Time-Averaged Heat Flux for a Recessed Tip, Lip, and Platform of a Transonic Turbine Blade

2000 ◽  
Vol 122 (4) ◽  
pp. 692-698 ◽  
Author(s):  
M. G. Dunn ◽  
C. W. Haldeman
Keyword(s):  
Heat Flux ◽  
Pressure Ratio ◽  
Design Flow ◽  
Turbine Stage ◽  
Suction Surface ◽  
Surface Measurements ◽  
Blade Tip ◽  
Flux Level ◽  
Transonic Turbine ◽  
Exit Mach Number

The results of an experimental research program determining the blade platform heat-flux level and the influence of blade tip recess on the tip region heat transfer for a full-scale rotating turbine stage at transonic vane exit conditions are described. The turbine used for these measurements was the Allison VBI stage operating in the closed vane position (vane exit Mach number≈1.1). The stage was operated at the design flow function, total to static pressure ratio, and corrected speed. Measurements were obtained at several locations on the platform and in the blade tip region. The tip region consists of the bottom of the recess, the lip region (on both the pressure and suction surface sides of the recess), and the 90 percent span location on the blade suction surface. Measurements were obtained for three vane/blade spacings; 20, 40, and 60 percent of vane axial chord and for a single value of the tip gap (the distance between the top of the lip and the stationary shroud) equal to 0.0012 m (0.046 in) or 2.27 percent of blade height. [S0889-504X(00)00604-8]

Time-Averaged Heat Flux for a Recessed Tip, Lip, and Platform of a Transonic Turbine Blade

2000 ◽  
Author(s):  
M. G. Dunn ◽  
C. W. Haldeman
Keyword(s):  
Heat Flux ◽  
Pressure Ratio ◽  
Design Flow ◽  
Turbine Stage ◽  
Suction Surface ◽  
Surface Measurements ◽  
Blade Tip ◽  
Flux Level ◽  
Transonic Turbine ◽  
Exit Mach Number

The results of an experimental research program determining the blade platform heat-flux level and the influence of blade tip recess on the tip region heat transfer for a full-scale rotating turbine stage at transonic vane exit conditions are described. The turbine used for these measurements was the Allison VBI stage operating in the closed vane position (vane exit Mach number _ 1.1). The stage was operated at the design flow function, total to static pressure ratio, and corrected speed. Measurements were obtained at several locations on the platform and in the blade tip region. The tip region consists of the bottom of the recess, the lip region (on both the pressure and suction surface sides of the recess), and the 90% span location on the blade suction surface. Measurements were obtained for three vane/blade spacings; 20%, 40%, and 60% of vane axial chord and for a single value of the tip gap (the distance between the top of the lip and the stationary shroud) equal to 0.0012-m (0.046-in) or 2.27% of blade height.

scholarly journals Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly Loaded Transonic Turbine Stage

2010 ◽  
Author(s):  
Vikram Shyam ◽  
Ali Ameri ◽  
Jen-Ping Chen
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Reynolds Number ◽  
Rotor Blade ◽  
High Speed ◽  
Pressure Ratio ◽  
Stagnation Temperature ◽  
Turbine Stage ◽  
Transonic Turbine ◽  
Highly Loaded

In a previous study, vane-rotor shock interactions and heat transfer on the rotor blade of a highly loaded transonic turbine stage were simulated. The geometry consists of a high pressure turbine vane and downstream rotor blade. This study focuses on the physics of flow and heat transfer in the rotor tip, casing and hub regions. The simulation was performed using the URANS (Unsteady Reynolds-Averaged Navier-Stokes) code MSU-TURBO. A low Reynolds number k-ε model was utilized to model turbulence. The rotor blade in question has a tip gap height of 2.1% of the blade height. The Reynolds number of the flow is approximately 3×106 per meter. Unsteadiness was observed at the tip surface that results in intermittent ‘hot spots’. It is demonstrated that unsteadiness in the tip gap is governed by inviscid effects due to high speed flow and is not strongly dependent on pressure ratio across the tip gap contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. In addition, an explanation is provided that shows the mechanism behind the rise in stagnation temperature on the casing to values above the absolute total temperature at the inlet. It is concluded that even in steady mode, work transfer to the near tip fluid occurs due to relative shearing by the casing. This is believed to be the first such explanation of the work transfer phenomenon in the open literature. The difference in pattern between steady and time-averaged heat flux at the hub is also explained.

