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Adiabatic film cooling effectiveness was measured on the endwall of a turbine vane under transonic flow conditions using binary pressure-sensitive paint. The combined effect of upstream and passage film cooling holes was evaluated. The experiments used a five-vane annular cascade in a blowdown wind tunnel. The mainstream velocity was set to exit isentropic Mach numbers of 0.7 and 0.9. The total coolant-to-mainstream mass flow ratio (MFR) was varied from 0.75 to 2.50%. Coolant-to-mainstream density ratios (DRs) of 1.0 and 2.0 were investigated. The first row was located upstream of the vanes, and the second and third rows were located in the passage between the vanes. Coolant was supplied by a divided plenum, allowing independent control of the MFR to the upstream and the passage holes. The coolant MFR was evenly split between the upstream and passage holes. Generally, increasing the MFR, the DR, and the Mach number increased the film cooling effectiveness. However, increasing the DR at MFRs of 0.625% and below decreased the laterally averaged film cooling effectiveness by as much as 60%, largely because mainstream flow was ingested through the second row. Upstream coolant provided some protection through the passage, with effectiveness as high as 0.2 at the second row.

Film cooling was measured on the endwall of a five-vane annular cascade in a blowdown wind tunnel at an exit Mach number of 0.9. The adiabatic film cooling effectiveness was calculated from the partial pressure of oxygen measured with binary pressure sensitive paint (BPSP). Cylindrical film cooling holes were located in the upstream and passage regions with the coolant-to-mainstream mass flow ratio (MFR) independently varied for each region. One row was located upstream of the vanes and supplied by an upstream plenum. Two rows were located in the passage between two vanes and supplied by a downstream plenum. Three total MFRs were investigated: 1%, 1.5%, and 2%. For a given total MFR, four combinations of upstream and downstream MFRs were compared to an even split of coolant. Coolant-to-mainstream density ratios (DRs) of 1.0 and 2.0 were investigated. The most efficient use of coolant hinged on balancing the downstream MFR for the second row due to the endwall pressure gradient preventing coolant from exiting the holes or a high jet velocity causing liftoff. For this row, selecting the optimum MFR increased the area-averaged film cooling effectiveness by up to 200% with a reduction in row 1 of less than 25%. At high downstream MFRs, increasing the density ratio delayed liftoff and increased film cooling effectiveness in row 2 by 65%. However, at low MFRs, increasing the density ratio reduced film cooling effectiveness in row 2 by 60%.

The effects of mainstream flow velocity, density ratio (DR), and coolant-to-mainstream mass flow ratio (MFR) were investigated on a vane endwall in a transonic, annular cascade. A blow down facility consisting of five vanes was used. The film cooling effectiveness was measured using binary pressure sensitive paint (BPSP). The mainstream flow was set using isentropic exit Mach numbers of 0.7 and 0.9. The coolant-to-mainstream density ratio varied from 1.0 to 2.0. The coolant to mainstream MFR varied from 0.75% to 1.25%. The endwall was cooled by eighteen discrete holes located upstream of the vane passage to provide cooling to the upstream half of the endwall. Due to the curvature of the vane endwall, the upstream holes provided uniform coverage entering the endwall passage. The coverage was effective leading to the throat of the passage, where the downstream holes could provide additional protection. Increasing the coolant flowrate increased the effectiveness provided by the film cooling holes. Increasing the density of the coolant increases the effectiveness on the endwall while enhancing the lateral spread of the coolant. Finally, increasing the velocity of the mainstream while holding the MFR constant also yields increased protection on the endwall. Over the range of flow conditions considered in this study, the binary pressure sensitive paint proved to be a valuable tool for obtaining detailed pressure and film effectiveness distributions.

This work focuses on the parametric study of film cooling effectiveness on the turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5, to 2.0. There are eight cylindrical holes on the endwall inside the passage. The pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall crossflow passage.

This work focuses on the parametric experimental study of film cooling effectiveness on the suction side of a scaled turbine vane. The experiments are performed in a five-vane annular sector cascad...

This work focuses on the parametric study of film cooling effectiveness on turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5 to 2.0. There are 8 cylindrical holes on the endwall inside the passage. Pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall cross-flow passage.

