Uncertainty Quantification of Aero-Thermal Performance of a Blade Endwall Considering Slot Geometry Deviation and Mainstream Fluctuation

2021 ◽  
pp. 1-48
Author(s):  
Zhi Tao ◽  
Zhendong Guo ◽  
Liming Song ◽  
Jun Li
Keyword(s):  
Uncertainty Quantification ◽  
Thermal Performance ◽  
Gas Turbines ◽  
Incidence Angle ◽  
Uncertain Parameters ◽  
Inlet Flow ◽  
Flow Angle ◽  
Actual Performance ◽  
Film Cooling Effectiveness ◽  
Input Uncertainties

Abstract With the ever-increasing aerodynamic and thermal loads, the endwalls of modern gas turbines have become critical areas that are susceptible to manufacturing and operational uncertainties, making them highly prone to thermal failures. Therefore, it is of vital importance to quantify the impacts of input uncertainties on the aero-thermal performance of endwalls. Firstly, based on the Kriging surrogate, an efficient uncertainty quantification (UQ) method suitable for expensive CFD problems is proposed. Using this method, the impacts of slot geometry deviations (slot width, endwall misalignment) and mainstream condition fluctuations (turbulence intensity, inlet flow angle) on the aero-thermal performance of endwalls are quantified. Results show that the actual performance of endwalls has a high probability of deviating from its nominal value. The maximum deviations of aerodynamic loss, area-averaged film cooling effectiveness, and area-averaged Nusselt number reach 0.33%, 45%, and 5.0%, respectively. The critical regions that are most sensitive to the input uncertainties are also identified. Secondly, a global sensitivity analysis method is also performed to pick out the significant uncertain parameters and explore the relationship between input uncertainties and performance output. The inlet flow angle is proved to be the most significant parameter among the four input uncertain parameters. Besides, a positive incidence angle is found to be detrimental to both the aerodynamic performance and the thermal management of endwalls. At last, the influence mechanisms of the inlet flow angle on endwall aero-thermal performance are clarified by a fundamental analysis of flow and thermal fields.

Uncertainty Quantification of Aero-Thermal Performance of a Blade Endwall Considering Slot Geometry Deviation and Mainstream Fluctuation

2020 ◽  
Author(s):  
Zhi Tao ◽  
Zhendong Guo ◽  
Liming Song ◽  
Jun Li
Keyword(s):  
Sensitivity Analysis ◽  
Nusselt Number ◽  
Uncertainty Quantification ◽  
Thermal Performance ◽  
Film Cooling ◽  
Gas Turbines ◽  
Cooling Effectiveness ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness

Abstract With the continuous increase of aerodynamic and thermal load, the endwall of modern gas turbines has become the critical region affected by the uncertainties in the manufacturing and operation process and thus is very likely to suffer performance degradation and thermal failure. Therefore, it is critical to understand and quantify the impacts of uncertainty factors on endwall aero-thermal performance. Based on Kriging surrogate, the frameworks of uncertainty quantification and global sensitivity analysis are constructed for a gas turbine blade endwall. The impacts of slot geometry deviations (slot width, endwall misalignment) and mainstream condition fluctuations (turbulence intensity, inlet flow angle) on endwall aero-thermal performance are quantified and analyzed. Results show that the actual performance of the endwall has a high probability of deviating from its nominal value. With respect to the nominal values, the maximum deviations of aerodynamic losses, averaged film cooling effectiveness and averaged Nusselt number reach up to 0.33%, 45% and 5.0%, respectively. The critical regions which are most sensitive to the input uncertainty parameters are identified. Furthermore, the inlet flow angle is proved to be the most significant parameter affecting the endwall aero-thermal performance through sensitivity analysis. The influence mechanisms of the inlet flow angle on endwall aero-thermal performance are clarified by detailed flow and thermal field analysis. Results show that the inlet flow angle significantly alters the size and strength of the secondary flow structures, resulting in a large variation of endwall aero-thermal performance. Quantitatively, a positive incidence angle of 2 degrees can lead to a 0.1% reduction of total pressure coefficient, a 12% reduction of averaged film cooling effectiveness and a 2% enhancement of averaged Nusselt number.