Parametrical Investigation of Turbine Stages With Open Tip Clearance for the Purpose of Stage Efficiency Increase

2010 ◽  
Author(s):  
Andrei Granovskiy ◽  
Mikhail Kostege ◽  
Nikolay Lomakin
Keyword(s):  
Pressure Ratio ◽  
Axial Flow ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Turbine Stage ◽  
Suction Surface ◽  
Tip Leakage ◽  
Degree Of Reaction ◽  
Blade Tip ◽  
Stage Efficiency

The aerodynamic loss due to tip leakage vortex of rotor blades represents a significant part of viscous losses in axial flow turbines. The mixing of leakage flow with the main rotor passage flow causes losses and reduces turbine stage efficiency. Many measures have been proposed to reduce the loss in the tip clearance area. In this paper the reduction of the tip clearance loss due to changes made to the blade tip section profile is presented. The blade tip profile was modified to decrease the pressure gradient between pressure surface and suction surface. This approach allows the reduction of tip leakage and tip vortex strength and consequently the reduction of tip clearance losses. A 3D Navier-Stokes solver with q-ω turbulence model is used to analyze the flow in the turbine with various tip section profiles. Test data of three single-stage experimental turbines have been used to validate analytical predictions: • Highly loaded turbine stage with a pressure ratio π0T = 3.2 and reaction degree ρmean = 0.5. • Two turbines with a pressure ratio π0T = 3.9. (One with high degree of reaction ρmean = 0.55; the other with low degree of reaction ρmean = 0.26). The numerical investigation of the influence of various tip section profiles on stage efficiency was carried out in the range of relative tip clearance 0.5%–2.4% with the objective of a decreasing the influence of the tip clearance on the stage efficiency.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

Investigation of the 3D Unsteady Rotor Pressure Field in a HP Turbine Stage

2002 ◽  
Author(s):  
E. Valenti ◽  
J. Halama ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Pressure Field ◽  
Pressure Ratio ◽  
Turbine Stage ◽  
Time Resolved ◽  
Mach Number Distribution ◽  
Rotor Surface ◽  
Exit Mach Number ◽  
Unsteady Pressure Measurements

This paper presents steady and unsteady pressure measurements at three span locations (15, 50 and 85%) on the rotor surface of a transonic turbine stage. The data are compared with the results of a 3D unsteady Euler stage calculation. The overall agreement between the measurements and the prediction is satisfactory. The effects of pressure ratio and Reynolds number are discussed. The rotor time-averaged Mach number distribution is very sensitive to the pressure ratio of the stage since the incidence of the flow changes as well as the rotor exit Mach number. The time-resolved pressure field is dominated by the vane trailing edge shock waves. The incidence and intensity of the shock strongly varies from hub to tip due to the radial equilibrium of the flow at the vane exit. The decrease of the pressure ratio attenuates significantly the amplitude of the fluctuations. An increase of the pressure ratio has less significant effect since the change in the vane exit Mach number is small. The effect of the Reynolds number is weak for both the time-averaged and the time-resolved rotor static pressure at mid-span, while it causes an increase of the pressure amplitudes at the two other spans.

Development of New Single and High-Density Heat-Flux Gauges for Unsteady Heat Transfer Measurements for a Rotating Transonic Turbine

2020 ◽  
Author(s):  
Richard Celestina ◽  
Spencer Sperling ◽  
Louis Christensen ◽  
Randall Mathison ◽  
Hakan Aksoy ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Pressure Ratio ◽  
High Density ◽  
Secondary Flows ◽  
Single Stage ◽  
High Pressure Turbine ◽  
Transonic Turbine ◽  
New Generation

Abstract This paper presents the development and implementation of a new generation of double-sided heat-flux gauges at The Ohio State University Gas Turbine Laboratory (GTL) along with heat transfer measurements for film-cooled airfoils in a single-stage high-pressure transonic turbine operating at design corrected conditions. Double-sided heat flux gauges are a critical part of turbine cooling studies, and the new generation improves upon the durability and stability of previous designs while also introducing high-density layouts that provide better spatial resolution. These new customizable high-density double-sided heat flux gauges allow for multiple heat transfer measurements in a small geometric area such as immediately downstream of a row of cooling holes on an airfoil. Two high-density designs are utilized: Type A consists of 9 gauges laid out within a 5 mm by 2.6 mm (0.20 inch by 0.10 inch) area on the pressure surface of an airfoil, and Type B consists of 7 gauges located at points of predicted interest on the suction surface. Both individual and high-density heat flux gauges are installed on the blades of a transonic turbine experiment for the second build of the High-Pressure Turbine Innovative Cooling program (HPTIC2). Run in a short duration facility, the single-stage high-pressure turbine operated at design-corrected conditions (matching corrected speed, flow function, and pressure ratio) with forward and aft purge flow and film-cooled blades. Gauges are placed at repeated locations across different cooling schemes in a rainbow rotor configuration. Airfoil film-cooling schemes include round, fan, and advanced shaped cooling holes in addition to uncooled airfoils. Both the pressure and suction surfaces of the airfoils are instrumented at multiple wetted distance locations and percent spans from roughly 10% to 90%. Results from these tests are presented as both time-average values and time-accurate ensemble averages in order to capture unsteady motion and heat transfer distribution created by strong secondary flows and cooling flows.

Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2009 ◽  
Vol 1 (2) ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

The Unsteady Effect of Passing Shocks on Pressure Surface Versus Suction Surface Heat Transfer in Film-Cooled Transonic Turbine Blades

2003 ◽  
Author(s):  
A. C. Smith ◽  
A. C. Nix ◽  
T. E. Diller ◽  
W. F. Ng
Keyword(s):  
Heat Transfer ◽  
Shock Wave ◽  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Surface Heat ◽  
Suction Surface ◽  
Time Resolved ◽  
Pressure Surface ◽  
Transonic Turbine

This paper documents the measurement of the unsteady effects of passing shock waves on film cooling heat transfer on both the pressure and suction surfaces of first stage transonic turbine blades with leading edge showerhead film cooling. Experiments were performed for several cooling blowing ratios with an emphasis on time-resolved pressure and heat flux measurements on the pressure surface. Results without film cooling on the pressure surface demonstrated that increases in heat flux were a result of shock heating (the increase in temperature across the shock wave) rather than shock interaction with the boundary layer or film layer. Time-resolved measurements with film cooling demonstrated that the relatively strong shock wave along the suction surface appears to retard coolant ejection there and causes excess coolant to be ejected from pressure surface holes. This actually causes a decrease in heat transfer on the pressure surface during a large portion of the shock passing event. The magnitude of the decrease is almost as large as the increase in heat transfer without film cooling. The decrease in coolant ejection from the suction surface holes did not appear to have any effects on suction surface heat transfer.

Influence of Rotor Blade Aerodynamic Loading on the Performance of a Highly Loaded Turbine Stage

1987 ◽  
Vol 109 (2) ◽  
pp. 155-161 ◽  
Author(s):  
S. H. Moustapha ◽  
U. Okapuu ◽  
R. G. Williamson
Keyword(s):  
Rotor Blade ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Single Stage ◽  
Turbine Stage ◽  
Blade Loading ◽  
Aerodynamic Loading ◽  
Transonic Turbine ◽  
Stage Performance ◽  
Highly Loaded

This paper describes the performance of a highly loaded single-stage transonic turbine with a pressure ratio of 3.76 and a stage loading factor of 2.47. Tests were carried out with three rotors, covering a range of blade Zweifel coefficient of 0.77 to 1.18. Detailed traversing at rotor inlet and exit allowed an assessment of rotor and stage performance as a function of blade loading under realistic operating conditions. The effect of stator endwall contouring on overall stage performance was also investigated using two different contours with the same vane design.

Heat-Flux Measurements for the Rotor of a Full-Stage Turbine: Part II—Description of Analysis Technique and Typical Time-Resolved Measurements

1986 ◽  
Vol 108 (1) ◽  
pp. 98-107 ◽  
Author(s):  
M. G. Dunn ◽  
W. K. George ◽  
W. J. Rae ◽  
S. H. Woodward ◽  
J. C. Moller ◽  
...  
Keyword(s):  
Thin Film ◽  
Heat Flux ◽  
Design Flow ◽  
Suction Surface ◽  
Stagnation Region ◽  
Time Resolved ◽  
Analysis Technique ◽  
Typical Time ◽  
Flux Measurements ◽  
Dominant Frequencies

This paper presents a detailed description of an analysis technique and an application of this technique to obtain time-resolved heat flux for the blade of a Garrett TFE 731-2 hp full-stage rotating turbine. A shock tube is used as a short-duration source of heated air and platinum thin-film gages are used to obtain the heat-flux measurements. To obtain the heat-flux values from the thin-film gage temperature histories, a finite-difference procedure has been used to solve the heat equation, with variable thermal properties. The data acquisition and the data analysis procedures are described in detail and then their application is illustrated for three midspan locations on the blade. The selected locations are the geometric stagnation point, 32.7 percent wetted distance on the suction surface, and 85.5 percent wetted distance on the suction surface. For these measurements, the turbine was operating at the design flow function and very near 100 percent corrected speed. The vane–blade axial spacing was consistent with the engine operating configuration. The results demonstrate that the magnitude of the heat-flux fluctuation resulting from the vane–blade interaction is large by comparison with the time-averaged heat flux at all locations investigated. The magnitude of the fluctuation is greatest in the stagnation region and decreases with increasing wetted distance along the surface. A Fourier analysis by FFT of a portion of the heat-flux record illustrates that the dominant frequencies occur at the wake-cutting frequency and its harmonics.

Sign in / Sign up

Export Citation Format

Share Document