EP3976934A1 - Procédé d'amélioration d'une turbine à gaz et turbine à gaz - Google Patents

Procédé d'amélioration d'une turbine à gaz et turbine à gaz

Info

Publication number
EP3976934A1
EP3976934A1 EP20739891.8A EP20739891A EP3976934A1 EP 3976934 A1 EP3976934 A1 EP 3976934A1 EP 20739891 A EP20739891 A EP 20739891A EP 3976934 A1 EP3976934 A1 EP 3976934A1
Authority
EP
European Patent Office
Prior art keywords
cooling air
guide vanes
gas turbine
guide
guide vane
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.)
Granted
Application number
EP20739891.8A
Other languages
German (de)
English (en)
Other versions
EP3976934B1 (fr
Inventor
Harald KUNTE
Robert Kunte
Karen Lee
Michael Wagner
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 Energy Global GmbH and Co KG
Original Assignee
Siemens Energy Global GmbH and Co KG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Energy Global GmbH and Co KG filed Critical Siemens Energy Global GmbH and Co KG
Publication of EP3976934A1 publication Critical patent/EP3976934A1/fr
Application granted granted Critical
Publication of EP3976934B1 publication Critical patent/EP3976934B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/80—Repairing, retrofitting or upgrading methods
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/15—Heat shield
    • 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
    • F05D2240/35—Combustors or associated equipment
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to a method for modernization of a gas turbine plant.
  • the invention also relates to a gas turbine system.
  • Gas turbine systems are known in the prior art in a wide variety of configurations. They comprise a combustion chamber lined with heat shield elements and a gas turbine arranged downstream of the combustion chamber and comprising guide vanes and rotor blades.
  • Those Hitzeschildele elements which are held in the downstream direction immediately in front of the gas turbine on the outside of a stationary annular support structure, and vane platforms of the guide vane in a first guide vane stage held on a stationary support structure define a radially inner and a radially outer ring gap between them due to the design.
  • Cooling air is introduced into these annular gaps via cooling air supply channels, which supply the guide vanes of the first guide vane stage with cooling air, in order to prevent overheating, in particular of the support structure, the support structure and the areas of said blade platforms facing the annular gap.
  • the introduction of the cooling air into the annular gap takes place mostly in the axial direction via cooling air openings which are formed on the face of the heat shield elements and are evenly distributed over the circumference of the annular gap.
  • the cooling air that is used to cool the heat shield elements is also used to cool the annular gap.
  • the cooling air passed through these cooling air channels thus enters the corresponding annular gap in the area of the pressure maxima and generates cooling air flows that prevent hot air from penetrating the annular gap in the area of the pressure maxima or in the area of the leading edges of the guide vanes.
  • the present invention creates a method for modernizing a gas turbine system, which has a combustion chamber lined with heat shield elements and a gas turbine arranged downstream of the combustion chamber, comprising guide vanes and rotor blades, the heat shield elements that are in the downstream direction immediately in front of the gas turbine on the The outside of a stationary support structure are held, and vane platforms of the guide vane held on a stationary support structure in a first guide vane stage define between them annular gaps, the method comprising the steps of: a) building out all guide vane in the first guide vane stage; b) Replacing the removed guide vane in the first guide vane stage with new or overhauled guide vane, with platforms of the new or refurbished guide vane being provided with cooling air bores which fluidly connect a cooling air supply duct that supplies the guide vane in the first guide vane stage with cooling air to one of the annular gaps and in open the corresponding annular gap, and where the Cooling air bores are arranged in such a way that more cooling air bore
  • the modernization method according to the invention is used in gas turbine systems that do not have any additional cooling in sections of the annular gaps in the area of the inflow edges of the guide vane in the first guide vane stage or the pressure maxima caused by this, there is a particular advantage in that to generate the cooling air holes no machining work in situ or on components that are difficult to dismantle, in particular on the support structure, which prevents unnecessary contamination of the gas turbine system during the implementation of the modernization process.
  • the cooling air ducts extending through the supporting structure are preferably at least partially closed after step a) and before step b), in particular all cooling air ducts being closed.
