scholarly journals An Experimental Study of the Effect of Wake Passing on Turbine Blade Film Cooling

1997 ◽  
Vol 123 (2) ◽  
pp. 214-221 ◽  
Author(s):  
James D. Heidmann ◽  
Barbara L. Lucci ◽  
Eli Reshotko
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Experimental Film ◽  
Turbine Cascade ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Nusselt Numbers ◽  
Wake Passing ◽  
Expected Increase ◽  
Effectiveness Data

The effect of wake passing on the showerhead film cooling performance of a turbine blade has been investigated experimentally. The experiments were performed in an annular turbine cascade with an upstream rotating row of cylindrical rods. Nickel thin-film gauges were used to determine local film effectiveness and Nusselt number values for various injectants, blowing ratios, and Strouhal numbers. Results indicated a reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated by a streamwise-constant decrement of 0.094St. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

scholarly journals An Experimental Study of the Effect of Wake Passing on Turbine Blade Film Cooling

1997 ◽  
Author(s):  
James D. Heidmann ◽  
Barbara L. Lucci ◽  
Eli Reshotko
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Experimental Film ◽  
Turbine Cascade ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Nusselt Numbers ◽  
Wake Passing ◽  
Expected Increase ◽  
Effectiveness Data

The effect of wake passing on the showerhead film cooling performance of a turbine blade has been investigated experimentally. The experiments were performed in an annular turbine cascade with an upstream rotating row of cylindrical rods. Nickel thin-film gauges were used to determine local film effectiveness and Nusselt number values for various injectants, blowing ratios, and Strouhal numbers. Results indicated a reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated by a streamwise-constant decrement of 0.094·St. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

Experimental Study of the Effects of an Oscillating Approach Flow on Overall Cooling Performance of a Simulated Turbine Blade Leading Edge

2009 ◽  
Author(s):  
Ross Johnson ◽  
Jonathan Maikell ◽  
David Bogard ◽  
Justin Piggush ◽  
Atul Kohli ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Oscillation Frequencies ◽  
Overall Effectiveness ◽  
Blade Leading Edge

When a turbine blade passes through wakes from upstream vanes it is subjected to an oscillation of the direction of the approach flow resulting in the oscillation of the position of the stagnation line on the leading edge of the blade. In this study an experimental facility was developed that induced a similar oscillation of the stagnation line position on a simulated turbine blade leading edge. The overall effectiveness was evaluated at various blowing ratios and stagnation line oscillation frequencies. The location of the stagnation line on the leading edge was oscillated to simulate a change in angle of attack between α = ± 5° at a range of frequencies from 2 to 20 Hz. These frequencies were chosen based on matching a range of Strouhal numbers typically seen in an engine due to oscillations caused by passing wakes. The blowing ratio was varied between M = 1, M = 2, and M = 3. These experiments were carried out at a density ratio of DR = 1.5 and mainstream turbulence levels of Tu ≈ 6%. The leading edge model was made of high conductivity epoxy in order to match the Biot number of an actual engine airfoil. Results of these tests showed that the film cooling performance with an oscillating stagnation line was degraded by as much as 25% compared to the performance of a steady flow with the stagnation line aligned with the row of holes at the leading edge.

Enhancement of Film Cooling Performance at the Leading Edge of Turbine Blade

2006 ◽  
Author(s):  
K.-S. Kim ◽  
Youn J. Kim ◽  
S.-M. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Cylindrical Body ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105, and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

Numerical Study on the Effects of V-Shaped Rib Angle on Film Cooling Performance for Turbine Blade Trailing Edge

2018 ◽  
Author(s):  
Lin Ye ◽  
Cun-liang Liu ◽  
Hai-yong Liu ◽  
Qi-jiao He ◽  
Gang Xie
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Film Cooling ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

