Detailed Film Cooling Effectiveness Distributions Research Articles

TPIV Experimental Investigation of Film Coolant-to-Mainstream Interaction From Shaped Cooling Holes With Various Inlet Geometries

Abstract In this investigation, multiple sets of time-averaged tomographic particle imaging velocimetry measurements are completed for laid-back, fan-shaped film cooling holes with “racetrack” shaped inlets. Traditional 10-10-10, laid-back, fan-shaped holes are inclined 30 deg to the mainstream flow on a flat plate. The inlet cross section varies from round to two elongated racetrack shapes. The cross-sectional area and the outlet-to-inlet area ratio for all the geometries are held constant. The flat plate is installed in a low-speed wind tunnel with a mainstream turbulence intensity of 8% and an average velocity of 21.6 m/s. The blowing ratios of the film jets range from 0.6 to 1.5 and the density ratio is 1. The Reynolds number of the cooling jet varies from 2600 to 8400. The characteristics of the resulting flowfield are coupled with the detailed film cooling effectiveness distributions. It can be noted from the results that the counter-rotating vortex pair generated by the 2:1 inlet is the closest to the surface and weakest in strength, likely caused by the minimum peak jet momentum of the three. The Reynolds stresses downstream of the 2:1 and 4:1 inlets are significantly lower than those downstream of the shaped hole with a round inlet. An inverse relation between volumetric turbulence accumulation (TA), and surface effectiveness (η), can be correlated for the blowing ratios considered. The turbulence accumulation term can thus be used to evaluate the performance of a film cooling hole design with flowfield data.

Transonic Turbine Vane Endwall Film Cooling Using the Pressure-Sensitive Paint Measurement Technique

Abstract 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.

Film Cooling Effectiveness From Two Rows of Compound Angled Cylindrical Holes Using Pressure-Sensitive Paint Technique

This study investigates the effects of blowing ratio, density ratio, and spanwise pitch on the flat plate film cooling from two rows of compound angled cylindrical holes. Two arrangements of two-row compound angled cylindrical holes are tested: (a) the first row and the second row are oriented in staggered and same compound angled direction (β = +45 deg for the first row and +45 deg for the second row); (b) the first row and the second row are oriented in inline and opposite direction (β = +45 deg for the first row and −45 deg for the second row). The cooling hole is 4 mm in diameter with an inclined angle of 30 deg. The streamwise row-to-row spacing is fixed at 3d, and the spanwise hole-to-hole (p) is varying from 4d, 6d to 8d for both designs. The film cooling effectiveness measurements were performed in a low-speed wind tunnel in which the turbulence intensity is kept at 6%. There are 36 cases for each design including four blowing ratios (M = 0.5, 1.0, 1.5, and 2.0), three density ratios (DR = 1.0, 1.5, and 2.0), and three hole-to-hole spacing (p/d = 4, 6, and 8). The detailed film cooling effectiveness distributions were obtained by using the steady-state pressure-sensitive paint (PSP) technique. The spanwise-averaged cooling effectiveness are compared over the range of flow parameters. Some interesting observations are discovered including blowing ratio effect strongly depending on geometric design; staggered arrangement of the hole with same orientation does not yield better effectiveness at higher blowing ratio. Currently, film cooling effectiveness correlation of two-row compound angled cylindrical holes is not available, so this study developed the correlations for the inline arrangement of holes with opposing angles and the staggered arrangement of holes with same angles. The results and correlations are expected to provide useful information for the two-row flat plate film cooling analysis.

Combined Effects of Freestream Pressure Gradient and Density Ratio on the Film Cooling Effectiveness of Round and Shaped Holes on a Flat Plate

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.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure-sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5% and 0.75% to 1% for the purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Influence of Unsteady Wake With Trailing Edge Coolant Ejection on Turbine Blade Film Cooling

