WO2015109040A1 - Système de refroidissement interne ayant un insert ondulé formant des canaux de refroidissement côté paroi proche pour profil aérodynamique de turbine à gaz - Google Patents

Système de refroidissement interne ayant un insert ondulé formant des canaux de refroidissement côté paroi proche pour profil aérodynamique de turbine à gaz Download PDF

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Publication number
WO2015109040A1
WO2015109040A1 PCT/US2015/011496 US2015011496W WO2015109040A1 WO 2015109040 A1 WO2015109040 A1 WO 2015109040A1 US 2015011496 W US2015011496 W US 2015011496W WO 2015109040 A1 WO2015109040 A1 WO 2015109040A1
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WO
WIPO (PCT)
Prior art keywords
wall
trailing edge
cooling channel
subpartition
partition walls
Prior art date
Application number
PCT/US2015/011496
Other languages
English (en)
Inventor
Ching-Pang Lee
Jae Y. Um
Gerald L. Hillier
Eric Schroeder
Mohamed Abdullah
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Publication of WO2015109040A1 publication Critical patent/WO2015109040A1/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
    • F01D5/14—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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F05D2240/127—Vortex generators, turbulators, or the like, for mixing
    • 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
    • F05D2250/60—Structure; Surface texture
    • F05D2250/61—Structure; Surface texture corrugated
    • 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