Inlet Flow Incidence Effects on the Thermal Performance of Showerhead Cooling in a Nozzle Vane Cascade

2019 ◽  
Author(s):  
H. Abdeh ◽  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Thermal Protection ◽  
Cooling System ◽  
Incidence Angle ◽  
Leading Edge ◽  
Intensity Level ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

Abstract In the present paper, the influence of inlet flow incidence on the thermal performance of a film cooled linear nozzle vane cascade is assessed. Tests have been carried out on a cooled cascade, featuring a showerhead cooling system made of 4 rows of cylindrical holes. The cascade was tested by varying the inlet flow angle in the range ±20° at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12. The thermal characterization of vane leading edge region was obtained through adiabatic film cooling effectiveness measurements. Vane load distributions supported the discussion of the results. Varying the incidence angle in either positive or negative angles, the thermal protection on the vane is reduced while the maximum protection happened at 0° incidence case.

scholarly journals Experimentally Measured Effects of Incidence Angle on the Adiabatic and Overall Effectiveness of a Fully Cooled Turbine Airfoil With Shaped Showerhead Holes

2017 ◽  
Vol 139 (9) ◽  
Author(s):  
Kyle Chavez ◽  
Thomas N. Slavens ◽  
David Bogard
Keyword(s):  
Incidence Angle ◽  
Density Ratio ◽  
Leading Edge ◽  
Inlet Flow ◽  
Flow Angle ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Inlet Conditions ◽  
Overall Effectiveness

Manufacturing and assembly variation can lead to shifts in the inlet flow incidence angles of a rotating turbine airfoil row. Understanding the sensitivity of the adiabatic film cooling effectiveness to a range of inlet conditions is necessary to verify the robustness of a cooling design. In order to investigate the effects of inlet flow incidence angles, adiabatic and overall effectiveness data were measured in a low speed linear cascade at 0 deg and 10 deg of the designed operating condition. Tests were completed at an inlet Reynolds number of Re = 120,000 and a turbulence intensity of Tu = 5% at the leading edge of the test article. Particle image velocimetry was used to verify the incident flow angle for each angle studied. The test section was first adjusted so that the pressure distribution and stagnation line of the airfoil matched those predicted by an aerodynamic computational fluid dynamics (CFD) model. IR thermography was then used to measure the adiabatic effectiveness levels of the fully cooled airfoil model with nine rows of shaped holes of varying construction and feed delivery. Measurements were taken over a range of blowing ratios and at a density ratio of DR = 1.23. This process was repeated for the two incidence angles measured, while the inlet pressure to the airfoil model was held constant for these incidence angle changes. Differences in laterally adiabatic effectiveness across the airfoil model were most evident in the showerhead, with changes as large as 0.2. The effect persisted most strongly at s/D = ±35 downstream of the stagnation row of holes, but was visible over the whole viewable area of 160 s/D. The effect was due to the stagnation line affecting the film at the showerhead row. Due to this effect, the showerhead was investigated in detail, including the effects of the stagnation line shift as well as the influence of the incidence angle on the overall effectiveness of the showerhead region. It was found that the stagnation line has the tendency to dramatically increase the near-hole adiabatic effectiveness levels when positioned within the breakout footprint of the hole. The effect persisted for the overall effectiveness study, since the hole spacing for this particular configuration was wide enough that the through hole convection was not completely dominant. This is the first study to present measured effectiveness values over both the pressure- and suction-side surfaces of a fully cooled airfoil for appreciably off-nominal incidence angles as well as examine adiabatic and overall effectiveness levels for a conical stagnation row of holes.

Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade

2018 ◽  
Author(s):  
H. Abdeh ◽  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Cooling System ◽  
Negative Impact ◽  
Surface Flow ◽  
Intensity Level ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
End Wall

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through 4 rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20°. The aero-thermal characterization of vane platform was obtained through 5-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20°.

