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.

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.

Boundary Layer Influence on the Unsteady Horseshoe Vortex Flow and Surface Heat Transfer

2007 ◽  
Author(s):  
D. R. Sabatino ◽  
C. R. Smith
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Bluff Body ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Surface Heat ◽  
Wall Region ◽  
Thermochromic Liquid Crystal ◽  
Surface Heat Transfer ◽  
End Wall

The spatial-temporal flow-field and associated surface heat transfer within the leading edge, end-wall region of a bluff body were examined using both particle image velocimetry and thermochromic liquid crystal temperature measurements. The horseshoe vortex system in the end-wall region is mechanistically linked to the upstream boundary layer unsteadiness. Hairpin vortex packets, associated with turbulent boundary layer bursting behavior, amalgamate with the horseshoe vortex resulting in unsteady strengthening and streamwise motion. The horseshoe vortex unsteadiness exhibits two different natural frequencies: one associated with the transient motion of the horseshoe vortex, and the other with the transient surface heat transfer. Comparable unsteadiness occurs in the end-wall region of the more complex airfoil geometry of a linear turbine cascade. To directly compare the horseshoe vortex behavior around a turning airfoil to that of a simple bluff body, a length scale based on the maximum airfoil thickness is proposed.

Linear Cascade and Wind Tunnel Development for Turbine Blade Tip Heat Transfer Investigations With and Without Film Cooling

2019 ◽  
Author(s):  
Michael Sampson ◽  
Avery Fairbanks ◽  
Jacob Moseley ◽  
Phillip M. Ligrani ◽  
Hongzhou Xu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Boundary Layer Thickness ◽  
Surface Heat ◽  
Heat Transfer Characteristics ◽  
Surface Heat Transfer ◽  
Linear Cascade ◽  
External Cooling ◽  
Blade Tip

Abstract Currently, there is a deficit of experimental data for surface heat transfer characteristics and thermal transport processes associated with tip gap flows, and a lack of understanding of performance and behavior of film cooling as applied to blade tip surfaces. As a result, many avenues of opportunity exist for development of creative tip configurations with innovative external cooling arrangements. Overall goals of the present investigations are to reduce cooling air requirements, and reduce thermal loading, with equivalent improvements of thermal protection and structural integrity. Described is the development of experimental facilities, including a Supersonic/Transonic Wind Tunnel and linear cascade, for investigations of surface heat transfer characteristics of transonic turbine blade tips with unique squealer geometries and innovative film cooling arrangements. Note that data from past investigations are used to illustrate some of the experimental procedures and approaches which will be employed within the investigation. Of interest is development of a two-dimensional linear cascade with appropriate cascade airfoil flow periodicity. Included are boundary layer flow bleed devices, downstream tailboards, and augmented cascade inlet turbulence intensity. The present linear cascade approach allows experimental configuration parameters to be readily varied. Tip gap magnitudes are scaled so that ratios of tip gap to inlet boundary layer thickness, ratios of tip gap to blade axial chord length, and ratios of tip gap magnitudes to blade true chord length match engine hardware configurations. Ratios of inlet boundary layer thickness to tip gap range from 3 to 5. Innovative film cooling configurations are utilized for one blade tip configuration, and scaled engine components are modelled and tested with complete external cooling arrangements. Blade tip and geometry characteristics are also considered, including squealer depth and squealer tip wall thickness. With these experimental components, results will be obtained with engine representative transonic Mach numbers, Reynolds numbers, and film cooling parameters, including density ratios, which are achieved using foreign gas injection with carbon dioxide. Transient, infrared thermography approaches will be employed to measure spatially-resolved distributions of surface heat transfer coefficients, adiabatic surface temperature, and adiabatic film cooling effectiveness.

