WO2016118136A1 - Profil aérodynamique de turbine - Google Patents

Profil aérodynamique de turbine Download PDF

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Publication number
WO2016118136A1
WO2016118136A1 PCT/US2015/012368 US2015012368W WO2016118136A1 WO 2016118136 A1 WO2016118136 A1 WO 2016118136A1 US 2015012368 W US2015012368 W US 2015012368W WO 2016118136 A1 WO2016118136 A1 WO 2016118136A1
Authority
WO
WIPO (PCT)
Prior art keywords
wall
airfoil
chordwise extending
grooves
chordwise
Prior art date
Application number
PCT/US2015/012368
Other languages
English (en)
Inventor
Ching-Pang Lee
Original Assignee
Siemens Energy, Inc.
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 Siemens Energy, Inc. filed Critical Siemens Energy, Inc.
Priority to PCT/US2015/012368 priority Critical patent/WO2016118136A1/fr
Publication of WO2016118136A1 publication Critical patent/WO2016118136A1/fr

Links

Classifications

    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • 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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F01D5/187—Convection cooling
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/181—Two-dimensional patterned ridged
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/182—Two-dimensional patterned crenellated, notched
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/183—Two-dimensional patterned zigzag
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/184—Two-dimensional patterned sinusoidal
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/28—Three-dimensional patterned
    • F05D2250/281—Three-dimensional patterned threaded
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/294—Three-dimensional machined; miscellaneous grooved
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • This invention is directed generally to turbine airfoils, and more particularly to leading edges in hollow turbine airfoils of gas turbine engines.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures.
  • turbine airfoils must be made of materials capable of withstanding such high temperatures.
  • the leading edge of a turbine airfoil is exposed to high temperatures and is subjected to higher heat loads than other aspects of the airfoil.
  • Turbine airfoils often contain cooling systems for prolonging the life of the airfoils and reducing the likelihood of failure as a result of excessive temperatures at the leading edge and in other areas of the airfoils.
  • the cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine airfoil at a relatively uniform temperature.
  • centrifugal forces and air flow at boundary layers often prevent some areas of the turbine airfoil from being adequately cooled, which results in the formation of localized hot spots.
  • Localized hot spots depending on their location, can reduce the useful life of a turbine airfoil and can damage a turbine airfoil to an extent necessitating replacement of the airfoil.
  • the leading edge of turbine airfoils includes a plurality of film cooling holes forming a showerhead. While the showerhead of film cooling holes cools the leading edge, conventional showerhead configurations are often inefficient.
  • a turbine airfoil usable in a turbine engine and having a leading edge formed from an outer wall with a nonlinear outer surface or a nonlinear inner surface, or both, for improved cooling is disclosed.
  • the outer surface of the leading edge may include one or more chordwise extending grooves that provide a thermal barrier coating with an improved ability to attach to the outer surface of the leading edge and to remain attached during turbine engine operation.
  • the inner surface of the leading edge may include one or more chordwise extending grooves that create a nonlinear surface which provides an increased surface area within an internal cooling system of the turbine airfoil.
  • the nonlinear surface increases the efficiency of the cooling system by increasing cooling capacity of the cooling system at the leading edge, thereby reducing the cooling air requirement of the cooling system.
  • the nonlinear outer surface may eliminate the need for showerhead cooling holes in the leading edge.
  • a turbine airfoil may include a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side, a suction side, and a cooling system positioned within interior aspects of the generally elongated hollow airfoil and formed from at least one cavity.
  • the outer wall forming the generally elongated hollow airfoil may form the leading edge of the generally elongated hollow airfoil.
  • An outer surface of the outer wall forming the leading edge may be a nonlinear surface formed by one or more chordwise extending grooves that are formed from one or more valleys defined on each side by one or more ridges.
  • At least a portion of the chordwise extending groove may extend onto the pressure side of the generally elongated hollow airfoil, and at least a portion of the chordwise extending groove may extend onto the suction side of the generally elongated hollow airfoil.
  • the valley of the chordwise extending groove may be curved, and one or more ridges of the chordwise extending groove may be curved.
  • the chordwise extending groove may be formed from a plurality of chordwise extending grooves. The plurality of chordwise extending grooves may all be aligned with each other. One or more valleys 42 of each groove of the plurality of grooves may be aligned with each other along a generally spanwise extending valley linear axis.
  • At least one ridge of each groove of the plurality of grooves may be aligned with each other along a generally spanwise extending ridge linear axis.
  • the turbine airfoil may include one or more thermal barrier coatings on the outer surface of the outer wall forming the leading edge. The thermal barrier coating may contact the valley defined on each side by at least one ridge of the chordwise extending groove.
  • the outer wall forming the generally elongated hollow airfoil and forming the leading edge of the generally elongated hollow airfoil may include one or more inner surfaces defining at least a portion of the cooling system positioned within interior aspects of the generally elongated hollow airfoil.
  • the inner surface of the outer wall forming the leading edge may be a nonlinear surface formed by one or more chordwise extending grooves that are formed from at least one valley defined on each side by at least one ridge.
