Heat-Flux Measurements for the Rotor of a Full-Stage Turbine: Part I—Time-Averaged Results

1986 ◽  
Vol 108 (1) ◽  
pp. 90-97 ◽  
Author(s):  
M. G. Dunn
Keyword(s):  
Heat Flux ◽  
Flat Plate ◽  
Guide Vane ◽  
Flow Direction ◽  
Leading Edge ◽  
Present Calculation ◽  
Suction Surface ◽  
Pressure Surface ◽  
Good Comparison ◽  
Flux Measurements

This paper describes time-averaged heat-flux distributions obtained for the blade of a Garrett TFE 731-2 hp full-stage rotating turbine. Blade measurements were obtained both with and without injection. The injected gas was supplied from a separate reservoir and was directed into the turbine gas path via nozzle guide vane (NGV) pressure surface slots located at approximately 63 percent of the wetted distance. Blade heat-flux measurements were performed for two different injection gas temperatures, Tc/T0 = 0.53 and Tc/T0 = 0.82. A shock tube is used as a short-duration source of heated air to which the turbine is subjected and thin-film gages are used to obtain the heat-flux measurements. Results are presented along the blade in the flow direction at 10, 50, and 90 percent span for both the pressure and suction surfaces. A sufficient number of measurements were obtained to also present span-wise distributions. At approximately the 50 percent span location, two contoured inserts containing closely spaced gages were installed in the blade so that the leading-edge region distribution could be resolved in detail. The blade results are compared with predictions obtained using a flat-plate technique and with predictions obtained using a version of STAN 5. The results suggest that: (1) The suction surface laminar flat-plate prediction is in reasonable agreement with the data from the stagnation point up to approximately 10 percent of the wetted distance. Beyond 10 percent, the laminar prediction falls far below the data and the turbulent flat-plate prediction falls above the data by about 60 percent. The laminar portion of the STAN 5 prediction as configured for the present calculation does not provide good comparison with the data. However, the turbulent flat-plate boundary-layer portion of STAN 5 does provide reasonably good comparison with the data. On the pressure surface, the turbulent flat-plate prediction is in good agreement with the data, but the laminar flat-plate and the STAN 5 predictions fall far low. (2) The influence of upstream NGV injection is to significantly increase the local blade heat flux in the immediate vicinity of the leading edge; i.e., up to 20 percent wetted distance on the suction surface and up to 10 percent on the pressure surface. (3) The effect on local heat flux of increasing the coolant-gas temperature was generally less than 10 percent.

Time-Averaged Heat-Flux Distributions and Comparison With Prediction for the Teledyne 702 HP Turbine Stage

1988 ◽  
Vol 110 (1) ◽  
pp. 51-56 ◽  
Author(s):  
M. G. Dunn ◽  
R. E. Chupp
Keyword(s):  
Thin Film ◽  
Heat Flux ◽  
Shock Tube ◽  
Flat Plate ◽  
Leading Edge ◽  
Turbine Stage ◽  
Heated Air ◽  
Edge Distribution ◽  
Flux Measurements ◽  
Good Agreement

Time-averaged heat-flux distributions are reported for the vane and blade of the Teledyne CAE 702 HP full-stage rotating turbine. A shock tube is used as a short-duration source of heated air to which the turbine is subjected and thin-film gages are used to obtain the heat-flux measurements. The thin-film gages were concentrated on the midspan region from the leading edge to near the trailing edge. The blade contained two contoured inserts wtih gages spaced very close together so that the leading edge distribution could be resolved. The NGV and blade results are compared with predictions obtained using a flat-plate technique, an eddy-diffusing model (STAN 5), and a k–ε model. The results of the comparison between data and prediction suggest that: (a) first, the vane data are bounded by the turbulent flat plate and the fully turbulent STAN 5 prediction. For the vane, the k–ε prediction is in relatively good agreement with the STAN 5 prediction and (b) secondly, the blade data are acceptably predicted by the k–ε prediction on both the pressure and the suction surfaces. The STAN 5 fully turbulent calculation for the blade falls above the data (essentially in agreement with the turbulent flat-plate calculation) and the STAN 5 fully laminar falls substantially below the data. With the exception of the pressure loadings and the geometry, the code inputs used for these predictions were identical to those previously used to predict the Garrett TFE 731-2 HP turbine and the Garrett LART HP turbine.

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.

