EP3875735A1 - Surface portante pour turbine à gaz - Google Patents

Surface portante pour turbine à gaz Download PDF

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Publication number
EP3875735A1
EP3875735A1 EP20161261.1A EP20161261A EP3875735A1 EP 3875735 A1 EP3875735 A1 EP 3875735A1 EP 20161261 A EP20161261 A EP 20161261A EP 3875735 A1 EP3875735 A1 EP 3875735A1
Authority
EP
European Patent Office
Prior art keywords
side wall
aerofoil
rib
pressure side
trailing edge
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.)
Withdrawn
Application number
EP20161261.1A
Other languages
German (de)
English (en)
Inventor
David Carter
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
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 AG filed Critical Siemens AG
Priority to EP20161261.1A priority Critical patent/EP3875735A1/fr
Publication of EP3875735A1 publication Critical patent/EP3875735A1/fr
Withdrawn legal-status Critical Current

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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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F05D2230/31—Layer deposition
    • 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
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing 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
    • F05D2240/304—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 trailing 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer

Definitions

  • the present invention relates to an internally cooled aerofoil for a gas turbine engine.
  • Aerofoils It is necessary to cool aerofoils of a gas turbine engine to prevent overheating and prolong service life. Aerofoils have internal cooling passages through which coolant is channelled during engine operation. In highly optimised aerofoils it is necessary to maintain a constant flow of coolant and have a constant external wall thickness to ensure a constant heat transfer coefficient and hence a constant cooling effect.
  • an aerofoil for a gas turbine engine.
  • the aerofoil comprising a suction side wall, a pressure side wall, a leading edge and a trailing edge, a first radial end and a second radial end.
  • the suction side wall and the pressure side wall extend from the first radial end to the second radial end and meet to define the leading edge and the trailing edge.
  • the suction side wall and the pressure side form a cavity therebetween for the flow of a coolant therethrough and a converge towards the trailing edge to form a converging portion.
  • Within the converging portion is a rib that extends radially between first radial end and the second radial end and defines a first passageway between the rib and the pressure side wall and a second passageway between the rib and the suction side wall.
  • the suction side wall and the pressure side wall may have a thickness that tapers towards the trailing edge.
  • the sum of the widths or sectional areas of the passageway between the rib and the suction side wall and the passageway between the rib and the pressure side wall may be approximately equal to the width or sectional area from the suction side wall to pressure side wall immediately on the trailing edge side of the rib.
  • the rib may have a tapering portion which tapers towards the trailing edge.
  • the tapering portion may taper at approximately the same rate as the converging portion converges.
  • the aerofoil may comprise at least one chordal rib that spans the cavity from the suction side wall to the pressure side wall and is generally perpendicular to the suction side wall and the pressure side wall.
  • the chordal rib may extend in the chordal direction and may be at least partly within the converging portion.
  • the rib may have a leading portion, a main body potion and a trailing portion.
  • the leading portion may extend towards the leading edge from and beyond the chordal rib.
  • a mean camber line is defined as the locus of points mid-way between the suction side wall and pressure side wall.
  • the rib may have a centre line.
  • the rib may be curved along the mean camber such that its centre line may be along the camber line.
  • a coolant flows through the converging portion in a generally chordal direction towards the trailing edge.
  • the trailing edge may comprise at least one aperture for the release of coolant from the cavity.
  • the aerofoil may be monolithically formed in a casting step or in an additive manufacturing process.
  • the aerofoil may be part of a vane.
  • the vane may comprise radially inner and radially outer platforms, the aerofoil spanning between the platforms.
  • the aerofoil may be part of a blade.
  • the blade may comprise a platform and a tip, the aerofoil spanning between the platform and the tip.
  • the tip comprises any one of the group a shroud, a squealer and a winglet.
  • FIG. 1 shows an example of a gas turbine engine 10 in a sectional view.
  • the gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20.
  • the gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
  • air 24 which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28.
  • the combustion chambers 28 and the burners 30 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22.
  • two discs 36 each carry an annular array of turbine blades 38.
  • the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.
  • the combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48.
  • the guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
  • the casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
  • the present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine.
  • the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated.
  • forward and rearward refer to the general flow of gas through the engine.
  • axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.
  • aerofoil is intended to apply to a stator vane or a rotor blade.
  • FIG. 2 shows part of the turbine section 18 of the gas turbine engine 10.
  • the working gas flow 34 through the turbine section 18 comprises an annular array of inlet guide vanes 44 and then an annular array of blades 38.
  • the working gas flow 34 is channeled between a radially inner annulus 60 and a radially outer annulus 62 which is referred to as the 'gas path' and denoted 61.