  • the one or more valleys of the chordwise extending groove on the inner surface of the outer wall may be curved, and the at least one ridge of the chordwise extending groove on the inner surface of the outer wall may be curved.
  • the chordwise extending groove on the inner surface of the outer wall may be formed from a plurality of chordwise extending grooves, each formed by one or more valleys defined on each side by at least one ridge.
  • the plurality of chordwise extending grooves on the inner surface of the outer wall may all be aligned with each other. At least one valley of each groove of the plurality of grooves on the inner surface of the outer wall may be aligned with each other along a generally spanwise extending valley linear axis, and wherein at least one ridge of each groove of the plurality of grooves on the inner surface of the outer wall may be aligned with each other along a generally spanwise extending ridge linear axis.
  • a distance between the generally spanwise extending valley linear axis and the generally spanwise extending ridge linear axis of the chordwise extending grooves on the inner surface of the outer wall may be less than a distance between a generally spanwise extending valley linear axis and generally spanwise extending ridge linear axis of chordwise extending grooves on the outer surface of the outer wall.
  • a distance between adjacent ridges of the plurality of chordwise extending grooves on the inner surface of the outer wall may be less than a distance between adjacent ridges of a plurality of chordwise extending grooves on the outer surface of the outer wall.
  • cooling fluids such as, but not limited to, air
  • the cooling fluid may enter the cavity and flow into contact with the inner surface of the outer wall.
  • the chordwise extending grooves increase the surface area of the inner surface, thereby increasing the overall convection at the leading edge within the internal cooling system.
  • the increased convection at the leading edge provides increased cooling capacity, thereby requiring less cooling fluid at the leading edge, which increases the efficiency of the internal cooling system.
  • the chordwise extending grooves on the outer surface may provide an enhanced attachment system for attaching a thermal barrier coating to the leading edge of the turbine airfoil.
  • the chordwise extending grooves on the outer surface provide physical structure to the thermal barrier coating that enables to the thermal barrier coating to better resist the hot gas path within the turbine engine downstream from the combustor.
  • Figure 1 is a perspective view of a turbine airfoil, in particular, a turbine vane, with an internal cooling system including chordwise extending grooves on the leading edge.
  • Figure 2 is a perspective view of a turbine airfoil, in particular, a turbine blade, with an internal cooling system including chordwise extending grooves on the leading edge.
  • Figure 3 is a cross-sectional view of the turbine vane of Figure 1 taken at section line 3-3 in Figure 1 .
  • Figure 4 is a cross-sectional view of the turbine blade of Figure 2 taken at section line 4-4 in Figure 2.
  • Figure 5 is a cross-sectional view of the leading edge of the turbine vane of Figure 3 taken at section line 5-5 in Figure 3 and the leading edge of the turbine blade of Figure 4 taken at section line 5-5 in Figure 4.
  • Figure 6 is a perspective view of a turbine airfoil with chordwise extending grooves on the leading edge and extending onto portions of the pressure side of the airfoil and onto portions of the suction side of the airfoil.
  • a turbine airfoil 10 usable in a turbine engine and having a leading edge 14 formed from an outer wall 16 with a nonlinear outer surface 18 or a nonlinear inner surface 20, or both, for improved cooling is disclosed.
  • the outer surface 18 of the leading edge 14 may include one or more chordwise extending grooves 22 that provide a thermal barrier coating 24 with an improved ability to attach to the outer surface 18 of the leading edge 14 and to remain attached during turbine engine operation.
  • the inner surface 20 of the leading edge 14 may include one or more chordwise extending grooves 24 that create a nonlinear surface 20 which provides an increased surface area within an internal cooling system 26 of the turbine airfoil.
  • the nonlinear surface 20 increases the efficiency of the cooling system 26 by increasing cooling capacity of the cooling system 26 at the leading edge 14, thereby reducing the cooling air requirement of the cooling system 26.
  • the nonlinear outer surface 18 may eliminate the need for showerhead cooling holes in the leading edge 14.
  • the turbine airfoil 10 may be formed from a generally elongated hollow airfoil 30 formed from an outer wall 16, and having a leading edge 14, a trailing edge 34, a pressure side 36, a suction side 38, and a cooling system 26 positioned within interior aspects of the generally elongated hollow airfoil 30 and formed from one or more cavities 40.
  • the configuration of the turbine airfoil 10 is not limited to a particular configuration but may have any appropriate configuration including a leading edge 14.
  • the turbine airfoils 10 may be, but is not limited to being, a stationary turbine vane or a rotatable turbine blade.
  • the outer wall 16 forming the generally elongated hollow airfoil 30 may form the leading edge 14 of the generally elongated hollow airfoil 30.
  • An outer surface 18 of the outer wall 16 forming the leading edge 14 may be a nonlinear surface 18 formed by one or more chordwise extending grooves 22 that is formed from one or more valleys 42 defined on each side by one or more ridges 44. At least a portion 46 of the chordwise extending groove 22 may extend onto the pressure side 36 of the generally elongated hollow airfoil 30. At least a portion 48 of the chordwise extending groove 22 may extend onto the suction side 38 of the generally elongated hollow airfoil 30.
  • a valley 42 of the chordwise extending groove 22 may be curved, and a ridge 44 of the chordwise extending groove 22 may be curved.
  • the curved portion of the valley 42 may extend for all or a portion of a length of the chordwise extending groove 22.
  • the curved portion of the ridge 44 may extend for all or a portion of a length of the chordwise extending groove 22.
  • the leading edge 14 of the turbine airfoil 10 may include a plurality of chordwise extending grooves 22.
  • the plurality of chordwise extending grooves 22 may be aligned with each other.