The trailing edge of the high-pressure turbine blade presents significant challenges to cooling structure design. To achieve better cooling performance at turbine blade trailing edge, a novel ribbed cutback structure is proposed for trailing edge cooling, which has rib structures on the cutback surface for heat transfer enhancement. In this study, numerical simulations have been performed on the effects of V-shaped rib angle on the film cooling characteristics and flow physics. Three V-shaped rib angles of 30°, 45° and 60° are studied. The distributions of adiabatic cooling effectiveness and heat transfer coefficient are obtained under blowing ratios with the value of 0.5, 1.0 and 1.5 respectively. Due to the relatively small rib height, the effect of V-shaped ribs on the film cooling effectiveness is not notable. The disadvantage of V-shaped ribs mainly exhibits at the downstream area of cutback surface. With the increase of V-shaped rib angle, the film cooling effectiveness becomes lower, but the values are still above 0.9. The V-shaped ribs obviously enhance the heat transfer on trailing edge cutback surface. The area-averaged heat transfer coefficient of the V-rib case is higher than that of the smooth case by 26.3–41.2%. The 45° V-rib case has higher heat transfer intensity than the other two V-shaped rib cases under all the three blowing ratios. However, the heat transfer coefficient distribution of the 60° V-rib case is more uniform. The heat transfer intensity of the 30° V-rib case is higher in the downstream region than the other two cases, but lower in the upstream region in which the difference becomes smaller with the increase of blowing ratio. The 45° V-rib case and the 60° V-rib case both reach the maximum value of area-averaged heat transfer intensity under blowing ratio is 1.0. Under higher blowing ratio, the 30° V-rib case and the 45° V-rib case outperform 2.1% and 6.7% higher value relative to the 60° V-rib case respectively due to the smaller velocity gradient in the 60° V-rib case in the downstream.

Experimental Investigation of Dust-Pan Shaped Hole Film Cooling Characteristics on Pressure Side of a Turbine Blade in a Linear Transonic Cascade

2017 ◽  
Author(s):  
Zhong-yi Fu ◽  
Hui-ren Zhu ◽  
Cong Liu ◽  
Zheng Li
Keyword(s):  
Pressure Gradient ◽  
Turbine Blade ◽  
Film Cooling ◽  
Stationary Condition ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Favorable Pressure Gradient ◽  
Shaped Hole

An experimental research of film cooling performance of three single dust-pan shaped hole rows in different positions of a turbine blade was carried out in the short-duration transonic linear cascade at stationary condition, which can model realistic engine aerodynamic conditions. The effects of inlet Reynolds number (Rein = 2.5 × 105∼7.5 × 105), isentropic exit Mach number (Mais = 0.71∼0.91) and coolant blowing ratio (M = 0.8∼2.6) on film cooling effectiveness are investigated. Three single hole rows are located at 11.7%, 36.3% and 55.6% relative arc on the pressure sides of three enlarged blade models respectively. The adiabatic film cooling effectiveness are derived from the surface temperatures based on transient heat transfer measurement method. The results show that in the range of blowing ratios studied in the present paper, for location 3 the cooling effectiveness decreases a lot with blowing ratio increasing due to the lift-off of coolant at high blowing ratios, while for location 1 and 2, the film cooling effectiveness increases with blowing ratio increasing, because the strong favorable pressure gradient and high concave curvature near the leading edge lead to a good attachment of coolant on the surface. At M≤1.0 conditions, the film cooling effectiveness of location 1 and 2 is lower than that of location 3, which reflects that strong favorable pressure gradient and high concave curvature weaken film cooling performance at low blowing ratio conditions, while the effect is opposite when M is greater than 1.0. For location 1, the highest general cooling performance is obtained at Rein = 2.5 × 105 condition, and for location 2, the change of Rein has different effects on cooling effectiveness in different regions. In the range of Mais studied in this paper, the change of Mais has little effect on film cooling effectiveness.