Definitions

  • This invention is directed generally to gas turbine engines, and more particularly to internal cooling systems for airfoils in 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 vane and blade assemblies to high temperatures. As a result, turbine vanes and blades must be made of materials capable of
  • Turbine engines typically include a plurality of rows of stationary turbine vanes extending radially inward from a shell and include a plurality of rows of rotatable turbine blades attached to a rotor assembly for turning the rotor.
  • the turbine vanes are exposed to high temperature combustor gases that heat the airfoil.
  • the airfoils include internal cooling systems for reducing the temperature of the airfoils. While there exist many configurations of cooling systems, there exists a need for improved cooling of gas turbine airfoils.
  • An airfoil for a gas turbine engine in which the airfoil includes an internal cooling system formed from one or more midchord cooling channels with a
  • corrugated insert positioned therein and creating nearwall leading edge, pressure side and suction side nearwall cooling channels.
  • the corrugated insert may be formed from a wall that oscillates in a repeating pattern between peaks and valleys, such that the peaks are closer to an inner surface of an outer wall forming a generally elongated hollow airfoil of the airfoil.
  • the corrugated insert may work in concert with the rows of partition walls to create periodic impingement on the inner surface of the outer wall, specifically the surfaces leading from the valleys to the peaks moving toward a trailing edge of the generally elongated hollow airfoil direct the fluids toward the outer wall.
  • Such cooling system may provide adequate cooling for use in environments in which few, if any, cooling holes are desired, such as in crude oil engine applications.
  • the turbine airfoil may be formed from 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.
  • the cooling system may include one or more midchord cooling channels with one or more corrugated inserts positioned therein and creating a leading edge nearwall cooling channel between the leading edge and the corrugated insert, a suction side nearwall cooling channel between the suction side and the corrugated insert and a pressure side nearwall cooling channel between the pressure side and the corrugated insert.
  • the corrugated insert may be formed from a wall that oscillates in a repeating pattern between peaks and valleys, such that the peaks are closer to an inner surface of the outer wall forming the generally elongated hollow airfoil.
  • a plurality of rows of partition walls may extend from the inner surface forming the outer wall and into the midchord cooling channel.
  • the partition walls may include gaps therein, and the rows of partition walls may be generally aligned with a direction of cooling fluid flow from the leading edge chordwise toward the trailing edge. At least a portion of the peaks in the corrugated insert extending radially outward may be aligned with the gaps within at least a portion of at least one partition wall.
  • one or more partition walls may include gaps therein that extend for a portion of or in a repeating pattern for an entire length of the at least one partition wall.
  • one or more partition walls may be linear.
  • each of the partition walls may be linear and each of the partition walls may be parallel with each other.
  • one or more partition walls may be formed from first and second subpartition walls, whereby the second subpartition wall may be offset laterally from first subpartition wall, and the first and second subpartition walls may be staggered in an alternating manner moving downstream in a direction from the leading edge toward the trailing edge.
  • each partition wall may be formed from a first and second subpartition walls, whereby the second subpartition wall may be offset laterally from first subpartition wall, and the first and second subpartition walls may be staggered in an alternating manner moving in a direction from the leading edge toward the trailing edge.
  • a second subpartition wall may be positioned along an axis that is equidistant from a first subpartition wall within a same partition wall as the second subpartition wall and from a first subpartition wall within an adjacent partition wall.
  • the turbine airfoil may also include a plurality of miniribs extending from the inner surface forming the outer wall and extending between adjacent partition walls.
  • the miniribs may be nonparallel with the partition walls and may be shorter in height extending from the inner surface than adjacent partition ribs.
  • the plurality of miniribs may be aligned with each other and may be orthogonal to the adjacent partition walls.
  • the cooling system may also include one or more trailing edge cooling channels positioned between the midchord cooling channel and the trailing edge of the generally elongated hollow airfoil.
  • the trailing edge cooling channel may include a plurality of pin fins extending from the outer wall forming the pressure side to the suction side and forming zigzag cooling flow channels within the trailing edge cooling channel. At least a portion of the pin fins may have cross-sectional areas with a leading edge and a trailing edge that is positioned on a downstream corner and separated from the leading edge by a concave side surface and a convex side surface.