  • the chordwise extending grooves 22 may be aligned, such as parallel to each other, when viewing the airfoil, as shown in Figures 5 and 6.
  • the valley 42 of each groove 22 of the plurality of grooves 22 may be aligned with each other along a generally spanwise extending valley linear axis 50
  • a ridge 44 of each groove 22 of the plurality of grooves 22 may be aligned with each other along a generally spanwise extending ridge linear axis 52.
  • the turbine airfoil 10 may include one or more thermal barrier coatings 24 on the outer surface 18 of the outer wall 16 forming the leading edge 14.
  • the thermal barrier coating 24 may contact the valley 42 defined on each side by one or more ridges 44 of the chordwise extending groove 22.
  • the thermal barrier coating 24 is not limited to any particular coating but may be formed from any appropriate material currently existing or heretofore have yet to be invented that are capable of withstanding the harsh, high temperature environment of the hot gas path in a turbine engine downstream of a combustor.
  • the outer wall 16 forming the generally elongated hollow airfoil 30 and forming the leading edge 14 of the generally elongated hollow airfoil 30 may include one or more inner surfaces 20 defining at least a portion of the cooling system 26 positioned within interior aspects of the generally elongated hollow airfoil 30.
  • the inner surface 20 of the outer wall 16 forming the leading edge 14 may be a nonlinear surface 20 formed by one or more chordwise extending grooves 56 that is formed from one or more valleys 58 defined on each side by one or more ridges 60.
  • the valley 58 of the chordwise extending groove 56 on the inner surface 20 of the outer wall 16 may be curved.
  • the ridge 60 of the chordwise extending groove 56 on the inner surface 20 of the outer wall 16 may be curved.
  • the turbine airfoil 10 may include a plurality of chordwise extending grooves 56, each formed by one or more valleys 58 defined on each side by one or more ridges 44.
  • the plurality of chordwise extending grooves 56 on the inner surface 20 of the outer wall 16 may all aligned with each other.
  • the plurality of chordwise extending grooves 56 on the inner surface 20 of the outer wall 16 may be aligned similarly to the chordwise extending grooves 22 on the outer surface 18 of the outer wall 16.
  • the valley 42 of each groove 56 of the plurality of grooves 56 on the inner surface 20 of the outer wall 16 may be aligned with each other along a generally spanwise extending valley linear axis 62, and a ridge 60 of each groove 56 of the plurality of grooves 56 on the inner surface 20 of the outer wall 16 may be aligned with each other along a generally spanwise extending ridge linear axis 64.
  • a distance between the generally spanwise extending valley linear axis 62 and the generally spanwise extending ridge linear axis 64 of the chordwise extending grooves 56 on the inner surface 20 of the outer wall 16 may be less than a distance between a generally spanwise extending valley linear axis 50 and generally spanwise extending ridge linear axis 52 of chordwise extending grooves 56 on the outer surface 18 of the outer wall 16.
  • the configuration of the grooves 22 on the outer surface 18 relative to the grooves 56 on the inner surface 20 may be such that a depth of the grooves 56 on the inner surface 20 is less than a depth of the grooves 22 on the outer surface 18.
  • a distance between adjacent ridges 60 of the plurality of chordwise extending grooves 56 on the inner surface 20 of the outer wall 16 may be less than a distance between adjacent ridges 44 of a plurality of chordwise extending grooves 22 on the outer surface 18 of the outer wall 16.
  • cooling fluids such as, but not limited to, air
  • the cooling fluid may enter the cavity 40 and flow into contact with the inner surface 20 of the outer wall 16.
  • the chordwise extending grooves 56 increase the surface area of the inner surface 20, thereby increasing the overall convection at the leading edge 14 within the internal cooling system 26.
  • the increased convection at the leading edge 14 provides increased cooling capacity, thereby requiring less cooling fluid at the leading edge 14, which increases the efficiency of the internal cooling system 26.
  • the chordwise extending grooves 22 on the outer surface 18 may provide an enhanced attachment system for attaching a thermal barrier coating 24 to the leading edge 14 of the turbine airfoil 10.
  • the chordwise extending grooves 22 on the outer surface 18 provide physical structure to the thermal barrier coating 24 that enables to the thermal barrier coating 24 to better resist the hot gas path within the turbine engine downstream from the combustor.
  • the chordwise extending grooves 22 on the outer surface 18 limits the propagation of thermal barrier coating spallation and provide better thermal barrier coating attachment due to increased bonding surface area.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne un profil aérodynamique (10) de turbine pouvant être utilisé dans une turbine à vapeur et présentant une paroi extérieure (16) formée à partir d'une paroi extérieure (16) de surface extérieure (18) non linéaire ou de surface intérieure (20) non linéaire, ou les deux, pour un refroidissement amélioré. Dans au moins un mode de réalisation, la surface extérieure (18) de la paroi extérieure (16) peut comporter une ou plusieurs rainures (22) s'étendant dans le sens de la corde fournissant un revêtement de barrière thermique (24) ayant une capacité améliorée à se fixer à la surface extérieure (18) de la paroi extérieure (16) et à rester fixé durant le fonctionnement de la turbine à vapeur. La surface intérieure (20) de la paroi extérieure (16) peut comporter une ou plusieurs rainures (58) s'étendant dans le sens de la corde qui créent une surface non linéaire ce qui fournit une superficie accrue à l'intérieur d'un système de refroidissement (26) interne du profil aérodynamique de turbine. La surface non linéaire augmente l'efficacité du système de refroidissement (26) en augmentant la capacité de refroidissement du système de refroidissement (26) au niveau de la paroi externe (16), ce qui permet de réduire les besoins en air de refroidissement du système de refroidissement (26).
PCT/US2015/012368 2015-01-22 2015-01-22 Profil aérodynamique de turbine WO2016118136A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2015/012368 WO2016118136A1 (fr) 2015-01-22 2015-01-22 Profil aérodynamique de turbine

