Effects of Shock Wave Passage on Heat Transfer in a Transonic Turbine Cascade

1994 ◽  
Author(s):  
D. G. Holmberg ◽  
T. Reid ◽  
T. Kiss ◽  
H. L. Moses ◽  
W. F. Ng ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Shock Wave ◽  
Heat Flux ◽  
Leading Edge ◽  
Main Flow ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Blow Down ◽  
Long Run ◽  
Transonic Turbine

Results from a new facility for measuring heat transfer in transonic turbine cascades are repotted. An air heater has been built into the blow-down wind tunnel to heat the main flow for a 20 second run time. This allows control of the direction and magnitude of the heat transfer into the blade throughout the tests. A Heat Flux Microsensor was inserted into the blade to measure simultaneous surface heat flux and temperature. Measurements were made on the suction surface of the blades toward the trailing edge. Because of the long run times (20 s), the adiabatic wall temperature could be determined directly from the measured surface temperature and heat flux. Simultaneous pressure measurements were made with a Kulite transducer at the same distance from the leading edge to document shock passage. A separate shock tube was used to generate a shock wave which was introduced into the test section in front of the cascade. This shock was carried over the blade by the main flow. The resulting changes in heat flux correlated strongly with the unsteady pressure changes. An overall increase of 1.5 W/cm2 in heat flux was recorded for a pressure increase of 7 kPa during the initial passage of the shock.

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.

Numerical Study on Aero-Thermal Performance of HP Turbine Endwalls Under Influence of Hot Streak and High Mainstream Turbulence

2016 ◽  
Author(s):  
Zhiduo Wang ◽  
Wenhao Zhang ◽  
Zhaofang Liu ◽  
Chen Zhang ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Turbulence Intensity ◽  
Wall Temperature ◽  
Numerical Study ◽  
Leading Edge ◽  
Heat Transfer Characteristics ◽  
Adiabatic Wall ◽  
Transfer Characteristics ◽  
Hot Streak

In this paper, unsteady RANS simulations were performed at two hot streak (HS) circumferential positions with inlet turbulence intensity of 5% and 20%. The interacted HS and high mainstream turbulence effects on endwall heat transfer characteristics of a high-pressure (HP) turbine were discussed by analyzing the flow structures and presenting the endwall adiabatic wall temperature, heat transfer coefficient (HTC) and heat flux distributions. The results indicate that both the wall temperature and HTC increase with the turbulence intensity at most stator endwall regions. In addition, the increase of wall temperature plays a greater role than HTC of influencing the wall heat flux. However, higher turbulence intensity decreases the intensity of the stator passage horse-shoe vortex, also the corresponding region HTC and heat flux are reduced. In rotor passage, the variation of HS circumferential position would alter the hub and casing endwall temperature, however, the discrepancy is weakened at higher turbulence. The elevated HS attenuation at higher turbulence results in temperature augmentation at the leading edge of rotor hub and casing endwalls, while temperature decrease after 50% axial chord, thus obtains more uniform temperature distributions on the endwalls. However, the rotor endwall HTC is only augmented significantly at the leading edge on hub endwall, and pressure side and downstream of trailing edge on casing endwall. Variation of HTC and adiabatic wall temperature jointly determines the rotor hub and casing endwall heat flux, and the temperature variation has dominant effects in the most regions. In general, the variation of adiabatic wall temperature and HTC should be considered simultaneously when analyzing the turbine endwall heat transfer characteristics.

Influence of Mainstream Turbulence Intensity on Heat Transfer Characteristics of a High Pressure Turbine Stage With Inlet Hot Streak

2015 ◽  
Vol 138 (4) ◽  
Author(s):  
Zhiduo Wang ◽  
Zhaofang Liu ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Turbulence Intensity ◽  
Computational Study ◽  
Leading Edge ◽  
Adiabatic Wall ◽  
Airfoil Surface ◽  
Turbulence Decay ◽  
Hot Streak

An unsteady computational study was carried out on GE-E3 high pressure (HP) turbine at inflow turbulence intensities of 5%, 10%, and 20% accompanying with inlet hot streak (HS) at two circumferential positions (impinging and nonimpinging relative to vane leading edge) to analyze the interacted turbulence and HS influences. Turbulence decay mechanisms in turbine passage were presented, and the airfoil heat transfer behaviors were explored by means of adiabatic wall temperature, heat transfer coefficient (HTC), and wall heat flux. The results indicate that the elevated turbulence leads to favorable turbine airfoil temperature distributions, and turbulence induced HS attenuation mainly occurs in vane passage. In addition, the HS dispersion is related directly to the temperature gradients. Although the endwall temperature increases by more than 20 K (2.8% inlet mass-averaged temperature) and midregion temperature decreases by 16 K at blade leading edge, the hot region on blade pressure surface (PS) is only weakened by about 7 K, when turbulence intensity is increased from 5% to 20%. Higher turbulence significantly affects the airfoil HTC, excepting the regions secondary and leakage flow effects are dominating. Therefore, the tip and blade suction surface (SS) trailing edge heat flux is decreased for the temperature decline at higher turbulence, which is beneficial to tip cooling. HS position not only affects the airfoil surface temperature variations but also slightly affects the vane and blade midspan HTC for the variation of heat transfer driving temperature.

