WO2017121689A1 - Profil aérodynamique de turbine à gaz - Google Patents

Profil aérodynamique de turbine à gaz Download PDF

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Publication number
WO2017121689A1
WO2017121689A1 PCT/EP2017/050276 EP2017050276W WO2017121689A1 WO 2017121689 A1 WO2017121689 A1 WO 2017121689A1 EP 2017050276 W EP2017050276 W EP 2017050276W WO 2017121689 A1 WO2017121689 A1 WO 2017121689A1
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WO
WIPO (PCT)
Prior art keywords
section
aerofoil
cross
cavity
sectional area
Prior art date
Application number
PCT/EP2017/050276
Other languages
English (en)
Inventor
Anthony Davis
Original Assignee
Siemens Aktiengesellschaft
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 filed Critical Siemens Aktiengesellschaft
Publication of WO2017121689A1 publication Critical patent/WO2017121689A1/fr

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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
    • 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
    • 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
    • F05D2240/301—Cross-sectional characteristics
    • 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
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to gas turbine aerofoils and, more particularly, to gas turbine aerofoil provided with internal cooling passages.
  • air is pressurized in a compressor and mixed with fuel in a combustor for generating hot
  • combustion gases are then channelled towards a turbine which transforms the energy from the hot gases into work for powering the compressor and other devices which converts power, for example an upstream fan in a typical aircraft turbofan engine application, or a generator in power generation application.
  • the turbine stages include stationary turbine nozzles having a row of vanes which channel the combustion gases into a corresponding row of rotor blades extending radially
  • the vanes and blades may have corresponding hollow aerofoils. Aerofoils may be designed and manufactured hollow in order to save weight, to change its eigenfrequency or to include a cooling circuit therein. In the latter case, the cooling gas which circulates inside the cooling circuit or circuits is typically bleed air from the compressor discharge.
  • the present invention relates to a turbine hollow aerofoil which can be used in a gas turbine vane or blade and which include an internal passage for cooling the aerofoil.
  • the external surface of a turbine hollow aerofoil is also typically provided with a plurality of outlet cooling holes.
  • the cooling gas from the inside of the hollow aerofoil flows through the outlet cooling holes forming a cooling film on the external surface of the turbine hollow .
  • the pressure ratio between inlet and outlet of the cooling hole is the principle factor that determines the mass flow rate and hence the effectiveness of the cooling.
  • the exit pressure is the determined by
  • the inlet pressure is determined by the cooling air supply pressure to the component and by the pressure losses encountered by the flow up to the entry to the cooling hole.
  • the inlet pressure may be further
  • the size and number of the cooling holes in that region may be restricted by casting, machining, or stress limitations.
  • EP2902589A1 discloses an aerofoil for a gas turbine
  • US7955053B1 discloses a turbine blade having an aft flowing triple pass serpentine cooling circuit with all convection cooled blade.
  • the three passes or legs of the serpentine flow circuit are formed by a leading edge rib and a trailing edge rib that are both slanted in order to provide decreasing flow cross sectional areas in the three passes of legs.
  • EP0207799A2 discloses a combustion turbine rotor blade provided with an airfoil portion having a plurality of coolant holes extending radially outwardly therethrough. The coolant holes are tapered to a smaller flow cross-section in the radially outward direction to produce more uniform cooling action over the length of the holes over which tapering is provided.
  • an aerofoil for a gas turbine comprises:
  • an internal passage for channelling a cooling medium to a plurality of outlet cooling holes the internal passage comprising a cavity communicating with at least a portion of the outlet cooling holes.
  • the cavity has a variable cross- sectional area.
  • Variations of the cross-sectional area at a specific portion of the outlet cooling holes is used to increase or decrease the internal pressures and hence the film cooling mass flow rate in a region of the aerofoil affected by that specific portion of the outlet cooling holes. Therefore, the cooling effectiveness in that region can be controlled without changing the size and number of the outlet cooling holes.
  • the cavity extends radially between the base and the tip of the aerofoil.
  • the outlet cooling holes may be provided on the external surface of the hollow aerofoil, in particular close to the trailing edge and/or to the leading edge .
  • this permits to control the cooling film on the external side of the hollow aerofoil.
  • the internal passage comprises a plurality of cavities in series, the last cavity of the plurality of cavities communicating with at least a portion of the outlet cooling holes provided on the external surface of the hollow aerofoil .
  • the cavity may comprise at least a first section having a first cross-sectional area and a second section having a second cross-sectional area, the second cross-sectional area being greater than the first cross-sectional area.
  • the second section may be closer to the tip of the aerofoil and the first section may be closer to the base, in such a way that the increase of the cross- sectional area compensates for the pressure losses along the flow path of the cooling gas.
  • the values of the first, second and third cross-sectional areas may be defined in such a way that the flow rate through at least a portion of the outlet cooling holes is constant and equal to a predefined desired value.
  • the cavity comprises a third section having a third cross-sectional area, the second cross-sectional area being greater than the first cross-sectional area and greater than the third cross- sectional area.