Detailed film cooling effectiveness distributions along a modeled turbine rotor blade under the combined effects of an upstream trailing edge unsteady wake with coolant ejection are presented using the pressure sensitive paint (PSP) mass transfer analogy method. The experiment is conducted in a low speed wind tunnel facility with a five blade linear cascade. The exit Reynolds number based on the axial chord is 370,000. Unsteady wakes and trailing edge coolant jets are produced by a spoked wheel-type wake generator with hollow rods equipped with several coolant ejections from holes. The coolant-to-mainstream density ratios for both the blade and trailing edge coolant ejection range from 1.5 to 2.0 for simulating realistic engine conditions. Blade blowing ratio studies are 0.5 and 1.0 on the suction surface and 1.0 and 2.0 on the pressure surface. The trailing edge jet blowing ratio and Strouhal numbers are 1.0 and 0.12, respectively. The results show that the unsteady wake reduces the overall film cooling effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances the overall effectiveness. The results also show that the overall filming cooling effectiveness increases by using heavier coolant for trailing edge ejection and for blade surface film cooling.

Effect of Freestream Turbulence Intensity on Film Cooling Jet Structure and Surface Effectiveness Using PIV and PSP

An experimental investigation of film cooling jet structure using two-dimensional particle image velocimetry (PIV) has been completed for cylindrical, simple angle (θ=35 deg) film cooling holes. The PIV measurements are coupled with detailed film cooling effectiveness distributions on the flat plate obtained using a steady state, pressure sensitive paint (PSP) technique. Both the flow and surface measurements were performed in a low speed wind tunnel where the freestream turbulence intensity was varied from 1.2% to 12.5%. With this traditional film cooling configuration, the blowing ratio was varied from 0.5 to 1.5 to compare the jet structure of relatively low and high momentum cooling flows. Velocity maps of the coolant flow (in the streamwise direction) are obtained on three planes spanning a single hole: centerline, 0.25D, and 0.5D (outer edge of the film cooling hole). From the seeded jets, time averaged, mean velocity distributions of the film cooling jets are obtained near the cooled surface. In addition, turbulent fluctuations are obtained for each flow condition. Combining the detailed flow field measurements obtained using PIV (both instantaneous and time averaged) with detailed film cooling effectiveness distributions on the surface (PSP) provides a more complete view of the coolant jet-mainstream flow interaction. Near the edge of the film cooling holes, the turbulent mixing increases, and as a result the film cooling effectiveness decreases. Furthermore, the PIV measurements show the increased mixing of the coolant jet with the mainstream at the elevated freestream turbulence level resulting in a reduction in the jet to effectively protect the film cooled surface.

Effect of Density Ratio on Flat Plate Film Cooling With Shaped Holes Using PSP

Detailed film-cooling effectiveness distributions are obtained on a flat plate using the pressure sensitive paint (PSP) technique. The applicability of the PSP technique is expanded to include a coolant-to-mainstream density ratio of 1.4. The effect of density ratio on the film-cooling effectiveness is coupled with varying blowing ratio (M=0.25–2.0), freestream turbulence intensity (Tu=1–12.5%), and film hole geometry. The effectiveness distributions are obtained on three separate flat plates containing either simple angle, cylindrical holes, simple angle, fanshaped holes (α=10 deg), or simple angle, laidback, fanshaped holes (α=10 deg and γ=10 deg). In all three cases, the film-cooling holes are angled at θ=35 deg from the mainstream flow. Using the PSP technique, the combined effects of blowing ratio, turbulence intensity, and density ratio are captured for each film-cooling geometry. The detailed film-cooling effectiveness distributions, for cylindrical holes, clearly show that the effectiveness at the lowest blowing ratio is enhanced at the lower density ratio (DR=1). However, as the blowing ratio increases, a transition occurs, leading to increased effectiveness with the elevated density ratio (DR=1.4). In addition, the PSP technique captures an upstream shift of the coolant jet reattachment point as the density ratio increases or the turbulence intensity increases (at moderate blowing ratios for cylindrical holes). With the decreased momentum of the shaped film-cooling holes, the greatest film-cooling effectiveness is obtained at the lower density ratio (DR=1.0) over the entire range of blowing ratios considered. In all cases, as the freestream turbulence intensity increases, the film effectiveness decreases; this effect is reduced as the blowing ratio increases for all three film hole configurations.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

This paper is focused on the effect of film-hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film-cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film-cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film-cooling holes are divided into four rows and loosely distributed on the platform. Four film-cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete-hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film-cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

