A Review of Shaped Hole Turbine Film-Cooling Technology

2005 ◽  
Vol 127 (4) ◽  
pp. 441-453 ◽  
Author(s):  
Ronald S. Bunker
Keyword(s):  
Film Cooling ◽  
High Efficiency ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Hole Shaping ◽  
Cooling Hole ◽  
Gas Blowing ◽  
Parameter Ranges

Film cooling represents one of the few game-changing technologies that has allowed the achievement of today’s high firing temperature, high-efficiency gas turbine engines. Over the last 30 years, only one major advancement has been realized in this technology, that being the incorporation of exit shaping to the film holes to result in lower momentum coolant injection jets with greater surface coverage. This review examines the origins of shaped film cooling and summarizes the extant literature knowledge concerning the performance of such film holes. A catalog of the current literature data is presented, showing the basic shaping geometries, parameter ranges, and types of data obtained. Specific discussions are provided for the flow field and aerodynamic losses of shaped film hole coolant injection. The major fundamental effects due to coolant-to-gas blowing ratio, compound angle injection, cooling hole entry flow character, and mainstream turbulence intensity are each reviewed with respect to the resulting adiabatic film effectiveness and heat transfer coefficients for shaped holes. A specific example of shaped film effectiveness is provided for a production turbine inlet vane with comparison to other data. Several recent unconventional forms of film hole shaping are also presented as a look to future potential improvements.

Full-Coverage Film Cooling: Film Effectiveness and Heat Transfer Coefficients for Dense Hole Arrays at Different Hole Angles, Contraction Ratios, and Blowing Ratios

2013 ◽  
Vol 135 (3) ◽  
Author(s):  
Phil Ligrani ◽  
Matt Goodro ◽  
Mike Fox ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Test Surface ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Contraction Ratio ◽  
Cooling Hole ◽  
Film Cooling Holes

Experimental results are presented for a full-coverage film cooling arrangement which simulates a portion of a gas turbine engine, with appropriate streamwise static pressure gradient. The test surface utilizes varying blowing ratio (BR) along the length of the contraction passage which contains the cooling hole arrangement. For the different experimental conditions examined, film cooling holes are sharp-edged and streamwise inclined either at 20 deg or 30 deg with respect to the liner surface. The film cooling holes in adjacent streamwise rows are staggered with respect to each other. Data are provided for turbulent film cooling, contraction ratios of 1, 3, 4, and 5, blowing ratios (at the test section entrance) of 2.0, 5.0, and 10.0, coolant Reynolds numbers Refc of 10,000–12,000, freestream temperatures from 75 °C to 115 °C, a film hole diameter of 7 mm, and density ratios from 1.15 to 1.25. Nondimensional streamwise and spanwise film cooling hole spacings, X/D and Y/D, are 6, and 5, respectively. When the streamwise hole inclination angle is 20 deg spatially averaged and line-averaged adiabatic effectiveness values at each x/D location are about the same as the contraction ratio varies between 1, 3, and 4, with slightly higher values at each x/D location when the contraction ratio Cr is 5. For each contraction ratio, there is a slight increase in effectiveness when the blowing ratio is increased from 2.0 to 5.0 but there is no further substantial improvement when the blowing ratio is increased to 10.0. Overall, line-averaged and spatially averaged-adiabatic film effectiveness data, and spatially averaged heat transfer coefficient data are described as they are affected by contraction ratio, blowing ratio, hole angle α, and streamwise location x/D. For example, when α = 20 deg, the detrimental effects of mainstream acceleration are apparent since heat transfer coefficients for contraction ratios Cr of 3 and 5 are often higher than values for Cr = 1, especially for x/D > 100.

