Interactions of Film Cooling Rows: Effects of Hole Geometry and Row Spacing on the Cooling Performance Downstream of the Second Row of Holes

Staggered Arrangement ◽

A comprehensive set of generic experiments is conducted to investigate the interaction of film cooling rows. Five different film cooling configurations are considered on a large scale basis each consisting of two rows of film cooling holes in staggered arrangement. The hole pitch to diameter ratio within each row is kept constant at P/D = 4. The spacing between the rows is either x/D = 10, 20, or 30. Fanshaped holes or simple cylindrical holes with an inclination angle of 30 deg. and a hole length of 6 hole diameters are used. With a hot gas Mach number of Mam = 0.3, an engine like density ratio of ρc/ρm = 1.75, and a freestream turbulence intensity of Tu = 5.1% are established. Operating conditions are varied in terms of blowing ratio for the upstream and, independently, the downstream row in the range 0.5<M<2.0. The results illustrate the importance of considering ejection into an already film cooled boundary layer. Adiabatic film cooling effectiveness and heat transfer coefficients are significantly increased. The decay of effectiveness with streamwise distance is much less pronounced downstream of the second row primarily due to pre-cooling of the boundary layer by the first row of holes. Additionally, a comparison of measured effectiveness data with predictions according to the widely used superposition model of Sellers [11] is given for two rows of fanshaped holes.

Interaction of Film Cooling Rows: Effects of Hole Geometry and Row Spacing on the Cooling Performance Downstream of the Second Row of Holes

10.1115/1.1731395 ◽

2004 ◽

Vol 126 (2) ◽

pp. 237-246 ◽

Cited By ~ 29

Author(s):

Christian Saumweber

Keyword(s):

Boundary Layer ◽

Film Cooling ◽

Large Scale ◽

Density Ratio ◽

Operating Conditions ◽

Transfer Coefficients ◽

Staggered Arrangement ◽

A comprehensive set of generic experiments is conducted to investigate the interaction of film cooling rows. Five different film cooling configurations are considered on a large-scale basis each consisting of two rows of film cooling holes in staggered arrangement. The hole pitch to diameter ratio within each row is kept constant at P/D=4. The spacing between the rows is either x/D=10, 20, or 30. Fan-shaped holes or simple cylindrical holes with an inclination angle of 30 deg and a hole length of 6-hole diameters are used. With a hot gas Mach number of Mam=0.3, an engine like density ratio of ρc/ρm=1.75, and a freestream turbulence intensity of Tu=5.1% are established. Operating conditions are varied in terms of blowing ratio for the upstream and, independently, the downstream row in the range 0.5<M<2.0. The results illustrate the importance of considering ejection into an already film-cooled boundary layer. Adiabatic film cooling effectiveness and heat transfer coefficients are significantly increased. The decay of effectiveness with streamwise distance is much less pronounced downstream of the second row primarily due to pre-cooling of the boundary layer by the first row of holes. Additionally, a comparison of measured effectiveness data with predictions according to the widely used superposition model of Sellers is given for two rows of fanshaped holes.

Aspects of Vane Film Cooling With High Turbulence: Part I — Heat Transfer

Volume 3: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/97-gt-239 ◽

1997 ◽

Cited By ~ 3

Author(s):

Forrest E. Ames

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Film Cooling ◽

Large Scale ◽

Velocity Ratio ◽

Transfer Coefficients ◽

Suction Surface ◽

Stagnation Region ◽

High Turbulence

A four vane subsonic cascade was used to investigate the influence of film injection on vane heat transfer distributions in the presence of high turbulence. The influence of high turbulence on vane film cooling effectiveness and boundary layer development was also examined in part II of this paper. A high level, large scale inlet turbulence was generated for this study with a mock combustor (12 %) and was used to contrast results with a low level (1 %) of inlet turbulence. The three geometries chosen to study in this investigation were one row and two staggered rows of downstream cooling on both the suction and pressure surfaces in addition to a showerhead array. Film cooling was found to have only a moderate influence on the heat transfer coefficients downstream from arrays on the suction surface where the boundary layer was turbulent. However, film cooling was found to have a substantial influence on heat transfer downstream from arrays in laminar regions of the vane such as the pressure surface, the stagnation region, and the near suction surface. Generally, heat transfer augmentation was found to scale on velocity ratio. In relative terms, the augmentation in the laminar regions for the low turbulence case was found to be higher than the augmentation for the high turbulence case. The absolute levels of heat transfer were always found to be the highest for the high turbulence case.

