Experimental and Numerical Investigation of Optimized Blade Tip Shapes—Part I: Turbine Rainbow Rotor Testing and Numerical Methods

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
Bogdan C. Cernat ◽  
Marek Pátý ◽  
Cis De Maesschalck ◽  
Sergio Lavagnoli
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
High Performance ◽  
Numerical Data ◽  
Secondary Flows ◽  
Design Strategies ◽  
Flow Physics ◽  
Leakage Flows ◽  
Numerical Predictions ◽  
Blade Tip

Blade tip design and tip leakage flows are crucial aspects for the development of modern aero-engines. The inevitable clearance between stationary and rotating parts in turbine stages generates high-enthalpy unsteady leakage flows that strongly reduce the engine efficiency and can cause thermally induced blade failures. An improved understanding of the tip flow physics is essential to refine the current design strategies and achieve increased turbine aerothermal performance. However, while past studies have mainly focused on conventional tip shapes (flat tip or squealer geometries), the open literature suffers from a shortage of experimental and numerical data on advanced blade tip configurations of unshrouded rotors. This work presents a complete numerical and experimental investigation on the unsteady flow field of a high-pressure turbine, adopting three different blade tip profiles. The aerothermal characteristics of two novel high-performance tip geometries, one with a fully contoured shape and the other presenting a multicavity squealer-like tip with partially open external rims, are compared against the baseline performance of a regular squealer geometry. The turbine stage is tested at engine-representative conditions in the high-speed turbine facility of the von Karman Institute. A rainbow rotor is mounted for simultaneous aerothermal testing of multiple blade tip geometries. On the rotor disk, the blades are arranged in sectors operating at two different clearance levels. A numerical campaign of full-stage simulations was also conducted on all the investigated tip designs to model the secondary flows development and identify the tip loss and heat transfer mechanisms. In the first part of this work, we describe the experimental setup, instrumentation, and data processing techniques used to measure the unsteady aerothermal field of multiple blade tip geometries using the rainbow rotor approach. We report the time-average and time-resolved static pressure and heat transfer measured on the shroud of the turbine rotor. The experimental data are compared against numerical predictions. These numerical results are then used in the second part of the paper to analyze the tip flow physics, model the tip loss mechanisms, and quantify the aero-thermal performance of each tip geometry.

scholarly journals Experimental and Numerical Investigation of Optimized Blade Tip Shapes: Part I — Turbine Rainbow Rotor Testing and CFD Methods

2018 ◽  
Author(s):  
Bogdan Cernat ◽  
Marek Pátý ◽  
Cis De Maesschalck ◽  
Sergio Lavagnoli
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
High Performance ◽  
Numerical Data ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Design Strategies ◽  
Flow Physics ◽  
Leakage Flows ◽  
Blade Tip

Blade tip design and tip leakage flows are crucial aspects for the development of modern aero-engines. The inevitable clearance between stationary and rotating parts in turbine stages generates high-enthalpy unsteady leakage flows that strongly reduce the engine efficiency and can cause thermally induced blade failures. An improved understanding of the tip flow physics is essential to refine the current design strategies and achieve increased turbine aerothermal performance. However, while past studies have mainly focused on conventional tip shapes (flat tip or squealer geometries), the open literature suffers from a shortage of experimental and numerical data on advanced blade tip configurations of unshrouded rotors. This work presents a complete numerical and experimental investigation on the unsteady flow field of a high-pressure turbine, adopting three different blade tip profiles. The aerothermal characteristics of two novel high-performance tip geometries, one with a fully contoured shape and the other presenting a multi-cavity squealer-like tip with partially open external rims, are compared against the baseline performance of a regular squealer geometry. The turbine stage is tested at engine-representative conditions in the high-speed turbine facility of the von Karman Institute. A rainbow rotor is mounted for simultaneous aerothermal testing of multiple blade tip geometries. On the rotor disk, the blades are arranged in sectors operating at two different clearance levels. A numerical campaign of full-stage simulations was also conducted on all the investigated tip designs to model the secondary flows development and identify the tip loss and heat transfer mechanisms. In the first part of this work, we describe the experimental setup, instrumentation and data processing techniques used to measure the unsteady aerothermal field of multiple blade tip geometries using the rainbow rotor approach. We report the time-average and time-resolved static pressure and heat transfer measured on the shroud of the turbine rotor. The experimental data are compared against CFD predictions. These numerical results are then used in the second part of the paper to analyze the tip flow physics, model the tip loss mechanisms and quantify the aero-thermal performance of each tip geometry.

scholarly journals Characterization of the Unsteady Aerodynamics of Optimized Turbine Blade Tips through Modal Decomposition Analysis