Full-Coverage Film Cooling: Film Effectiveness and Heat Transfer Coefficients for Dense Hole Arrays at Different Hole Angles, Contraction Ratios, and Blowing Ratios

2012 ◽  
Author(s):  
Matt Goodro ◽  
Phil Ligrani ◽  
Mike Fox ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Test Surface ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Contraction Ratio ◽  
Cooling Hole ◽  
Film Cooling Holes

Experimental results are presented for a full coverage film cooling arrangement which simulates a portion of a gas turbine engine, with appropriate streamwise static pressure gradient. The test surface utilizes varying blowing ratio along the length of the contraction passage which contains the cooling hole arrangement. For the different experimental conditions examined, film cooling holes are sharp-edged and streamwise inclined either at 20° or 30° with respect to the liner surface. The film cooling holes in adjacent streamwise rows are staggered with respect to each other. Data are provided for turbulent film cooling, contraction ratios of 1, 3, 4, and 5, blowing ratios (at the test section entrance) of 2.0, 5.0, and 10.0, coolant Reynolds numbers Refc of 10,000 to 12,000, freestream temperatures from 75°C to 115°C, a film hole diameter of 7 mm, and density ratios from 1.15 to 1.25. Non-dimensional streamwise and spanwise film cooling hole spacings, X/D and Y/D, are 6, and 5, respectively. When the streamwise hole inclination angle is 20°, spatially-averaged and line-averaged adiabatic effectiveness values at each x/D location are about the same as the contraction ratio varies between 1, 3, and 4, with slightly higher values at each x/D location when the contraction ratio Cr is 5. For each contraction ratio, there is a slight increase in effectiveness when the blowing ratio is increased from 2.0 to 5.0 but there is no further substantial improvement when the blowing ratio is increased to 10.0. Overall, line-averaged and spatially-averaged adiabatic film effectiveness data, and spatially-averaged heat transfer coefficient data are described as they are affected by contraction ratio, blowing ratio, hole angle α, and streamwise location x/D. For example, when α = 20°, the detrimental effects of mainstream acceleration are apparent since heat transfer coefficients for contraction ratios Cr of 3 and 5 are often higher than values for Cr = 1, especially for x/D > 100.

Full-Coverage Film Cooling: Film Effectiveness and Heat Transfer Coefficients for Dense and Sparse Hole Arrays at Different Blowing Ratios

2011 ◽  
Author(s):  
Matt Goodro ◽  
Phil Ligrani ◽  
Mike Fox ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Hole Array ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Cooling Hole ◽  
Hole Arrays ◽  
Film Cooling Holes

Experimental results are presented for a full coverage film cooling arrangement which simulates a portion of a gas turbine engine, with appropriate streamwise static pressure gradient and varying blowing ratio along the length of the contraction passage which contains the cooling hole arrangement. Film cooling holes are sharp-edged, streamwise inclined at 20° with respect to the liner surface, and are arranged with a length to diameter ratio of 8.35. The film cooling holes in adjacent streamwise rows are staggered with respect to each other. Data are provided for turbulent film cooling, contraction ratios of 1 and 4, blowing ratios (at the test section entrance) of 2.0, 5.0, and 10.0, coolant Reynolds numbers Refc from 10,000 to 12,000, freestream temperatures from 75°C to 115°C, a film hole diameter of 7 mm, and density ratios from 1.15 to 1.25. Changes to X/D and Y/D, non-dimensional streamwise and spanwise film cooling hole spacings, with Y/D of 3, 5, and 7, and with X/D of 6 and 18, are considered. For all X/D = 6 hole spacings, only a slight increase in effectiveness (local, line-averaged, and spatially-averaged) values are present as the blowing ratio increases from 2.0 to 5.0, with no significant differences when the blowing ratio increases from 5.0 to 10.0. This lack of dependence on blowing ratio indicates a condition where excess coolant is injected into the mainstream flow, a situation not evidenced by data obtained with the X/D = 18 hole spacing arrangement. With this sparse array configuration, local and spatially-averaged effectiveness generally increase continually as the blowing ratio becomes larger. Line-averaged and spatially-averaged heat transfer coefficients are generally higher at each streamwise location, also with larger variations with streamwise development, with the X/D = 6 hole array, compared to the X/D = 18 array.