Experimentally Measured Effects of Incidence Angle on the Adiabatic and Overall Effectiveness of a Fully Cooled Turbine Airfoil With Showerhead Shaped Holes

2016 ◽  
Author(s):  
Kyle Chavez ◽  
Thomas N. Slavens ◽  
David Bogard
Keyword(s):  
Incidence Angle ◽  
Density Ratio ◽  
Leading Edge ◽  
Inlet Flow ◽  
Flow Angle ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Inlet Conditions ◽  
Overall Effectiveness

Manufacturing and assembly variation can lead to shifts in the inlet flow incidence angles of a rotating turbine airfoil row. Understanding the sensitivity of the adiabatic film cooling effectiveness to a range of inlet conditions is necessary to verifying the robustness of a cooling design. In order to investigate the effects of inlet flow incidence angles, adiabatic and overall effectiveness data were measured in a low speed linear cascade at 0° and 10° of the designed operating condition. Tests were completed at an inlet Reynolds number of Re=120000 and a turbulence intensity of Tu = 5% at the leading edge of the test article. Particle Image Velocimetry was used to verify the incident flow angle for each angle studied. The test section was first adjusted so that the pressure distribution and stagnation line of the airfoil matched those predicted by an aerodynamic CFD model. IR thermography was then used to measure the adiabatic effectiveness levels of the fully-cooled airfoil model with nine rows of shaped holes of varying construction and feed delivery. Measurements were taken over a range of blowing ratios and at a density ratio of DR=1.23. This process was repeated for the two incidence angles measured, while the inlet pressure to the airfoil model was held constant for these incidence angle changes. Differences in laterally adiabatic effectiveness across the airfoil model were most evident in the showerhead, with changes as large as 0.2. The effect persisted most strongly at s/D=±35 downstream of the stagnation row of holes, but was visible over the whole viewable area of 160 s/D. The effect was due to the stagnation line affecting the film at the showerhead row. Due to this effect, the showerhead was investigated in detail, including effects of the stagnation line shift as well as the influence of the incidence angle on the overall effectiveness of the showerhead region. It was found that the stagnation line has the tendency to dramatically increase the near-hole adiabatic effectiveness levels when positioned within the breakout footprint of the hole. The effect persisted for the overall effectiveness study, since the hole spacing for this particular configuration was wide enough that the through hole convection was not completely dominant. This is the first study to present measured effectiveness values over both the pressure- and suction-side surfaces of a fully-cooled airfoil for appreciably off-nominal incidence angles as well as examine adiabatic and overall effectiveness levels for a conical stagnation row of holes.

Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade

2019 ◽  
Vol 141 (5) ◽  
Author(s):  
H. Abdeh ◽  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Cooling System ◽  
Negative Impact ◽  
Surface Flow ◽  
Intensity Level ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
End Wall

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through four rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20 deg. The aero-thermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20 deg.

Aero-Thermal Performance of a Nozzle Vane Cascade With a Generic Non Uniform Inlet Flow Condition: Part II — Influence of Purge and Film Cooling Injection

2016 ◽  
Author(s):  
G. Barigozzi ◽  
H. Abdeh ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Axial Position ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Inlet Flow ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, the influence of the presence of an inlet flow non uniformity on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out with platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. An obstruction was installed upstream of the cascade at variable tangential and axial position to generate a flow non uniformity. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition. Aero-thermal characterization of vane platform was obtained through 5-hole probe and end wall adiabatic film cooling effectiveness measurements. Results show a relevant negative impact of inlet flow non uniformity on the cooled cascade aerodynamic and thermal performance. Higher film cooling effectiveness and lower aerodynamic losses are obtained when the inlet flow non uniformity is located at mid pitch, while lower effectiveness and higher losses are obtained when it is aligned to the vane leading edge. Moving the non uniformity axially or changing its blockage only marginally influences the platform thermal protection.

Effect of Inlet Flow Angle on Gas Turbine Blade Tip Film Cooling

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Intensity Level ◽  
Pressure Side ◽  
Trailing Edge ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The influence of incidence angle on film-cooling effectiveness is studied for a cutback squealer blade tip. Three incidence angles are investigated −0 deg at design condition and ±5 deg at off-design conditions. Based on mass transfer analogy, the film-cooling effectiveness is measured with pressure sensitive paint techniques. The film-cooling effectiveness distribution on the pressure side near tip region, squealer cavity floor, and squealer rim tip is presented for the three incidence angles at varying blowing ratios. The average blowing ratio is controlled to be 0.5, 1.0, 1.5, and 2.0. One row of shaped holes is provided along the pressure side just below the tip; two rows of cylindrical film-cooling holes are arranged on the blade tip in such a way that one row is offset to the suction side profile and the other row is along the camber line. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade. Test is done in a five-blade linear cascade in a blow-down facility with a tip gap clearance of 1.5% of the blade span. The Mach number and turbulence intensity level at the cascade inlet were 0.23 and 9.7%, respectively. It is observed that the incidence angle affects the coolant jet direction on the pressure side near tip region and the blade tip. The film-cooling effectiveness distribution is also altered. The peak of laterally averaged effectiveness is shifted upstream or downstream depending on the off-design incidence angle. The film cooling effectiveness inside the tip cavity can increase by 25% with the positive incidence angle. However, in general, the overall area-averaged film-cooling effectiveness is not significantly changed by the incidence angles in the range of study.