  • the positions of the cooling air outlet openings of the cooling channels formed in the support structure no longer correspond to the positions of the pressure maxima, which is why penetration of hot gas into the annular gaps in the area of the leading edges of the guide vanes can no longer be reliably prevented.
  • the present invention proposes to replace the cooling previously effected by these cooling channels at least partially, preferably completely, by cooling via the cooling air bores of the new guide vanes installed in step b). In this way, the advantage is also achieved that machining operations in situ or on components of the gas turbine system that are difficult to dismantle are avoided.
  • the cooling air bores formed in the vane platforms of the new or overhauled guide vanes define groups of cooling air bores arranged circumferentially at a distance from one another, which simplifies the manufacture of the guide vanes.
  • surfaces of the vane platforms facing in the radial direction of the guide vanes removed in step a) are provided with film cooling holes that are fluidically connected to one of the cooling air supply ducts when the vanes are installed, and surfaces of the vane platforms facing in the radial direction the new guide vanes installed in step b) are provided with film cooling holes which, when the guide vanes are installed, are fluidically connected to one of the cooling air supply ducts, the number of film cooling holes in the new or overhauled guide vanes being less than the number of film cooling holes in step a ) removed guide vane.
  • baffle plates are arranged on the vane platforms of the new or overhauled guide vanes provided with through holes, which are designed and arranged in such a way that they have to be passed by the cooling air coming from the corresponding cooling air supply duct in order to get to the film cooling holes. With such baffles an improved cooling is achieved.
  • each of the baffle plates is designed and arranged in such a way that an intermediate space remains between it and the film cooling holes (16).
  • some of the cooling air bores formed in the blade platforms of the new or overhauled guide vanes are arranged in such a way that they open into the intermediate space.
  • the present invention creates a gas turbine system which has a combustion chamber lined with heat shield elements and a gas turbine arranged downstream of the combustion chamber, comprising guide vanes and rotor blades, with those heat shield elements that are located in the downstream direction immediately in front of the gas turbine on the outside of a stationary one Support structure are held, and vane platforms of the guide vane held on a stationary support structure in a first guide vane stage define between them annular gaps, the vane platforms of the guide vanes being provided with cooling air bores which fluidically connect a cooling air supply duct that supplies the guide vane in the first guide vane stage with cooling air to one of the annular gaps and open into the corresponding annular gap.
  • cooling air bores open into areas of an annular gap which are arranged radially inward of the leading edges of the guide vanes than into other areas of the Ringspal th.
  • the cooling air bores formed in the vane platforms of the guide vane in the first guide vane stage define groups of cooling air bores that are circumferentially spaced from one another.
  • the cooling air holes of each cooling air hole group are positioned identically.
  • surfaces of the vane platforms of the vane in the first vane stage facing in the radial direction are provided with film cooling holes which, when the vane is installed, are fluidically connected to one of the cooling air supply channels.
  • the first Leit vane stage are advantageously versehe ne baffles with through holes arranged, which are designed and arranged so that they must be passed by the cooling air coming from one of the cooling air ducts to get to the film cooling holes.
  • Each of the baffle plates is preferably designed and arranged in such a way that there is an intermediate space between it and the film cooling holes.
  • Some of the cooling air bores are advantageously arranged in such a way that some of the cooling air bores open into the space.
  • FIG. 1 is a schematic sectional view of a partial area of a gas turbine plant
  • FIG. 2 is a partial view in the direction of the arrows II in FIG
  • Figure 3 is a perspective view of a guide vane
  • FIG. 1 a first guide vane stage of the gas turbine system shown in FIG. 1, in which a baffle plate is not shown;
  • FIG. 4 shows a perspective view of a new or overhauled guide vane, in which a baffle plate is not shown;
  • FIG. 5 shows a sectional view along the sectional plane V in
  • FIG. 6 is a view analogous to FIG. 4, which shows a new or overhauled guide vane with alternative patterns of cooling air bores.
  • the gas turbine system 1 shown in FIG. 1 comprises a combustion chamber 3 lined with heat shield elements 2 and a gas turbine 6 arranged downstream of the combustion chamber 3 and comprising guide vanes 4 and rotor blades 5.
  • Those heat shield elements 2 which, in the downstream direction immediately in front of the gas turbine 6 on the outside of a stationary Support structure 7, 8 are held, and on the stationary support structure 7 on the one hand and on a further stationary support structure 10 on the other hand held vane platforms 11 of the guide vanes 4 of the first guide vane stage define between each other annular gaps 12.
