US20110097191A1 - Method and structure for cooling airfoil surfaces using asymmetric chevron film holes - Google Patents

Method and structure for cooling airfoil surfaces using asymmetric chevron film holes Download PDF

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Publication number
US20110097191A1
US20110097191A1 US12/607,586 US60758609A US2011097191A1 US 20110097191 A1 US20110097191 A1 US 20110097191A1 US 60758609 A US60758609 A US 60758609A US 2011097191 A1 US2011097191 A1 US 2011097191A1
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United States
Prior art keywords
film
region
trough
airfoil
trough region
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
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US12/607,586
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English (en)
Inventor
Ronald Scott Bunker
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General Electric Co
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General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US12/607,586 priority Critical patent/US20110097191A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BUNKER, RONALD SCOTT
Assigned to ENERGY, UNITED STATES DEPARTMENT OF reassignment ENERGY, UNITED STATES DEPARTMENT OF CONFIRMATORY LICENSE (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL ELECTRIC COMPANY
Assigned to UNITED STATES DEPARTMENT OF ENERGY reassignment UNITED STATES DEPARTMENT OF ENERGY CONFIRMATORY LICENSE (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL ELECTRIC COMPANY
Priority to DE102010037050A priority patent/DE102010037050A1/de
Priority to JP2010184516A priority patent/JP5738555B2/ja
Priority to CH01344/10A priority patent/CH702107B1/de
Priority to CN201010272834.XA priority patent/CN102052092B/zh
Publication of US20110097191A1 publication Critical patent/US20110097191A1/en
Abandoned legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates generally to film-cooled parts and more particularly to a method of film cooling common locations on airfoil surfaces using asymmetric chevron film holes.
  • Film cooling refers to a technique for cooling a part in which cool air is discharged through a plurality of small holes in the external walls of the part to provide a thin, cool barrier along the external surface of the part and prevent or reduce direct contact with hot gasses.
  • film cool vane and blade airfoils include, among others, the leading edge, pressure side, and suction side, as well as endwall film cooling, including the inner and outer vane endwalls and the blade platforms.
  • Film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces. This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience. There is significant migration of film cooling across the flow passage, making injection and efficient cooling very difficult.
  • Injection film holes are generally either round or diffuser shaped. These holes are oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling in others.
  • a film-cooled airfoil or airfoil region is configured with one or more asymmetric chevron film cooling holes.
  • a film-cooled turbine structure comprises at least one asymmetric chevron film cooling hole such that one side of the chevron is dominant over the other side of the chevron with respect to directing a portion of injected coolant onto a surface of the film-cooled turbine structure.
  • FIG. 1 is a top view illustrating a chevron film-cooling hole known in the art
  • FIG. 2 is a side view of the film-cooling hole depicted in FIG. 1 ;
  • FIG. 3 is a frontal view of the film-cooling hole depicted in FIG. 1 ;
  • FIG. 4 is a perspective view illustrating film-cooling holes disposed in the endwall region of a turbine vane
  • FIG. 5 is a perspective view illustrating film-cooling holes disposed in the endwall region of a turbine blade
  • FIG. 6 illustrates injection film holes oriented with injection along the approximate direction of the local surface streamline for the endwall region of a turbine vane
  • FIG. 7 illustrates an asymmetric chevron film hole suitable for use on airfoils or endwalls
  • FIG. 8 illustrates two asymmetric chevron regions in which each chevron region comprises a pair of wing troughs with dissimilar geometries respect to one another.
  • Chevron film holes have proven beneficial for improving film effectiveness on airfoil surfaces.
  • Present chevron film holes are always based on symmetric designs about the film hole centerline.
  • Film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces, as stated herein before.
  • This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience.
  • the airfoil surfaces do however experience such secondary flow effects, but generally to a much lesser degree, except in regions where the airfoils meet the endwall regions.
  • Injection film holes are generally either round or diffuser shaped. These holes are usually oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling in others.
  • Asymmetric chevron film hole embodiments achieving similar fluid flow benefits in regions and uses where the surface fluid streamline curvature is significant are described herein. These embodiments, while still using a single circular through-hole, alter the two halves of the chevron footprint to have differing sizes or orientations of troughs. This asymmetry advantageously makes one side of the chevron dominant over the other side with respect to directing a portion of the injected coolant onto the surface to be cooled. The dominant or major side of the chevron should be directed/oriented to counteract the streamline curvature imposed by the hot gases.
  • FIG. 1 is a top view illustrating a symmetric chevron film-cooling hole 10 known in the art.
  • the ridge 12 is outwardly convex laterally in depth between the two wing toughs 14 .
  • the convex ridge 12 is arcuate and generally triangular in profile, and diverges in the downstream direction between the inlet bore 16 and the junction of its downstream end with the outer surface 18 .
  • the trailing edge of the ridge 12 blends flush with the outer surface 18 along a laterally arcuate downstream end of the chevron outlet, with the convex trailing edge being bowed upstream toward the inlet hole 16 .