The effect of film cooling hole inlet geometry is experimentally investigated in this study. Detailed film cooling effectiveness distributions are obtained on a flat plate using Pressure Sensitive Paint (PSP). The inlet of a traditional 12°-12°-12°, laidback, fanshaped hole varies from a traditional “round” opening to an oblong, racetrack shaped opening. In this study, a single racetrack inlet with an aspect ratio of 2:1 is compared to the round inlet. For both designs, the holes are inclined at θ = 30° relative to the mainstream. Blowing ratios of 0.5, 1.0, and 1.5 are considered as the coolant–to–mainstream density ratio varies between 1.0 and 4.0. For all cases, the freestream turbulence intensity is maintained at 7.5%. With the introduction of the racetrack shaped inlet, the coolant spreads laterally across the diffuse, laidback fanshaped outlet. The centerline film cooling effectiveness is reduced with the enhanced lateral spread of the coolant. However, the benefit of the shaped inlet is also observed with an increase in the area averaged film cooling effectiveness, compared to the traditional round inlet. Not only does the shaped inlet promote spreading of the coolant, it is also believed the racetrack shape suppresses turbulence within the hole allowing for enhanced film cooling protection near the film cooling holes.

This work focuses on the parametric experimental study of film cooling effectiveness on the suction side of a scaled turbine vane under transonic flow condition. The experiments were performed in a five-vane annular sector cascade blowdown facility. The controlled exit Mach numbers were 0.7, 0.9, and 1.1, from high subsonic to transonic conditions. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of coolant-to-mainstream density ratios, ranging from 1.0, 1.5 to 2.0. Three row-averaged coolant-to-mainstream blowing ratios in the range 0.7, 1.0, and 1.6 are studied. The test vane includes three rows of radial-angle cylindrical holes around the leading edge and two rows of compound-angle shaped holes on the suction side. All the cooling holes are active in order to study the resultant film cooling on suction side as well as from leading edge. Pressure sensitive paint (PSP) technique was used to obtain the film cooling effectiveness distributions from suction side holes and the contribution from leading edge showerhead holes. This work shows the effects of coolant-to-mainstream blowing ratio, density ratio, and exit Mach number on the film cooling effectiveness as well as its interaction with a potential shock wave. The results indicate that when the cooling holes are located in a critical region on the vane suction surface, the parametric effect on film cooling performance will significantly deviate from the common trend for a typical hole geometry.

Hot section turbine engine components are often cooled through the use of a cool film of air on the component wall. The source of the air used for film cooling comes from the compressor of the gas turbine engine and may be 800°C, or more, cooler than the hot gas path air. The temperature differential between the hot mainstream gas and the film coolant results in a large difference in density between the two gases. In order to investigate the effect of high density ratios on film cooling performance, a traditional, round hole (θ = 30°) and a laidback, fan shaped hole (θ = 30°, α = γ = 10°) were observed using Stereo-Particle Image Velocimetry (S-PIV). Flowfield measurements were performed on various planes downstream of the film cooling hole (x/d = 0, 1, 3 and 10 for the round hole and x/d = 0, 3, and 10 for the shaped hole). At each location the coolant-to-mainstream interaction was captured at multiple density ratios (DR = 1, 2, 3, 4) and blowing ratios (M = 0.5, 1.0, 1.5). Using S-PIV, the three-dimensional flow field was measured. Distributions of the flow vorticity were derived from the high speed velocity measurements taken during S-PIV testing. For the simple angle, round holes, the results show at the elevated density ratios, the coolant spreads laterally near the hole; while at DR = 1, the coolant trace is limited to the width of the film cooling hole. Furthermore, as the cooling jet exits from the round hole, the vorticity within the jet is very strong, leading to increased mixing with the mainstream. However, as the density ratio increases (at a given blowing ratio), this mixing was reduced. For a given flow condition, the rotation was reduced with the jet exiting the shaped hole (compared to the round hole), and this led to enhanced protection on the surface. While investigating both round and shaped holes, it was shown the S-PIV method is a valuable tool to observe and quantify the jet–to–mainstream interactions near the film cooled surface.

