An Experimental Study of Film Cooling in a Rotating Transonic Turbine

1994 ◽  
Vol 116 (1) ◽  
pp. 63-70 ◽  
Author(s):  
R. S. Abhari ◽  
A. H. Epstein
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Test Facility ◽  
Turbine Stage ◽  
Suction Surface ◽  
Time Resolved ◽  
Pressure Surface ◽  
Transonic Turbine ◽  
Percent Decrease ◽  
Time Resolved Measurements

Time-resolved measurements of heat transfer on a fully cooled transonic turbine stage have been taken in a short duration turbine test facility, which simulates full engine nondimensional conditions. The time average of this data is compared to uncooled rotor data and cooled linear cascade measurements made on the same profile. The film cooling reduces the time-averaged heat transfer compared to the uncooled rotor on the blade suction surface by as much as 60 percent, but has relatively little effect on the pressure surface. The suction surface rotor heat transfer is lower than that measured in the cascade. The results are similar over the central 3/4 of the span, implying that the flow here is mainly two dimensional. The film cooling is shown to be much less effective at high blowing ratios than at low ones. Time-resolved measurements reveal that the cooling, when effective, both reduced the dc level of heat transfer and changed the shape of the unsteady waveform. Unsteady blowing is shown to be a principal driver of film cooling fluctuations, and a linear model is shown to do a good job in predicting the unsteady heat transfer. The unsteadiness results in a 12 percent decrease in heat transfer on the suction surface and a 5 percent increase on the pressure surface.

scholarly journals An Experimental Study of Film Cooling in a Rotating Transonic Turbine

1992 ◽  
Author(s):  
Reza S. Abhari ◽  
A. H. Epstein
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Test Facility ◽  
Turbine Stage ◽  
Unsteady Heat Transfer ◽  
Suction Surface ◽  
Time Resolved ◽  
Pressure Surface ◽  
Transonic Turbine ◽  
Time Resolved Measurements

Time-resolved measurements of heat transfer on a fully cooled transonic turbine stage have been taken in a short duration turbine test facility which simulates full engine non-dimensional conditions. The time average of this data is compared to uncooled rotor data and cooled linear cascade measurements made on the same profile. The film cooling reduces the time-averaged heat transfer compared to the uncooled rotor on the blade suction surface by as much as 60%, but has relatively little effect on the pressure surface. The suction surface rotor heat transfer is lower than that measured in the cascade. The results are similar over the central 3/4 of the span implying that the flow here is mainly two-dimensional. The film cooling is shown to be much less effective at high blowing ratios than at low ones. Time-resolved measurements reveal that the cooling, when effective, both reduced the d.c. level of heat transfer and changed the shape of the unsteady waveform. Unsteady blowing is shown to be a principal driver of film cooling fluctuations, and a linear model is shown to do a good job in predicting the unsteady heat transfer. The unsteadiness results in a 12% decrease in heat transfer on the suction surface and a 5% increase on the pressure surface.

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.

Unsteady Rotor Heat Transfer in a Transonic Turbine Stage

2002 ◽  
Author(s):  
F. Didier ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Strong Dependence ◽  
Convective Heat Transfer Coefficient ◽  
Operating Conditions ◽  
Turbine Stage ◽  
Trailing Edge ◽  
Time Resolved ◽  
Number Distribution ◽  
Transonic Turbine

This experimental investigation reports the convective heat transfer coefficient around the rotor of a transonic turbine stage. Both time-resolved and time-averaged aspects are addressed. The measurements are performed around the rotor blade at 15%, 50% and 85% span as well as on the rotor tip and the hub platform. Four operating conditions are tested covering two Reynolds numbers and three pressure ratios. The tests are performed in the compression tube turbine test rig CT3 of the von Karman Institute, allowing a correct simulation of the operating conditions encountered in modern aero-engines. The time-averaged Nusselt number distribution shows the strong dependence on both blade Mach number distribution and Reynolds number. The time-resolved heat transfer rate is mostly dictated by the vane trailing edge shock impingement on the rotor boundary layer. The shock passage corresponds to a sudden heat transfer increase. The effects are more pronounced in the leading edge region. The increase of the stage pressure ratio causes a stronger vane trailing edge shock and thus larger heat transfer fluctuations. The influence of the Reynolds number is hardly visible.

Comparison of Steady and Unsteady Coupled Heat-Transfer Simulations of a High-Pressure Turbine Blade

2015 ◽  
Author(s):  
Markus Schmidt ◽  
Christoph Starke
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Steady State ◽  
Flow Field ◽  
Film Cooling ◽  
Strong Increase ◽  
Pressure Side ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Time Resolved

This article presents results for the coupled simulation of a high-pressure turbine stage in consideration of unsteady hot gas flows. A semi-unsteady coupling process was developed to solve the conjugate heat transfer problem for turbine components of gas turbines. Time-resolved CFD simulations are coupled to a finite element solver for the steady state heat conduction inside of the blade material. A simplified turbine stage geometry is investigated in this paper to describe the influence of the unsteady flow field onto the time-averaged heat transfer. Comparisons of the time-resolved results to steady state results indicate the importance of a coupled simulation and the consideration of the time-dependent flow-field. Different film-cooling configurations for the turbine NGV are considered, resulting in different temperature and pressure deficits in the vane wake. Their contribution to non-linear effects causing the time-averaged heat load to differ from a steady result is discussed to further highlight the necessity of unsteady design methods for future turbine developments. A strong increase in the pressure side heat transfer coefficients for unsteady simulations is observed in all results. For higher film-cooling mass flows in the upstream row, the preferential migration of hot fluid towards the pressure side of a turbine blade is amplified as well, which leads to a strong increase in material temperature at the pressure side and also in the blade tip region.

