scholarly journals Experimental Investigation of Disc Cavity Leakage Flow and Hub Endwall Contouring in a Linear Rotor Cascade

2011 ◽  
Author(s):  
Ryan D. Erickson ◽  
Terrence W. Simon ◽  
Luzeng Zhang ◽  
Hee-Koo Moon
Keyword(s):  
Thermal Performance ◽  
Film Cooling ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Temperature Fields ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Endwall Contouring

An experimental study is carried out in a stationary linear cascade which simulates a turbine rotor to compare the thermal performance of two new axisymmetric endwall contour geometries. Measurements of endwall adiabatic film cooling effectiveness and near-endwall passage temperature fields are made for this purpose. In addition to documenting endwall contouring effects, a range of disc cavity leakage flow rates is investigated. This information is meant to quantify, over the range tested, the benefits and penalties of introducing leakage flow into the passage using the designated endwall contouring. Special attention is paid to determine whether the endwall curvature has any effect on the interaction between mainstream and secondary flows within the passage. Results indicate improved thermal performance when strong endwall curvature exists near the blade leading edge. The strong curvature causes cavity leakage flow to remain closer to the endwall, thereby increasing cooling effectiveness.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

An Experimental Study of the Leakage Flow Effect on the Film Cooling Effectiveness of a Gas Turbine Shroud

2021 ◽  
pp. 1-18
Author(s):  
Gi Mun Kim ◽  
Soo In Lee ◽  
Jin Young Jeong ◽  
Jae Su Kwak ◽  
Seokbeom Kim ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Density Ratio ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
High Heat Transfer ◽  
Hole Configuration ◽  
Cooling Hole

Abstract In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and this results in a locally non-uniform and high heat transfer on the surfaces. The present study experimentally investigates the effects of leakage flow through the slot between the gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment. The experiment is carried out in a linear cascade with five blades. Instead of the vane, a row of rods at the location of the vane trailing edge is installed to consider the wake effect. The leakage flow is introduced through the slot between the vane and blade rows, and additional coolant air is injected from the cooling holes installed at the vane's outer zone. The effects of the slot geometry, cooling hole configuration, and blowing ratio on the film cooling effectiveness are experimentally investigated using the pressure sensitive paint (PSP) technique. CO2 gas and a mixture of SF6 and N2 (25%+75%) are used to simulate the leakage flow to the mainstream density ratios of 1.5 and 2.0, respectively. The results indicate that the area averaged film cooling effectiveness is affected more by the slot width than by the cooling hole configuration at the same injection conditions, and the lower density ratio cases show higher film cooling effectiveness than the higher density ratio case at the same cooling configuration.

Experimental Study of Leakage Flow on Flow Field and Film Cooling of High Pressure Turbine Blade Platform

2015 ◽  
Author(s):  
Kenichiro Takeishi ◽  
Yutaka Oda ◽  
Shintaro Kozono
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Temperature Fields ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
Film Cooling Effectiveness ◽  
Injection Angle

An experiment has been conducted to study stator/rotor disc cavity leakage flow on the platform of a highly loaded stationary linear blade cascade. The linear cascade consists of a scaled-up model of the high-pressure turbine blades of an E3 (Energy efficient engine) and leakage slot models installed under the platform. Experiments have been conducted to investigate the effect of the slot injection angle, leakage flow rates, distance between the leading edge of the blade and the slot, and spacing of the blades. The film-cooling effectiveness was measured by pressure sensitive paint (PSP), and the temperature fields and flow fields were investigated using laser-induced fluorescence (LIF) and particle image velocimetry (PIV), respectively. It was observed from the experiments that the leakage flow covered the surface of the blade platform when the distance between the leading edge and the slot was zero; however, with increasing distance, the horseshoe vortex dominates near the junction of the blade leading edge, and the leakage flow could not cover the region. It was also found that the leakage flow has an effect that promotes the formation of the horseshoe vortex for some experimental conditions.

Assessment of URANS and DES for Prediction of Leading Edge Film Cooling

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
Toshihiko Takahashi ◽  
Ken-ichi Funazaki ◽  
Hamidon Bin Salleh ◽  
Eiji Sakai ◽  
Kazunori Watanabe
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Turbulence Modeling ◽  
Turbine Blades ◽  
Leading Edge ◽  
Vortex Structures ◽  
Temperature Fields ◽  
Detached Eddy Simulation ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

This paper describes the assessment of CFD simulations for the film cooling on the blade leading edge with circular cooling holes in order to contribute durability assessment of the turbine blades. Unsteady RANS applying a k-ε-v2-f turbulence model and the Spalart and Allmaras turbulence model and detached-eddy simulation (DES) based on the Spalart and Allmaras turbulence model are addressed to solve thermal convection. The CFD calculations were conducted by simulating a semicircular model in the wind tunnel experiments. The DES and also the k-ε-v2-f model evaluate explicitly the unsteady fluctuation of local temperature by the vortex structures, so that the predicted film cooling effectiveness is comparatively in agreement with the measurements. On the other hand, the predicted temperature fields by the Spalart and Allmaras model are less diffusive than the DES and the k-ε-v2-f model. In the present turbulence modeling, the DES only predicts the penetration of main flow into the film cooling hole but the Spalart and Allmaras model is not able to evaluate the unsteadiness and the vortex structures clearly, and overpredict film cooling effectiveness on the partial surface.