2019 ◽  
Vol 4 (2) ◽  
pp. 12
Author(s):  
Bogdan C. Cernat ◽  
Sergio Lavagnoli
Keyword(s):  
High Performance ◽  
Decomposition Analysis ◽  
Unsteady Aerodynamics ◽  
Tip Leakage Flow ◽  
Modal Decomposition ◽  
Time Resolved ◽  
Leakage Flows ◽  
Numerical Predictions ◽  
Blade Tip ◽  
Generation Mechanisms

The present research focused on the analysis of the leakage flows developing from advanced blade tip geometries. The aerodynamic field of a contoured blade tip and of a high-performance rimmed blade were investigated against a baseline squealer rotor. Time-resolved numerical predictions were combined with high-frequency pressure measurements to characterize the tip leakage flow of each tip design. High spatial and temporal resolution measurements provided a detailed representation of the unsteady flow in the near-tip region and at the stage outlet. Numerical computations, based on the nonlinear harmonic method, were employed to assess the unsteady blade row interactions and identify the loss generation mechanisms depending on the tip design. The space- and time-resolved flow field was analysed by modal decomposition to identify the main periodicities of the near-tip and outlet flow and classify the most relevant sources of aerodynamic unsteadiness and entropy generation across the stage.

Impact of Wall Temperature on Turbine Blade Tip Aerothermal Performance

2014 ◽  
Vol 136 (5) ◽  
Author(s):  
Q. Zhang ◽  
L. He
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Flow Field ◽  
Temperature Dependence ◽  
Turbine Blade ◽  
Wall Temperature ◽  
High Speed ◽  
Convective Heat Transfer Coefficient ◽  
Secondary Flows ◽  
Blade Tip

Currently the aerodynamics and heat transfer over a turbine blade tip tend to be analyzed separately with the assumption that the wall thermal boundary conditions do not affect the over-tip-leakage (OTL) flow field. There are some existing correlations for correcting the wall temperature effect on heat transfer when scaled to engine realistic conditions. But they were either developed to account for the temperature dependence of fluid properties largely empirically, or based on a boundary-layer model. It would be difficult (if not impossible) to define a boundary layer in many parts of a realistic blade passage with marked three-dimensional (3D) end wall and secondary flows (including those within a blade tip and around it). The questions to be asked here are: is the OTL aerodynamics significantly affected by the wall thermal condition? And if it is, how can we count this effect consistently in turbine blade tip design and analysis using modern CFD methods? In the present study the problem has been examined for typical high-pressure turbine blade tip configurations. An extensively developed RANS code (HYDRA) is employed and validated against the experimental data from a high speed linear cascade testing rig. The numerical analysis reveals that the wall–gas temperature ratio could greatly affect the transonic OTL flow field and there is a strong two-way coupling between aerodynamics and heat transfer. The feedbacks of the thermal boundary condition to aerodynamics behave differently at different flow regimes over the tip, clearly indicating a highly localized dependence of the convective heat transfer coefficient (HTC) upon wall temperatures. This implies that to use HTC for blade metal temperature predictions without resorting a fully conjugate solution, the temperature dependence needs to be corrected locally. A nonlinear correction approach has been adopted in the present work, and the results demonstrate its effectiveness for the transonic turbine tip configurations studied.

Transonic Turbine Blade Tip Aero-Thermal Performance With Different Tip Gaps: Part I—Tip Heat Transfer

2010 ◽  
Author(s):  
Q. Zhang ◽  
D. O. O’Dowd ◽  
L. He ◽  
M. L. G. Oldfield ◽  
P. M. Ligrani
Keyword(s):  
Heat Transfer ◽  
Thermal Performance ◽  
High Speed ◽  
Hot Spot ◽  
Structural Characteristics ◽  
Surface Heat ◽  
Tip Leakage Flow ◽  
Surface Heat Transfer ◽  
Numerical Predictions ◽  
Blade Tip

A closely combined experimental and CFD study on a transonic blade tip aero-thermal performance at engine representative Mach and Reynolds numbers (Mexit = 1, Reexit = 1.27×106) is presented in this and its companion paper (Part II). The present paper considers surface heat transfer distributions on tip surfaces, and on suction and pressure side surfaces (near-tip region). Spatially-resolved surface heat transfer data are measured using infrared thermography and transient techniques within the Oxford University High Speed Linear Cascade research facility. The Rolls-Royce PLC HYDRA suite is employed for numerical predictions for the same tip configuration and flow conditions. The CFD results are generally in good agreement with experimental data, and show that the flow over a large portion of the blade tip is supersonic for all three tip gaps investigated. Mach numbers within the tip gap become lower as the tip gap decreases. For the flow regions near the leading edge of the tip gap, surface Nusselt numbers decrease as the tip gap decreases. Opposite trends are observed for the trailing edge region. Several ‘hot spot’ features on blade tip surfaces are attributed to enhanced turbulence thermal diffusion in local regions. Other surface heat transfer variations are attributed to flow variations induced by shock waves. Flow structure and surface heat transfer variations are also investigated numerically when a moving casing is present. The inclusion of moving casing leads to notable changes to flow structural characteristics and associated surface heat transfer variations. However, significant portions of the tip leakage flow remain transonic with clearly identifiable shock wave structures.

