Showerhead film cooling injection orientation design on the turbine vane leading edge considering representative lean burn combustor outflow

2021 ◽  
pp. 095441002199656
Author(s):  
Zhuang Wu ◽  
Hui-ren Zhu ◽  
Cun-liang Liu ◽  
Lin Li ◽  
Xu-yang Liu
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Pressure Gradient ◽  
Film Cooling ◽  
Design Principle ◽  
Leading Edge ◽  
Radial Pressure ◽  
Mass Ratio ◽  
Lean Burn ◽  
Radial Pressure Gradient

The heat transfer performance of showerhead film cooling on the vane leading edge was numerically investigated considering representative lean burn combustor swirling outflow. Three cases with different inflow conditions (uniform inflow, positive swirling inflow, and negative swirling inflow) and three cases with different film injection angles (45°, 90°, and 135°) were studied. As the first study to explore the showerhead film design principle under swirling inflow, a newly designed asymmetrical counter-inclined (45° and 135°) film cooling was also proposed. To examine the design principles, the cooling effectiveness, heat transfer augmentation, and heat flux reduction of the newly designed asymmetrical case were evaluated compared with the traditional symmetrical case. The results show that the swirling inflow introduces obvious radial pressure gradient on the vane. The radial pressure gradient is the key influence factor to deflect the coolant migration, decrease the cooling effect, and degrade the homogeneity. The film with opposite orientation to the radial pressure gradient can weaken the deflect effect. The radial pressure gradient direction differs in different regions, making it impossible for the film with congruent injection orientation to simultaneously resist the pressure gradient on the entire vane. For the new design, the boundary line of the counter-inclined holes is consistent with the twisted stagnation line to guarantee that the injection orientation of all the film holes is opposite to the radial pressure gradient. As expected, the new design can effectively weaken the deflection effect and show uniform film distribution. The higher coolant mass ratio provides more obvious enhancement effect. At coolant mass ratio 3.71% and 4.56%, the overall area-averaged heat flux reduction (Δ q) is increased by 0.311 and 0.576, and the overall area-averaged relative standard deviation is reduced by 12.17 and 11.66 compared with the traditional design. The results have confirmed the adaptability of the film design principle under swirling inflow.

Net Heat Flux Reduction and Overall Effectiveness for a Turbine Blade Leading Edge

2005 ◽  
Author(s):  
Brian D. Mouzon ◽  
Elon J. Terrell ◽  
Jason E. Albert ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Cooling Performance ◽  
Net Heat Flux ◽  
Laboratory Models ◽  
Overall Effectiveness ◽  
Blade Leading Edge

The external cooling performance of a film cooled turbine airfoil can be quantified as a net reduction in heat transfer relative to the turbine airfoil without film cooling. This quantification is generally accomplished by using measurements of the adiabatic effectiveness and the change in heat transfer coefficients (hf/h0) for the film cooled surface to determine the net heat flux reduction (Δqr). Although measurement of Δqr for laboratory models give an indication of the ultimate film cooling performance, this does not show how much the surface temperature of the airfoil is reduced by film cooling. Measurement of scaled surface temperatures can be accomplished by using laboratory models constructed so that the Biot number is matched with that of the actual airfoil. These measurements provide a scaled temperature distribution on the airfoil that is referred to as the overall effectiveness, φ. For the current study, measurements of Δqr and φ have been made for a simulated turbine blade leading edge. The simulated leading edge incorporated shaped coolant holes, and had three rows of coolant holes. Improvements due to the shaped holes were determined by comparisons with previously measured round hole configurations. Spatially distributed hf/h0 show increases of 5% to 15% for M = 1.0 and 10% to 30% for M = 2.0. Results show that local variation in Δqr much greater than variation in φ, but laterally averaged Δqr distributions are reasonable predictors of the laterally averaged φ distributions.

Heat Transfer Analysis Over a Film Cooled Platform of a Vane Cascade With a Non-Uniform Inlet Flow

2017 ◽  
Author(s):  
G. Barigozzi ◽  
S. Ravelli ◽  
H. Abdeh ◽  
A. Perdichizzi ◽  
M. Henze ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Film Cooling ◽  
Cooling System ◽  
Convective Heat Transfer Coefficient ◽  
Leading Edge ◽  
Heat Transfer Analysis ◽  
Flow Distortion ◽  
Inlet Flow ◽  
Inlet Flow Distortion