Unsteady Rotor Heat Transfer in a Transonic Turbine Stage

2002 ◽  
Vol 124 (4) ◽  
pp. 614-622 ◽  
Author(s):  
F. Didier ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Strong Dependence ◽  
Convective Heat Transfer Coefficient ◽  
Operating Conditions ◽  
Turbine Stage ◽  
Trailing Edge ◽  
Time Resolved ◽  
Number Distribution ◽  
Transonic Turbine

This experimental investigation reports the convective heat transfer coefficient around the rotor of a transonic turbine stage. Both time-resolved and time-averaged aspects are addressed. The measurements are performed around the rotor blade at 15, 50, and 85% span as well as on the rotor tip and the hub platform. Four operating conditions are tested covering two Reynolds numbers and three pressure ratios. The tests are performed in the compression tube turbine test rig CT3 of the von Karman Institute, allowing a correct simulation of the operating conditions encountered in modern aero-engines. The time-averaged Nusselt number distribution shows the strong dependence on both blade Mach number distribution and Reynolds number. The time-resolved heat transfer rate is mostly dictated by the vane trailing edge shock impingement on the rotor boundary layer. The shock passage corresponds to a sudden heat transfer increase. The effects are more pronounced in the leading edge region. The increase of the stage pressure ratio causes a stronger vane trailing edge shock and thus larger heat transfer fluctuations. The influence of the Reynolds number is hardly visible.

scholarly journals Unsteady Flow in a Single Stage Turbine

1998 ◽  
Author(s):  
Mary A. Hilditch ◽  
Graham C. Smith ◽  
Udai K. Singh
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Fluctuations ◽  
Pressure Waves ◽  
Turbine Stage ◽  
Unsteady Heat Transfer ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Pressure Surface ◽  
Wake Interaction

This paper presents unsteady pressure and heat transfer measurements made on a high pressure turbine stage at DERA Pyestock, and compares them with numerical simulations made using the 2D unsteady code UNSFLO. The aim of the work was to evaluate the performance of the code, and to use the predictions to allow a fuller interpretation of the flow physics than could have been achieved from the measurements alone. The unsteady heat transfer and pressure fluctuations around the mid height section of the rotor blade have been examined in detail. Agreement between measured and predicted pressure fluctuations on the rotor was excellent. Interaction with the ngv potential field dominated the pressure surface, while the suction surface showed pressure waves moving forward over the blade, possibly as a result of shock/wake interaction.

Aerothermal Impact of Stator-Rim Purge Flow and Rotor-Platform Film Cooling on a Transonic Turbine Stage

2008 ◽  
Author(s):  
M. Pau ◽  
G. Paniagua ◽  
D. Delhaye ◽  
A. de la Loma ◽  
P. Ginibre
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Fluctuations ◽  
Pressure Sensors ◽  
Suction Side ◽  
Turbine Stage ◽  
Purge Flow ◽  
Rotor Disk ◽  
Transonic Turbine ◽  
The Impact

This paper describes the effects on the mainstream flow of two types of cooling techniques in a transonic turbine stage: purge gas ejected out of the cavity between the stator rim and the rotor disk, as well as film cooling gas discharged from the rotor-platform. The tests were carried out in a full annular stage fed by a compression tube, at M2is = 1.1, Re = 1.1×106, and at temperature ratios reproducing engine conditions. The stator outlet was instrumented to allow the aerothermal characterization of the purge flow. The rotor blade was heavily instrumented with fast-response pressure sensors and double-layer thin film gauges. The tests are coupled with numerical calculations performed using the ONERA’s code elsA. The stator-rotor interaction is seen to be significantly affected by the stator-rim seal, both in terms of heat transfer and pressure fluctuations. The flow exchange between the rotor disk cavity and the mainstream passage is mainly governed by the vane shock patterns. The purge flow leads to the appearance of a large coherent vortex structure on the suction side of the blade which enhances the overall heat transfer coefficient due to the blockage effect created. Secondly, the impact of the platform cooling is observed to be restricted to the platform, with negligible effects on the blade suction side. The platform cooling results in a clear attenuation of pressure pulsations at some specific locations. Finally the turbine performance was analyzed, comparing measured and CFD results. A detailed loss breakdown analysis has been done using correlations, in order to isolate the different loss component contributions. The presented results should help designers improve the protection of the rotor platform and minimize the amount of coolant used.