Film Cooling Performance on Turbine Blade Suction Side With Various Film Cooling Hole Arrangements

2020 ◽  
Author(s):  
Zhiyu Zhou ◽  
Haiwang Li ◽  
Gang Xie ◽  
Ruquan You
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Stokes Equations ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blowing Ratio

Abstract Numerical simulations were carried out to study the film cooling effectiveness distributions of different hole arrangements on the suction side of a high pressure turbine blade under rotating condition. The chord length and the height of the blade are 60mm and 80mm, respectively. Totally 12 models with different hole arrangements and different injection angles were studied. Each blade model has three rows of round holes with diameter of 0.9mm on the suction surface. The first row and the third row are fixed at streamwise location of 12.4% and 34% respectively. Three injection angles, 30°, 45°, and 60°, were investigated. Simulations were conducted under three rotational speeds, 600rpm, 800rpm, 1000rpm, with blowing ratio varying from 0.5 to 2.0. The Mainstream Reynolds numbers corresponding to the rotational speeds are 40560, 54080, and 67600 respectively. The temperature of the mainstream and the coolant is set at 463K and 303K so as to control the density ratio at 1.47. Simulations were performed by using SST turbulence model and were solved by using the three-dimensional Reynolds-averaged Navier–Stokes equations. Results showed that on the rotating turbine blade suction surface, film trajectories are drawn toward the midspan. The film trajectory arrangement may be different from the hole arrangement. Inline film trajectory arrangement can achieve higher film cooling effectiveness with slightly larger injection angle. Staggered film trajectory arrangement is better for uniform film cooling effectiveness distribution in spanwise and can achieve higher film cooling effectiveness with smaller injection angle. A smaller distance between the first row and the second row can achieve better film cooling performance at the downstream. With the increase of rotational speed, the mainstream Reynolds number increases, which improves the film cooling performance with smaller blowing ratio.

Experimental Study on the Film Cooling Performance at the Leading Edge of Turbine Blade Using Infrared Thermography

2006 ◽  
Vol 326-328 ◽  
pp. 1161-1164
Author(s):  
Kwang Su Kim ◽  
Youn Jea Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Turbine Blades ◽  
Density Ratio ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Cylindrical Body ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

In order to protect turbine blades from high temperature, film cooling can be applied to gas turbine engine system since it can prevent corrosion and facture of material. To enhance the film cooling performance in the vicinity of the turbine blade leading edge, flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and the mainstream turbulence intensities were about 0.2%. CO2 was used as coolant to simulate the effect of coolant-tomainstream density ratio. The effect of coolant flow rates was studied for various blowing ratios of 0.5, 0.8, 1.1 and 1.4, respectively. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Film Cooling From Novel Sister Shaped Single-Holes

2014 ◽  
Author(s):  
Siavash Khajehhasani ◽  
Bassam Jubran
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Lateral Spreading ◽  
Navier Stokes ◽  
The Novel ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
Lift Off

A numerical investigation of the film cooling performance from novel sister shaped single-holes (SSSH) is presented in this paper and the obtained results are compared with a single cylindrical hole, a forward diffused shaped hole, as well as discrete sister holes. Three types of the novel sister shaped single-hole schemes namely downstream, upstream and up/downstream SSSH, are designed based on merging the discrete sister holes to the primary hole in order to reduce the jet lift-off effect and increase the lateral spreading of the coolant on the blade surface as well as a reduction in the amount of coolant in comparison with discrete sister holes. The simulations are performed using three-dimensional Reynolds-Averaged Navier Stokes analysis with the realizable k–ε model combined with the standard wall function. The upstream SSSH demonstrates similar film cooling performance to that of the forward diffused shaped hole for the low blowing ratio of 0.5. While it performs more efficiently at M = 1, where the centerline and laterally averaged effectiveness results improved by 70% and 17%, respectively. On the other hand, the downstream and up/downstream SSSH schemes show a considerable improvement in film cooling performance in terms of obtaining higher film cooling effectiveness and less jet lift-off effect as compared with the single cylindrical and forward diffused shaped holes for both blowing ratios of M = 0.5 and 1. For example, the laterally averaged effectiveness for the downstream SSSH configuration shows an improvement of approximately 57% and 110% on average as compared to the forward diffused shaped hole for blowing ratios of 0.5 and 1, respectively.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

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