  • the inlet guide vanes 44 comprise at least one aerofoil 64 having a suction side wall 68 and a pressure side wall 66, a leading edge 70 and a trailing edge 72, a first radial end 74 and a second radial end 76.
  • the suction side wall 68 and the pressure side wall 66 extend from the first radial end 74 to the second radial end 76 and meet to define the leading edge 70 and the trailing edge 72.
  • the first radial end 74 is attached to a radially inner platform 80 and the second radial end 76 is attached to a radially outer platform 82.
  • the radially inner platform 80 and the radially outer platform 82 form part of the gas path 61.
  • the annular array of inlet guide vanes 44 comprises single vanes where there is a single aerofoil 64 mounted between the two platforms 80, 82.
  • the annular array of inlet guide vanes 44 may comprise double, triple or more vane segments where there are two, three of more aerofoils 64 mounted between common platforms 80, 82.
  • the blades 38 comprise one aerofoil 84 having a pressure side wall 86 and a suction side wall 88, a leading edge 90 and a trailing edge 92, a first radial end 94 and a second radial end 96.
  • the suction side wall 88 and the pressure side wall 86 extend from the first radial end 94 to the second radial end 96 and meet to define the leading edge 90 and the trailing edge 92.
  • the first radial end 94 is attached to a radially inner platform 100 and the second radial end 76 is a blade tip 96 and is free.
  • the blade tip 96 forms a tip gap 102 with a casing 104.
  • the radially inner platform 100 and the casing 102 form part of the gas path 61.
  • the tip 96 is a squealer configuration as is known in the art, but in other examples the blade tip 96 may comprise a shroud or a winglet as is known.
  • the blade 38 comprises a root portion 106 which comprises a firtree configuration that engages a complimentary slot formed in the disc 36 for retaining the blades 38 in known manner.
  • the root portion may be a dovetail configuration that engages a complimentary slot formed in the disc 36 for retaining the blades 38 in known manner.
  • the present invention is applicable to the aerofoils 64, 84 of the vane 44 and blade 38.
  • the suction side wall 68, 88 and the pressure side wall 66, 86 form a cavity 108 (see Fig. 6 ) therebetween.
  • the cavity 108 is formed for the flow of a coolant therethrough.
  • the cavity 108 may have many different configurations for optimal cooling depending on each application.
  • FIG.3 to FIG.5 Each of these figures show a portion of a conventional aerofoil 200 in cross-section and only the downstream portion including a trailing edge 202 formed by a suction side wall 204 and a pressure side wall 206 in conventional manner.
  • a cavity 212 is formed between the suction side wall 204 and the pressure side wall 206.
  • a coolant flow 208 flows through the cavity 212 and in the chordal direction towards the trailing edge 202 and out through an aperture or apertures 210.
  • the cavity 212, the suction side wall 204 and the pressure side wall 206 all taper towards the trailing edge 202.
  • the width W of the cavity and the thicknesses Ts and Tp of the suction side wall 204 and the pressure side wall 206 all reduce the nearer to the trailing edge.
  • this tapering causes the coolant flow 208 to accelerate towards the trailing edge 202 and therefore provide lessening cooling of the suction side wall 204 and the pressure side wall 206.
  • the aerofoil 200 comprises the same configuration as described in FIG. 3 and comprises a chordal rib 214 running axially and at a prescribed pitch from hub to tip.
  • the chordal rib 214 spans the cavity 212 from the suction side wall 204 to the pressure side wall 206 and is generally perpendicular to the suction side wall 204 and the pressure side wall 206.
  • the chordal rib 214 extends in the chordal direction, that being generally in the plane of the figure and generally flowing the camber line of the aerofoil 200.
  • the chordal rib 214 has a leading edge 216 and a trailing edge 218. In this example, there is a smooth transition of the wall thicknesses Ts and Tp.
  • the cavity 212 has a width W2 at the leading edge 216 of the chordal rib 214 and a width W1 at the trailing edge 218.
  • W2 is greater than W1 meaning the coolant flow must accelerate from the leading edge 216 of the chordal rib to its trailing edge 218.
  • the coolant flow accelerates it picks up less heat from the walls and leads to reduced cooling of the walls 204, 206 and the chordal rib 214 in the direction towards the trailing edge 202.
  • the aerofoil 200 comprises the same configuration as described in FIG. 4 except that the suction side wall 204 and the pressure side wall 206 have thickened portions 220, 222, respectively, in the region of the chordal rib 214.
  • the thickening portions 220, 222 create a constant width portion 224 of the cavity 212.
  • Coolant flow 208 passing through the constant width portion 224 has a constant velocity and therefore cools the walls 204, 206 and the chordal rib 214 more evenly and has a reduced temperature gradient in the walls 204, 206 in the vicinity of and in the chordal rib 214.
  • the relatively abrupt increase in thickness of the thickened portions 220, 222, particularly in zones 226, causes relatively poor solidification during the casting process. This can lead to a relatively high scrap rate, shortened service life and even failure.
  • FIG. 6 depicts one exemplary embodiment of the present invention.
  • the aerofoil 84 comprises the suction side wall 88, the pressure side wall 86, the leading edge 90 and the trailing edge 92 and configured as previously described.
  • the suction side wall 88 and the pressure side wall 86 form the cavity 108.
  • FIG. 6 shows a trailing edge portion 92R which extends from the trailing edge 92 towards the leading edge 90 a distance 10-30%, preferably 15-20%, of the length of a mean camber line 99 in the case of a blade and a distance 25-50%, preferably 35-40%, of the length of a mean camber line 99 in the case of a vane.
  • the mean camber line 99 is defined as the locus of points halfway between the external surfaces of the suction side wall 88 and the pressure side wall 86.
