Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Hans-Ju¨rgen Rehder
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Coolant Flow ◽  
Coolant Fluid ◽  
Pressure Side ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Number Range ◽  
Mixing Losses

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade: Aerodynamic Experiments and Loss Analysis

2009 ◽  
Author(s):  
Hans-Ju¨rgen Rehder
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Coolant Flow ◽  
Coolant Fluid ◽  
Pressure Side ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Number Range ◽  
Mixing Losses

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2%. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure and suction side boundary layer, trailing edge and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part II: Independent Influences of Vane Trailing Edge and Purge Cooling

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Flow Rate ◽  
Pressure Field ◽  
Its Region ◽  
Coolant Flow ◽  
Coolant Flow Rate ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Trailing Edge

The independent influences of vane trailing edge and purge cooling are studied in detail for a one-and-one-half stage transonic high-pressure turbine operating at design-corrected conditions. This paper builds on the conclusions of Part I, which investigated the combined influence of all cooling circuits. Heat-flux measurements for the airfoil, platform, tip, and root of the turbine blade, as well as the shroud and the vane side of the purge cavity, are used to track the influence of cooling flow. By independently varying the coolant flow rate through the vane trailing edge or purge circuit, the region of influence of each circuit can be isolated. Vane trailing edge cooling is found to create the largest reductions in blade heat transfer. However, much of the coolant accumulates on the blade suction surface and little influence is observed for the pressure surface. In contrast, the purge cooling is able to cause small reductions in heat transfer on both the suction and pressure surfaces of the airfoil. Its region of influence is limited to near the hub, but given that the purge coolant mass flow rate is 1/8 that of the vane trailing edge, it is impressive that any impact is observed at all. The cooling contributions of these two circuits account for nearly all of the cooling reductions observed for all three circuits in Part I, indicating that the vane inner cooling circuit that feeds most of the vane film-cooling holes has little impact on the downstream blade heat transfer. Time-accurate pressure measurements provide further insight into the complex interactions in the purge region that govern purge coolant injection. While the pressures supplying the purge coolant and the overall coolant flow rate remain fairly constant, the interactions of the vane pressure field and the rotor pressure field create moving regions of high pressure and low pressure at the exit of the cavity. This results in pulsing regions of injection and ingestion.

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 for the Blade of a Cooled Stage and One-Half High-Pressure Turbine: Part II—Independent Influences of Vane Trailing Edge and Purge Cooling

2010 ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Flow Rate ◽  
Pressure Field ◽  
Its Region ◽  
Coolant Flow ◽  
Coolant Flow Rate ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Trailing Edge

The independent influences of vane trailing edge and purge cooling are studied in detail for a one-and-one-half stage transonic high-pressure turbine operating at design corrected conditions. This paper builds on the conclusions of Part I, which investigated the combined influence of all cooling circuits. Heat-flux measurements for the airfoil, platform, tip, and root of the turbine blade as well as the shroud and the vane side of the purge cavity are used to track the influence of cooling flow. By independently varying the coolant flow rate through the vane trailing edge or purge circuit, the region of influence of each circuit can be isolated. Vane trailing edge cooling is found to create the largest reductions in blade heat transfer. However, much of the coolant accumulates on the blade suction surface and little influence is observed for the pressure surface. In contrast, the purge cooling is able to cause small reductions in heat transfer on both the suction and pressure surfaces of the airfoil. Its region of influence is limited to near the hub, but given that the purge coolant mass flow rate is 1/8th that of the vane trailing edge, it is impressive that any impact is observed at all. The cooling contributions of these two circuits account for nearly all of the cooling reductions observed for all three circuits in Part I, indicating that the vane inner cooling circuit that feeds most of the vane film-cooling holes has little impact on the downstream blade heat transfer. Time-accurate pressure measurements provide further insight into the complex interactions in the purge region that govern purge coolant injection. While the pressures supplying the purge coolant and the overall coolant flow rate remain fairly constant, the interactions of the vane pressure field and the rotor pressure field create moving regions of high pressure and low pressure at the exit of the cavity. This results in pulsing regions of injection and ingestion.

High-Fidelity Simulations of a High-Pressure Turbine Vane With End Walls: Impact of Secondary Structures and Spanwise Temperature Profiles on Losses

2021 ◽  
Author(s):  
Yaomin Zhao ◽  
Richard D. Sandberg
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Suction Side ◽  
High Fidelity ◽  
Temperature Profiles ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Vortical Structures ◽  
Wall Boundary ◽  
End Wall

Abstract In the present study, highly resolved large-eddy simulations of a VKI LS-89 high-pressure turbine (HPT) blade (T. Arts et al., 1990) with spanwise end-walls are performed at a Reynolds number of 0.57 million and an exit Mach number of 0.9. Two different spanwise temperature profiles, one uniform as baseline and one asymmetric profile extracted at the combustor exit from the public literature, are set at the inlet boundary. The high-fidelity data generated by the present cases are analyzed to investigate the end-wall secondary features, and show that the vortical structures that form near the leading edge of the vane include pressure-side and suction-side legs. While the end-wall vortical structures show no obvious effects on the time-averaged behavior of the pressure-side blade boundary layer, the suction-side structures induce counter rotating vortical structures and trigger rapid transition in the end-wall boundary layers. Furthermore, the cases with different inlet temperature profiles are directly compared, and the effects of the inflow on the aerothermal performance of the HPT vane are discussed. Based on the recently proposed entropy loss analysis (Zhao and Sandberg, GT2019-90126), we have been able to quantitatively show that the end-wall effects contribute significantly to the total loss of the turbine, with the main contributors being the extra viscous dissipation and turbulence production in the end-wall boundary layer.

