distribution on platform and phantom cooling effectiveness on blade suction surface. Experiments have completed in a low speed wind tunnel facility with a five blade linear cascade. The inlet Reynolds number based on the chord length is 250,000. Swirl purge flow is simulated by coolant injection through fifty inclined cylindrical holes ahead of the blade leading edge row. Coolant injections from cylindrical holes go through nozzle endwall and a dolphin nose axisymmetric contour before reach platform and blade suction surface. Different “coolant injection angles” and “coolant injection velocity to cascade inlet velocity” results in various swirl ratios to simulate real engine conditions. Simulated swirl purge flow uses coolant injection angles of 30, 45, and 60 degrees to produce swirl ratios of 0.4, 0.6, and 0.8, respectively. Traditional purge flow has coolant injection angle of 90 degree to generate swirl ratio of 1. Coolant to mainstream mass flow rate ratio (MFR) is 0.5%, 1.0% and 1.5% for all swirl ratios. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Most of the swirl purge and purge coolant approaches platform, but small amount of the coolant migrates to blade suction surface. Swirl ratio of 0.4 has highest relative motion between rotor and coolant and severely decreases film cooling and phantom cooling effectiveness. Higher MFR of 1% and 1.5% cases suffer from apparent decrement of the effectiveness while increasing relative motion.

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