  • the first section is closer to the base, the third section is closer to the tip and the second section is intermediate between the first section and the third section.
  • the outlet cooling holes are provided on an external surface of an impingement tube extending radially inside the hollow aerofoil, the cavity, which communicates with at least a portion of the outlet cooling holes,
  • An insert may be provided inside the impingement tube, in such a way that the cavity is comprised between the insert and the external surface of the impingement tube, the insert having a shape such that the cavity has a variable cross-sectional area.
  • this permits to locally increasing or
  • At least a part of the cavity may vary in cross-sectional area at a constant rate between or from the first section and the second section and/or where applicable between the second section and the third section.
  • At least a part of the cavity may vary in cross-sectional area at a non-constant rate between or from the first section and the second section and/or where applicable between the second section and the third section.
  • the non-constant rate of change of area of the cavity may comprise a sudden or step changes in the area.
  • the internal passage may be at least partly formed by the aerofoil and the cooling holes extend from the internal passage to the external surface.
  • FIG. 1 shows part of a turbine engine in a sectional view and in which the present inventive aerofoil is incorporated
  • FIG. 2 shows a longitudinal sectional view of a gas turbine aerofoil (a blade) according to the present invention
  • FIG. 3 shows a longitudinal sectional view of another
  • FIG. 4 shows a cross-sectional view of the gas turbine aerofoil of fig. 3, taken along the sectional line IV-IV in Fig. 3,
  • FIG. 5 shows an isometric partially-sectioned view of yet another embodiment of a gas turbine aerofoil according to the present invention
  • FIG. 6 shows a detailed view of the gas turbine aerofoil of fig. 5. Detailed Description
  • 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
  • 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
  • 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, 44 (a first stage of guiding vanes 44 and a second stage of 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, 60 (a first stage of turbine blades 60 and a second stage of turbine blades 38) .
  • inlet guiding vanes 40, 44 are provided and turn the flow of working gas onto the turbine blades 38, 60.
  • 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.
  • Inventive embodiments of a gas turbine aerofoil 100, 101, 102 are shown in Figs. 2 to 6.
  • the inventive turbine aerofoil 100 may be used, in general, in the turbine blades 38, 60 or in the turbine guiding vanes 40, 44.
  • a first embodiment of a gas turbine aerofoil 100 is shown, applied to a turbine blade 38.
  • the rotor blade 38 comprises:
  • a platform 124 having a lower surface 131, from which the root 122 extends and an upper surface 130 opposite to the lower surface 131,
  • the aerofoil 100 has an external surface 128.
  • the external surface 128 axially extends between a leading edge 133 and a trailing edge 135 and radially extends between a base 134 and a tip 136.
  • a first plurality of outlet cooling holes 140 are provided in a portion of the hollow aerofoil 100 including the trailing edge 135.
  • the cooling medium passes through the external surface 128 via the cooling holes to form a film of cooling medium over the external surface of the aerofoil.
  • These cooling holes 140 and the other cooling holes in the external surface of the aerofoil are known as film cooling holes or effusion cooling holes .
  • the aerofoil 100 is hollow and further comprises an internal passage 150 for channelling a cooling medium (for example compressed air from the compressor section 14) to the first plurality of outlet cooling holes 140.
  • a cooling medium for example compressed air from the compressor section 14
  • the cooling medium flows exiting the first plurality outlet cooling holes 140 is represented by a respective first plurality of arrows 145a, 145b, 145c, 145d, 145e.
  • the internal passage 150 of the aerofoil 100 is connected to a root passage 123, provided in the root 122 of the rotor blade 38.
  • the root passage 123 provides a connection between the internal passage 150 and a source (not shown) of the cooling medium.
  • the internal passage 150 comprises a plurality of cavities (three cavities 151, 152, 153) disposed in series between the root passage 123 and the first plurality of outlet cooling holes 140. Each of the cavities 151, 152, 153 extends
  • the first cavity 151 is connected to the root passage 123 and extends radially from the base 134 to the tip 136, adjacently to the leading edge 133.
  • the second cavity 152 is connected to the first cavity 151 by means of a first curve 151a of the passage 150, provided at the tip 136.
  • the second cavity 152 extends radially from the tip 136 to the base 134.
  • the third cavity 153 is connected to the second cavity 152 by means of a second curve 152a of the passage 150, provided at the base 134.
  • the third cavity 153 extends radially from the base 134 to the tip 136, adjacently to the trailing edge 135.
  • the third cavity 153 communicates with the first plurality of outlet cooling holes 140, thus permitting the cooling medium to exit the external surface 128 for cooling the aerofoil 100 at the trailing edge 135.
  • the third cavity 153 has a variable cross-sectional area.
  • the third cavity 153 comprises a first section 154 and a second section 155, both orthogonal to the radial direction.
  • the first section 154 and the second section 155 are disposed relatively to each other in such a way that the first section 154 is closer to the base 134 and the second section 155 is closer to the tip 136 of the aerofoil 100.
  • the first section 154 has a first cross-sectional area and the second section 155 having a second cross-sectional area, the second cross-sectional area being greater than the first cross-sectional area.
