Deposition Near Film Cooling Holes on a High Pressure Turbine Vane

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
Weiguo Ai ◽  
Nathan Murray ◽  
Thomas H. Fletcher ◽  
Spencer Harding ◽  
Scott Lewis ◽  
...  
Keyword(s):  
Film Cooling ◽  
Elevated Temperatures ◽  
Coal Ash ◽  
Test Facility ◽  
Barrier Coating ◽  
Turbine Vane ◽  
Turbine Component ◽  
Shaped Hole ◽  
Density Ratios ◽  
Film Cooling Holes

Deposition on film-cooled turbine components was studied in an accelerated test facility. The accelerated deposition facility seeds a natural-gas burning combustor with finely ground coal ash particulate at 1180°C and 180 m/s (M=0.25). Both cylindrical and shaped holes, with and without thermal barrier coating, were studied over a range of blowing ratios from 0.5 to 4.0. Coolant density ratios were maintained at values from 2.1 to 2.4. Deposition patterns generated with the cylindrical film cooling holes indicated regions of low deposition in the path of the coolant with heightened deposition between film holes. This distinctive pattern was more accentuated at higher blowing ratios. Optical temperature measurements of the turbine component surface during deposition showed elevated temperatures between coolant paths. This temperature nonuniformity became more accentuated as deposition increased, highlighting a mechanism for deposition growth that has been documented on in-service turbines as well. The shaped-hole components exhibited little or no deposition in the region just downstream of the holes due to the distributed coolant film. Close cylindrical hole spacing of 2.25d displayed similar behavior to the shaped-hole configuration.

Deposition Near Film Cooling Holes on a High Pressure Turbine Vane

2008 ◽  
Author(s):  
Weiguo Ai ◽  
Nathan Murray ◽  
Thomas H. Fletcher ◽  
Spencer Harding ◽  
Scott Lewis ◽  
...  
Keyword(s):  
Film Cooling ◽  
Elevated Temperatures ◽  
Coal Ash ◽  
Test Facility ◽  
Turbine Vane ◽  
Hole Configuration ◽  
Turbine Component ◽  
Shaped Hole ◽  
Density Ratios ◽  
Film Cooling Holes

Deposition on film-cooled turbine components was studied in an accelerated test facility. The accelerated deposition facility seeds a natural-gas burning combustor with finely-ground coal ash particulate at 1180°C and 180 m/s (M = 0.25). Both cylindrical and shaped holes, with and without TBC coating, were studied over a range of blowing ratios from 0.5 to 4.0. Coolant density ratios were maintained at values from 2.1 to 2.4. Deposition patterns generated with the cylindrical film cooling holes indicated regions of low deposition in the path of the coolant, with heightened deposition between film holes. This distinctive pattern was more accentuated at higher blowing ratios. Optical temperature measurements of the turbine component surface during deposition showed elevated temperatures between coolant paths. This temperature non-uniformity became more accentuated as deposition increased, highlighting a mechanism for deposition growth that has been documented on in-service turbines as well. The shaped-hole components exhibited little or no deposition in the region just downstream of the holes, due to the distributed coolant film. Close cylindrical hole spacing of 2.25d displayed similar behavior to the shaped hole configuration.

Adiabatic Effectiveness Measurements for a Baseline Shaped Film Cooling Hole

2014 ◽  
Author(s):  
Robert P. Schroeder ◽  
Karen A. Thole
Keyword(s):  
Film Cooling ◽  
Freestream Turbulence ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Cylindrical Holes ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Density Ratios ◽  
Film Cooling Holes

Film cooling on airfoils is a crucial cooling method as the gas turbine industry seeks higher turbine inlet temperatures. Shaped film cooling holes are widely used in many designs given the improved performance over that of cylindrical holes. Although there have been numerous studies of shaped holes, there is no established baseline shaped hole to which new cooling hole designs can be compared. The goal of this study is to offer the community a shaped hole design, representative of proprietary and open literature holes that serves as a baseline for comparison purposes. The baseline shaped cooling hole design includes the following features: hole inclination angle of 30° with a 7° expansion in the forward and lateral directions; hole length of 6 diameters; hole exit-to-inlet area ratio of 2.5; and lateral hole spacing of 6 diameters. Adiabatic effectiveness was measured with this new shaped hole and was found to peak near a blowing ratio of 1.5 at density ratios of 1.2 and 1.5 as well as at both low and moderate freestream turbulence of 5%. Reductions in area-averaged effectiveness due to freestream turbulence at low blowing ratios were as high as 10%.

