WO2014107217A1 - Distributeur de turbine hybride - Google Patents

Distributeur de turbine hybride Download PDF

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Publication number
WO2014107217A1
WO2014107217A1 PCT/US2013/065598 US2013065598W WO2014107217A1 WO 2014107217 A1 WO2014107217 A1 WO 2014107217A1 US 2013065598 W US2013065598 W US 2013065598W WO 2014107217 A1 WO2014107217 A1 WO 2014107217A1
Authority
WO
WIPO (PCT)
Prior art keywords
band
blind
structural
vanes
pockets
Prior art date
Application number
PCT/US2013/065598
Other languages
English (en)
Inventor
Joshua Brian JAMISON
Steven STRANG
Aaron Todd Williams
Original Assignee
General Electric Company
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 General Electric Company filed Critical General Electric Company
Priority to CN201380065235.2A priority Critical patent/CN104870754A/zh
Priority to BR112015014809A priority patent/BR112015014809A2/pt
Priority to EP13850000.4A priority patent/EP2935798A1/fr
Priority to CA2894854A priority patent/CA2894854A1/fr
Priority to JP2015549376A priority patent/JP2016505103A/ja
Publication of WO2014107217A1 publication Critical patent/WO2014107217A1/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
    • 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
    • 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
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/21—Oxide ceramics
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making