  • the curved form of the compound chevron film cooling hole 10 enjoys the advantages of compound inclination angles A, B as illustrated in further detail in FIG. 2 that depicts a side view of chevron film-cooling hole 10 relative to the flow of hot gas 20 in which the chevron outlet diverges aft from the inlet hole 16 differently inclined at an inclination angle A. More specifically, inclination angles A and B are the two limiting angles, one along the centerline and the other in each trough.
  • FIG. 3 is a frontal view of the symmetric film-cooling hole 10 depicted in FIG. 1 . This frontal view is in the direction of the hot gas 20 shown in FIG. 2 .
  • Chevron film hole 10 is based on a symmetric design about the film hole 10 centerline shown in FIG. 1 , and illustrates further details of ridge 12 and wing troughs 14 .
  • FIG. 4 is a perspective view illustrating film-cooling holes disposed in the endwall regions 22 of a turbine vane 24
  • FIG. 5 is a perspective view illustrating film-cooling holes 26 disposed in the endwall region of a turbine blade 28 .
  • film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces, as stated herein before.
  • the airfoil surfaces do however experience such secondary flow effects, but generally to a much lesser degree, except in regions where the airfoils meet the endwall regions, also stated above.
  • This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience, causing significant migration of film cooling across the flow passage, making injection and efficient cooling very difficult.
  • FIG. 6 illustrates round injection film holes oriented with injection along the approximate direction of the local surface streamline for the endwall region of a turbine vane 29 .
  • round injection film holes are shown in FIG. 6
  • the injection film holes are generally either round or diffuser shaped. These holes are oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling other regions, as stated above.
  • FIG. 7 is a top view illustrating an asymmetric chevron film hole 30 suitable for use in the endwall regions depicted in FIGS. 4 and 5 , according to one embodiment of the invention.
  • a flat ridge 32 is increasing laterally in width between the two wing troughs 34 , 36 .
  • the ridge 32 is flat and generally triangular in profile, and diverges in the downstream direction between an inlet bore 38 and the junction of its downstream end with the outer surface 40 .
  • the trailing edge of the ridge 32 blends flush with the outer surface 40 along a laterally flat downstream end of the chevron outlet.
  • the size of wing trough 34 is different from the size of wing trough 36 such that the wing troughs 34 , 36 each blend into the surrounding portions of the inlet bore 38 differently with respect to one another.
  • a curved form of the asymmetric chevron film cooling hole 30 not shown, having compound inclination angles in which the chevron outlet diverges aft from the inlet bore 38 will enjoy similar advantages to those described above for compound inclination angels B, C shown in FIG. 2 for a symmetric chevron film-cooling hole 10 .
  • Asymmetric chevron film cooling hole 30 is substantially different from symmetric film cooling hole 10 in that particular asymmetric film cooling hole 30 embodiments my include differing tough depths, differing trough widths, differing trough diffusion angles, differing trough shaping, and so on such as depicted in FIG. 8 .
  • FIG. 8 illustrates two asymmetric chevron regions in which each chevron region comprises a pair of wing troughs with dissimilar geometries respect to one another.
  • wing trough 34 has a depth that is different from the depth of wing trough 36 .
  • wing trough 34 has a width that is different from the width of wing trough 36 .
  • wing trough 34 has a diffusion angle B that is different from diffusion angle C.
  • wing trough 34 has a shape that is different from the shape of wing trough 36 .
  • asymmetric chevron film cooling hole 30 employs a single round through-hole feeding the coolant to the chevron region.
  • an asymmetric chevron film cooling hole is described herein for improving film cooling for a variety of airfoil surfaces, particularly in regions and applications where the surface fluid streamline curvature is significant.
  • This use of asymmetric film cooling holes allows for more efficient injection of film cooling on a surface in the presence of a strong lateral pressure gradient seeking to move the film coolant away from the intended region to be protected, and keeps the film coolant in the desired region(s) without creating undue mixing losses. Higher mixing losses result from conventional film holes that simply inject the flow crossways to the main hot gas to attempt to counteract the pressure gradients. More efficient coolant use leads to higher efficiency engines such as industrial engines with longer lives.
  • Asymmetric chevron film cooling holes offer advantages beyond those achievable using known trial and error placement of film holes until a compromise of cooling adequacy and losses is found, or beyond those achievable simply by adding diffuser shaping to round film holes, and possibly a compound angle on the diffuser to help direct coolant in the desired direction.
  • Asymmetric chevron film cooling holes further offer advantages beyond those achievable by simply altering the shape of the endwall itself to help mitigate secondary flows and pressure gradients, rather than modifying the film holes.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US12/607,586 2009-10-28 2009-10-28 Method and structure for cooling airfoil surfaces using asymmetric chevron film holes Abandoned US20110097191A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US12/607,586 US20110097191A1 (en) 2009-10-28 2009-10-28 Method and structure for cooling airfoil surfaces using asymmetric chevron film holes
DE102010037050A DE102010037050A1 (de) 2009-10-28 2010-08-18 Verfahren und Struktur zum Kühlen von Schaufelblattflächen unter Anwendung asymmetrischer Winkel-Filmlöcher
JP2010184516A JP5738555B2 (ja) 2009-10-28 2010-08-20 非対称シェブロンフィルム孔を使用して翼形部表面を冷却するための方法及び構造
CH01344/10A CH702107B1 (de) 2009-10-28 2010-08-23 Turbinenbauteil mit asymmetrischen, winkelförmigen Filmkühlungslöchern.
CN201010272834.XA CN102052092B (zh) 2009-10-28 2010-08-27 使用非对称人字形薄膜孔来冷却翼型表面的方法和结构