  • At least one of the zigzag cooling flow channels may be formed from a plurality of pin fins aligned such that the concave and convex side surfaces alternate moving towards the trailing edge.
  • a first upstream pin fin may have a cross-sectional area formed from an upstream section and a downstream section.
  • the upstream section may be generally linear with a constant width
  • the downstream section may be generally linear with a tapered width that reduces in width moving towards the trailing edge.
  • the upstream and downstream sections may be nonlinear and nonorthogonal with each other.
  • cooling fluids are supplied from a compressor or other such source to the inner aspect of the corrugated insert of the internal cooling system.
  • Cooling fluids are passed through the inlets into the leading edge nearwall cooling channel where the fluids separate with a portion flowing into the pressure side nearwall cooling channel and a portion flowing into the suction side nearwall cooling channel.
  • the peaks and valleys of the corrugated insert create periodic
  • the partition walls also direct the cooling fluids towards the trailing edge, and in at least come embodiments, create a nonlinear flow path toward the trailing edge to increase convection of heat from the outer wall to the cooling fluid.
  • the cooling fluid may be exhausted through film cooling holes at downstream ends of the pressure and suction side nearwall cooling channels.
  • Cooling fluids may also be passed to the trailing edge cooling channel via one or more trailing edge supply channels.
  • the cooling fluids may be metered through metering holes into the trailing edge cooling channel.
  • the cooling fluids may flow into the zigzag cooling flow channels where the pin fins direct the cooling fluids in a nonlinear motion to the trailing edge, where the cooling fluids are exhausted.
  • the cooling fluids in the trailing edge cooling channel receive heat from the outer wall forming the pressure and suction sides and from the pin fins.
  • An advantage of the internal cooling system is that the corrugated insert creates periodic impingement and sufficient cooling fluid mixing to cool the airfoil.
  • Figure 1 is a perspective view of a turbine vane with aspects of this invention.
  • Figure 2 is a cross-sectional view of the turbine vane taken at section line 2-2 in Figure 1 .
  • Figure 3 is a detailed view of the pressure side nearwall cooling channel.
  • Figure 4 is a detailed view of the trailing edge cooling channel of the cooling system shown in Figure 2.
  • Figure 5 is a detailed view of the midchord cooling channel of the cooling system shown in Figure 2.
  • Figure 6 is a side view of a three dimensional model of the cooling system within the turbine vane in Figure 1 .
  • Figure 7 is a perspective view of the three dimensional model of the cooling system shown in Figure 6.
  • Figure 8 is schematic, cross-sectional diagram of a portion of the cooling system of Figure 2.
  • Figure 9 is a detail view of a portion of the cooling system taken at the detail line in Figure 8.
  • Figure 10 is a perspective view of a portion of the cooling system of Figure 2.
  • Figure 1 1 is a detail view of a portion of the cooling system taken at detail line 1 1 -1 1 in Figure 10.
  • Figure 12 is a cross-sectional view of an alternative embodiment of the turbine vane taken at section line 2-2 in Figure 1 .
  • Figure 13 is a detailed view of an alternative embodiment of the pressure side nearwall cooling channel.
  • Figure 14 is a detailed view of an alternative embodiment of the trailing edge cooling channel of the cooling system shown in Figure 2.
  • Figure 15 is a detailed view of an alternative embodiment of the midchord cooling channel of the cooling system shown in Figure 2.
  • Figure 16 is a perspective view of a turbine blade with aspects of this invention.
  • an airfoil 10 for a gas turbine engine in which the airfoil 10 includes an internal cooling system 14 formed from one or more midchord cooling channels 16 with a corrugated insert 18 positioned therein and creating nearwall leading edge, pressure side and suction side nearwall cooling channels 20, 22, 24 is disclosed.
  • the corrugated insert 18 may be formed from a wall 26 that oscillates in a repeating pattern between peaks 28 and valleys 30, such that the peaks 28 are closer to an inner surface 50 of an outer wall 32 forming a generally elongated hollow airfoil 34 of the airfoil 10.
  • the corrugated insert 18 may work in concert with the rows 36 of partition walls 38 to create periodic impingement on the inner surface 50 of the outer wall 32, specifically the surfaces 40 leading from the valleys 30 to the peaks 28 moving toward a trailing edge 42 of the generally elongated hollow airfoil 34 direct the fluids toward the outer wall 32.
  • Such cooling system 14 may provide adequate cooling for use in environments in which few, if any, cooling holes are desired, such as in crude oil engine applications.
  • the turbine airfoil 10 may be formed from a generally elongated hollow airfoil 34 formed from an outer wall 32, and having a leading edge 44, a trailing edge 42, a pressure side 46, a suction side 48, and the internal cooling system 14 positioned within interior aspects of the generally elongated hollow airfoil 34.
  • the turbine airfoil 10 may be a stationary vane, as shown in Figure 1 , such as a turbine vane, with inner and outer endwalls.