Improved Methodologies for Time Resolved Heat Transfer Measurements, Demonstrated on an Unshrouded Transonic Turbine Casing

2015 ◽  
Author(s):  
Matthew Collins ◽  
Kam Chana ◽  
Thomas Povey
Keyword(s):  
Heat Transfer ◽  
Wall Temperature ◽  
High Frequency ◽  
Processing Technique ◽  
Temperature Cycling ◽  
Test Facility ◽  
Inlet Temperature ◽  
Data Set ◽  
Time Resolved ◽  
Transonic Turbine

The HP rotor tip and over-tip casing are often life-limiting features in the turbine stages of current gas turbine engines. This is due to the high thermal load, and high temperature cycling both at low and high frequency. In the last few years there have been numerous studies of turbine tip heat transfer. Comparatively fewer studies have considered the over-tip casing heat transfer. This is in part, no doubt, due to the more onerous test facility requirements to validate computational simulations. Because the casing potential field is dominated by the passing rotor, to perform representative over-tip measurements a rotating experiment is an essential requirement. In this paper we describe improved methodologies for time resolved heat transfer measurements. Specifically we show that: 1. Changes in driving temperature (within limits) can be accounted for in both time-resolved and steady heat transfer measurement processing. This allows useful data to be extracted even under varying inlet temperature. 2. Superposition of several runs with different starting wall temperatures can be used to improve the accuracy of time resolved regressions by extending the wall temperature range over which the unsteady regressions are conducted. 3. A new time-resolved data processing technique that can be applied to data sets involving changes in wall temperature has been developed and is applied to experimental measurements to compute time resolved TAW and Nu. These improvements are demonstrated using unsteady heat transfer measurements conducted on the stationary casing above an unshrouded transonic turbine. The measurements were taken in the Oxford Turbine Research Facility (OTRF), an engine-scale rotating turbine facility which replicates engine-representative conditions of Mach number, Reynolds number, and gas-to-wall temperature ratio. High density arrays of miniature thin-film heat-flux gauges were used with a spatial resolution of 0.8 mm and temporal resolution of ∼120 kHz. The small size of the gauges, the high frequency response, and the improved processing methods allowed very detailed measurements of the heat transfer in this region. Time-resolved measurements of TAW and Nu are presented for the casing region (−30 % to +125% CAX) and compared to other results in the literature. The results provide an almost unique data set for calibrating CFD tools for heat transfer prediction in this highly unsteady environment dominated by the rotor over-tip flow.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part I: Experimental Configuration and Data Review With Inlet Temperature Profile Effects

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Harika S. Kahveci ◽  
Charles W. Haldeman ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Turbine Stage ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Pressure Surface ◽  
Cooling Flow ◽  
The Impact

This paper investigates the vane airfoil and inner endwall heat transfer for a full-scale turbine stage operating at design corrected conditions under the influence of different vane inlet temperature profiles and vane cooling flow rates. The turbine stage is a modern 3D design consisting of a cooled high-pressure vane, an un-cooled high-pressure rotor, and a low-pressure vane. Inlet temperature profiles (uniform, radial, and hot streaks) are created by a passive heat exchanger and can be made circumferentially uniform to within ±5% of the bulk average inlet temperature when desired. The high-pressure vane has full cooling coverage on both the airfoil surface and the inner and outer endwalls. Two circuits supply coolant to the vane, and a third circuit supplies coolant to the rotor purge cavity. All of the cooling circuits are independently controlled. Measurements are performed using double-sided heat-flux gauges located at four spans of the vane airfoil surface and throughout the inner endwall region. Analysis of the heat transfer measured for the uncooled downstream blade row has been reported previously. Part I of this paper describes the operating conditions and data reduction techniques utilized in this analysis, including a novel application of a traditional statistical method to assign confidence limits to measurements in the absence of repeat runs. The impact of Stanton number definition is discussed while analyzing inlet temperature profile shape effects. Comparison of the present data (Build 2) to the data obtained for an uncooled vane (Build 1) clearly illustrates the impact of the cooling flow and its relative effects on both the endwall and airfoils. Measurements obtained for the cooled hardware without cooling applied agree well with the solid airfoil for the airfoil pressure surface but not for the suction surface. Differences on the suction surface are due to flow being ingested on the pressure surface and reinjected on the suction surface when coolant is not supplied for Build 2. Part II of the paper continues this discussion by describing the influence of overall cooling level variation and the influence of the vane trailing edge cooling on the vane heat transfer measurements.

The Effects of Blade Passing on the Heat Transfer Coefficient of the Overtip Casing in a Transonic Turbine Stage

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
Steven J. Thorpe ◽  
Roger W. Ainsworth
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Turbine Stage ◽  
Transfer Coefficients ◽  
Time Resolved ◽  
Turbine Engine ◽  
Transonic Turbine

In a modern gas turbine engine, the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an overtip casing design that has an acceptable service life, it is essential for manufacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the overtip casing wall of a fully scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade tip is immediately above the point in question, while the lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the overtip casing wall. The data obtained from this fully scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade-style tests of similar high-pressure blade geometries.

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