This paper reports on heat transfer measurements performed on the film cooled platform of a linear nozzle vane cascade, subject to non-uniform inlet flow conditions. An obstruction, installed upstream of the cascade at different tangential positions, was responsible for inlet flow distortion. The platform cooling system included both purge flow from a slot located upstream of the leading edge and coolant ejection from a row of cylindrical holes located upstream of the slot. Testing was performed at inlet Mach number of Ma1 = 0.12 with both slot and combustor holes blowing at nominal conditions. Measured values of adiabatic film cooling effectiveness on the platform were used to obtain a detailed map of the convective heat transfer coefficient. The final goal was to compute the net heat flux reduction (NHFR), due to film cooling, when varying the relative position between obstruction and airfoil. Aligning the inflow non uniformity with the vane leading edge leads to a detrimental increase in the heat flux into the platform, within the vane passage. Conversely, positive NHFR values are observed over most of the platform surface if the inlet flow distortion is moved toward the suction side of the adjacent vane.

scholarly journals Highly Turbulent Mainstream Effects on Film Cooling of a Simulated Airfoil Leading Edge

1999 ◽  
Author(s):  
Christopher A. Johnston ◽  
David G. Bogard ◽  
Marcus A. McWaters
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Leading Edge ◽  
Relative Increase ◽  
Stagnation Line ◽  
Net Heat Flux ◽  
Adiabatic Effectiveness

The influence of a high mainstream turbulence was examined in an experimental study of film cooling on a simulated turbine blade leading edge. Detailed heat transfer coefficient and adiabatic effectiveness values were measured under conditions representative of actual environments in a gas turbine engine. The two parameters were also combined for a net heat flux reduction analysis. Turbulence levels of Tu = 17% were achieved by modifying a cross-jets turbulence generator with a large cylinder element. A quarter cylinder geometry was used to simulate the turbine blade leading edge. Two staggered rows of nine holes each were incorporated with a geometry consistent with current industry design practices. One row was positioned nominally on the stagnation line, x/d = 0, while the other was located 25° from the stagnation line. The holes were spaced at S/d = 7.64 with a shallow injection angle of 20° and oriented at 90° to the streamwise direction. Comparisons were made to previous studies of heat transfer rates and adiabatic effectiveness values under low turbulence (Tu < 0.5%) conditions. Adiabatic effectiveness was generally decreased by about 20% due to the high mainstream turbulence, although a much greater decrease occurred at the stagnation line at lower blowing rates. The relative increase in heat transfer coefficient due the coolant injection was found to be significantly smaller for the high mainstream turbulence case compared to the low mainstream turbulence case. This was particularly important when evaluating the overall performance of this film cooling hole configuration, since the much smaller relative increase in heat transfer coefficient resulted in good performance in terms of net heat flux reduction.

scholarly journals Effect of Coolant Injection on Heat Transfer for a Simulated Turbine Airfoil Leading Edge

1998 ◽  
Author(s):  
Ushio M. Yuki ◽  
David G. Bogard ◽  
J. Michael Cutbirth
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Film Cooling ◽  
Leading Edge ◽  
Streamwise Vortex ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Net Heat Flux ◽  
Overall Performance ◽  
Adiabatic Effectiveness

This paper presents an experimental study of the heat transfer on the leading edge of a simulated film cooled turbine airfoil. Previous studies have shown that use of film cooling on the leading edge of an airfoil can significantly increase the heat transfer coefficients around the leading edge which counter-acts the benefits of the adiabatic effectiveness provided by the coolant film. These heat transfer results complement our earlier study of the adiabatic effectiveness for this leading edge and film cooling hole geometry. Heat transfer and adiabatic effectiveness results were combined to determine the overall performance of the film cooling in terms of the net heat flux reduction. Heat transfer coefficients were found to be significantly increased by the film cooling flow in a narrow region which followed the path of the coolant flow. However, heat transfer coefficients were maximum to one side of the coolant jet, consistent with a streamwise vortex flow which is believed to be generated by the interaction of the mainstream with the coolant jet. Overall performance in terms of the net heat flux reduction was found to be unaffected by the large heat transfer coefficients in the vicinity of the holes, but was significantly diminished farther downstream.

