Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes, as well as elliptical, diffuser, and teardrop shaped dimples, simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of 2 increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2008 ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes as well as elliptical, diffuser, and teardrop shaped dimples simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of two increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

Heat Transfer Coefficients on the Squealer Tip and Near Tip Regions of a Gas Turbine Blade With Single or Double Squealer

2003 ◽  
Author(s):  
Jae Su Kwak ◽  
Jaeyong Ahn ◽  
Je-Chin Han ◽  
C. Pang Lee ◽  
Robert Boyle ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Heat Transfer Coefficients on the Squealer Tip and Near-Tip Regions of a Gas Turbine Blade With Single or Double Squealer

2003 ◽  
Vol 125 (4) ◽  
pp. 778-787 ◽  
Author(s):  
Jae Su Kwak ◽  
Jaeyong Ahn ◽  
Je-Chin Han ◽  
C. Pang Lee ◽  
Ronald S. Bunker ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid-crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Turbine Blade Tip Film-Cooling and Heat Transfer Measurements at High Blowing Ratios

2014 ◽  
Author(s):  
Diganta Narzary ◽  
Kevin Liu ◽  
Je-Chin Han ◽  
Shantanu Mhetras ◽  
Kenneth Landis
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Side Rail ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling and heat transfer characteristics of a gas turbine blade tip with a suction side rail was investigated in a stationary 3-blade rectilinear cascade. Mounted at the end of a blow-down facility the cascade operated at inlet and exit Mach numbers of 0.29 and 0.75, respectively. The rail was marginally offset from the suction side edge of the tip and extended from the leading to the trailing edge. A total of 17 film-cooling holes were placed along the near-tip pressure side surface and 3 on the near-tip leading edge surface with the objective of providing coolant to the tip. The tip surface itself did not carry any film-cooling holes. Relatively high blowing ratios of 2.0, 3.0, 4.0, and 4.5 and three tip gaps of 0.87%, 1.6%, and 2.3% of blade span made up the test matrix. Pressure sensitive paint (PSP) and Thermo-Chromic Liquid Crystal (TLC) were the experimental techniques employed to measure film-cooling effectiveness and heat transfer coefficient, respectively. Results indicated that when the tip gap was increased, film-cooling effectiveness on the tip surface decreased and heat transfer to the tip surface increased. On the other hand, when the blowing ratio was increased, film effectiveness increased but the effect on heat transfer coefficient was relatively small. The highest heat transfer coefficient levels were found atop the suction side rail, especially in the downstream two-thirds of its length whereas the lowest levels were found on the tip floor in the widest section of the blade.

Freestream Flow Effects on Film Effectiveness and Heat Transfer Coefficient Augmentation for Compound Angle Shaped Holes

2017 ◽  
Author(s):  
Joshua B. Anderson ◽  
John W. McClintic ◽  
David G. Bogard ◽  
Thomas E. Dyson ◽  
Zachary Webster
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Layer Thickness ◽  
Film Cooling ◽  
Gas Turbines ◽  
Boundary Layer Thickness ◽  
Adiabatic Effectiveness ◽  
Compound Angle

The use of compound-angled shaped film cooling holes in gas turbines provides a method for cooling regions of extreme curvature on turbine blades or vanes. These configurations have received surprisingly little attention in the film cooling literature. In this study, a row of laid-back fanshaped holes based on an open-literature design, were oriented at a 45-degree compound angle to the approaching freestream flow. In this study, the influence of the approach flow boundary layer thickness and character were experimentally investigated. A trip wire and turbulence generator were used to vary the boundary layer thickness and freestream conditions from a thin laminar boundary layer flow to a fully turbulent boundary layer and freestream at the hole breakout location. Steady-state adiabatic effectiveness and heat transfer coefficient augmentation were measured using high-resolution IR thermography, which allowed the use of an elevated density ratio of DR = 1.20. The results show adiabatic effectiveness was generally lower than for axially-oriented holes of the same geometry, and that boundary layer thickness was an important parameter in predicting effectiveness of the holes. Heat transfer coefficient augmentation was highly dependent on the freestream turbulence levels as well as boundary layer thickness, and significant spatial variations were observed.

scholarly journals Heat Transfer on a Film-Cooled Rotating Blade

1999 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox’s k-ω model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and ω distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

A Novel Antivortex Turbine Film-Cooling Hole Concept

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
James D. Heidmann ◽  
Srinath Ekkad
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Circular Cross Section ◽  
Round Hole ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Film Cooling Holes

A novel turbine film-cooling hole shape has been conceived and designed at NASA Glenn Research Center. This “antivortex” design is unique in that it requires only easily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The hole design is intended to counteract the detrimental vorticity associated with standard circular cross-section film-cooling holes. This vorticity typically entrains hot freestream gas and is associated with jet separation from the turbine blade surface. The antivortex film-cooling hole concept has been modeled computationally for a single row of 30 deg angled holes on a flat surface using the 3D Navier–Stokes solver GLENN-HT. A blowing ratio of 1.0 and density ratios of 1.05 and 2.0 are studied. Both film effectiveness and heat transfer coefficient values are computed and compared to standard round hole cases for the same blowing rates. A net heat flux reduction is also determined using both the film effectiveness and heat transfer coefficient values to ascertain the overall effectiveness of the concept. An improvement in film effectiveness of about 0.2 and in net heat flux reduction of about 0.2 is demonstrated for the antivortex concept compared to the standard round hole for both blowing ratios. Detailed flow visualization shows that as expected, the design counteracts the detrimental vorticity of the round hole flow, allowing it to remain attached to the surface.

scholarly journals Impingement/Effusion Cooling: Overall Wall Heat Transfer

1988 ◽  
Author(s):  
G. E. Andrews ◽  
A. A. Asere ◽  
C. I. Hussain ◽  
M. C. Mkpadi ◽  
A. Nazari
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Transfer Coefficients ◽  
Film Cooling Effectiveness ◽  
Impingement Cooling ◽  
Heat Transfer Coefficients ◽  
Impingement Heat Transfer ◽  
Combined Heat Transfer

Experimental results are presented for the overall heat transfer coefficient within an impingement/effusion wall, using a transient cooling technique. This was previously used for determining the effusion hole heat transfer alone. Two impingement/effusion geometries were used with an 8 mm gap and the same impingement wall with an X/D of 11. The separate impingement and effusion short hole heat transfer coefficients were also determined. The impingement/effusion overall heat transfer was 45% and 30% higher than the impingement heat transfer alone for the two test geometries. The greater increase was for the higher pressure loss effusion wall. It was shown that the combined heat transfer was predominantly the addition of the impingement and effusion heat transfer coefficients but the interaction effects were significant and resulted in an approximately 15% deterioration in the combined heat transfer coefficient. Overall film cooling effectiveness was obtained that showed a significant improvement with the addition of the impingement cooling, but still had a major effusion film cooling contribution.

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