The combined effects of a favorable, mainstream pressure gradient and coolant-to-mainstream density ratio have been investigated. Detailed film cooling effectiveness distributions have been obtained on a flat plate with either cylindrical (θ = 30 deg) or laidback, fan-shaped holes (θ = 30 deg and β = γ = 10 deg) using the pressure-sensitive paint (PSP) technique. In a low-speed wind tunnel, both nonaccelerating and accelerating flows were considered, while the density ratio varied from 1 to 4. In addition, the effect of blowing ratio was considered, with this ratio varying from 0.5 to 1.5. The film produced by the shaped hole outperformed the round hole under the presence of a favorable pressure gradient for all the blowing and density ratios. At the lowest blowing ratio, in the absence of freestream acceleration, the round holes outperformed the shaped holes. However, as the blowing ratio increases, the shaped holes prevent lift-off of the coolant and offer enhanced protection. The effectiveness afforded by both the cylindrical and shaped holes, with and without freestream acceleration, increased with density ratio.

This paper experimentally studies the effect of transonic flow on local heat transfer coefficients and film cooling effectiveness distributions of a turbine vane’s suction surface with compound-angle shaped-hole configuration. A Temperature Sensitive Paint (TSP) method is used to determine the local heat transfer coefficients and film cooling effectiveness simultaneously. Tests were performed in a five-vane annular-sector cascade blow-down facility. The exit Mach numbers are controlled to be 0.7 and 0.9, from subsonic to transonic conditions. Compressed air is used as coolant with a coolant-to-mainstream density ratio 0.91 on film cooling and heat transfer study. Three averaged coolant-to-mainstream blowing ratios in the range, 0.7, 1.0, and 1.6 are investigated. The test vane features three rows of radial-angle cylindrical holes around the leading edge, and two rows of compound-angle shaped holes on the suction side. Effects of blowing ratio and exit Mach number on the vane suction surface heat transfer and film cooling effectiveness distributions are obtained, and the results are presented and explained in this investigation.

This paper studies the effect of transonic flow velocity on local film cooling effectiveness distribution of turbine vane suction side, experimentally. A conduction-free Pressure Sensitive Paint (PSP) method is used to determine the local film cooling effectiveness. Tests were performed in a five-vane annular cascade at Texas A&M Turbomachinery laboratory blow-down flow loop facility. The exit Mach numbers are controlled to be 0.7, 0.9, and 1.1, from subsonic to transonic flow conditions. Three foreign gases N2, CO2 and Argon/SF6 mixture are selected to study the effects of three coolant-to-mainstream density ratios, 1.0, 1.5, and 2.0 on film cooling. Four averaged coolant blowing ratios in the range, 0.7, 1.0, 1.3 and 1.6 are investigated. The test vane features 3 rows of radial-angle cylindrical holes around the leading edge, and 2 rows of compound-angle shaped holes on the suction side. Results suggest that the PSP technique is capable of producing clear and detailed film cooling effectiveness contours at transonic condition. The effects of coolant to mainstream blowing ratio, density ratio, and exit Mach number on the vane suction-surface film cooling distribution are obtained, and the consequence results are presented and explained in this investigation.Copyright © 2015 by ASME