Numerical Study on the Influence of Trench Width on Film Cooling Characteristics of Double-Wave Trench

2017 ◽  
Author(s):  
Bo-lun Zhang ◽  
Li Zhang ◽  
Hui-ren Zhu ◽  
Jian-sheng Wei ◽  
Zhong-yi Fu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Trench Width

Film cooling performance of the double-wave trench was numerically studied to improve the film cooling characteristics. Double-wave trench was formed by changing the leading edge and trailing edge of transverse trench into cosine wave. The film cooling characteristics of transverse trench and double-wave trench were numerically studied using Reynolds Averaged Navier Stokes (RANS) simulations with realizable k-ε turbulence model and enhanced wall treatment. The film cooling effectiveness and heat transfer coefficient of double-wave trench at different trench width (W = 0.8D, 1.4D, 2.1D) conditions are investigated, and the distribution of temperature field and flow field were analyzed. The results show that double-wave trench effectively improves the film cooling effectiveness and the uniformity of jet at the downstream wall of the trench. The span-wise averaged film cooling effectiveness of the double-wave trench model increases 20–63% comparing with that of the transverse trench at high blowing ratio. The anti-counter-rotating vortices which can press the film on near-wall are formed at the downstream wall of the double-wave trench. With the double-wave trench width decreasing, the film cooling effectiveness gradually reduces at the hole center-line region of the downstream trench. With the increase of the blowing ratio, the span-wise averaged heat transfer coefficient increases. The span-wise averaged heat transfer coefficient of the double-wave trench with 0.8D and 2.1D trench width is higher than that of the double-wave trench with 1.4D trench width at the high blowing ratio conditions.

scholarly journals Heat Transfer on a Film-Cooled Rotating Blade

1999 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox’s k-ω model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and ω distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

scholarly journals Highly Resolved Distribution of Heat Transfer for Turbine Leading Edge Film Cooling Including Reynolds Number and Blowing Rate Effects

1998 ◽  
Author(s):  
K. Jung ◽  
D. K. Hennecke
Keyword(s):  
Heat Transfer ◽  
Mass Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Heat Mass Transfer ◽  
Leading Edge ◽  
Transfer Coefficients ◽  
Complex Flow ◽  
Long Distance ◽  
Naphthalene Sublimation Technique

The effect of leading edge film cooling on heat transfer was experimentally investigated using the naphthalene sublimation technique. The experiments were performed on a symmetrical model of the leading edge suction side region of a high pressure turbine blade with one row of film cooling holes on each side. Two different lateral inclinations of the injection holes were studied: 0° and 45°. In order to build a data base for the validation and improvement of numerical computations, highly resolved distributions of the heat/mass transfer coefficients were measured. Reynolds numbers (based on hole diameter) were varied from 4000 to 8000 and blowing rate from 0.0 to 1.5. For better interpretation, the results were compared with injection-flow visualizations. Increasing the blowing rate causes more interaction between the jets and the mainstream, which creates higher jet turbulence at the exit of the holes resulting in a higher relative heat transfer. This increase remains constant over quite a long distance dependent on the Reynolds number. Increasing the Reynolds number keeps the jets closer to the wall resulting in higher relative heat transfer. The highly resolved heat/mass transfer distribution shows the influence of the complex flow field in the near hole region on the heat transfer values along the surface.

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.

Clocking Effects of Inlet Non-Uniformities in a Fully Cooled High-Pressure Vane: A Conjugate Heat Transfer Analysis

2015 ◽  
Author(s):  
Duccio Griffini ◽  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Cooling System ◽  
Hot Spot ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Heat Transfer Analysis

A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations have been performed. A preliminary grid sensitivity analysis with uniform inlet flow has been used to quantify the effect of spatial discretization. Turbulence model has been assessed in comparison with available experimental data. The effects of the relative alignment between combustion chamber and high-pressure vanes are then investigated considering realistic inflow conditions in terms of hot spot and swirl. The inlet profiles used are derived from the EU-funded project TATEF2. Two different clocking positions are considered: the first one where hot spot and swirl core are aligned with passage and the second one where they are aligned with the leading edge. Comparisons between metal temperature distributions obtained from conjugate heat transfer simulations are performed evidencing the role of swirl in determining both the hot streak trajectory within the passage and the coolant redistribution. The leading edge aligned configuration is resulted to be the most problematic in terms of thermal load, leading to increased average and local vane temperature peaks on both suction side and pressure side with respect to the passage aligned case. A strong sensitivity of both injected coolant mass flow and heat removed by heat sink effect has also been highlighted for the showerhead cooling system.

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Imran Qureshi ◽  
Andy D. Smith ◽  
Thomas Povey
Keyword(s):  
Heat Transfer ◽  
Cooling System ◽  
Three Dimensional ◽  
Leading Edge ◽  
Good Deal ◽  
Vortex Center ◽  
Research Facility ◽  
Lean Burn ◽  
Numerical Predictions ◽  
The Impact

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design.

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