  • the turbine airfoil 10 may be a rotary blade, as shown in Figure 16, such as a turbine blade.
  • the turbine airfoil 10 may include one or more midchord cooling channels 16 with one or more corrugated inserts 18 positioned therein and creating a leading edge nearwall cooling channel 20 between the leading edge 44 and the corrugated insert 18, a suction side nearwall cooling channel 24 between the suction side 48 and the corrugated insert 18 and a pressure side nearwall cooling channel 22 between the pressure side 46 and the corrugated insert 18.
  • the corrugated insert 18 may be formed from a wall 26 that oscillates in a repeating pattern between peaks 28 and valleys 30, such that the peaks 28 are closer to an inner surface 50 of the outer wall 32 forming the generally elongated hollow airfoil 34.
  • the peaks 28 and valleys 30 may be generally curved. Portions of the corrugated insert 18 extending between the peaks 28 and valleys 30 may be generally linear.
  • the turbine airfoil 10 may also include a plurality of rows 36 of partition walls 38 extending from the inner surface 50 forming the outer wall 32 and into the midchord cooling channel 16. On or more partition walls 38 may include gaps 52 therein. The gaps 52 maybe shorter in length than the individual sections of the partition walls 38.
  • the rows 36 of partition walls 38 may be generally aligned with a direction 54 of cooling fluid flow from the leading edge 44 chordwise toward the trailing edge 42.
  • the partition walls 38 may be
  • the partition walls 38 may be orthogonal with the leading edge 44 or the trailing edge 42, or both. At least a portion of the peaks 28 in the corrugated insert 18 extending radially outward may be aligned with the gaps 52 within at least a portion of at least one partition wall 38. Such alignment of the gaps 52 with the peaks 28 creates periodic impingement upon the impingement surfaces 40 and partial lateral, spanwise movement and mixing of the cooling fluids.
  • the corrugated insert 18 functions in concert with the rows 36 of partition walls 38 to create periodic impingement on the corrugated insert 18, specifically on those surfaces 40 leading from the valleys 30 to the peaks 28 moving toward the trailing edge 42.
  • the plurality of rows 36 of partition walls 38 may extend from the inner surface 50 into the pressure side nearwall cooling channel 22. Similarly, the plurality of rows 36 of partition walls 38 may extend from the inner surface 50 into the suction side nearwall cooling channel 24. The rows of partition walls 38 may extend in a direction from the leading edge 44 chordwise toward the trailing edge 42. As shown in Figures 2 and 12, the leading edge nearwall cooling channel 20 may not include partition walls 38 or other components. In another embodiment, the leading edge nearwall cooling channel 20 may include partition walls 38 or other impingement components.
  • One or more inlets 56 may be positioned in the corrugated insert 18 at the leading edge nearwall cooling channel 20 to pass cooling fluid from interior aspects of the corrugated insert 18 to the leading edge nearwall cooling channel 20.
  • the inlets 56 may function as an impingement holes as the compressed air flows from inner aspects of the corrugated insert 18 through the inlets 56 and impinges on a backside, inner surface of the outer wall 32 forming the leading edge 44.
  • the partition wall gaps 52 may extend for a partial length or an entire length of the partition wall 38.
  • One or more of the partition walls 38 may be linear.
  • each of the partition walls 38 may be linear and each of the partition walls 38 may be parallel with each other.
  • the partition walls 38 may have a cross-sectional area with one or more of the following shapes: square, rectangular and trapezoidal , or other appropriate shape.
  • one or more partition walls 38 may be formed from first and second subpartition walls 58, 60.
  • the second subpartition wall 60 may be offset laterally from first subpartition wall 58, and the first and second subpartition walls 58, 60 may be staggered in an alternating manner moving in a direction from the leading edge 44 toward the trailing edge 42.
  • each partition wall 38 may be formed from the first and second subpartition walls 58, 60 with the same configuration set forth immediately above. Adjacent partition walls 38 may be spaced from each other equidistant, randomly, in a pattern or in another manner.
  • a second subpartition wall 60 may be positioned along an axis 62 that is equidistant from the first subpartition wall 58 within a same partition wall 64 as the second subpartition wall 60 and a first subpartition wall 58 within an adjacent partition wall 66.
  • the cooling system 14 may also include one or more miniribs 68, as shown in Figures 3, 5, 9-1 1 , 13 and 15, extending from the inner surface 50 forming the outer wall 32 and extending between adjacent partition walls 38.
  • the miniribs 68 may be nonparallel with the partition walls 38 and may be shorter in height extending from the inner surface 50 than adjacent partition ribs 38.
  • the miniribs 68 may have a cross-sectional area with one or more of the following shapes: square, rectangular, and trapezoidal, or other appropriate shape.
  • a cross-sectional area of a minirib 68 may be less than a cross-sectional area of the partition wall 38.