  • the values of the cross-sectional areas along the third cavity 153 may be preferably arranged in such a way that the flow rate of the cooling medium flows 145a, 145b, 145c, 145d, 145e through the first plurality of cooling holes 140 is constant and equals a predefine desired value.
  • the flow rate of the cooling medium flows 145a, 145b, 145c, 145d, 145e through the first plurality of cooling holes 140 may vary as required by design needs.
  • the gas turbine aerofoil 100 of Fig. 2 may be applied to a guiding vane 40.
  • a second embodiment of a gas turbine aerofoil 101 is shown, applied to a turbine blade 38.
  • a second plurality of outlet cooling holes 141 are provided in a portion of the hollow aerofoil 101 including the leading edge 133.
  • the internal passage 150 comprises a cavity 163 directly connected to the root passage 123 and extending radially between the base 134 and the tip 136 of the aerofoil 101, adjacently to the leading edge 133.
  • the cavity 163 communicates with the second plurality of outlet cooling holes 141, thus permitting the cooling medium to exit the external surface 128 for cooling the aerofoil 101 at the leading edge 133.
  • the cavity 163 has a variable cross-sectional area.
  • the cavity 163 comprises a first section 154 having a first cross-sectional area, a second section 155 having a second cross-sectional area and a third section 156 having a third cross-sectional area.
  • the first, second and third section 154, 155, 156 are all orthogonal to the radial direction.
  • the first, second and third section 154, 155, 156 are
  • the second cross-sectional area is greater than the first cross-sectional area and greater than the third cross-sectional area.
  • the values of the cross-sectional areas along the cavity 163 are arranged in such a way that the flow rate of the cooling medium flows 145c through holes of the second plurality of cooling holes 141 which are intermediate between the base 134 and the tip 136 is greater than the flow rate of the cooling medium flows 145a, 145b which are closer to the base 134 or than the flow rate of the cooling medium flows 145d, 145e which are closer to the tip 136.
  • flow rates through holes 145a, 145b, 145c, 145d, 145e may be varied as necessary for optimum design by varying cross-section areas.
  • the gas turbine aerofoil 101 of Figs. 3 and 4 may be applied to a guiding vane 40.
  • the gas turbine aerofoil 100 and 101 of Figs. 2 and 2 to 4 respectively may be combined in the same the turbine blade 38 or guiding vane 40, i.e. both the cavities 153 and 163 may be present in the same aerofoil for
  • a third embodiment of a gas turbine aerofoil 102 is shown, applied to a turbine blade 38.
  • the turbine aerofoil 102 comprises, inside the external surface 128, an impingement tube 170 radially extending between the base 134 and the tip 136.
  • the internal passage 150 comprises a cavity 173 inside the impingement tube 170 directly connected to the root passage 123 and extending radially between the base 134 and the tip 136 of the aerofoil 102.
  • the cavity 173 communicates with the second plurality of outlet cooling holes 142, thus permitting the cooling medium to exit the external surface 174 of the impingement tube 170 for cooling from the inside a target portion of the external surface 128 of the aerofoil 102.
  • the cavity 173 has a variable cross-sectional area.
  • an insert 171 is provided inside the impingement tube 170, in such a way that the cavity 173 is therefore comprised between the insert 171 and the external surface 174 of the impingement tube 170.
  • the insert 171 has a smaller thickness in a region
  • the cavity 173 comprises a first section 154 having a first cross-sectional area, a second section 155 having a second cross-sectional area and a third section 156 having a third cross-sectional area.
  • the first, second and third section 154, 155, 156 are all orthogonal to the radial direction.
  • the first, second and third section 154, 155, 156 are disposed relatively to each other in such a way that the first section 154 is closer to the base 134, the third section 156 is closer to the tip 136 and the second section 155 is intermediate between the first section 154 and the third section 156.
  • the second cross-sectional area is greater than the first cross-sectional area and greater than the third cross-sectional area.
  • the values of the cross-sectional areas along the cavity 173 are arranged in such a way that the flow rate of the cooling medium flows 145c through holes of the second plurality of cooling holes 142 which are intermediate between the base 134 and the tip 136 is greater than the flow rate of the cooling medium flows 145a, 145b which are closer to the base 134 or than the flow rate of the cooling medium flows 145d, 145e which are closer to the tip 136.
  • flow rates through holes 145a, 145b, 145c, 145d, 145e may be varied as necessary for optimum design by varying cross-section areas.
  • the gas turbine aerofoil 102 of Figs. 5 and 6 may be applied to a guiding vane 40.
  • values of the cross-sectional areas along the cavities 153 or 163 or 173 of the turbine aerofoil 100, 101, 102, respectively are arranged in order to increase or decrease the flow rate of the cooling medium flows through a specific portion of the outlet cooling holes in order to respectively increase or decrease the cooling effectiveness in a specific region of the turbine aerofoil 100, 101, 102.
  • the direction of cooling medium is in a direction from the first section 154 towards the second section 155 and that the cross-section area of the cavity increases between the first section 154 towards the second section 155.
  • the direction of cooling medium is in a direction from the second section 155 towards the third section 156 and that the cross- section area of the cavity decreases between the second section 155 towards the third section 156.
  • the cooling medium travels in an increasing cross-sectional area cavity its velocity decreases and therefore its static pressure increases. Where there in an increase in static pressure in the cavity there is a corresponding increase in the amount of cooling medium that is ejected through the film cooling holes where there is an increase in static pressure. Similarly, where there is a decrease in the cross-sectional area of the cavity the cooling medium increases in velocity and