  • the suction side wall 88 and the pressure side wall 86 converge towards one another in a direction towards the trailing edge to form a converging portion 112.
  • the cavity 108 decreases in width W in the direction towards the trailing edge 92 in the converging portion 112.
  • the cavity 108 decreases in sectional area i.e. the area of the cavity 108 through which the coolant flow passes through or the area of the plane perpendicular to the coolant flow direction and between the suction side wall 88, the pressure side wall 86 and radially inner and out boundaries of the cavity.
  • the converging portion 112 extends over a part of the trailing edge portion 92R but may extend over all the trailing edge portion 92R.
  • a rib 114 that extends radially between the first radial end 94 and the second radial end 96 of the aerofoil and preferably extends radially from the first radial end 94 to the second radial end 96.
  • the rib 114 is not directly in contact with the suction side wall 88 or the pressure side wall 86.
  • a first passageway 116, having a width W3, is defined between the rib 114 and the pressure side wall 86 and a second passageway 118, having a width W4, is defined between the rib 114 and the suction side wall 88.
  • the pressure side wall 86 and the suction side wall 88 have respective thicknesses Tp and Ts that preferably taper towards the trailing edge 92.
  • coolant 110 flows through the cavity 108 in the direction of the arrow 110 and portions of the coolant 110 pass through the first passageway 116 and second passageway 118.
  • the width W3 of the first passageway 116 is equal to the width W4 of the second passageway 118 in this exemplary embodiment.
  • the rib 114 may be offset from the mean camber line 99 towards one or the other of the pressure side wall 86 or the suction side wall 88 and therefore W3 may be greater than W4 and vice versa. Nonetheless, the sum of the widths W3 and W4 of the first and second passageways is approximately equal to the width W5 from the suction side wall 88 to pressure side wall 86 immediately on the trailing edge side 92 of the rib 114 or immediately downstream of the rib 114.
  • the sectional areas of the passageways 116, 118 and the cavity 108 for the coolant to flow through, in a direction from the leading edge to the tailing edge of the aerofoil is proportional to the widths such that the widths W3, W4 and W5 are also indicative of the sectional areas at the indicated positions.
  • the radial height of the cavity 108 is constant; however, in other embodiments the radial height of the cavity may vary and therefore the sectional areas indicated at the widths W3, W4 and W5 are relevant and here W3, W4 and W5 are indicative of the sectional areas rather than merely the widths.
  • the rib 114 has a leading edge portion 120 and a tapering portion 122 which tapers towards the trailing edge 92.
  • the tapering portion 122 tapers at approximately the same rate as the converging portion 112 converges such that the width W3 of the first passageway 116 and the width W4 of the second passageway 118 remain approximately constant over the length of the tapering portion 122.
  • the aerofoil 84 comprises a chordal rib 124 that spans the cavity 108 from the suction side wall 88 to the pressure side wall 86 and is generally perpendicular to the suction side wall 86 and the pressure side wall 88.
  • the chordal rib 124 extends in the chordal direction, i.e. the direction from the leading edge 90 towards the trailing edge 92 of the aerofoil 86.
  • the chordal rib 124 is relatively thin, i.e. its thickness is relatively small into the figure or radially, and is approximately the same thickness as the pressure or suction side walls but may be less.
  • the chordal rib 124 is at least partly located within the converging portion 112.
  • the leading edge portion 120 of the rib 114 extends and tapers towards the leading edge 90 of the aerofoil 84 from the leading edge of the chordal rib 114. In other applications the leading edge portion 120 may be aligned with or start from the leading edge of the chordal rib 114.
  • the tapering portion 122 extends approximately 50% of the axial length of the chordal rib 114. Alternatively, the tapering portion 122 may extend anywhere up to the trailing edge of the chordal rib 114.
  • the design of the rib 114 is based upon maintaining a desired velocity of the coolant passing through the cavity 108 thereby maximising cooling and minimising any temperature gradients in the aerofoil.
  • chordal rib 124 are integral with one another such that the rib 114 is supported to prevent bending, vibrations and lateral movements towards either the suction or pressure side walls. Although only one chordal rib 124 is shown, there may be more than one chordal rib 124. The chordal ribs 124 may be equally spaced over the radial extent of the aerofoil or otherwise positioned to optimise stability of the rib 114.
  • the mean camber line 99 is defined as the locus of points mid-way between the suction side wall 88 and pressure side wall 86 and in particular their external surfaces.
  • the rib 114 has a centre line 126 and as seen the rib 114 is curved along the mean camber 99 such that its centre line 126 is along or coincident with the mean camber line 99.
  • some aerofoils 84 may have a constant overall thickness in any radial line or the overall thickness may taper in the radial direction. Where the overall thickness tapers in the radial direction, the width W5 may also taper radially and therefore either one or both the widths W3 and W4 are intended to taper and by virtue of the sectional thickness 118 of the rib 114 also tapering in the radial direction.
  • the aerofoil 84 and host vane or blade is monolithically formed in a casting step during manufacturing.
  • the aerofoil is monolithically formed in an additive manufacturing process.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP20161261.1A 2020-03-05 2020-03-05 Surface portante pour turbine à gaz Withdrawn EP3875735A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP20161261.1A EP3875735A1 (fr) 2020-03-05 2020-03-05 Surface portante pour turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP20161261.1A EP3875735A1 (fr) 2020-03-05 2020-03-05 Surface portante pour turbine à gaz