Formation and Transport of Secondary Flows Caused by Vortex-Blade Interaction in a High Pressure Turbine

2015 ◽  
Author(s):  
Zuo-Jun Wei ◽  
Wei-Yang Qiao ◽  
Ping-Ping Chen ◽  
Jian Liu
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Gradient ◽  
Suction Side ◽  
Secondary Flows ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Blade Loading ◽  
Passage Vortex

As modern turbines are designed with low aspect ratio and high blade loading, secondary flow interactions become more important. In the present work, numerical simulation is performed in a two-stage high-pressure turbine with divergent meridional passage to investigate the transport and interaction of secondary vortex from the first stage rotor within the second stage’s stator. Scale-Adaptive Simulation model coupled with Shear Stress Transport model (SAS-SST turbulence model) is used to capture the flow structures caused by the interaction in the second stator. Coupled with the passage vortex of the first rotor, the shed vortex rotates opposite in the direction and has comparable strength. As both of these vortices convect downstream to the stator bladerow, each deforms into two legs on the pressure and suction sides in the passage. In the passage due to the cross pressure gradient by blade loading, all the low-momentum fluid contained in these vortices moves towards the suction side. Besides, with the existing static pressure gradient in radial direction and vortex dynamics, the suction-side leg and the pressure-side leg move in different radial directions. The suction side leg of incoming passage vortex moves towards the endwall along the suction surface and interacts with the developing passage vortex of the second stator. The incoming shed vortex moves towards the midspan and rolls up the boundary layer fluid from suction surface. Due to the interactions between the incoming shed vortices from the hub and casing and the boundary layer of second stator, two counter-rotating vortices are formed near the midspan. Additional high loss is found there at the outlet plane, which has a comparable magnitude to the endwall secondary loss. The pressure side leg of the incoming passage vortex remains in a certain span with that of the incoming shed vortex and is not engulfed by the developing passage vortex.

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.

Validation of MISES Two-Dimensional Boundary Layer Code for High-Pressure Turbine Aerodynamic Design

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Philip L. Andrew ◽  
Harika S. Kahveci
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Operating Cost ◽  
Design Tool ◽  
Operating Conditions ◽  
Aerodynamic Design ◽  
High Pressure Turbine ◽  
Reduce Operating Cost ◽  
Wide Range ◽  
Dimensional Boundary

Avoiding aerodynamic separation and excessive shock losses in gas turbine turbomachinery components can reduce fuel usage and thus reduce operating cost. In order to achieve this, blading designs should be made robust to a wide range of operating conditions. Consequently, a design tool is needed—one that can be executed quickly for each of many operating conditions and on each of several design sections, which will accurately capture loss, turning, and loading. This paper presents the validation of a boundary layer code, MISES, versus experimental data from a 2D linear cascade approximating the performance of a moderately loaded mid-pitch section from a modern aircraft high-pressure turbine. The validation versus measured loading, turning, and total pressure loss is presented for a range of exit Mach numbers from ≈0.5 to 1.2 and across a range of incidence from −10 deg to +14.5 deg relative to design incidence.

APPLICATION OF SCALE ADAPTIVE SIMULATION MODEL TO STUDYING COOLING CHARACTERISTICS OF A HIGH PRESSURE TURBINE BLADE CUTBACK TRAILING EDGE WITH DIFFERENT COOLING CONFIGURATIONS

2021 ◽  
pp. 1-53
Author(s):  
Yuefeng Li ◽  
Huazhao Xu ◽  
Jianhua Wang ◽  
Wei Song ◽  
Ming Wang ◽  
...  
Keyword(s):  
High Pressure ◽  
High Pressure Turbine ◽  
Baseline Model ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Adaptive Simulation ◽  
Rear Part ◽  
Pin Fins ◽  
A Minor ◽  
Better Than

Abstract This paper adopted Scale Adaptive Simulation (SAS) to investigate fluid flow and cooling characteristics in detail downstream of a high pressure turbine (HPT) blade trailing edge (TE) cutback region. The effects of typical TE configurations on cutback cooling performance are investigated including three types of internal turbulators, the cutback with/without land extensions and three kinds of ejection lip profiles. The elliptic pin fins with streamwise orientation significantly improve ηaw at the rear part of the cutback surface over the baseline model with cylindrical pin fins and slightly increase Cd. However, the elliptic pin fins with spanwise orientation drastically reduce the ηaw and Cd. Downstream of the cutback, the coherent structures are strongly disturbed and become chaotic compared to the TE with cylindrical and streamwise oriented elliptic pin fins. The application of land extensions only causes an evident change to the coherent structure immediate downstream of the lip, and slightly improves ηaw and reduces Cd over the baseline model on the rear part of the cutback surface. Rounded lip shapes B and C also show an obvious increase in ηaw on the rear part of the cutback surface but only a minor increase in Cd compared to the straight lip shape A. The rounded lip helps the coolant diffuse into the TE cutback and reduce the intensity of mixing. Due to larger rounding radius of shape B, the cooling effectiveness predicted by shape B is slightly better than shape C.

Modeling and Impact of High-pressure Turbine Blade Trailing Edge Film Cooling Hole Variations

2020 ◽  
Author(s):  
Jan Kamenik ◽  
David J. Toal ◽  
Andy Keane ◽  
Lars Högner ◽  
Marcus Meyer ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Hole
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