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Application Number Priority Date Filing Date Title
PCT/US2015/012368 WO2016118136A1 (fr) 2015-01-22 2015-01-22 Profil aérodynamique de turbine

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Publication Number Publication Date
WO2016118136A1 true WO2016118136A1 (fr) 2016-07-28

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20190024520A1 (en) * 2017-07-19 2019-01-24 Micro Cooling Concepts, Inc. Turbine blade cooling
TWI761815B (zh) * 2019-06-05 2022-04-21 日商三菱動力股份有限公司 燃氣輪機的葉片的修補方法及燃氣輪機的葉片

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5337568A (en) * 1993-04-05 1994-08-16 General Electric Company Micro-grooved heat transfer wall
US20090087312A1 (en) * 2007-09-28 2009-04-02 Ronald Scott Bunker Turbine Airfoil Concave Cooling Passage Using Dual-Swirl Flow Mechanism and Method
US20130101806A1 (en) * 2011-10-19 2013-04-25 General Electric Company Method for adhering a coating to a substrate structure

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5337568A (en) * 1993-04-05 1994-08-16 General Electric Company Micro-grooved heat transfer wall
US20090087312A1 (en) * 2007-09-28 2009-04-02 Ronald Scott Bunker Turbine Airfoil Concave Cooling Passage Using Dual-Swirl Flow Mechanism and Method
US20130101806A1 (en) * 2011-10-19 2013-04-25 General Electric Company Method for adhering a coating to a substrate structure

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20190024520A1 (en) * 2017-07-19 2019-01-24 Micro Cooling Concepts, Inc. Turbine blade cooling
TWI761815B (zh) * 2019-06-05 2022-04-21 日商三菱動力股份有限公司 燃氣輪機的葉片的修補方法及燃氣輪機的葉片

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