  • the cross-sectional area of a minirib 68 may be less than one quarter of the cross-sectional area of the partition wall 38. In yet another embodiment, the cross-sectional area of a minirib 68 may be less than one eight of the cross-sectional area of the partition wall 38.
  • the plurality of miniribs 68 may be aligned with each other and may be orthogonal to the adjacent partition walls 38. The miniribs 68 may extend from a partition wall 38 to an adjacent partition wall 38.
  • neither end of the miniribs 68 contacts the partition walls 38, as shown in Figures 5 and 15.
  • a plurality of miniribs 68 may extend between each adjacent partition walls 38 between gaps 52.
  • at least four miniribs 68 extend between adjacent partition walls 38 between gaps 52.
  • the cooling system 14 may also include one or more trailing edge cooling channels 70, as shown in Figures 4 and 14, positioned between the midchord cooling channel 16 and the trailing edge 42 of the generally elongated hollow airfoil 34.
  • the trailing edge cooling channel 70 may include a plurality of pin fins 72 extending from the outer wall 32 forming the pressure side 46 to the suction side 48 and forming zigzag cooling flow channels 74 within the trailing edge cooling channel 70.
  • At least a portion of the pin fins 72 may have cross-sectional areas with a leading edge 76 and a trailing edge 78 that is positioned on a downstream corner 80 and separated from the leading edge 76 by a concave side surface 82 and a convex side surface 84 on an opposite side of the pin fin 72 from the concave side 82.
  • One or more of the zigzag cooling flow channels 74 may be formed from a plurality of pin fins 72 aligned such that the concave and convex side surfaces 82, 84 alternate moving towards the trailing edge 42.
  • a first upstream pin fin 86 may have a cross-sectional area formed from an upstream section 88 and a downstream section 90.
  • the upstream section 88 may be generally linear with a constant width, and the downstream section 90 may be generally linear with a tapered width that reduces in width moving towards the trailing edge 42.
  • the upstream and downstream sections 88, 90 may be nonlinear and nonorthogonal with each other.
  • cooling fluids may be supplied from a compressor or other such source to the inner aspect of the corrugated insert 18 of the internal cooling system 14. Cooling fluids are passed through the inlets 56 into the leading edge nearwall cooling channel 20 where the fluids separate with a portion flowing into the pressure side nearwall cooling channel 22 and a portion flowing into the suction side nearwall cooling channel 24.
  • the peaks 28 and valleys 30 of the corrugated insert 18 create periodic impingement on the inner surface 50 of the outer wall 32.
  • the partition walls 38 also direct the cooling fluids towards the trailing edge 42, and in at least some embodiments, create a nonlinear flow path toward the trailing edge to increase convection of heat from the outer wall to the cooling fluid.
  • the cooling fluid may be exhausted through film cooling holes 92 at downstream ends of the pressure and suction side nearwall cooling channels 22, 24.
  • Cooling fluids may also be passed to the trailing edge cooling channel 70 via one or more trailing edge supply channels 94.
  • the cooling fluids may be metered through metering holes 96 into the trailing edge cooling channel 70.
  • the cooling fluids may flow into the zigzag cooling flow channels 74 where the pin fins 72 direct the cooling fluids in a nonlinear, back and forth motion to the trailing edge 42, where the cooling fluids are exhausted.
  • the cooling fluids in the trailing edge cooling channel 70 receive heat from the outer wall 32 forming the pressure and suction sides 46, 48 and from the pin fins 72.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un profil aérodynamique (10) pour une turbine à gaz, le profil aérodynamique comprenant un système de refroidissement interne (14) formé à partir d'un ou de plusieurs canaux de refroidissement à mi-corde (16) ayant un insert ondulé (18) positionné dans celui-ci et créant un bord d'attaque côté paroi proche, des canaux de refroidissement côté paroi proche côté pression et côté aspiration (20, 22, 24). L'insert ondulé peut être formé à partir d'une paroi (26) qui oscille en une configuration répétée entre des crêtes (28) et des vallées (30), de sorte que les crêtes soient plus proches de la surface intérieure (50) de la paroi extérieure (32) formant le profil aérodynamique creux généralement allongé. L'insert ondulé peut fonctionner de manière coordonnée avec des rangées (36) de parois de séparation (38) pour créer un contact périodique sur la surface intérieure de la paroi extérieure. Un tel système de refroidissement procure un refroidissement adéquat à des fins d'utilisation dans des environnements dans lesquels seuls quelques trous de refroidissement sont souhaités, voire aucun, comme dans des applications de moteur à pétrole brut.
PCT/US2015/011496 2014-01-15 2015-01-15 Système de refroidissement interne ayant un insert ondulé formant des canaux de refroidissement côté paroi proche pour profil aérodynamique de turbine à gaz WO2015109040A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US14/155,422 US20150198050A1 (en) 2014-01-15 2014-01-15 Internal cooling system with corrugated insert forming nearwall cooling channels for airfoil usable in a gas turbine engine
US14/155,422 2014-01-15