Effect of Upstream Wake With Vortex on Turbine Blade Platform Film Cooling With Simulated Stator-Rotor Purge Flow

Detailed film cooling effectiveness distributions were experimentally obtained on a turbine blade platform within a linear cascade. The film cooling effectiveness distributions were obtained on the platform with upstream disturbances used to simulate the passing vanes. Cylindrical rods, placed upstream of the blades, simulated the wake created by the trailing edge of the stator vanes. The rods were placed at four locations to show how the film cooling effectiveness was affected relative to the vane location. In addition, delta wings were placed upstream of the blades to model the effect of the passage vortex (generated in the vane passage) on the platform film cooling effectiveness. The delta wings create a vortex similar to the passage vortex as it exits the upstream vane passage. The film cooling effectiveness was measured with the delta wings placed at four location, to investigate the effect of the passing vanes. Finally, the delta wings were coupled with the cylindrical rods to examine the combined effect of the upstream wake and passage vortex on the platform film cooling effectiveness. The detailed film cooling effectiveness distributions were obtained using pressure sensitive paint in the five blade linear cascade. An advanced labyrinth seal was placed upstream of the blades to simulate purge flow from a stator-rotor seal. The coolant flow rate varied from 0.5% to 2.0% of the mainstream flow, while the Reynolds number of the mainstream flow remained constant at 3.1×105 (based on the inlet velocity and chord length of the blade). The film cooling effectiveness was not significantly affected with the upstream rod. However, the vortex generated by the delta wings had a profound impact on the film cooling effectiveness. The vortex created more turbulent mixing within the blade passage, and the result is reduced film cooling effectiveness through the entire passage. When the vane induced secondary flow is included, the need for additional platform cooling becomes very obvious.

Film Cooling Effectiveness Distributions on a Turbine Blade Cascade Platform With Stator-Rotor Purge and Discrete Film Hole Flows

An experimental investigation to obtain detailed film cooling effectiveness distributions on a cooled turbine blade platform within a linear cascade has been completed. The Reynolds number of the freestream flow is 3.1×105, and the platform has a labyrinthlike seal upstream of the blades to model a realistic stator-rotor seal configuration. An additional coolant is supplied to the downstream half of the platform via discrete film cooling holes. The coolant flow rate through the upstream seal varies from 0.5% to 2.0% of the mainstream flow, while the blowing ratio of the coolant through the discrete holes varies from 0.5 to 2.0 (based on the mainstream velocity at the exit of the cascade). Detailed film cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique under a wide range of coolant flow conditions and various freestream turbulence levels (0.75% or 13.4%). The PSP technique clearly shows how adversely the coolant is affected by the passage induced flow. With only purge flow from the upstream seal, the coolant flow rate must exceed 1.5% of the mainstream flow in order to adequately cover the entire passage. However, if discrete film holes are used on the downstream half of the passage, the platform can be protected while using less coolant (i.e., the seal flow rate can be reduced).

Effect of rotation on leading edge region film cooling of a gas turbine blade with three rows of film cooling holes

Effect of rotation on detailed film cooling effectiveness distributions in the leading edge region of a gas turbine blade with three showerhead rows of radial-angle holes were measured using the Pressure Sensitive Paint (PSP) technique. Tests were conducted on the first-stage rotor blade of a three-stage axial turbine at three rotational speeds. The effect of the blowing ratio was also studied. The Reynolds number based on the axial chord length and the exit velocity was 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and exit Mach number was 0.1 and 0.3, respectively. The film cooling effectiveness distributions were presented along with the discussions on the influences of rotational speed, blowing ratio, and vortices around the leading edge region. Results showed that different rotation speeds significantly change the film cooling traces with the average film cooling effectiveness in the leading edge region increasing with blowing ratio.