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EP3875735A1 true EP3875735A1 (fr) 2021-09-08

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EP20161261.1A Withdrawn EP3875735A1 (fr) 2020-03-05 2020-03-05 Surface portante pour turbine à gaz

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20220220857A1 (en) * 2021-01-11 2022-07-14 Doosan Heavy Industries & Construction Co., Ltd. Turbine airfoil and turbine including same

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4526512A (en) * 1983-03-28 1985-07-02 General Electric Co. Cooling flow control device for turbine blades
EP1561901A2 (fr) * 2004-02-04 2005-08-10 United Technologies Corporation Amortisseur des vibrations pour aubes rotoriques refroidies de turbines
US7527474B1 (en) * 2006-08-11 2009-05-05 Florida Turbine Technologies, Inc. Turbine airfoil with mini-serpentine cooling passages
WO2011050025A2 (fr) * 2009-10-20 2011-04-28 Siemens Energy, Inc. Plan de sustentation incorporant des structures de refroidissement effilées définissant des passages de refroidissement
JP2011111946A (ja) * 2009-11-25 2011-06-09 Mitsubishi Heavy Ind Ltd 翼体及びこの翼体を備えたガスタービン

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4526512A (en) * 1983-03-28 1985-07-02 General Electric Co. Cooling flow control device for turbine blades
EP1561901A2 (fr) * 2004-02-04 2005-08-10 United Technologies Corporation Amortisseur des vibrations pour aubes rotoriques refroidies de turbines
US7527474B1 (en) * 2006-08-11 2009-05-05 Florida Turbine Technologies, Inc. Turbine airfoil with mini-serpentine cooling passages
WO2011050025A2 (fr) * 2009-10-20 2011-04-28 Siemens Energy, Inc. Plan de sustentation incorporant des structures de refroidissement effilées définissant des passages de refroidissement
JP2011111946A (ja) * 2009-11-25 2011-06-09 Mitsubishi Heavy Ind Ltd 翼体及びこの翼体を備えたガスタービン

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20220220857A1 (en) * 2021-01-11 2022-07-14 Doosan Heavy Industries & Construction Co., Ltd. Turbine airfoil and turbine including same
US11448074B2 (en) * 2021-01-11 2022-09-20 Doosan Enerbility Co., Ltd. Turbine airfoil and turbine including same

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