Introducing a New Test Rig for Film Cooling Measurements With Realistic Hole Inflow Conditions

2017 ◽  
Author(s):  
Marc Fraas ◽  
Tobias Glasenapp ◽  
Achmed Schulz ◽  
Hans-Jörg Bauer
Keyword(s):  
Film Cooling ◽  
Gas Flow ◽  
Test Rig ◽  
Hot Gas ◽  
Cooling Hole ◽  
Measurement Principle ◽  
Density Ratios ◽  
Operational Range ◽  
Film Cooling Holes

Further improvements in film cooling require an in-depth understanding of the influencing parameters. Therefore, a new test rig has been designed and commissioned for the assessment of novel film cooling holes under realistic conditions. The test rig is designed for generic film cooling studies. External hot gas flow as well as internal coolant passage flow are simulated by two individual flow channels connected to each other by the cooling holes. Based on a similarity analysis, the geometry of the test rig is scaled up by a factor of about 20. It furthermore offers the possibility to conduct experiments at high density ratios and realistic approach flow conditions at both cooling hole exit and inlet. The operational range of the new test rig is presented and compared to real engine conditions. It is shown that the important parameters are met and the transfer-ability of the results is ensured. Special effort is put onto the uniformity of the approaching hot gas flow, which will be demonstrated by temperature and velocity profiles. A first measurement of the heat transfer coefficient without film cooling is used to demonstrate the quality of the measurement principle.

Shock Wave - Film Cooling Interactions in Transonic Flows

2001 ◽  
Author(s):  
P. M. Ligrani ◽  
C. Saumweber ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Cross Flow ◽  
Injection Velocity ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Density Ratios ◽  
Film Cooling Holes

Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.

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.

scholarly journals Study on the Turbine Vane and Blade for a 1500°C Class Industrial Gas Turbine

1993 ◽  
Author(s):  
S. Amagasa ◽  
K. Shimomura ◽  
M. Kadowaki ◽  
K. Takeishi ◽  
H. Kawai ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
High Efficiency ◽  
Development Program ◽  
Test Results ◽  
Directionally Solidified ◽  
Barrier Coating ◽  
Turbine Vane ◽  
Nickel Base ◽  
Industrial Gas Turbine

This paper describes the summary of a three year development program for the 1st stage stationary vane and rotating blade for the next generation, 1500°C Class, high efficiency gas turbine. In such a high temperature gas turbine, the 1st turbine vane and blade are the most important hot parts. Full coverage film cooling (FCFC) is adopted for the cooling scheme, and directionally solidified (DS) nickel base super-alloy and thermal barrier coating (TBC) will be used to prolong the creep and thermal fatigue life. The concept of the cooling configuration, fundamental cascade test results and material test results will be presented.

Experimental and Computational Investigation of Shaped Film Cooling Holes Designed to Minimize Inlet Separation

2020 ◽  
Author(s):  
Fraser B. Jones ◽  
Dale W. Fox ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Velocity Ratio ◽  
Tangential Velocity ◽  
Velocity Fields ◽  
Radius Of Curvature ◽  
Large Radius ◽  
Combustion Gases ◽  
Shaped Hole ◽  
Manufacturing Technologies ◽  
Film Cooling Holes