Measurement and Analyses of Heat Flux Data in a Turbine Stage: Part II—Discussion of Results and Comparison With Predictions

1984 ◽  
Vol 106 (1) ◽  
pp. 234-240 ◽  
Author(s):  
M. G. Dunn ◽  
W. J. Rae ◽  
J. L. Holt
Keyword(s):  
Boundary Layer ◽  
Heat Flux ◽  
Flat Plate ◽  
Wall Temperature ◽  
Rotor Blade ◽  
Guide Vane ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Pressure Surface ◽  
Experimental Apparatus

Parts I and II of this paper report the measurement and analysis of detailed heat flux distributions obtained on the nozzle guide vane (NGV), rotor, and shroud of the full-stage rotating turbine of the Garrett TFE 731-2 engine. Part I is devoted to a description of the experimental apparatus and the data analysis procedure. Part II is concerned with the experimental results and their comparison with predictions obtained using a flat-plate technique and with predictions obtained using STAN5. Measurements were performed for values of the ratio of wall temperature to total temperature equal to 0.21, 0.33, and 0.53. For the NGV airfoil and rotor blade, the influence of wall temperature ratio on the Stanton number is shown to be generally 10 percent or less. For the NGV airfoil, and the rotor blade pressure surface, the flat-plate turbulent-flow prediction provides a reasonably good comparison with the data. With the rotor operating, the state of the boundary layer generally appears to be turbulent for both the NGV airfoil and the rotor blade. Predictions obtained using STAN5 confirm that boundary layer transition occurs at a value of Reθ substantially less than 200.

Film Cooling Performance of Tripod Antivortex Injection Holes Over the Pressure and Suction Surfaces of a Nozzle Guide Vane

2017 ◽  
Vol 9 (2) ◽  
Author(s):  
Sridharan Ramesh ◽  
Christopher LeBlanc ◽  
Diganta Narzary ◽  
Srinath Ekkad ◽  
Mary Anne Alvin
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Nozzle Guide

Film cooling performance of the antivortex (AV) hole has been well documented for a flat plate. The goal of this study is to evaluate the same over an airfoil at three different locations: leading edge suction and pressure surface and midchord suction surface. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low-speed linear cascade wind tunnel. Steady-state infrared (IR) technique was employed to measure the adiabatic film cooling effectiveness. The study has been divided into two parts: the initial part focuses on the performance of the antivortex tripod hole compared to the cylindrical (CY) hole on the leading edge. Effects of blowing ratio (BR) and density ratio (DR) on the performance of cooling holes are studied here. Results show that the tripod hole clearly provides higher film cooling effectiveness than the baseline cylindrical hole case with overall reduced coolant usage on the both pressure and suction sides of the airfoil. The second part of the study focuses on evaluating the performance on the midchord suction surface. While the hole designs studied in the first part were retained as baseline cases, two additional geometries were also tested. These include cylindrical and tripod holes with shaped (SH) exits. Film cooling effectiveness was found at four different blowing ratios. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations, thus providing cooling in the important trailing edge portion of the airfoil.

Numerical Simulation of Clocking Effect on Wake Transportation and Interaction in a 1.5-Stage Axial Turbine

2010 ◽  
Author(s):  
Wei Li ◽  
Hua Ouyang ◽  
Zhao-hui Du
Keyword(s):  
Stokes Equations ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Maximum Efficiency ◽  
Suction Surface ◽  
Navier Stokes Equations ◽  
Pressure Surface ◽  
Turbine Efficiency ◽  
Small Influence

To give insight into the clocking effect and its influence on the wake transportation and its interaction, the unsteady three-dimensional flow through a 1.5-stage axial low pressure turbine is simulated numerically using a density-correction based, Reynolds-Averaged Navier-Stokes equations commercial CFD code. The 2nd stator clocking is applied over ten equal tangential positions. The results show that the harmonic blade number ratio is an important factor affecting the clocking effect. The clocking effect has a very small influence on the turbine efficiency in this investigation. The efficiency difference between the maximum and minimum configuration is nearly 0.1%. The maximum efficiency can be achieved when the 1st stator wake enters the 2nd stator passage near blade suction surface and its adjacent wake passes through the 2nd stator passage close to blade pressure surface. The minimum efficiency appears if the 1st stator wake impinges upon the leading edge of the 2nd stator and its adjacent wake of the 1st stator passed through the mid-channel in the 2nd stator.

Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First Stage Vane

2004 ◽  
Author(s):  
D. G. Knost ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Extreme Environment ◽  
Wall Region ◽  
Pressure Side ◽  
Stagnation Region ◽  
Pressure Surface ◽  
Adiabatic Effectiveness ◽  
Endwall Film Cooling

In gas turbine development, the direction has been towards higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

Study on Application of CAE in a Centrifugal Compressor Impeller

2013 ◽  
Vol 787 ◽  
pp. 594-599 ◽  
Author(s):  
Bao Long Gong ◽  
Xiu Jie Jia ◽  
Guang Cun Wang ◽  
Zi Wu Liu
Keyword(s):  
Centrifugal Compressor ◽  
Equivalent Stress ◽  
Leading Edge ◽  
Middle Part ◽  
Turbulent Intensity ◽  
Erosion Wear ◽  
Suction Surface ◽  
Trailing Edge ◽  
Wear Area ◽  
Pressure Surface

