Controlling Leakage Flows Over a Rotor Blade Tip Using Air-Curtain Injection: Part II — Rotor Casing Film Cooling

2019 ◽  
Author(s):  
Qiang Zhao ◽  
Xing Yang ◽  
Zhao Liu ◽  
Zhenping Feng ◽  
Terrence W. Simon
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Air Curtain ◽  
Film Cooling Effectiveness ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Blade Tip ◽  
Tip Injection ◽  
Tip Leakage Flows

Abstract In modern gas turbine engines, the rotor casing region experiences high thermal loads due to complex flow structures and aerothermal effects. Thus, casing cooling is one of essential measures to ensure turbine service lifetime and performance. However, studies on heat transfer and cooling over the rotor casing with tip leakage flows are limited in the open literature during the past decades. The present work aims at controlling leakage flows over the blade tip and decreasing heat loads on the rotor casing. A novel approach proposed in a companion paper (GT2019-90232) is adopted in this paper as Part II by introducing an air-curtain injection from the rotor casing through a pair of inclined rows of discrete holes positioned in the range of 30% and 50% axial chord downstream of the blade leading edge in the casing. This air-curtain injection approach is applied to flat and recessed tips with and without tip injection to evaluate its sealing capability on tip leakage flows and film cooling effectiveness on the casing for two injection ratios of 0.7% and 1.0%. In this paper, Reynolds-averaged Navier-Stokes (RANS) simulations with Shear Stress Transport (SST) k-ω turbulence model and γ-Reθ transition model, which are validated with relevant experimental data, are performed to investigate tip leakage flows and film cooling effectiveness on the casing in a single-stage, high-pressure gas turbine engine. Results show that casing injection can reduce tip leakage mass flow effectively by changing the development and migration of tip leakage mass flows, especially when the recessed tip is applied. Adding tip injection would further reduces the tip leakage. The casing injection also provides an excellent cooling effect on the casing across rotor middle chord through trailing edge regions. In the presence of the recessed tip, coolant spreads out well on the rotor tip and the casing surfaces, resulting in better film cooling effectiveness on the casing over rotor tip leading edge. In addition, the tip injection could provide an extra cooling effect in some other regions of the casing.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Heat Transfer Coefficient and Film Cooling Effectiveness on the Partial Cavity Tip of a Gas Turbine Blade

2018 ◽  
Author(s):  
Jin Young Jeong ◽  
Woobin Kim ◽  
Jae Su Kwak ◽  
Jung Shin Park
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cavity Shape ◽  
Blade Tip ◽  
The Heat Transfer Coefficient

Leakage flow between the rotating turbine blade tip and the fixed casing causes high heat loads and thermal stress on the tip and near the tip region. For this study, new squealer tips called partial cavity tips, which combine the advantages of plane and squealer tips, were suggested, and the effects of the cavity shape on the tip heat transfer coefficient and film cooling effectiveness were investigated experimentally in a low speed linear cascade. The suggested blade tips had a flat surface near the leading edge and a squealer cavity from the mid-chord to trailing edge region to achieve the advantages of both blade tip types. The heat transfer coefficient was measured via the 1-D transient heat transfer technique using an IR camera, and the film cooling effectiveness was obtained via the pressure sensitive paint (PSP) technique. Results showed that the heat transfer coefficient and film cooling effectiveness on the partial cavity tips strongly depended on the cavity shape. Near the leading edge, the heat transfer coefficients for the partial cavity tip cases were lower than that for the squealer tip case. However, the heat transfer coefficient on the cavity surface was higher for the partial cavity tip cases. The D10 tip showed a similar distribution of film cooling effectiveness to that of the PLN tip near the leading edge and the DSS tip near the mid-chord region. However, the overall averaged film cooling effectiveness of the DSS tip was higher than that of the D10 tip.

Turbine Blade Tip Cooling With Blade Rotation: Part II — Shroud Coolant Injection

2015 ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Complementary Method ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip

In this paper, blade-tip cooling is investigated with coolant injection from the shroud alone and a combination of shroud coolant injection and tip cooling. The blade rotates at a nominal speed of 1200 RPM, and consists of a cut back squealer tip with a tip clearance of 1.7% of the blade span. The blade consists of tip holes and pressure side shaped holes, while the shroud has an array of angled holes and a circumferential slot upstream of the rotor section. Different combinations of the three cooling configurations are utilized to study the effectiveness of shroud cooling as a complementary method of cooling the blade tip. The measurements are done using liquid crystal thermography. Blowing ratios of 0.5, 1.0, 2.0, 3.0 and 4.0 are studied for shroud slot cooling and blowing ratios of 1.0, 2.0, 3.0, 4.0 and 5.0 are studied for shroud hole cooling. For cases with coolant injection from the tip, the blowing ratios used are 1.0, 2.0, 3.0 and 4.0. The results show an increase in film cooling effectiveness with increasing blowing ratio for shroud hole cooling. The increased effectiveness from shroud hole cooling is concentrated mainly in the tip-region below the shroud holes and towards the blade suction side and the suction side squealer rim. Slot cooling injection results in increased effectiveness on the blade tip near the blade leading edge up to a maximum blowing ratio, after which the cooling effectiveness decreases with increasing blowing ratio. The combination of the different cooling methods results in better overall cooling coverage of the blade tip with the shroud hole and blade tip cooling combination being the most effective. The level of coolant protection is strongly dependent on the blowing ratio and combination of blowing ratios.