Full-Coverage Film Cooling: Film Effectiveness and Heat Transfer Coefficients for Dense and Sparse Hole Arrays at Different Blowing Ratios

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Phil Ligrani ◽  
Matt Goodro ◽  
Mike Fox ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Hole Array ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Cooling Hole ◽  
Hole Arrays ◽  
Film Cooling Holes

Experimental results are presented for a full coverage film cooling arrangement which simulates a portion of a gas turbine engine, with appropriate streamwise static pressure gradient and varying blowing ratio along the length of the contraction passage which contains the cooling hole arrangement. Film cooling holes are sharp-edged, streamwise inclined at 20 deg with respect to the liner surface, and are arranged with a length to diameter ratio of 8.35. The film cooling holes in adjacent streamwise rows are staggered with respect to each other. Data are provided for turbulent film cooling, contraction ratios of 1 and 4, blowing ratios (at the test section entrance) of 2.0, 5.0, and 10.0, coolant Reynolds numbers Refc from 10,000 to 12,000 (for a blowing ratio of 5.0), freestream temperatures from 75 °C to 115 °C, a film hole diameter of 7 mm, and density ratios from 1.15 to 1.25. Changes to X/D and Y/D, nondimensional streamwise and spanwise film cooling hole spacings, with Y/D of 3, 5, and 7, and with X/D of 6 and 18, are considered. For all X/D=6 hole spacings, only a slight increase in effectiveness (local, line-averaged, and spatially-averaged) values are present as the blowing ratio increases from 2.0 to 5.0, with no significant differences when the blowing ratio increases from 5.0 to 10.0. This lack of dependence on blowing ratio indicates a condition where excess coolant is injected into the mainstream flow, a situation not evidenced by data obtained with the X/D=18 hole spacing arrangement. With this sparse array configuration, local and spatially-averaged effectiveness generally increase continually as the blowing ratio becomes larger. Line-averaged and spatially-averaged heat transfer coefficients are generally higher at each streamwise location, also with larger variations with streamwise development, with the X/D=6 hole array, compared to the X/D=18 array.

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.

Numerical Simulation of Film Cooling on the Tip of a Gas Turbine Blade

2002 ◽  
Author(s):  
Sumanta Acharya ◽  
Huitao Yang ◽  
Srinath V. Ekkad ◽  
Chander Prakash ◽  
Ron Bunker
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Transfer Coefficients ◽  
Local Pressure ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Coolant Delivery

Numerical simulations of flow and heat transfer are presented for a GE-E3 turbine blade with a film-cooled tip. Results are presented for both a flat tip and a squealer tip. Straight-through coolant holes are considered, and the calculation domain includes the flow development in the coolant delivery tubes. Results are presented with three different tip gaps representing 1%, 1.5% and 2.5% of blade span, a blowing ratio (ratio of coolant-jet-exit velocity to average passage flow velocity) of 1, and an inlet turbulence intensity of 6.1%. On a flat tip, film coolant injection is shown to lower the local pressure ratio and alters the nature of the leakage vortex. High film cooling effectiveness and low heat transfer coefficients are obtained along the coolant trajectory; these values increase slightly with increasing tip clearances. For a squealer tip, the flow inside the squealer cavity exhibits streamwise directed flow, which alters the trajectory of the coolant jets and reduces their effectiveness.

Film-Cooling Characteristics of Pressure-Side Discharge Slots in an Accelerating Mainstream Flow

2005 ◽  
Author(s):  
Yong W. Kim ◽  
Chad Coon ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Mainstream Flow

Pressure-side discharge is commonly employed in turbine blades and nozzle guide vanes to keep the trailing edge metal temperatures within an allowable limit while minimizing aerodynamic penalties. Despite its widespread use, film-cooling data of the discharge slot are scarce in open literature. The objectives of the present experimental study were to measure detailed local heat transfer and film-cooling effectiveness from a 10x scale trailing-edge model of an industrial gas turbine airfoil in a low speed wind tunnel. To simulate the mainstream flow acceleration in vane and blade row passages, a linear velocity gradient was imposed using an adjustable top wall. The present work employed the composite slab quasi-steady liquid crystal method that allows measurements of local heat transfer coefficients and film-cooling effectiveness from two related tests. With this technique, the heat transfer measurement can be performed in a cold wind tunnel. The coolant-to-mainstream blowing ratio was varied between 0.25 and 1.0. The slot hydraulic diameter based Reynolds number ranged from 4,760 to 19,550. The coolant-to-mainstream density ratio was fixed at 0.95. Slot discharge coefficients were also measured with mainstream acceleration. Both local heat transfer coefficients and film-cooling effectiveness displayed a strong dependency on blowing ratio and mainstream acceleration. However, the discharge coefficients showed little dependency on the mainstream acceleration.