Effect of Unheated Starting Lengths on Film Cooling Experiments

10.1115/1.2184355 ◽

2006 ◽

Vol 128 (3) ◽

pp. 579-588 ◽

Cited By ~ 13

Author(s):

Sarah M. Coulthard ◽

Ralph J. Volino ◽

Karen A. Flack

Keyword(s):

Film Cooling ◽

Infrared Camera ◽

Stanton Number ◽

Flow Structures ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Single Row ◽

The effect of an unheated starting length upstream of a row of film cooling holes was studied experimentally to determine its effect on heat transfer coefficients downstream of the holes. Cases with a single row of cylindrical film cooling holes inclined at 35deg to the surface of a flat plate were considered at blowing ratios of 0.25, 0.5, 1.0, and 1.5. For each case, experiments were conducted to determine the film-cooling effectiveness and the Stanton number distributions in cases with the surface upstream of the holes heated and unheated. Measurements were made using an infrared camera, thermocouples, and hot and cold-wire anemometry. Ratios were computed of the Stanton number with film cooling (Stf) to corresponding Stanton numbers in cases without film cooling (Sto), but the same surface heating conditions. Contours of these ratios were qualitatively the same regardless of the upstream heating conditions, but the ratios were larger for the cases with a heating starting length. Differences were most pronounced just downstream of the holes and for the lower blowing rate cases. Even 12 diameters downstream of the holes, the Stanton number ratios were 10–15% higher with a heated starting length. At higher blowing rates the differences between the heated and unheated starting length cases were not significant. The differences in Stanton number distributions are related to jet flow structures, which vary with blowing rate.

Experimental Investigation of Rotor Tip Film Cooling at an Axial Turbine with Swirling Inflow Using Pressure Sensitive Paint

International Journal of Turbomachinery Propulsion and Power ◽

10.3390/ijtpp4030023 ◽

2019 ◽

Vol 4 (3) ◽

pp. 23 ◽

Cited By ~ 1

Author(s):

Manuel Wilhelm ◽

Heinz-Peter Schiffer

Keyword(s):

Film Cooling ◽

Large Scale ◽

Density Ratio ◽

Pressure Sensitive Paint ◽

Axial Turbine ◽

Pressure Sensitive ◽

Inlet Swirl ◽

Blade Tip ◽

Rotor tip film cooling is investigated at the Large Scale Turbine Rig, which is a 1.5-stage axial turbine rig operating at low speeds. Using pressure sensitive paint, the film cooling effectiveness η at a squealer-type blade tip with cylindrical pressure-side film cooling holes is obtained. The effect of turbine inlet swirl on η is examined in comparison to an axial inflow baseline case. Coolant-to-mainstream injection ratios are varied between 0.45% and 1.74% for an engine-realistic coolant-to-mainstream density ratio of 1.5. It is shown that inlet swirl causes a reduction in η for low injection ratios by up to 26%, with the trailing edge being especially susceptible to swirl. For injection ratios greater than 0.93%, however, η is increased by up to 11% for swirling inflow, while for axial inflow a further increase in coolant injection does not transfer into a gain in η .

Enhanced Film Cooling Effectiveness With Surface Trenches

Volume 3B: Heat Transfer ◽

10.1115/gt2013-94530 ◽

2013 ◽

Cited By ~ 1

Author(s):

K. Vighneswara Rao ◽

Jong S. Liu ◽

Daniel C. Crites ◽

Luis A. Tapia ◽

Malak F. Malak ◽

...