Experimental Investigation of Gas Turbine Axial Diffuser Performance: Part II — Effect of Inlet Flow Profiles at On- and Off-Design Conditions

2021 ◽  
Author(s):  
Kenneth Brown ◽  
Stephen Guillot ◽  
Wing Ng ◽  
Lee Iksang ◽  
Kim Dongil ◽  
...  
Keyword(s):  
Experimental Investigation ◽  
Flow Visualization ◽  
Gas Turbines ◽  
Full Scale ◽  
Flow Conditions ◽  
Inlet Flow ◽  
Flow Angle ◽  
Companion Article ◽  
Pressure Distributions ◽  
Physical Insight

Abstract An experimental investigation of the effect of inlet flow conditions and improved geometries on the performance of modern axial exhaust diffusers of gas turbines has been completed. The first article in the two-part series [1] leveraged a scaled model to examine parametric variations in both diffuser geometry and inlet flow conditions with the latter having significant consequences for diffuser performance. This second article pivots on the conclusions of the companion article and offers findings and physical insight on diffuser performance for on- and off-design inlet flow conditions. Using a high-performing diffuser design from the companion article, an experimental investigation is carried out with tailored distributions of inlet Mach distribution, inlet swirl angle, and inlet radial flow angle which are designed to replicate conditions of an industry diffuser at various loads. Six different inlet distributions were investigated including a design condition and five other conditions which feature mass flows both greater than and less than the design condition. The measurements were taken at near full-scale turbine exit Reynolds number (ReH roughly 39% of the value for an H-class diffuser) and at full-scale turbine exit Mach number. The study was accomplished in a blow-down, cold-flow wind tunnel facility, and measurements included 5-hole probe traverses at planes of interest, axial pressure distributions, strut pressure distributions, and oil-flow visualization. Over the range of inlet conditions studied, pressure recovery at the exit varied by up to 68.5% from that of on-design operation. Tracking of performance coefficients along the axial direction suggested the existence of flow phenomena which were in some cases able to be confirmed with on-strut pressure measurements and flow visualization. In addition to physical insight, the results presented here offer an experimental benchmark for the sensitivity of diffuser performance to inlet flow conditions.

Effect of Incidence Angle on Gas Turbine First-Stage Nozzle Guide Vane Leading Edge and Gill Region Film Cooling

2012 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Inlet Flow ◽  
Flow Angle ◽  
Cooling Performance ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Compound Angle

The nonuniformity of the Hp turbine inlet flow field put forward higher requirements for NGV (Nozzle Guide Vanes) leading edge and gill region film cooling. The assumption of design condition in most of the experiments couldn’t reflect the true operation environment in the Hp turbine NGV. The factor of off-design condition was incorporated into the experiment in this research. The GE-E3 Hp turbine nozzle guide vanes were used in the experiment to investigate the cooling performance of injection from leading edge and gill region with inlet Reynolds numbers of Re = 3.5×105 and inlet Mach number of Ma = 0.1. The compound angle fan-shaped film cooling hole configuration was applied. The cooling characteristics at off-design condition were analyzed and compared in the paper. The leading edge and gill region film cooling performance was assessed with the incidence angle varying from i = −10deg to i = +10deg. The blowing ratio varying from M = 0.7 to M = 1.3, was also selected as an experimental variable. Film cooling effectiveness distribution was measured using PSP (Pressure Sensitive Paint) technique. The film cooling performance of the compound angle fan-shaped holes was assessed at both design and off-design conditions. The object of this research is to change the concept that NGV leading edge film cooling experiment only needs the data at design condition. Through the comparative analysis of experimental results at different inlet flow angle, the influence of off-design condition on NGV leading edge and gill region film cooling could be illustrated at a reasonable level.

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