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.

An Experimental Study on the Leakage Flow on the Film Cooling Effectiveness of a Gas Turbine Shroud

2020 ◽  
Author(s):  
Gi Mun Kim ◽  
Soo In Lee ◽  
Jin Young Jeong ◽  
Jae Su Kwak ◽  
Seokbeom Kim ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Density Ratio ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Hole Size ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
High Heat Transfer

Abstract In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and results in locally non-uniform and high heat transfer on the surfaces. In this study, the effects of leakage flow through the slot between gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment were experimentally investigated. The experiment was carried out in a linear cascade with five blades. Instead of vane, a row of rods at the location of the vane trailing edge was installed to consider the wake effect. The leakage flow was introduced through the slot between vane and blade rows and additional coolant air was injected from the cooling holes installed at the vane outer zone. The effects of the slot geometry, hole size, and blowing ratio on the film cooling effectiveness were experimentally investigated by using a pressure sensitive paint technique. CO2 gas and the mixture of SF6 and N2 (25%+75%) were used as leakage flow in order to simulate leakage flow to mainstream density ratios of 1.5 and 2.0, respectively. Results showed that the area averaged film cooling effectiveness was more affected by the slot width than the cooling hole size at the same blowing ratio, and the lower density ratio cases showed higher film cooling effectiveness than that of higher density ratio case at the same cooling configuration.

Effects of Endwall 3D Contouring on Film Cooling Effectiveness of Cylindrical Hole Injections at Different Locations on Vane Endwall

2018 ◽  
Author(s):  
Pingting Chen ◽  
Hongyu Gao ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Secondary Flows ◽  
Cylindrical Hole ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Endwall Contouring

With the development of gas turbine, the secondary flow loss in vane passage is getting higher. To reduce the strength of secondary flows within vane passage, endwall 3D contouring is an effective design. Endwall 3D contouring can lead to significant changes in the secondary flow vortices, which lead to changes on jet-to-secondary flow interaction and then changes on the film cooling effectiveness. Meanwhile, the geometry configuration of the contoured endwall, such as the rising and falling on the endwall, can also have an impact on film cooling performance. As a result, the film cooling performance on contoured endwall differs from that on flat endwall. Understanding the difference in film cooling characteristics on the contoured endwall and flat endwall may help to make better endwall contouring design and better endwall film cooling arrangement. The present experiment compares the film cooling effectiveness of cylindrical hole injections at different locations on 3D contoured endwall versus flat endwall in an NGV (nozzle guide vane) passage. The measurement is performed in a low speed wind tunnel with a F-class annular sector NGV cascade. The cylindrical hole injections are located as 4 different rows at −30% axial chord, 30% axial chord, 50% axial chord and 70% axial chord. Endwall pressure distribution is measured with pressure taps by pressure sensor while film cooling effectiveness is measured using PSP (Pressure Sensitive Paint). Two density ratios with 1.0 and 1.5 and several average blowing ratios are investigated. Effects of endwall contouring, density ratio and blowing ratio on film cooling effectiveness are obtained and the results are presented and explained in this investigation.

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.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Vol 139 (3) ◽  
Author(s):  
Andrew F Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure-sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5% and 0.75% to 1% for the purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Assessment of URANS and DES for Prediction of Leading Edge Film Cooling

2010 ◽  
Author(s):  
Toshihiko Takahashi ◽  
Ken-ich Funazaki ◽  
Hamidon Bin Salleh ◽  
Eiji Sakai ◽  
Kazunori Watanabe
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Turbine Blades ◽  
Leading Edge ◽  
Vortex Structures ◽  
Temperature Fields ◽  
Detached Eddy Simulation ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Circular Model

This paper describes assessment of CFD simulations for the film-cooling on the blade leading edge with circular cooling holes in order to contribute durability assessment of the turbine blades. Unsteady RANS (URANS) applying a k–ε–v2–f turbulence model and the Spallart and Allmaras turbulence model, and Detached-Eddy Simulation (DES) based on the Spallart and Allmaras turbulence model are addressed to solve thermal convection. The CFD calculations were conducted by simulating a semi-circular model in the wind tunnel experiments. The DES and also the k–ε–v2–f model evaluate explicitly unsteady fluctuation of local temperature by the vortex structures, so that the predicted film cooling effectiveness comparatively in agreement with measurements. On the other hand, the predicted temperature fields by the Spallart and Allmaras model are less diffusive than the DES and the k–ε–v2–f model. In the present turbulence modeling, the DES only predicts penetration of main flow into the film cooling hole, but the Spallart and Allmaras model is not able to evaluate the unsteadiness and the vortex structures clearly, and over-predict film cooling effectiveness on the partial surface.

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