Keyword(s):

Film Cooling ◽

Density Ratio ◽

Operating Conditions ◽

Airfoil Surface ◽

Cooling Effectiveness ◽

Ansys Cfx ◽

Cooling Hole ◽

Computational Fluid Dynamics Cfd ◽

In this study, cylindrical and fan shaped film cooling holes are evaluated on the blade surface numerically, using the Computational Fluid Dynamics (CFD) tool ANSYS-CFX, with the objective of improving cooling effectiveness by understanding the flow pattern at the cooling hole exit. The coolant flow rates are adjusted for blowing ratios of 0.5, 1.0 & 1.5 (momentum flux ratios of 0.125, 0.5 & 1.125 respectively). The density ratio is maintained at 2.0. New shaped holes viz. straight, concave and convex trench holes are introduced and are evaluated under similar operating conditions. Results are presented in terms of surface temperatures and adiabatic effectiveness at three different blowing ratios for the different film cooling hole shapes analyzed. Comparison is made with reference to the fan shaped film cooling hole to bring out relative merits of different shapes. The new trench holes improved the film cooling effectiveness by allowing more residence time for coolant to spread laterally while directing smoothly onto the airfoil surface. While convex trench improved the centre-line effectiveness, straight trench improved the laterally-averaged and overall effectiveness at all blowing ratios. Concave trench improved the effectiveness at blowing ratios 0.5 and 1.0.

Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film Cooling Array: Part II—Temperature and Heat Transfer Measurements

Volume 5: Heat Transfer, Parts A and B ◽

10.1115/gt2011-46498 ◽

2011 ◽

Cited By ~ 4

Author(s):

Douglas R. Thurman ◽

Lamyaa A. El-Gabry ◽

Philip E. Poinsatte ◽

James D. Heidmann

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Large Scale ◽

Local Heat Transfer ◽

Local Heat ◽

Surface Heat ◽

Transfer Coefficients ◽

Surface Heat Transfer ◽

The second of a two-part paper, this study focuses on the temperature field and surface heat transfer measurements on a large-scale models of an inclined row of film cooling holes. Detailed surface and flow field measurements were taken and presented in Part I. The model consists of three holes of 1.9-cm diameter that are spaced 3 hole diameters apart and inclined 30° from the surface. Additionally, another model with an anti-vortex adaptation to the film cooling holes is also tested. The coolant stream is metered and cooled to 20°C below the mainstream temperature. A thermocouple is used to obtain the flow temperatures along the jet centerline and at various streamwise locations. Steady state liquid crystal thermography is used to obtain surface heat transfer coefficients. Results are obtained for blowing ratios of up to 2 in order to capture off-design conditions in which the jet is lifted. Film cooling effectiveness values of 0.4 and 0.15 were found along the centerline for blowing ratios of 1 and 2 respectively. In addition, an anti-vortex design was tested and found to have improved film effectiveness. This paper presents the detailed temperature contours showing the extent of mixing between the coolant and freestream and the local heat transfer results.

Aspects of Vane Film Cooling With High Turbulence: Part I—Heat Transfer

10.1115/1.2841788 ◽

1998 ◽

Vol 120 (4) ◽

pp. 768-776 ◽

Cited By ~ 22

Author(s):

F. E. Ames

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Film Cooling ◽

Large Scale ◽

Velocity Ratio ◽

Transfer Coefficients ◽

Suction Surface ◽

Stagnation Region ◽

High Turbulence

A four-vane subsonic cascade was used to investigate the influence of film injection on vane heat transfer distributions in the presence of high turbulence. The influence of high turbulence on vane film cooling effectiveness and boundary layer development was also examined in part II of this paper. A high-level, large-scale inlet turbulence was generated for this study with a mock combustor (12 percent) and was used to contrast results with a low level (1 percent) of inlet turbulence. The three geometries chosen to study in this investigation were one row and two staggered rows of downstream cooling was found to have only a moderate influence on the heat transfer coefficients downstream from arrays on the suction surface where the boundary layer was turbulent. However, film cooling was found to have a substantial influence on heat transfer downstream from arrays in laminar regions of the vane such as the pressure surface, the stagnation region, and the near-suction surface. Generally, heat transfer augmentation was found to scale on velocity ratio. In relative terms, the augmentation in the laminar regions for the low turbulence case was found to be the highest for the high turbulence case.