Film Cooling Effectiveness on the Leading Edge Region of a Rotating Turbine Blade With Two Rows of Film Cooling Holes Using Pressure Sensitive Paint

Detailed film cooling effectiveness distributions are measured on the leading edge of a rotating gas turbine blade with two rows (pressure-side row and suction-side row from the stagnation line) of holes aligned to the radial axis using the pressure sensitive paint (PSP) technique. Film cooling effectiveness distributions are obtained by comparing the difference of the measured oxygen concentration distributions with air and nitrogen as film cooling gas respectively and by applying the mass transfer analogy. Measurements are conducted on the first-stage rotor blade of a three-stage axial turbine at 2400rpm (positive off-design), 2550rpm (design), and 3000rpm (negative off-design), respectively. The effect of three blowing ratios is also studied. The blade Reynolds number based on the axial chord length and the exit velocity is 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and outlet Mach numbers are 0.1 and 0.3, respectively. The film cooling effectiveness distributions are presented along with discussions on the influence of rotational speed (off design incidence angle), blowing ratio, and upstream nozzle wakes around the leading edge region. Results show that rotation has a significant impact on the leading edge film cooling distributions with the average film cooling effectiveness in the leading edge region decreasing with an increase in the rotational speed (negative incidence angle).

Effect of Hole Geometry on the Thermal Performance of Fan-Shaped Film Cooling Holes

This study evaluates the impact of typical cooling hole shape variations on the thermal performance of fan-shaped film holes. A comprehensive set of experimental test cases featuring 16 different film-cooling configurations with different hole shapes have been investigated. The shape variations investigated include hole inlet-to-outlet area ratio, hole coverage ratio, hole pitch ratio, hole length, and hole orientation (compound) angle. Flow conditions applied cover a wide range of film blowing ratios M=0.5 to 2.5 at an engine-representative density ratio DR=1.7. An infrared thermography data acquisition system is used for highly accurate and spatially resolved surface temperature mappings. Accurate local temperature data are achieved by an in situ calibration procedure with the help of thermocouples embedded in the test plate. Detailed film-cooling effectiveness distributions and discharge coefficients are used for evaluating the thermal performance of a row of fan-shaped film holes. An extensive variation of the main geometrical parameters describing a fan-shaped film-cooling hole is done to cover a wide range of typical film-cooling applications in current gas turbine engines. Within the range investigated, laterally averaged film-cooling effectiveness was found to show only limited sensitivity from variations of the hole geometry parameters. This offers the potential to tailor the hole geometry according to needs beyond pure cooling performance, e.g., manufacturing facilitations.

Investigation of Detailed Film Cooling Effectiveness and Heat Transfer Distributions on a Gas Turbine Airfoil

In the present study film cooling effectiveness and heat transfer were systematically investigated on a turbine NGV airfoil employing the transient liquid crystal technique and a multiple regression procedure. Tests were conducted in a linear cascade at exit Reynolds numbers of 0.52e6, 1.02e6 and 1.45e6 and exit Mach numbers of 0.33, 0.62 and 0.8, at two mainstream turbulence intensities of 5.5 and 10 percent. The film cooling geometry consisted of a single compound angle row on the pressure side (PS), and a single or a double row on the suction side (SS). Foreign gas injection was used to obtain a density ratio of approximately 1.65, while air injection yielded a density ratio of unity. Tests were conducted for blowing ratios of 0.25 to 2.3 on the SS, and 0.55 to 7.3 on the PS. In general film cooling injection into a laminar BL showed considerably higher effectiveness in the near-hole region, as compared to a turbulent BL. While mainstream turbulence had only a weak influence on SS cooling, higher effectiveness was noted on the PS at high turbulence due to increased lateral spreading of the coolant. Effects of mainstream Mach and Reynolds number were attributed to changes of the BL thickness and flow acceleration. Higher density coolant yielded higher effectiveness on both SS and PS, whereas heat transfer ratios were increased on the SS and decreased on the PS. Comparison of the single and double row cooling configurations on the SS revealed a better film cooling performance of the double row due to an improved film coverage and delayed jet separation.

New Techniques for Measuring Film Cooling Effectiveness

Techniques to measure local film effectiveness distributions on large-scale models of turbine blade and vane sections were developed for use in subsonic, room-temperature wind-tunnel environments. Highly detailed film cooling effectiveness distributions were obtained on near-adiabatic, cast rigid foam test models using infrared radiometer scanning techniques. Sample results from both flat-plate film cooling studies and airfoil leading-edge film cooling studies are presented.