Transonic Turbine Blade Tip Aerothermal Performance With Different Tip Gaps—Part I: Tip Heat Transfer

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
Q. Zhang ◽  
D. O. O’Dowd ◽  
L. He ◽  
M. L. G. Oldfield ◽  
P. M. Ligrani
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
Hot Spot ◽  
Structural Characteristics ◽  
Leading Edge ◽  
Surface Heat ◽  
Tip Leakage Flow ◽  
Surface Heat Transfer ◽  
Numerical Predictions ◽  
Blade Tip

A closely combined experimental and computational fluid dynamics (CFD) study on a transonic blade tip aerothermal performance at engine representative Mach and Reynolds numbers (Mexit=1,Reexit=1.27×106) is presented here and its companion paper (Part II). The present paper considers surface heat-transfer distributions on tip surfaces and on suction and pressure-side surfaces (near-tip region). Spatially resolved surface heat-transfer data are measured using infrared thermography and transient techniques within the Oxford University high speed linear cascade research facility. The Rolls-Royce PLC HYDRA suite is employed for numerical predictions for the same tip configuration and flow conditions. The CFD results are generally in good agreement with experimental data and show that the flow over a large portion of the blade tip is supersonic for all three tip gaps investigated. Mach numbers within the tip gap become lower as the tip gap decreases. For the flow regions near the leading edge of the tip gap, surface Nusselt numbers decrease as the tip gap decreases. Opposite trends are observed for the trailing edge region. Several “hot spot” features on blade tip surfaces are attributed to enhanced turbulence thermal diffusion in local regions. Other surface heat-transfer variations are attributed to flow variations induced by shock waves. Flow structure and surface heat-transfer variations are also investigated numerically when a moving casing is present. The inclusion of moving casing leads to notable changes to flow structural characteristics and associated surface heat-transfer variations. However, significant portions of the tip leakage flow remain transonic with clearly identifiable shock wave structures.

scholarly journals A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

1994 ◽  
Author(s):  
Y. W. Kim ◽  
W. Abdel-Messeh ◽  
J. P. Downs ◽  
F. O. Soechting ◽  
G. D. Steuber ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Secondary Flows ◽  
Test Surface ◽  
Leakage Flow ◽  
Test Technique ◽  
Coolant Injection ◽  
Blade Tip

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger who has made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work that has been performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow which is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

Effects of Guide Vanes on the Tip Heat Transfer Enhancement of a Turbine Blade

2011 ◽  
Author(s):  
Gongnan Xie ◽  
Bengt Sunde´n
Keyword(s):  
Heat Transfer ◽  
Fluid Flow ◽  
Heat Transfer Enhancement ◽  
Turbine Blade ◽  
Pressure Loss ◽  
Transfer Enhancement ◽  
Guide Vanes ◽  
Flow And Heat Transfer ◽  
Numerical Predictions ◽  
Blade Tip

Gas turbine blade tips encounter large heat load as they are exposed to the high temperature gas. A common way to cool the blade and its tip is to design serpentine passages with 180-deg turns under the blade tip-cap inside the turbine blade. Improved internal convective cooling is therefore required to increase the blade tip life time. This paper presents numerical predictions of turbulent fluid flow and heat transfer through two-pass channels with and without guide vanes placed in the turn regions using RANS turbulence modeling. The effects of adding guide vanes on the tip-wall heat transfer enhancement and the channel pressure loss were analyzed. The guide vanes have a height identical to that of the channel. The inlet Reynolds numbers are ranging from 100,000 to 600,000. The detailed three-dimensional fluid flow and heat transfer over the tip-walls are presented. The overall performances of several two-pass channels are also evaluated and compared. It is found that the tip heat transfer coefficients of the channels with guide vanes are 10∼60% higher than that of a channel without guide vanes, while the pressure loss might be reduced when the guide vanes are properly designed and located, otherwise the pressure loss is expected to be increased severely. It is suggested that the usage of proper guide vanes is a suitable way to augment the blade tip heat transfer and improve the flow structure, but is not the most effective way compared to the augmentation by surface modifications imposed on the tip-wall directly.