Controlling Secondary-Flow Structure by Leading-Edge Airfoil Fillet and Inlet Swirl to Reduce Aerodynamic Loss and Surface Heat Transfer

2002 ◽  
Author(s):  
T. I.-P. Shih ◽  
Y.-L. Lin
Keyword(s):  
Heat Transfer ◽  
Stokes Equations ◽  
Leading Edge ◽  
Surface Heat ◽  
Secondary Flows ◽  
Maximum Height ◽  
Surface Heat Transfer ◽  
Aerodynamic Loss ◽  
Swirl Angle ◽  
Inlet Swirl

Computations, based on the ensemble-averaged compressible Navier-Stokes equations closed by the shear-stress transport (SST) turbulence model, were performed to investigate the effects of leading-edge airfoil fillet and inlet-swirl angle on the flow and heat transfer in a turbine-nozzle guide vane. Three fillet configurations were simulated: no fillet (baseline), a fillet whose thickness fades on the airfoil, and a fillet whose thickness fades on the endwall. For both fillets, the maximum height above the endwall is positioned along the stagnation zone/line on the airfoil under the condition of no swirl. For each configuration, three inlet swirls were investigated: no swirl (baseline) and two linearly varying swirl angle from one endwall to the other (+30° to −30° and −30° to +30°). Results obtained show that both leading-edge fillet and inlet swirl can reduce aerodynamic loss and surface heat transfer. For the conditions of this study, the difference in stagnation pressure from the nozzle’s inlet to its exit were reduced by more than 40% with swirl or with fillet without swirl. Surface heat transfer was reduced by more than 10% on the airfoil and by more than 30% on the endwalls. When there is swirl, leading-edge fillets became less effective in reducing aerodynamic loss and surface heat transfer, because the fillets were not optimized for swirl angles imposed. Since the intensity and size of the cross flow were found to increase instead of decrease by inlet swirl and by the type of fillet geometries investigated, the results of this study indicate that the mechanisms responsible for aerodynamic loss and surface heat transfer are more complex than just the intensity and the magnitude of the secondary flows. This study shows their location and interaction with the main flow to be more important, and this could be exploited for positive results.

Flowfield Measurements for a Highly Turbulent Flow in a Stator Vane Passage

1999 ◽  
Author(s):  
R. W. Radomsky ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbulent Flow ◽  
Reynolds Shear Stress ◽  
Correlation Coefficients ◽  
Leading Edge ◽  
Surface Heat ◽  
Laser Doppler Velocimeter ◽  
Surface Heat Transfer ◽  
Turbine Vanes ◽  
Stator Vane

Turbine vanes experience high convective surface heat transfer as a consequence of the turbulent flow exiting the combustor. Before improvements to vane heat transfer predictions through boundary layer calculations can be made, we need to understand how the turbulent flow in the inviscid region of the passage reacts as it passes between two adjacent turbine vanes. In this study, a scaled-up turbine vane geometry was used in a low-speed wind tunnel simulation. The test section included a central airfoil with two adjacent vanes. To generate the 20% turbulence levels at the entrance to the cascade, which simulates levels exiting the combustor, an active grid was used. Three-component laser Doppler velocimeter measurements of the mean and fluctuating quantities were measured in a plane at the vane mid-span. Coincident velocity measurements were made to quantify Reynolds shear stress and correlation coefficients. The energy spectra and length scales were also measured to give a complete set of inlet boundary conditions that can be used for numerical simulations. The results show that the turbulent kinetic energy throughout the inviscid region remained relatively high. The surface heat transfer measurements indicated high augmentation near the leading edge as well as the pressure side of the vane as a result of the elevated turbulence levels.

Aerothermal Measurements and Predictions of an Intermediate Turbine Duct Turning Vane

2015 ◽  
Author(s):  
Martin Johansson ◽  
Jonathan Mårtensson ◽  
Hans Abrahamsson ◽  
Thomas Povey ◽  
Kam Chana
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
High Speed ◽  
Test Facility ◽  
Spanwise Direction ◽  
Tip Leakage Flow ◽  
Current Configuration ◽  
Speed Test ◽  
Intermediate Turbine Duct ◽  
Numerical Predictions

Flow in a turbine duct is highly complex, influenced by the upstream turbine stage flow structures, including tip leakage flow and non-uniformities originating from the upstream HPT vane and rotor. The complexity of the flow makes the prediction using numerical methods difficult, hence there exists a need for experimental validation. This paper presents experimental data including both aerodynamic and heat transfer measurements within an intermediate turbine duct. These have been conducted in the Oxford Turbine Research Facility, a short duration high speed test facility enabling the use of an engine sized turbine, operating at the correct non-dimensional parameters relevant for aerodynamic and heat transfer measurements. The current configuration consists of a HPT stage and a downstream duct including a turning vane, for use in a counter rotating turbine configuration. With a stator-to-stator vane count of 32-to-24, instrumentation was installed on three adjacent intermediate turbine duct vanes and endwalls to investigate its influence. Flow phenomena such as trailing edge wakes and vortex structures from the upstream HPT vane travels through the rotor and forms an inlet condition to the intermediate turbine duct with tangential variations. Time-averaged experimental data show this effect to be distinguishable although varying in the spanwise direction. Comparisons with results from numerical predictions are included to further analyse the flow through the 1.5 stage.