Publications (1)

Publication Number Publication Date
WO2015109040A1 true WO2015109040A1 (fr) 2015-07-23

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US (1) US20150198050A1 (fr)
WO (1) WO2015109040A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3181823A1 (fr) * 2015-12-18 2017-06-21 United Technologies Corporation Procédé et appareil de refroidissement d'un composant de moteur de turbine à gaz
EP3354852A3 (fr) * 2017-01-26 2018-09-19 United Technologies Corporation Composants de moteurs à refroidissement interne

Families Citing this family (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6245740B2 (ja) * 2013-11-20 2017-12-13 三菱日立パワーシステムズ株式会社 ガスタービン翼
US9840930B2 (en) * 2014-09-04 2017-12-12 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil
JP6407414B2 (ja) * 2014-09-04 2018-10-17 シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ ガスタービン翼の後方冷却キャビティ内に壁近傍冷却通路を形成する挿入体を有する内部冷却システム
CA2935398A1 (fr) * 2015-07-31 2017-01-31 Rolls-Royce Corporation Profils aerodynamiques de turbine dotes de fonctionnalites de micro refroidissement
JP2017089601A (ja) * 2015-11-17 2017-05-25 株式会社東芝 冷却構造及びガスタービン
WO2017105379A1 (fr) * 2015-12-14 2017-06-22 Siemens Aktiengesellschaft Profil aérodynamique de turbine ayant un élément de blocage d'écoulement profilé permettant un meilleur refroidissement de paroi proche
US10408073B2 (en) * 2016-01-20 2019-09-10 General Electric Company Cooled CMC wall contouring
US9759073B1 (en) * 2016-02-26 2017-09-12 Siemens Energy, Inc. Turbine airfoil having near-wall cooling insert
US10436037B2 (en) * 2016-07-22 2019-10-08 General Electric Company Blade with parallel corrugated surfaces on inner and outer surfaces
US10465520B2 (en) 2016-07-22 2019-11-05 General Electric Company Blade with corrugated outer surface(s)
US10443399B2 (en) 2016-07-22 2019-10-15 General Electric Company Turbine vane with coupon having corrugated surface(s)
US10450868B2 (en) 2016-07-22 2019-10-22 General Electric Company Turbine rotor blade with coupon having corrugated surface(s)
US10465525B2 (en) 2016-07-22 2019-11-05 General Electric Company Blade with internal rib having corrugated surface(s)
US10830060B2 (en) * 2016-12-02 2020-11-10 General Electric Company Engine component with flow enhancer
US10830056B2 (en) * 2017-02-03 2020-11-10 General Electric Company Fluid cooling systems for a gas turbine engine
US10494948B2 (en) * 2017-05-09 2019-12-03 General Electric Company Impingement insert
KR102048863B1 (ko) * 2018-04-17 2019-11-26 두산중공업 주식회사 인서트 지지부를 구비한 터빈 베인
US10767492B2 (en) 2018-12-18 2020-09-08 General Electric Company Turbine engine airfoil
US11371360B2 (en) * 2019-06-05 2022-06-28 Raytheon Technologies Corporation Components for gas turbine engines
US11988110B2 (en) * 2019-06-28 2024-05-21 Siemens Energy Global GmbH & Co. KG Turbine airfoil incorporating modal frequency response tuning
US11268392B2 (en) * 2019-10-28 2022-03-08 Rolls-Royce Plc Turbine vane assembly incorporating ceramic matrix composite materials and cooling
US11725526B1 (en) 2022-03-08 2023-08-15 General Electric Company Turbofan engine having nacelle with non-annular inlet