Controlling Leakage Flows Over a Rotor Blade Tip Using Air-Curtain Injection: Part I — Aero-Thermal Performance of Rotor Tips

2019 ◽  
Author(s):  
Xing Yang ◽  
Qiang Zhao ◽  
Zhao Liu ◽  
Zhenping Feng ◽  
Terrence W. Simon
Keyword(s):  
Film Cooling ◽  
Leakage Flow ◽  
Efficiency Improvement ◽  
Tip Leakage Flow ◽  
Air Curtain ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Blade Tip ◽  
Cooling Air ◽  
Blade Tips

Abstract The rotor casing of gas turbine engines is generally cooled with cooling air from compressors and then the cooling air is discharged into the passage flow of the rotor. In this paper, a novel design both for the blade tip leakage flow control and for the rotor casing and tip cooling is proposed. Cooling air is injected through a pair of inclined rows of discrete holes positioned between 30% and 50% axial chord downstream of the blade leading edge in the casing. The casing injection forms as air-curtain within the blade tip gap, and inhibits the development of the tip leakage flows and provides secondary-order cooling for the rotor tip. Air injection from the rotor casing onto flat and recessed blade tips is investigated using numerical simulations that is validated by extensive aerodynamic and heat transfer experimental data. Flow and film cooling over the blade tip and turbine overall aerodynamic performance are examined in detail for two casing injection rates. Comparisons between flat tip without casing injection (baseline) case and the casing injection cases show that the air-curtain injection significantly alters the flow structures near the casing by modifying the development and migration of the tip leakage flow. The air-curtain injection over the flat and recessed tips both generates turbine stage overall aerodynamic efficiency improvement due to the sealing effects of the casing injection, but the efficiency gain depends on the competing results between the sealing effects and the “over-blown” effects of the air-curtain injection. Applying a recess to the blade tip is generally detrimental to the efficiency improvement by the air-curtain injection. In addition to efficiency improvement, secondary-order cooling effects from the casing injection are found to provide considerable thermal protection for the blade tips. However, increasing injection rate reduces the film cooling performance over the rotor tip surfaces. The recessed tip could present better film cooling effectiveness than the flat tip in the presence of the air-curtain.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2010 ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Heat Transfer Coefficient and Film Cooling Effectiveness on the Partial Cavity Tip of a Gas Turbine Blade

2019 ◽  
Vol 141 (7) ◽  
Author(s):  
Jin Young Jeong ◽  
Woobin Kim ◽  
Jae Su Kwak ◽  
Jung Shin Park
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cavity Shape ◽  
Blade Tip ◽  
The Heat Transfer Coefficient

Leakage flow between the rotating turbine blade tip and the fixed casing causes high heat loads and thermal stress on the tip and near the tip region. For this study, new squealer tips called partial cavity tips, which combine the advantages of plane and squealer tips, were suggested, and the effects of the cavity shape on the tip heat transfer coefficient and film cooling effectiveness were investigated experimentally in a low-speed linear cascade. The suggested blade tips had a flat surface near the leading edge and a squealer cavity from the mid-chord to trailing edge region to achieve the advantages of both blade tip types. The heat transfer coefficient was measured via the 1-D transient heat transfer technique using an IR camera, and the film cooling effectiveness was obtained via the pressure-sensitive paint (PSP) technique. Results showed that the heat transfer coefficient and film cooling effectiveness on the partial cavity tips strongly depended on the cavity shape. Near the leading edge, the heat transfer coefficients for the partial cavity tip cases were lower than that for the squealer tip case. However, the heat transfer coefficient on the cavity surface was higher for the partial cavity tip cases. The D10 tip showed a similar distribution of film cooling effectiveness to that of the plane (PLN) tip near the leading edge and the double side squealer (DSS) tip near the mid-chord region. However, the overall average film cooling effectiveness of the DSS tip was higher than that of the D10 tip.

Turbine Blade Tip Cooling With Blade Rotation—Part II: Shroud Coolant Injection

2016 ◽  
Vol 138 (9) ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Complementary Method ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip

In this paper, blade tip cooling is investigated with coolant injection from the shroud alone and a combination of shroud coolant injection and tip cooling. With a nominal rotation speed of 1200 rpm, each blade consists of a cut back squealer tip with a tip clearance of 1.7% of the blade span. The blades also consist of tip holes and pressure side (PS) shaped holes, while the shroud has an array of angled holes and a circumferential slot upstream of the rotor section. Different combinations of the three cooling configurations (tip and PS holes, shroud angled holes, and shroud circumferential slot) are utilized to study the effectiveness of coolant injected from the shroud as a complementary method of cooling the blade tip. The measurements are done using liquid crystal thermography. Blowing ratios of 0.5, 1.0, 2.0, 3.0, and 4.0 are studied for shroud slot cooling, and blowing ratios of 1.0, 2.0, 3.0, 4.0, and 5.0 are studied for shroud hole cooling. For cases with coolant injection from the blade tip, the blowing ratios used are 1.0, 2.0, 3.0, and 4.0. The results show an increase in film cooling effectiveness with increasing blowing ratio for shroud hole coolant injection. The increased effectiveness from shroud hole coolant is concentrated mainly in the tip region below the shroud holes and toward the blade suction side and the suction side squealer rim. Slot coolant injection results in increased effectiveness on the blade tip near the blade leading edge up to a maximum blowing ratio, after which the cooling effectiveness decreases with increasing blowing ratio. The combination of the different cooling methods results in better overall cooling coverage of the blade tip with the shroud hole and blade tip coolant combination being the most effective. The level of coolant protection is strongly dependent on the blowing ratio and combination of blowing ratios.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

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