An Experimental and Numerical Investigation of Near Cooling Hole Heat Fluxes on a Film-Cooled Turbine Blade

1989 ◽  
Vol 111 (1) ◽  
pp. 63-70 ◽  
Author(s):  
C. Camci
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Internal Heat ◽  
External Heat ◽  
Heat Fluxes ◽  
Temperature Gradients ◽  
Transfer Coefficients ◽  
External Heat Transfer ◽  
Heat Transfer Coefficients ◽  
Cooling Hole

Discrete hole film cooling on highly curved surfaces of a gas turbine blade produces very significant wall temperature gradients and wall heat flux variations near downstream and upstream of rows of circular cooling holes. In this study a set of well-defined external heat transfer coefficient distributions in the presence of discrete hole film cooling is presented. Heat transfer coefficients are measured on the suction side of an HP rotor blade profile in a short-duration facility under well-simulated gas turbine flow conditions. The main emphasis of the study is to evaluate the internal heat flux distributions in a detailed way near the cooling holes by using a computational technique. The method uses the measured external heat transfer coefficients as boundary conditions in addition to available internal heat transfer correlations for the internal passages. The study shows the details of the near hole temperature gradients and heat fluxes. The convective heat transfer inside the circular film cooling holes is shown to be very significant even with their relatively small diameter and lengths compared to the chord length. The study also indicates a nonnegligible wall temperature reduction at near upstream of discrete cooling holes. This is explained with the elliptic nature of the internal conduction field of the blade and relatively low coolant temperature levels at the exit of a film cooling hole compared to the mean blade temperature.

Comparison the Cooling Performance of Cylindrical and Fan-Shaped Cooling Holes With Special Emphasis on the Effect of Internal Coolant Cross-Flow

2008 ◽  
Author(s):  
Christian Saumweber ◽  
Achmed Schulz
Keyword(s):  
Film Cooling ◽  
Symmetry Plane ◽  
Cross Flow ◽  
Transfer Coefficients ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Heat Transfer Coefficients ◽  
Expansion Angle ◽  
Cooling Hole

Cooling holes in real gas turbine applications are prevalently exposed to cross-flow in the coolant passage. The majority of the studies available in literature do not consider the effects of flow in the coolant passage. Our own studies however reveal that especially diffuser holes are very susceptible in respect to cross-flow at the hole entrance, if the orientation of the cross-flow is perpendicular to the symmetry plane of the cooling hole. The effect of coolant cross-flow will be discussed in detail. The superordinate target is to identify the dominating mechanisms, which determine the flow field within the diffuser hole and hence limit the potential of cooling performance augmentation. For this reason a fan-shaped hole with 14° expansion angle will be compared to a simple cylindrical hole. Both holes have a length-to-diameter ratio of 6 and an inclination angle of 30°. The comparison will be performed by means of experimentally gained discharge coefficients, local and laterally averaged adiabatic film cooling effectiveness, and heat transfer coefficients. Numerical simulations of the cooling flow will support the interpretation of the experimental results.

scholarly journals Detailed Film Cooling Measurements over a Gas Turbine Blade Using a Transient Liquid Crystal Image Technique

2001 ◽  
Vol 7 (6) ◽  
pp. 415-424 ◽  
Author(s):  
Hui Du ◽  
Srinath V. Ekkad ◽  
Je-Chin Han ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Liquid Crystal ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Image Technique

Detailed heat transfer coefficient and film effectiveness distributions over a gas turbine blade with film cooling are obtained using a transient liquid crystal image technique. The test blade has three rows of film holes on the leading edge and two rows each on the pressure and suction surfaces. A transient liquid crystal technique maps the entire blade midspan region, and helps provide detailed measurements, particularly near the film hole. Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity is5.3×105. Two different coolants (air andCo2) were used to simulate coolant density effect. Coolant blowing ratio was varied between 0.8 and 1.2 for air injection and 0.4–1.2 forCo2injection. Results show that film injection promotes earlier laminar-turbulent boundary layer transition on the suction surface and also enhances local heat transfer coefficients (up to 80%) downstream of injection. An increase in coolant blowing ratio produces higher heat transfer coefficients for both coolants. This effect is stronger immediately downstream of injection holes. Film effectiveness is highest at a blowing ratio of 0.8 for air injection and at a blowing ratio of 1.2 forCo2injection. Such detailed results will help provide insight into the film cooling phenomena on a gas turbine blade.

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