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
US3846041A (en) * 1972-10-31 1974-11-05 Avco Corp Impingement cooled turbine blades and method of making same
JPS6380004A (ja) * 1986-09-22 1988-04-11 Hitachi Ltd ガスタ−ビン静翼
US5246341A (en) * 1992-07-06 1993-09-21 United Technologies Corporation Turbine blade trailing edge cooling construction
EP0698725A2 (fr) * 1994-08-26 1996-02-28 ABB Management AG Refroidissement d'une partie de paroi par jets d'air
EP1221538A2 (fr) * 2001-01-05 2002-07-10 General Electric Company Aube de guidage refroidie
US20060210399A1 (en) * 2003-11-21 2006-09-21 Tsuyoshi Kitamura Turbine cooling vane of gas turbine engine
EP2233693A1 (fr) * 2008-01-08 2010-09-29 IHI Corporation Structure de refroidissement d'aube de turbine
EP2233694A1 (fr) * 2009-03-26 2010-09-29 United Technologies Corporation Mesure d'éléments d'espacement pour déflecteur de surface portante
US20120070302A1 (en) * 2010-09-20 2012-03-22 Ching-Pang Lee Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
CA2859154A1 (fr) * 2011-12-15 2013-06-20 Ihi Corporation Mecanisme de refroidissement par contact, aube de turbine et chambre de combustion

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2590457A (en) * 1946-04-20 1952-03-25 Sigma Jet propulsion device for airscrews or rotary wings
US2994124A (en) * 1955-10-03 1961-08-01 Gen Electric Clad cermet body
US3446481A (en) * 1967-03-24 1969-05-27 Gen Electric Liquid cooled turbine rotor
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure
DE3003347A1 (de) * 1979-12-20 1981-06-25 BBC AG Brown, Boveri & Cie., Baden, Aargau Gekuehlte wand
DE3327659A1 (de) * 1983-07-30 1985-02-14 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Verfahren zur herstellung eines verbundkoerpers aus keramik oder faserverstaerkter keramik sowie ein nach diesem verfahren gefertigtes sandwich-gebilde
US5704763A (en) * 1990-08-01 1998-01-06 General Electric Company Shear jet cooling passages for internally cooled machine elements
US6018950A (en) * 1997-06-13 2000-02-01 Siemens Westinghouse Power Corporation Combustion turbine modular cooling panel
US6142734A (en) * 1999-04-06 2000-11-07 General Electric Company Internally grooved turbine wall
US20110110772A1 (en) * 2009-11-11 2011-05-12 Arrell Douglas J Turbine Engine Components with Near Surface Cooling Channels and Methods of Making the Same

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
US3846041A (en) * 1972-10-31 1974-11-05 Avco Corp Impingement cooled turbine blades and method of making same
JPS6380004A (ja) * 1986-09-22 1988-04-11 Hitachi Ltd ガスタ−ビン静翼
US5246341A (en) * 1992-07-06 1993-09-21 United Technologies Corporation Turbine blade trailing edge cooling construction
EP0698725A2 (fr) * 1994-08-26 1996-02-28 ABB Management AG Refroidissement d'une partie de paroi par jets d'air
EP1221538A2 (fr) * 2001-01-05 2002-07-10 General Electric Company Aube de guidage refroidie
US20060210399A1 (en) * 2003-11-21 2006-09-21 Tsuyoshi Kitamura Turbine cooling vane of gas turbine engine
EP2233693A1 (fr) * 2008-01-08 2010-09-29 IHI Corporation Structure de refroidissement d'aube de turbine
EP2233694A1 (fr) * 2009-03-26 2010-09-29 United Technologies Corporation Mesure d'éléments d'espacement pour déflecteur de surface portante
US20120070302A1 (en) * 2010-09-20 2012-03-22 Ching-Pang Lee Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
CA2859154A1 (fr) * 2011-12-15 2013-06-20 Ihi Corporation Mecanisme de refroidissement par contact, aube de turbine et chambre de combustion

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3181823A1 (fr) * 2015-12-18 2017-06-21 United Technologies Corporation Procédé et appareil de refroidissement d'un composant de moteur de turbine à gaz
US10156147B2 (en) 2015-12-18 2018-12-18 United Technologies Corporation Method and apparatus for cooling gas turbine engine component
EP3354852A3 (fr) * 2017-01-26 2018-09-19 United Technologies Corporation Composants de moteurs à refroidissement interne
US10344607B2 (en) 2017-01-26 2019-07-09 United Technologies Corporation Internally cooled engine components

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