Effects of Inlet Swirl on Hot Streak Migration Across Tip Clearance and Heat Transfer on Rotor Blade Tip

2015 ◽  
Author(s):  
Zhaofang Liu ◽  
Zhiduo Wang ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Heat Load ◽  
Leading Edge ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Inlet Swirl ◽  
Blade Tip ◽  
Hot Streak

This paper presents an investigation on the hot streak migration across tip clearance and heat transfer on blade tip in a high pressure (HP) gas turbine with different inlet swirl directions and clocking positions. The geometry is taken from the first stage of GE-E3 turbine engine. Two swirl directions (positive and negative) and two circumferential clocking positions (aligning with S1 nozzle leading edge and mid passage) for inlet hot streak and swirl have been employed and investigated, respectively. Two cases with only hot streak at different inlet circumferential positions are adopted as the baseline in this study. By solving the unsteady compressible Reynolds-averaged Navier-Stokes equations, the time dependent solutions were obtained. The results indicate that the influence of inlet swirl on pressure distribution focuses on the suction side. Positive swirl attracts more hot fluid to the upper endwall, when it aligns with nozzle stator leading edge. Because of the squeezing mechanism between positive swirl and leakage flow, the heat transfer on rotor blade tip is more uniform. While negative swirl increases tip leakage flow and the heat load at the first half on tip surface. In all cases with swirl, the heat load at the second half on blade tip is effectively reduced, which is good for cooling rotor blade tip. If the stator is cooled effectively, inlet positive swirl aligning with nozzle vane leading edge will be the best choice for protecting rotor blade tip. By comparing with the results of previous literature, it is concluded that whatever arrangement the blade rows locate, the swirl direction which is opposite to the leakage flow should be chosen for protecting not only blade surface but also blade tip when the inlet swirl exists.

Effect of Two-Scale Roughness on Boundary Layer Transition Over a Heated Flat Plate: Part 1—Surface Heat Transfer

1999 ◽  
Vol 122 (2) ◽  
pp. 301-307 ◽  
Author(s):  
Mark W. Pinson ◽  
Ting Wang
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Leading Edge ◽  
Step Change ◽  
Surface Heat ◽  
Boundary Layer Transition ◽  
Test Surface ◽  
Layer Transition ◽  
Surface Heat Transfer ◽  
Scale Roughness

An experimental study was conducted to investigate surface heat transfer and boundary layer development associated with flow over a flat test surface covered with two roughness scales. Two-scale roughness was used because in-service aeroengines commonly display larger roughness concentrated at the leading edge with smaller roughness distributed downstream. The first scale, covering up to the first 5 cm of the test surface, was in the form of a sandpaper strip, an aluminum strip, or a cylinder. The second roughness scale covered the remainder of the test surface (2 m) in the form of sandpaper or a smooth surface. In Part 1, the surface heat transfer results are examined. Even though the roughness scales were hydraulically smooth, they induced significantly earlier transition onset, with the two-dimensional roughness causing earlier transition than three-dimensional roughness. All of the rough/smooth cases unexpectedly triggered earlier transition than rough/rough cases. This indicated that the scale of the step-change at the joint between two roughness scales was predominant over the downstream roughness on inducing early transition. Reducing the overall height of the step change was shown to have a greater effect on transition than the specific geometry of the roughness scale. [S0889-504X(00)00701-7]

Experimental and Numerical Characterization of Droplet-Induced Spreading-Splashing Transition in Surface Cooling