Film Cooling From Spanwise-Oriented Holes in Two Staggered Rows

10.1115/1.2841158 ◽

1997 ◽

Vol 119 (3) ◽

pp. 562-567 ◽

Cited By ~ 19

Author(s):

P. M. Ligrani ◽

A. E. Ramsey

Keyword(s):

Film Cooling ◽

Density Ratio ◽

Test Surface ◽

Spanwise Direction ◽

Transfer Coefficients ◽

Blowing Ratio ◽

Streamwise Location ◽

Adiabatic Effectiveness ◽

Adiabatic effectiveness and iso-energetic heat transfer coefficients are presented from measurements downstream of film-cooling holes inclined at 30 deg. with respect to the test surface in spanwise/normal planes. With this configuration, holes are spaced 3d apart in the spanwise direction and 4d in the streamwise direction in two staggered rows. Results are presented for an injectant to free-stream density ratio near 1.0, and injection blowing ratios from 0.5 to 1.5. Spanwise-averaged adiabatic effectiveness values downstream of the spanwise/normal plane holes are significantly higher than values measured downstream of simple angle holes for x/d < 25–70(depending on blowing ratio) when compared for the same normalized streamwise location, blowing ratio, and spanwise and streamwise hole spacings. Spanwise-averaged iso-energetic Stanton number ratios range between 1.0 and 1.41, increase with blowing ratio at each streamwise station, and show little variation with streamwise location for each value of blowing ratio tested.

Effect of Unheated Starting Lengths on Film Cooling Experiments

Heat Transfer: Volume 3 ◽

10.1115/ht2005-72392 ◽

2005 ◽

Cited By ~ 1

Author(s):

Sarah M. Coulthard ◽

Ralph J. Volino ◽

Karen A. Flack

Keyword(s):

Film Cooling ◽

Infrared Camera ◽

Stanton Number ◽

Flow Structures ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Single Row ◽

The effect of an unheated starting length upstream of a row of film cooling holes was studied experimentally to determine its effect on heat transfer coefficients downstream of the holes. Cases with a single row of cylindrical film cooling holes inclined at 35 degrees to the surface of a flat plate were considered at blowing ratios of 0.25, 0.5, 1.0 and 1.5. For each case experiments were conducted to determine the film cooling effectiveness and the Stanton number distributions in cases with the surface upstream of the holes heated and unheated. Measurements were made using an infrared camera, thermocouples, and hot and cold wire anemometry. Ratios were computed of the Stanton number with film cooling (Stf) to corresponding Stanton numbers in cases without film cooling (Sto) but the same surface heating conditions. Contours of these ratios were qualitatively the same regardless of the upstream heating conditions, but the ratios were larger for the cases with a heating starting length. Differences were most pronounced just downstream of the holes and for the lower blowing rate cases. Even 12 diameters downstream of the holes the Stanton number ratios were 10 to 15% higher with a heated starting length. The differences in Stanton number distributions are related to jet flow structures which vary with blowing rate.

Validation of a First Vane Platform Cooling Design

Volume 5: Heat Transfer, Parts A and B ◽

10.1115/gt2011-45252 ◽

2011 ◽

Cited By ~ 3

Author(s):

Joerg Krueckels ◽

William Colban ◽

Michael Gritsch ◽

Martin Schnieder

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Gas Turbines ◽

Secondary Flows ◽

Operating Conditions ◽

Transfer Coefficients ◽

Design Data ◽

High Lift ◽

Heat Transfer Coefficients

Low emission requirements for large industrial gas turbines can be achieved with flat combustor temperature profiles reducing the combustor peak temperature. As a result the heat load on the first stage vane platforms increases and platform film cooling is an important requirement. Furthermore, high lift airfoils generate stronger secondary flows including complex vortex flows over the platforms, which impacts heat transfer coefficients and film cooling. Cascade tests have been performed on a high lift profile with a platform film configuration and will be presented. The linear cascade was operated at engine representative Mach numbers. Pressure measurements are compared to design data to ensure correct operating conditions and periodicity of the cascade. The thermochromic liquid crystal measurement technique is used to obtain adiabatic film cooling effectiveness. The upstream gap (corresponding to the gap between the combustor and turbine) and the purge air exiting this gap are included in the investigations. The effect of the purge air on the recovery temperature is very strong and needs to be taken into account for the layout of the cooling scheme. The heat transfer coefficient distribution on the platform is obtained for an uncooled configuration using a transient infrared imaging technique with heat flux reconstruction. Computational fluid dynamics (CFD) assessments are used to support the validation results. Heat transfer coefficients and the effect of the purge air on adiabatic wall temperatures are compared with experimental results.