  • cooling air openings 13 which receive cooling air via cooling air supply channels 14, 15.
  • Thisdeluftöff openings 13 of the annular gap 12 is cooled during operation of the gas turbine system 1 with cooling air that was previously used for cooling the heat shield elements 2 was used.
  • cooling air channels 17 extending from the corresponding cooling air supply channel 14, 15 through the support structures 7 and 8 open out.
  • These cooling air channels 17 are used for this purpose to prevent the entry of hot gas into the annular gap 12 due to an inhomogeneous pressure distribution in the area of the annular gap 12.
  • This inhomogeneous pressure distribution is caused by the fact that the hot gas accumulates on the leading edges 16 of the guide vanes 4 of the first guide vane stage when it enters the gas turbine 6, whereby pressure maxima are generated in the area of the flow edges 16, due to which the hot gas is pressed into the annular gaps 12 becomes.
  • the guide vanes 4 of the first guide vane stage must be replaced.
  • all Leitschaufein 4 of the first Leitschaufeistu fe expanded are replaced by new guide vanes 4.
  • a problem that is associated with the fact that fewer new guide vanes 4 are installed than were previously installed is that the positions of the leading edges 16 of the guide vanes 4 and thus the positions of the pressure maxima of the inhomogeneous pressure distribution change.
  • the cooling air channels 17 extending through the support structures 7, 8 also no longer open in the correct positions in order to be able to effectively counteract penetration of hot gas into the annular gaps 12 in the area of the leading edges 16 of the guide vanes 4.
  • the blade platforms 11 of the new guide vanes 4, one of which is shown in FIGS. 4 and 5 are provided with cooling air bores 22, which fluidly connect the cooling air supply ducts 14, 15 to the annular gaps 12 and into the annular gaps 12 open.
  • These cooling air holes 22 are arranged in such a way that more cooling air holes 22 open into areas of the Ringspal te 12, which are arranged radially with respect to the leading edges 16 of the guide vanes 4 than in other areas of the annular gaps 12.
  • cooling air holes 22 take over the function of the cooling air channels 17.
  • six cooling air bores 22 are provided on each blade platform 11.
  • the other three cooling air bores 22 close with the axial direction here at an angle ß in the range between 15 ° -28 ° and open in the flow direction of the cooling air in front of the baffle 20.
  • the angle and ß depending on the construction of the guide vane have values in the range between 0 ° and 30 °.
  • the new guide vanes 4 are provided with film cooling holes 19, but their number is less than the number of film cooling holes 19 of the removed ones Guide vane 4 is.
  • the new guide screens 4 have fewer film cooling holes 19 than the old guide screens 4, as can be seen from the comparison of FIGS. 3 and 4. This has the advantage that part of the cooling air previously used for film cooling is now available for cooling the annular gaps 12, so that the total cooling air flow is not impaired due to the additional cooling air bores 22.
  • the cooling air ducts 17 extending through the support structures 7 and 8 can be left as they are. Alternatively, however, they can also be closed before the new guide vanes 4 are installed.
  • a major advantage associated with the design of the new guide vanes 4 is that no new cooling air ducts 17 have to be introduced into the support structures 7, 8 in order to feed the cooling air into the annular gaps 12 at the changing positions of the leading edges 16 of the Guide blades 4 and thus adjust the pressure maxima.
  • no machining operations need to be carried out in situ or on components of the gas turbine system 1 that are difficult to dismantle. Rather, the cooling air holes 22 can be taken gefer directly during the manufacture of the new guide vanes 4.
  • the method described above can also be carried out in gas turbine systems 1 that do not have any cooling air ducts 17 that counteract the penetration of hot gas into the annular gaps 12 in the area of the inflow edges 16 of the guide vanes 4.
  • a corresponding countermeasure against the ingress of hot air due to inhomogeneous pressure distribution is provided for the first time, regardless of whether the number of new or overhauled guide vanes 4 is less than, equal to or greater than the number of existing guide vanes 4 of the gas turbine plant 1 to be modernized.
  • the positions, orientations and the number of cooling air bores 22 of the new guide vanes 4 can vary. So FIG.
  • FIG. 6 shows an example of an alternative pattern of cooling air bores 22 opening into an annular gap 12 radially of a leading edge 13. It should also be pointed out that the new or overhauled guide vanes 4 can only be provided with cooling air bores 21 on one of their blade platforms 11, so that through the Leitschaufein 4 only cooling air is passed into one of the two annular gaps 12.