2016 ◽  
Author(s):  
Taolue Zhang ◽  
J. P. Muthusamy ◽  
Jorge Alvarado ◽  
Anoop Kanjirakat ◽  
Reza Sadr
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Surface Temperature ◽  
Weber Number ◽  
High Speed ◽  
Single Phase ◽  
Surface Heat ◽  
Droplet Impingement ◽  
Surface Heat Transfer ◽  
Dry Out

The effects of droplet train impingement on spreading-splashing transition and surface heat transfer were investigated experimentally and numerically. Experimentally, a single stream of HFE-7100 droplet train was generated using a piezo-electric droplet generator with the ability to adjust parameters such as droplet impingement frequency, droplet diameter and droplet impingement velocity. A thin layer of Indium Tin Oxide (ITO) was coated on a translucent sapphire substrate, which was used as heating element. High-speed and infrared imaging techniques were employed to characterize the hydrodynamics and heat transfer of droplet train impingement. Numerically, the high frequency droplet train impingement process was simulated using ANSYS-Fluent with the Volume of Fluid (VOF) method [1]. The heat transfer process was simulated by applying constant heat flux conditions on the droplet receiving surface. Droplet-induced spreading-splashing transition behavior was investigated by increasing the droplet Weber number while holding flow rate constant. High speed crown propagation images showed that at low-Weber number (We < 400), droplet impingements resulted in smooth spreading of the droplet-induced crown. However, within the transitional droplet Weber number range (We = 400–500), fingering and splashing (i.e. emergence of secondary droplets) could be observed at the crown’s rim. At high droplet Weber number (We > 800), breakup of the crown was observed during the crown propagation process in which the liquid film behaved chaotically. Droplet-induced spreading-splashing transition phenomena were also investigated numerically. Reasonable agreement was reached between the experimental and numerical results in terms of crown morphology at different droplet Weber number values. The effects of spreading-splashing transition on surface heat transfer were also investigated at fixed flow rate conditions. Time-averaged Infrared (IR) temperature measurements indicate that heat flux-surface temperature curves are linear at low surface temperatures and before the onset of dry-out, which indicate that single phase forced convection is the primary heat transfer mechanism under those conditions. Numerical heat transfer simulations were performed within the single phase forced convection regime only. Instantaneous numerical results reveal that droplet-induced crown propagation effectively convect heat radially outward within the droplet impingement zone. Under high heat flux conditions, a sharp increase in surface temperature was observed experimentally when dry-out appeared on the heater surface. It was also found that strong splashing (We > 800) is unfavorable for heat transfer at high surface temperature due to the onset of instabilities seen in the liquid film, which leads to dry-out conditions. In summary, the results indicate that droplet Weber number is a significant factor in the spreading-splashing transition and surface heat transfer.

Flowfield Measurements for a Highly Turbulent Flow in a Stator Vane Passage

1999 ◽  
Vol 122 (2) ◽  
pp. 255-262 ◽  
Author(s):  
R. W. Radomsky ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbulent Flow ◽  
Reynolds Shear Stress ◽  
Correlation Coefficients ◽  
Leading Edge ◽  
Surface Heat ◽  
Laser Doppler Velocimeter ◽  
Surface Heat Transfer ◽  
Turbine Vanes ◽  
Stator Vane

Turbine vanes experience high convective surface heat transfer as a consequence of the turbulent flow exiting the combustor. Before improvements to vane heat transfer predictions through boundary layer calculations can be made, we need to understand how the turbulent flow in the inviscid region of the passage reacts as it passes between two adjacent turbine vanes. In this study, a scaled-up turbine vane geometry was used in a low-speed wind tunnel simulation. The test section included a central airfoil with two adjacent vanes. To generate the 20 percent turbulence levels at the entrance to the cascade, which simulates levels exiting the combustor, an active grid was used. Three-component laser-Doppler velocimeter measurements of the mean and fluctuating quantities were measured in a plane at the vane midspan. Coincident velocity measurements were made to quantify Reynolds shear stress and correlation coefficients. The energy spectra and length scales were also measured to give a complete set of inlet boundary conditions that can be used for numerical simulations. The results show that the turbulent kinetic energy throughout the inviscid region remained relatively high. The surface heat transfer measurements indicated high augmentation near the leading edge as well as the pressure side of the vane as a result of the elevated turbulence levels. [S0889-504X(00)02302-3]

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