The amount of cooling air assigned to seal high pressure turbine (HPT) rim cavities is critical for performance as well as component life. Insufficient air leads to excessive hot annulus gas ingestion and its penetration deep into the cavity compromising disk or cover plate life. Excessive purge air, on the other hand, adversely affects performance. Experiments on a rotating turbine stage rig which included a rotor–stator forward disk cavity were performed at Arizona State University (ASU). The turbine rig has 22 vanes and 28 blades, while the cavity is composed of a single-tooth lab seal and a rim platform overlap seal. Time-averaged static pressures were measured in the gas path and the cavity, while mainstream gas ingestion into the cavity was determined by measuring the concentration distribution of tracer gas (carbon dioxide) under a range of purge flows from 0.435% (Cw = 1540) to 1.74% (Cw = 6161). Additionally, particle image velocimetry (PIV) was used to measure fluid velocity inside the cavity between the lab seal and the rim seal. The data from the experiments were compared to time-dependent computational fluid dynamics (CFD) simulations using fluent CFD software. The CFD simulations brought to light the unsteadiness present in the flow during the experiment which the slower response data did not fully capture. An unsteady Reynolds averaged Navier–Stokes (RANS), 360-deg CFD model of the complete turbine stage was employed in order to increase the understanding of the swirl physics which dominate cavity flows and better predict rim seal ingestion. Although the rotor–stator cavity is geometrically axisymmetric, it was found that the interaction between swirling flows in the cavity and swirling flows in the gas path create nonperiodic/time-dependent unstable flow patterns which at the present are not accurately modeled by a 360 deg full stage unsteady analysis. At low purge flow conditions, the vortices that form inside the cavities are greatly influenced by mainstream ingestion. Conversely at high purge flow conditions the vortices are influenced by the purge flow, therefore ingestion is minimized. The paper also discusses details of meshing, convergence of time-dependent CFD simulations, and recommendations for future simulations in a rotor–stator disk cavity such as assessing the observed unsteadiness in the frequency domain in order to identify any critical frequencies driving the system.

A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value including totally off. A wide range of stage pressure ratios, coolant to freestream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3% to 4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.

A TOBI (tangential on board injection), or preswirl, system is a critical component of a high pressure turbine cooling delivery system. Its efficient performance and characterization are critical because the blade and disk life depend on the accuracy of delivering the required flow at the correct temperature and pressure. This paper presents a TOBI flow discharge coefficient validation process applied to a low radius radial configuration starting from a 1dimensional (1D) network flow analysis to a 3 dimensional (3D) frozen rotor computational fluid dynamics (CFD) analysis of the rotor cooling air delivery system. The analysis domain commences in the combustor plenum stationary reference frame, includes the TOBI, transitions to the rotating reference frame as the flow travels through the rotating cover plate orifice, continues up the turbine disk into the slot bottom blade feed cavity, and terminates in the turbine blade. The present effort includes matching a 1D network model with 3D CFD results using simultaneous goal-matching of the pressure predictions throughout the circuit, defining test rig pressure measurements at critical “nondisturbing” locations for quanification of pressure ratio across the TOBI, and finally comparing the TOBI flow coefficient resulting from stationary cold flow tests with what was obtained from the 3D CFD results. An analysis of the results indicates that the discharge coefficient varies with the pressure ratio and that the traditional method of using a constant discharge coefficient extracted from a cold flow test run under choked conditions leads to overpredicting turbine cooling flows. The TOBI flow coefficient prediction for the present study compares well with the stationary data published by otherresearchers for the configuration under investigation and the process described in this paper is general for any TOBI configuration.

Numerical calculations are performed to examine the merits of a novel cooling scheme proposed to improve the durability of heat shields in a gas turbine combustor. The novel scheme is comprised of a combination of convective and impingement heat transfer. The impingement scheme incorporates three rows of holes two of which target the bore of the heat-shield while the third row targets the OD of the heat-shield and are all staggered with respect to each other. The convective heat transfer is provided through annular collection of the spent impingement flow from the two rows targeting the bore. The impingement height-to-hole diameter spacing is in the range of 1.4 ≤ Z/d ≤ 2.3, with the jet Reynolds number 12050 ≤ Red ≤ 13770. The spent impingement flow from the inner-diameter holes however, acts as a cross-flow on the impingement jets targeting the outer diameter of the heat-shield which reduces the impingement heat transfer rate of these outer-diameter holes. U-shaped ribs are introduced on the surface of heat-shield with the objective of first reducing the detrimental effects of cross-flow on impingement heat transfer and secondly, to increase the convective heat transfer area due to annular collection of the spent flow. It is shown that, provided the U-ribs are clocked with respect to the 3rd (outer-diameter) impingement row, the maximum heat-shield metal temperature is reduced by as much as 80 °C. The thermal gradients in both radial and “through-the-wall” directions are also reduced which in turn reduce the hoop (tangential) stresses. The combination of heat-shield temperature and stress reduction, due to the introduction of U-ribs, increase the heat-shield life.