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

Abstract

La présente invention concerne un profil aérodynamique (100, 101, 102) destiné à une turbine à gaz (10) et comprenant : une surface externe (128) s'étendant axialement entre un bord d'attaque (133) et un bord de fuite (135) et s'étendant radialement entre une base (134) et une pointe (136) ; un passage interne (150) permettant d'acheminer un fluide de refroidissement vers une pluralité de trous de refroidissement de sortie (140, 141, 142), le passage interne comprenant une cavité (153, 163, 173) présentant une section transversale variable et communiquant avec au moins une partie des trous de refroidissement de sortie (140, 141, 142).
PCT/EP2017/050276 2016-01-15 2017-01-06 Profil aérodynamique de turbine à gaz WO2017121689A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP16151604 2016-01-15
EP16151604.2 2016-01-15

Publications (1)

Publication Number Publication Date
WO2017121689A1 true WO2017121689A1 (fr) 2017-07-20

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WO (1) WO2017121689A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3456923A1 (fr) * 2017-09-11 2019-03-20 MTU Aero Engines GmbH Aube d'une turbomachine pourvue de canal de refroidissement et corps de refoulement agencé dans une telle aube d'une turbomachine ainsi que procédé de fabrication
EP3889392A1 (fr) * 2020-03-31 2021-10-06 General Electric Company Aube de rotor de turbomachine comportant un circuit de refroidissement ayant une nervure de décalage

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0207799A2 (fr) 1985-07-03 1987-01-07 Westinghouse Electric Corporation Canaux de refroidissement pour les aubes d'une turbine à gaz
US7955053B1 (en) 2007-09-21 2011-06-07 Florida Turbine Technologies, Inc. Turbine blade with serpentine cooling circuit
EP2902589A1 (fr) 2014-01-29 2015-08-05 Siemens Aktiengesellschaft Composant refroidi par impact pour une turbine à gaz

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0207799A2 (fr) 1985-07-03 1987-01-07 Westinghouse Electric Corporation Canaux de refroidissement pour les aubes d'une turbine à gaz
US7955053B1 (en) 2007-09-21 2011-06-07 Florida Turbine Technologies, Inc. Turbine blade with serpentine cooling circuit
EP2902589A1 (fr) 2014-01-29 2015-08-05 Siemens Aktiengesellschaft Composant refroidi par impact pour une turbine à gaz

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3456923A1 (fr) * 2017-09-11 2019-03-20 MTU Aero Engines GmbH Aube d'une turbomachine pourvue de canal de refroidissement et corps de refoulement agencé dans une telle aube d'une turbomachine ainsi que procédé de fabrication
EP3889392A1 (fr) * 2020-03-31 2021-10-06 General Electric Company Aube de rotor de turbomachine comportant un circuit de refroidissement ayant une nervure de décalage
US11629601B2 (en) * 2020-03-31 2023-04-18 General Electric Company Turbomachine rotor blade with a cooling circuit having an offset rib

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