Landscapes

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

Abstract

L'invention concerne un procédé d'amélioration d'une turbine à gaz (1). Ledit procédé comprend les étapes consistant à : a) retirer toutes les aubes directrices (4) du premier étage d'aube directrice ; b) remplacer les aubes directrices retirées (4) du premier étage d'aube directrice avec des aubes directrices nouvelles ou reconditionnées (4). Les plateformes d'aube (11) des aubes directrices nouvelles ou reconditionnées (4) sont pourvues d'alésages d'air de refroidissement (19) qui relient de façon fluidique un conduit d'alimentation en air de refroidissement (12) à l'espace annulaire (10) et débouchent dans l'espace annulaire (10), et les trous d'air de refroidissement (19) sont disposés de telle manière que plus de trous d'air de refroidissement (19) débouchent dans des zones de l'espace annulaire (10) qui sont disposées radialement vers l'intérieur à partir des bords d'attaque (13) des aubes directrices (4) que dans d'autres zones de l'espace annulaire (10). L'invention concerne en outre une turbine à gaz améliorée correspondante (1).
EP20739891.8A 2019-07-31 2020-06-29 Procédé de modernisation d'une installation de turbine à gaz et turbine à gaz Active EP3976934B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102019211418.0A DE102019211418A1 (de) 2019-07-31 2019-07-31 Verfahren zur Modernisierung einer Gasturbinenanlage sowie Gasturbinenanlage
PCT/EP2020/068226 WO2021018495A1 (fr) 2019-07-31 2020-06-29 Procédé d'amélioration d'une turbine à gaz et turbine à gaz

Publications (2)

Publication Number Publication Date
EP3976934A1 true EP3976934A1 (fr) 2022-04-06
EP3976934B1 EP3976934B1 (fr) 2023-06-14

Family

ID=71607923

Family Applications (1)

Application Number Title Priority Date Filing Date
EP20739891.8A Active EP3976934B1 (fr) 2019-07-31 2020-06-29 Procédé de modernisation d'une installation de turbine à gaz et turbine à gaz

Country Status (4)

Country Link
US (1) US11879346B2 (fr)
EP (1) EP3976934B1 (fr)
DE (1) DE102019211418A1 (fr)
WO (1) WO2021018495A1 (fr)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2023132236A1 (fr) * 2022-01-06 2023-07-13 三菱重工業株式会社 Aube statique de turbine, structure d'ajustement et turbine à gaz

Family Cites Families (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP3324256B2 (ja) * 1994-02-01 2002-09-17 石川島播磨重工業株式会社 タービン静翼の組立方法
EP0902164B1 (fr) * 1997-09-15 2003-04-02 ALSTOM (Switzerland) Ltd Refroidissement de la platte-forme dans les turbines à gas
US6154959A (en) * 1999-08-16 2000-12-05 Chromalloy Gas Turbine Corporation Laser cladding a turbine engine vane platform
US7775050B2 (en) * 2006-10-31 2010-08-17 General Electric Company Method and apparatus for reducing stresses induced to combustor assemblies
US8973374B2 (en) * 2007-09-06 2015-03-10 United Technologies Corporation Blades in a turbine section of a gas turbine engine
EP2229507B1 (fr) * 2007-12-29 2017-02-08 General Electric Technology GmbH Turbine à gaz
US8118554B1 (en) * 2009-06-22 2012-02-21 Florida Turbine Technologies, Inc. Turbine vane with endwall cooling
EP2634373A1 (fr) * 2012-02-28 2013-09-04 Siemens Aktiengesellschaft Agencement pour turbomachine
EP2754858B1 (fr) 2013-01-14 2015-09-16 Alstom Technology Ltd Dispositif pour étanchéifier une cavité ouverte contre un entraînement gazeux chaud
EP3085900B1 (fr) * 2015-04-21 2020-08-05 Ansaldo Energia Switzerland AG Lèvre abradable pour une turbine à gaz
EP3141702A1 (fr) * 2015-09-14 2017-03-15 Siemens Aktiengesellschaft Segment d'aube directrice de turbine à gaz et procédé de fabrication
US10252790B2 (en) * 2016-08-11 2019-04-09 General Electric Company Inlet assembly for an aircraft aft fan
DE102017212575A1 (de) * 2017-07-21 2019-01-24 Siemens Aktiengesellschaft Verfahren zur Erhöhung der Leistung einer Gasturbine

Also Published As

Publication number Publication date
WO2021018495A1 (fr) 2021-02-04
DE102019211418A1 (de) 2021-02-04
US20220268172A1 (en) 2022-08-25
EP3976934B1 (fr) 2023-06-14
US11879346B2 (en) 2024-01-23

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