CAE technology is the most common method to study properties of impeller in a centrifugal compressor. The fluid field was numerically simulated by CFX program to obtain the distribution rules of pressure, turbulence intensity and erosion wear. Based on fluid-solid interaction, stress and deformation were analyzed by Ansys program. According to the simulation results, the maximum deformation and equivalent stress of the impeller are all located on the junction between the blade trailing edge and the shroud. The most serious damaged part by erosion wear in impeller is on the pressure surface of long blade. The erosion wear area on blade pressure surface caused by particles impact primarily locates on blade trailing edge root and middle part. In the flow field, the turbulent intensity on suction surface is greater than that on pressure surface in the corresponding position and the greatest turbulent intensity is located on the leading edge of suction surface. There is backflow phenomenon around the suction surface of long blade and the short blade has significant effect to reduce backflow. The results of numerical simulation explain some actual impeller failure cases and can be applied to anti-wear impeller design and repair.

Numerical Investigation of the Effect of Inlet Flow Angle on Film Cooling Performance for the Blade Tip With Suction Surface Squealer Rim

2019 ◽  
Author(s):  
Zhiqiang Yu ◽  
Jianjun Liu ◽  
Chen Li ◽  
Baitao An
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Flow Ratio ◽  
Suction Surface ◽  
Cooling Performance ◽  
Tip Leakage ◽  
Pressure Surface ◽  
Mass Flow Ratio

Abstract Numerical investigations have been performed to study the effect of incidence angle on the aerodynamic and film cooling performance for the suction surface squealer tip with different film-hole arrangements at τ = 1.5% and BR = 1.0. Meanwhile, the full squealer tip as baseline is also investigated. Three incidence angles at design condition (0 deg) and off-design conditions (± 7 deg) are investigated. The suction surface, pressure surface, and the camber line have seven holes each, with an extra hole right at the leading edge. The Mach number at the cascade inlet and outlet are 0.24 and 0.52, respectively. The results show that the incidence angle has a significant effect on the tip leakage flow characteristics and coolant flow direction. The film cooling effectiveness distribution is altered, especially for the film holes near the leading edge. When the incidence angle changes from +7 deg to 0 and −7 deg, the ‘re-attachment line’ moves downstream and the total tip leakage mass flow ratio decreases, but the suction surface tip leakage mass flow ratio near leading edge increases. In general, the total tip leakage mass flow ratio for suction surface squealer tip is 1% greater than that for full squealer tip at the same incidence angle. The total pressure loss coefficient of suction surface squealer tip is larger than that for full squealer tip. The full squealer tip with film holes near suction surface and the suction surface squealer tip with film hole along camber line show high film cooling performance, and the area averaged film cooling effectiveness at positive incidence angle +7 deg is higher than that at 0 and −7 deg. The coolant discharged from film holes near pressure surface only cools narrow region near pressure surface.

Investigation of the separation bubble formed behind the sharp leading edge of a flat plate at incidence

2000 ◽  
Vol 214 (3) ◽  
pp. 157-176 ◽  
Author(s):  
M J Crompton ◽  
R V Barrett
Keyword(s):  
Reynolds Number ◽  
Flat Plate ◽  
Laser Doppler Anemometry ◽  
Reverse Flow ◽  
Separation Bubble ◽  
Flow Direction ◽  
Leading Edge ◽  
Laser Doppler ◽  
Static Pressure Distribution ◽  
Sharp Leading Edge

Detailed measurements of the separation bubble formed behind the sharp leading edge of a flat plate at low speeds and incidence are reported. The Reynolds number based on chord length ranged from 0.1 × 105 to 5.5 × 105. Extensive use of laser Doppler anemometry allowed detailed velocity measurements throughout the bubble. The particular advantages of laser Doppler anemometry in this application were its ability to define flow direction without ambiguity and its non-intrusiveness. It allowed the mean reattachment point to be accurately determined. The static pressure distribution along the plate was also measured. The length of the separation bubble was primarily determined by the plate incidence, although small variations occurred with Reynolds number because of its influence on the rate of entrainment and growth of the shear layer. Above about 105, the Reynolds number effect was no longer evident. The reverse flow boundary layer in the bubble exhibited signs of periodic stabilization before separating close to the leading edge, forming a small secondary bubble rotating in the opposite sense to the main bubble.

Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First-Stage Vane

2005 ◽  
Vol 127 (2) ◽  
pp. 297-305 ◽  
Author(s):  
D. G. Knost ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Extreme Environment ◽  
Wall Region ◽  
Pressure Side ◽  
Stagnation Region ◽  
Pressure Surface ◽  
Adiabatic Effectiveness ◽  
Endwall Film Cooling

In gas turbine development, the direction has been toward higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first-stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool, including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

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