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Application Number Priority Date Filing Date Title
US12/607,586 US20110097191A1 (en) 2009-10-28 2009-10-28 Method and structure for cooling airfoil surfaces using asymmetric chevron film holes

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US20110097191A1 true US20110097191A1 (en) 2011-04-28

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US12/607,586 Abandoned US20110097191A1 (en) 2009-10-28 2009-10-28 Method and structure for cooling airfoil surfaces using asymmetric chevron film holes

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US (1) US20110097191A1 (zh)
JP (1) JP5738555B2 (zh)
CN (1) CN102052092B (zh)
CH (1) CH702107B1 (zh)
DE (1) DE102010037050A1 (zh)

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US20130156549A1 (en) * 2011-12-15 2013-06-20 Jaime Javier Maldonado Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components
US8522558B1 (en) 2012-02-15 2013-09-03 United Technologies Corporation Multi-lobed cooling hole array
US8572983B2 (en) 2012-02-15 2013-11-05 United Technologies Corporation Gas turbine engine component with impingement and diffusive cooling
US8584470B2 (en) 2012-02-15 2013-11-19 United Technologies Corporation Tri-lobed cooling hole and method of manufacture
WO2014008016A1 (en) * 2012-07-02 2014-01-09 United Technologies Corporation Airfoil cooling arrangement
WO2014007961A1 (en) * 2012-07-02 2014-01-09 United Technologies Corporation Airfoil cooling arrangement
US8683813B2 (en) 2012-02-15 2014-04-01 United Technologies Corporation Multi-lobed cooling hole and method of manufacture
US8683814B2 (en) 2012-02-15 2014-04-01 United Technologies Corporation Gas turbine engine component with impingement and lobed cooling hole
US8689568B2 (en) 2012-02-15 2014-04-08 United Technologies Corporation Cooling hole with thermo-mechanical fatigue resistance
US20140099189A1 (en) * 2012-10-04 2014-04-10 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
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US20140271229A1 (en) * 2011-12-15 2014-09-18 Ihi Corporation Turbine blade
US8850828B2 (en) 2012-02-15 2014-10-07 United Technologies Corporation Cooling hole with curved metering section
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US9024226B2 (en) 2012-02-15 2015-05-05 United Technologies Corporation EDM method for multi-lobed cooling hole
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US9416971B2 (en) 2012-02-15 2016-08-16 United Technologies Corporation Multiple diffusing cooling hole
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US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
US9976746B2 (en) 2015-09-02 2018-05-22 General Electric Company Combustor assembly for a turbine engine
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US10401029B2 (en) * 2015-06-05 2019-09-03 Rolls-Royce Deutschland Ltd & Co Kg Device for cooling a wall of a component of a gas turbine
US10422230B2 (en) 2012-02-15 2019-09-24 United Technologies Corporation Cooling hole with curved metering section
US10563867B2 (en) 2015-09-30 2020-02-18 General Electric Company CMC articles having small complex features for advanced film cooling
US10605092B2 (en) 2016-07-11 2020-03-31 United Technologies Corporation Cooling hole with shaped meter
CN112031877A (zh) * 2020-08-21 2020-12-04 天津理工大学 一种展向非对称的凹坑气膜冷却孔型
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JP5738555B2 (ja) 2015-06-24
CN102052092B (zh) 2016-01-20
JP2011094609A (ja) 2011-05-12

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