Unsteady Heat Transfer and Pressure Measurements on the Airfoils of a Rotating Transonic Turbine With Multiple Cooling Configurations

2016 ◽  
Author(s):  
Jeremy B. Nickol ◽  
Randall M. Mathison ◽  
Michael G. Dunn ◽  
Jong S. Liu ◽  
Malak F. Malak
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Film Cooling ◽  
Experimental Results ◽  
Blow Down ◽  
Pressure Transducers ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Transonic Turbine ◽  
Small Benefit

Measurements are presented for a high-pressure transonic turbine stage operating at design-corrected conditions with forward and aft purge flow and blade film cooling in a short-duration blow-down facility. Four different film-cooling configurations are investigated: simple cylindrical-shaped holes, diffusing fan-shaped holes, an advanced-shaped hole, and uncooled blades. A rainbow turbine approach is used so each of the four blade types comprise a wedge of the overall bladed disk and are investigated simultaneously at identical speed and vane exit conditions. Double-sided Kapton heat-flux gauges are installed at midspan on all three film-cooled blade types, and single-sided Pyrex heat-flux gauges are installed on the uncooled blades. Kulite pressure transducers are installed at midspan on cooled blades with round and fan-shaped cooling holes. Experimental results are presented both as time-averaged values and as time-accurate encoder-averages. In addition, the results of a steady RANS CFD computation are compared to the time-averaged data. The computational and experimental results show that the cooled blades reduce heat transfer into the blade significantly from the uncooled case, but the overall differences in heat transfer among the three cooling configurations is small. This challenges previous conclusions for simplified geometries that show shaped cooling holes outperforming cylindrical holes by a great margin. It suggests that the more complicated flow physics associated with an airfoil operating in an engine-representative environment reduces the effectiveness of the shaped cooling holes. The experimental results appear to show a small benefit to the advanced cooling holes, but this is on the order of the variation caused by changes in the alignment of heat-flux gauges with cooling holes.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part I: Influence of Cooling Variation

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Flows ◽  
Cooling Flow

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design-corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film-cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the uncooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton number distribution on the blade.

Effects of a Wake on Turbine Blade Heat Transfer in a Transonic Cascade

1997 ◽  
Author(s):  
James H. Hale ◽  
Thomas E. Diller ◽  
Wing F. Ng
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Surface ◽  
Maximum Increase ◽  
Pressure Surface ◽  
Transonic Cascade

The effects of a wake generated by a stationary upstream strut on surface beat transfer to turbine blades were measured experimentally in a heated, transonic cascade tunnel. Five pitchwise locations of the upstream strut were tested, while maintaining a constant axial distance between the strut and the leading edge plane of the cascade. Time-resolved unsteady heat flux measurements were made with Heat Flux Microsensors (HFM) at three positions on the suction surface and one position on the pressure surface. In addition, hot-wire surveys were taken along the leading edge plane of the cascade to document the disturbance generated by the upstream strut. Results from the hot-wire surveys show that with the strut placed upstream and near the stagnation point of the turbine blade, the turbulence intensity in the wake was as high as 50%. This high level of turbulence intensity was due to the coupling of the strut wake with the potential flow around the blunt leading edge of the turbine blade. A strong influence on the heat transfer coefficient was seen from the relative pitchwise position of the strut with respect to the leading edge of the test blades. For the suction surface, the maximum increase in average heat transfer coefficient occurred when the upstream strut was placed near the stagnation point of the blade. The heat transfer coefficients were increased by 15, 20, and 10% for the gages located on 10, 22, and 50% chord positions of the suction surface, respectively, compared to the baseline case of no strut. For the pressure side, results show the maximum increase in heat transfer coefficient occurred when the upstream strut was placed along the pitchline near the middle of the blade passage. At 30% chord position on the pressure surface, the heat transfer coefficient was increased by 25 %. Attempts to correlate these increases in mean heat transfer with integral values of the measured unsteadiness of the flow or heat flux were not successful.

Accuracy of Conventional Adiabatic Effectiveness and Heat Transfer Augmentation Factors in Predicting Heat Flux Into a Turbine Blade Leading Edge

2010 ◽  
Author(s):  
Laurene D. Dobrowolski ◽  
David G. Bogard ◽  
Silvia Ravelli
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Wall Temperature ◽  
Conjugate Heat Transfer ◽  
Density Ratio ◽  
Leading Edge ◽  
Adiabatic Wall ◽  
Stagnation Line ◽  
Conducting Wall ◽  
Heat Transfer Simulation

This paper focuses on the legitimacy of using conventional predictions, based on adiabatic wall temperature (Taw) and heat transfer coefficient (HTC) augmentation, of heat flux into a film cooled leading edge. To answer this question, the heat flux predicted using Taw was compared to the heat flux into a conducting leading edge. The study used numerical simulations with the k-ε turbulence model of FLUENT. The model simulated was a three-row leading edge with one row of holes on the stagnation line and two additional rows located at ±25°, which has been experimentally studied extensively. The adiabatic wall temperature was obtained from an adiabatic simulation. External heat transfer coefficients were determined from a constant heat flux simulation using a density ratio of DR = 1.0, as is commonly done experimentally. A conjugate heat transfer simulation was also run to give the surface temperature and heat flux into the conducting leading edge. Overall, the heat transfer was well predicted with the use of Taw and HTC augmentation. However, between the holes, conventional predictions of heat transfer were poor, with disparity up to 30% when compared with the conducting wall heat flux obtained from the conjugate heat transfer simulation. Thermal boundary layer profiles were used to understand the disparity between the heat fluxes obtained from the conventional prediction and the conducting wall simulation.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine: Part I—Influence of Cooling Variation

2010 ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Flows ◽  
Cooling Flow

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the un-cooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton Number distribution on the blade.

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

2021 ◽  
Author(s):  
Richard Celestina ◽  
Spencer Sperling ◽  
Louis Christensen ◽  
Randall Mathison ◽  
Hakan Aksoy ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Density ◽  
Unsteady Heat Transfer ◽  
Transonic Turbine

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.

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