Abstract Film cooling is used to protect turbine components from the extreme temperatures by ejecting coolant through arrays of holes to create an air buffer from the hot combustion gases. Limitations in traditional machining meant film cooling holes universally have sharp inlets which create separation regions at the hole entrance. The present study uses experimental and computational data to show that these inlet separation are a major cause of performance variation in crossflow fed film cooling holes. Three hole designs were experimentally tested by independently varying the coolant velocity ratio (VR) and the coolant channel velocitty ratio (VRc) to isolate the effects of crossflow on hole performance. Leveraging additive manufacturing technologies, the addition of a 0.25D radius fillet to the inlet of a 7-7-7 shaped hole is shown to significantly improve diffuser usage and significantly reduce variation in performance with VRc. A second AM design used a very large radius of curvature inlet to reduce biasing caused by the inlet crossflow. Experiments showed that this “swept” hole design did minimize biasing of coolant flow to one side of the shaped hole and it significantly reduced variations due to varying VRc. RANS simulations at six VR and three VRc conditions were made for each geometry to better understand how the new geometries changed the velocity field within the hole. The sharp and rounded inlets were seen to have very similar tangential velocity fields and jet biasing. Both AM inlets created more uniform, slower velocity fields entering the diffuser. The results of this paper indicate large improvements in film cooling performance can be found by leveraging AM technology.

Application of a CFD-Based Film Cooling Model to a Gas Turbine Vane Cascade With Cylindrical and Shaped Hole Endwall Film Cooling

2011 ◽  
Author(s):  
Tilman auf dem Kampe ◽  
Stefan Vo¨lker
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Flows ◽  
Turbine Vane ◽  
Mesh Resolution ◽  
Shaped Hole ◽  
Node Count ◽  
Endwall Film Cooling ◽  
Cooling Model ◽  
Near Wall

This paper presents the application of a CFD-based film cooling model to a gas turbine vane cascade test rig. The experimental investigations feature aerodynamic and endwall film cooling measurements on a first stage gas turbine vane in a linear cascade. An extended version of a previously developed cylindrical hole film cooling model has been employed, which now includes modeling of shaped hole cooling flows. The computational domain extends approximately one axial chord length upstream of the leading edge and downstream of the trailing edge of the vane. Adjacent solid parts are included by means of a conjugate heat transfer analysis to account for conduction effects. A hybrid mesh with resolved boundary layers and high spatial mesh resolution in the near-wall region is being used. This meshing approach ensures that the near-wall mesh resolution requirements of the film cooling model are satisfied, while maintaining a manageable total node count. Results obtained using the film cooling model are compared to surface distributions of film cooling effectiveness from the experimental cascade. Due to the moderate node count (≈ 3.5 × 106), CFD calculations including film cooling flows can be performed at comparatively low computational cost. The film cooling model, which previously had been validated against flat plate measurement data and applied to single cooling hole configurations only, is therefore shown to be a viable tool for the thermal design of gas turbine components with film cooling.

Practical Slot Configurations for Turbine Film Cooling Applications

2009 ◽  
Author(s):  
Joshua E. Bruce-Black ◽  
Frederick T. Davidson ◽  
David G. Bogard ◽  
David R. Johns
Keyword(s):  
Film Cooling ◽  
Flow Characteristics ◽  
Suction Side ◽  
Operating Conditions ◽  
Coolant Flow ◽  
Film Cooling Effectiveness ◽  
Diffusion Technique ◽  
Turbine Vane ◽  
Structural Weakness ◽  
Turbine Component

Turbine component film cooling is most effective when using a continuous slot to introduce coolant to the surface. However, this is not practical due to the structural weakness that would be inherent with a continuous slot. In this study, several slot-like designs are investigated to establish the film cooling effectiveness. These slot configurations extended only a partial distance through the simulated turbine vane wall, and were fed with impinging cylindrical holes. The configurations were studied on the suction side of a scaled-up turbine vane. In this study varying slot widths, discrete and continuous slots, and diffusing the coolant flow within the slot prior to it being emitted onto the surface of the vane were investigated. Rows of discrete round and shaped holes were also tested for comparison with the slots. The study of varying slot geometries showed that decreasing the width of the slots led to a substantial increase in adiabatic effectiveness. An internal coolant diffusion technique showed promise by maintaining performance levels while potentially providing a design configuration that more readily meets structural demands in real world operating conditions. The coolant flow characteristics were also studied through the use of thermal profiles measurements. These thermal profiles showed significant mainstream ingestion on the top surface of the slot prior to the coolant emitting onto the surface of the vane.

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