Comparisons of Pins/Dimples/Protrusions Cooling Concepts for a Turbine Blade Tip-Wall at High Reynolds Numbers

2010 ◽  
Author(s):  
Gongnan Xie ◽  
Bengt Sunde´n ◽  
Weihong Zhang
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Computational Models ◽  
Reynolds Numbers ◽  
Low Reynolds Numbers ◽  
Gas Leakage ◽  
Leakage Flows ◽  
Blade Tip ◽  
High Reynolds ◽  
The Cost

The blade tip region encounters high thermal loads because of the hot gas leakage flows, and it must therefore be cooled to ensure a long durability and safe operation. A common way to cool a blade tip is to design serpentine passages with 180° turn under the blade tip-cap inside the turbine blade. Improved internal convective cooling is therefore required to increase blade tip lifetime. Pins, dimples and protrusions are well recognized as effective devices to augment heat transfer in various applications. In this paper, enhanced heat transfer of an internal blade tip-wall has been predicted numerically. The computational models consist of a two-pass channel with 180° turn and arrays of circular pins or hemispherical dimples or protrusions internally mounted on the tip-wall. Inlet Reynolds numbers are ranging from 100,000 to 600,000. The overall performance of the two-pass channels is evaluated. Numerical results show that the heat transfer enhancement of the pinned tip is up to a factor of 3.0 higher than that of a smooth tip while the dimpled-tip and protruded-tip provide about 2.0 times higher heat transfer. These augmentations are achieved at the cost of an increase of pressure drop by less than 10%. By comparing the present cooling concepts with pins, dimples and protrusions, it is shown that the pinned-tip exhibit best performance to improve the blade tip cooling. However, when disregarding the added active area and considering the added mechanical stress, it is suggested that the usage of dimples is more suitable to enhance blade tip cooling, especially at low Reynolds numbers.

Numerical Predictions of Turbine Cascade Secondary Flows and Heat Transfer With Inflow Turbulence

2020 ◽  
Author(s):  
Yousef Kanani ◽  
Sumanta Acharya ◽  
Forrest Ames
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Leading Edge ◽  
Suction Side ◽  
Lateral Spreading ◽  
Secondary Flows ◽  
Pressure Side ◽  
Initial Flow ◽  
Numerical Predictions ◽  
Inflow Turbulence

Abstract Turbine passage secondary flows are studied for a large rounded leading edge airfoil geometry considered in the experimental investigation of Varty et al. (J. Turbomach. 140(2):021010) using high resolution Large Eddy Simulation (LES). The complex nature of secondary flow formation and evolution are affected by the approach boundary layer characteristics, components of pressure gradients tangent and normal to the passage flow, surface curvature, and inflow turbulence. This paper presents a detailed description of the secondary flows and heat transfer in a linear vane cascade at exit chord Reynolds number of 5 × 105 at low and high inflow turbulence. Initial flow turning at the leading edge of the inlet boundary layer leads to a pair of counter-rotating flow circulation in each half of the cross-plane that drive the evolution of the pressure-side and suction side of the near-wall vortices such as the horseshoe and leading edge corner vortex. The passage vortex for the current large leading-edge vane is formed by the amplification of the initially formed circulation closer to the pressure side (PPC) which strengthens and merges with other vortex systems while moving toward the suction side. The predicted suction surface heat transfer shows good agreement with the measurements and properly captures the augmented heat transfer due to the formation and lateral spreading of the secondary flows towards the vane midspan downstream of the vane passage. Effects of various components of the secondary flows on the endwall and vane heat transfer are discussed in detail.

Evaluation of Different Squealer Cavity Configurations in a HPT Blade Tip Region and its Influence on the Heat Transfer

2014 ◽  
Author(s):  
Lucilene Moraes da Silva ◽  
Jesuino Takachi Tomita ◽  
João Roberto Barbosa ◽  
Cleverson Bringhenti
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Rotor Blade ◽  
High Performance ◽  
Flow Behavior ◽  
Academic Research ◽  
Cavity Depth ◽  
Depth Variation ◽  
Blade Tip

In high performance turbomachines the tip region is a key point to improve aiming at high pressure ratios without high penalties. In the case of HPT, several techniques are still in development by academic research laboratories and industry. Some geometrical configurations were created at the rotor tip region, as winglets and squealers geometries. In the case of squealers, the depth of their cavity is an important parameter to evaluate, because its values can cause different flow behavior on this region. Changing the heat transfer. In this work, the rotor blade of a HPT developed in the E3 program was changed, the aim is to study the influence of the squealer cavity depth variation on its performance. The flow within the turbine was calculated using a commercial CFD package. The details of the rotor geometrical changes, the differences between a simple flat rotor tip surface and squealer configurations are discussed and presented.

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