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.

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°.

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.

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.

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

2015 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Intensity Level ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without 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. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.

Aerothermal Performance of a Nozzle Vane Cascade With a Generic Nonuniform Inlet Flow Condition—Part II: Influence of Purge and Film Cooling Injection

2017 ◽  
Vol 139 (10) ◽  
Author(s):  
G. Barigozzi ◽  
H. Abdeh ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Film Cooling ◽  
Leading Edge ◽  
Axial Position ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Inlet Flow ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
Flow Nonuniformity

In the present paper, the influence of the presence of an inlet flow nonuniformity 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 nonuniformity. 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. Aerothermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Results show a relevant negative impact of inlet flow nonuniformity on the cooled cascade aerodynamic and thermal performance. Higher film cooling effectiveness and lower aerodynamic losses are obtained when the inlet flow nonuniformity is located at midpitch, while lower effectiveness and higher losses are obtained when it is aligned to the vane leading edge. Moving the nonuniformity axially or changing its blockage only marginally influences the platform thermal protection.

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.

scholarly journals Unsteady pressure measurement and numerical simulations in an end-wall region of a linear blade cascade

2021 ◽  
Vol 3 (8) ◽  
Author(s):  
Erik Flídr ◽  
Petr Straka ◽  
Milan Kladrubský ◽  
Tomáš Jelínek
Keyword(s):  
Numerical Simulations ◽  
Pressure Measurement ◽  
Power Spectra ◽  
Surface Flow ◽  
Wall Region ◽  
Suction Surface ◽  
Flow Angle ◽  
Unsteady Pressure ◽  
Flow Visualizations ◽  
End Wall

AbstractThis contribution describes experimental and numerical research of an unsteady behaviour of a flow in an end-wall region of a linear nozzle cascade. Effects of compressibility ($$M_\mathrm {2,is}$$ M 2 , is ) and inlet flow angle ($$\alpha _1$$ α 1 ) were investigated. Reynolds number ($$Re_\mathrm {2,is}$$ R e 2 , is $$=8.5\times 10^5$$ = 8.5 × 10 5 ) was held constant for all tested cases. Unsteady pressure measurement was performed at the blade mid-span in the identical position $${\mathfrak {s}}$$ s to obtain reference data. Surface flow visualizations were performed as well as the steady pressure measurement to support conclusions obtained from the unsteady measurements. Comparison of the surface Mach number distributions obtained from the experiments and from the numerical simulations are presented. Flow visualizations are then compared with calculated limiting streamlines on the blade suction surface. It was shown, that the flow structures in the end-wall region were not affected by the primary flow at the blade mid-span, even when the shock wave formed. This conclusion was made from the experimental, numerical, steady as well as unsteady points of view. Three significant frequencies in the power spectra suggested that there was a periodical interaction between the vortex structures in the end-wall region. Based on the data analyses, anisotropic turbulence was observed in the cascade.

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.

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.

Sign in / Sign up

Export Citation Format

Share Document