Definitions

  • This invention relates generally to gas turbine engines, and more particularly to turbine nozzles for such engines incorporating airfoils made of a low-ductility material.
  • a typical gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship.
  • the core is operable in a known manner to generate a primary gas flow.
  • the high pressure turbine also referred to as a gas generator turbine
  • the high pressure turbine includes one or more stages which extract energy from the primary gas flow. Each stage comprises a stationary turbine nozzle followed by a downstream rotor carrying turbine blades.
  • These components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life.
  • the air used for cooling is extracted (bled) from the compressor. Bleed air usage negatively impacts specific fuel consumption (“SFC”) and should generally be minimized.
  • SFC specific fuel consumption
  • CMCs ceramic matrix composites
  • the density of CMCs is approximately one-third of that of conventional metallic superalloys used in the hot section of turbine engines, so by replacing the metallic alloy with CMC while maintaining the same airfoil geometry, the weight of the component decreases.
  • the total weight of the assembly decreases, as well as the need for cooling air flow.
  • CMC and similar materials have unique mechanical properties that must be considered during design and application of an article such as a shroud segment.
  • CMC materials have relatively low tensile ductility or low strain to failure when compared with metallic materials.
  • CMCs have a coefficient of thermal expansion ("CTE") approximately one-third that of superalloys, which means that a rigid joint between the two different materials induces large strains and stresses with a change in temperature from the assembled condition.
  • CTE coefficient of thermal expansion
  • the allowable stress limits for CMCs are also lower than metal alloys which drives a need for simple and low stress design for CMC components.
  • a shroud apparatus for a gas turbine engine includes a turbine nozzle apparatus for a gas turbine engine includes: an annular inner band; an annular outer band circumscribing the inner band; a plurality of airfoil- shaped structural vanes extending between and interconnecting the inner band and the outer band; and a plurality of airfoil-shaped non-structural vanes extending between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.
  • a method for assembling a turbine nozzle for a gas turbine engine.
  • the method includes: providing an annular inner band, an annular outer band circumscribing the inner band, and a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; inserting an airfoil-shaped non-structural vane through an opening formed in one of the inner and outer bands; and closing the opening, such that the nonstructural vanes extend between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.
  • FIG. 1 is a schematic perspective view of a turbine nozzle assembly for a gas turbine engine, constructed according to an aspect of the present invention
  • FIG. 2 is an enlarged view of a portion of the turbine nozzle shown in FIG. 1;
  • FIG. 3 is a schematic perspective view of portion of an alternative turbine nozzle for a gas turbine engine
  • FIG. 4 is an enlarged view of a portion ofthe turbine nozzle shown in FIG. 3;
  • FIG. 5 is a cross-sectional view of a portion of a turbine nozzle; and
  • FIG. 6 is a view taken along lines 6-6 of FIG. 5.
  • FIGS. 1 and 2 depict an exemplary turbine nozzle 10 constructed according to an aspect ofthe present invention.
  • the turbine nozzle 10 is a stationary component forming part of a turbine section ofa gas turbine engine. It will be understood that the turbine nozzle 10 would be mounted in a gas turbine engine upstream of a turbine rotor with a rotor disk carrying an array of airfoil-shaped turbine blades, the nozzle and the rotor defining one stage ofthe turbine.
  • the primary function of the nozzle is to direct the combustion gas flow into the downstream turbine rotor stage.
  • a turbine is a known component of a gas turbine engine of a known type, and functions to extract energy from high-temperature, pressurized combustion gases from an upstream combustor (not shown) and to convert the energy to mechanical work, which is then used to drive a compressor, fan, shaft, or other mechanical load (not shown).
  • the principles described herein are equally applicable to turbofan, turbojet and turboshaft engines, as well as turbine engines used for other vehicles or in stationary applications.
  • axial or “longitudinal” refers to a direction parallel to an axis of rotation of a gas turbine engine
  • radial refers to a direction perpendicular to the axial direction
  • tangential or “circumferential” refers to a direction mutually perpendicular to the axial and tangential directions.
  • the turbine nozzle 10 includes an annular inner band 12 and an annular outer band 14, which define the inner and outer boundaries, respectively, of a hot gas flowpath through the turbine nozzle 10.
  • An array of airfoil-shaped turbine vanes is disposed between the inner band 12 and the outer band 14.
  • the array of vanes includes a group of structural vanes 16A (marked with an "x" in FIG. 1) alternating with a group of non-structural vanes 16B.
  • the turbine nozzle 10 may be considered a "hybrid" structure in that the structural and nonstructural vanes 16A and 16B are made from materials with different properties.
  • Each structural vane 16A has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 18 and a tip end 20.
  • a sufficient number of structural vanes 16A are provided so as to maintain a concentric relationship between the inner band 12 and the outer band 14 during engine operation and to control the relative thermal growth between the inner band 12 and the outer band 14.
  • the term "structural" identifies vanes 16A which are configured and mounted so as to transfer thermal and/or mechanical loads between the inner band 12 and the outer band 14.
  • the structural vanes 16A are functionally integral with the inner and outer bands 12 and 14, and may be part of a single cast or forged component, or may be welded, brazed, or mechanically fastened to the inner and outer bands 12 and 14.
  • the inner and outer bands 12 and 14 are each continuous annual structures, but alternatively one or both of the bands may be made up of a plurality of segments.
  • the structural vanes 16A are constructed from a strong, ductile material such as a metal alloy.
  • a strong, ductile material such as a metal alloy.
  • a known type of nickel-, iron-, or cobalt-based "superalloy" may be used for this purpose.
  • Each non-structural vane 16B has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 22 and a tip end 24.
  • One or more non-structural vanes 16B are disposed circumferentially between each pair of structural vanes 16A. In the specific example illustrated, there are 48 non-structural vanes 16B equally spaced around the circumference of the turbine nozzle 10, and the non-structural vanes 16B are disposed in groups of four.
  • a single structural vane 16A separates adjacent groups of non-structural vanes 16B.
  • non-structural identifies vanes 16B which are configured and mounted such that they do not transfer significant thermal and/or mechanical loads between the inner band 12 and the outer band 14.
  • vanes 16A and 16B are individually subject to significant aerodynamic (e.g. gas pressure) loads, and must have sufficient stiffness and yield strength to withstand these loads in operation.
  • Each of the non-structural vanes 16B is constructed from a low-ductility, high-temperature-capable material
  • a suitable material for the nonstructural vanes 16B is a ceramic matrix composite (CMC) material of a known type.
  • CMC materials include a ceramic type fiber for example SiC, forms of which are coated with a compliant material such as Boron Nitride (BN). The fibers are carried in a ceramic type matrix, one form of which is Silicon Carbide (SiC).
  • SiC Silicon Carbide
  • CMC type materials have a room temperature tensile ductility of no greater than about 1%, herein used to define and mean a low tensile ductility material.
  • CMC type materials have a room temperature tensile ductility in the range of about 0.4 to about 0.7%. This is compared with metals typically having a room temperature tensile ductility of at least about 5%, for example in the range of about 5 to about 15%.
  • the inner band 12 incorporates an array of airfoil-shaped blind root pockets 26 formed therein.
  • Each root pocket 26 receives the root end 22 of one of the non-structural vanes 16B.
  • the root pockets 26 are sized and shaped so that each permits a small gap between the root pocket 26 and the associated non-structural vane 16B.
  • the outer band 14 incorporates an array of apertures 28 formed therein.
  • Each aperture 28 is centered between adjacent structural vanes 16A, and each structural vane 16A carries an outer band segment 30 at its tip end 20.
  • An arcuate cover 32 is provided for each aperture 28.
  • the covers 32 are sized and shaped such that when installed in the apertures 28, they form a continuous annular structure in cooperation with the outer band segments 30.
  • each cover 32 has array of airfoil-shaped blind tip pockets 34 formed therein; alternatively, each cover 32 could have a fewer number of tip pockets 34, or an individual cover could be provided for each non-structural vane 16B.
  • Each tip pocket 34 receives the tip end 24 of one of the non-structural vanes 16B.
  • the tip pockets 34 are sized and shaped such that each permits a small gap between the tip pocket 34 and the associated non-structural vane 16B.
  • the turbine nozzle 10 is assembled as follows. First, the non-structural vanes 16B are inserted from radially outside the outer band 14, through the apertures 28, until their root ends 22 engage the root pockets 26. Next, a cover 32 is installed into each aperture 28. The tip ends 24 of the non-structural vanes 16B are then manipulated to enter the tip pockets 34 of the covers 32.
  • covers 32 are secured in the apertures 28. This could be done, for example, using known brazing or welding techniques, or by using mechanical fasteners (not shown). After engine service, the covers 32 may optionally be removed, permitting the non-structural vanes 16B to be replaced as needed, without replacing the entire nozzle 10.
  • the non-structural vanes 16B are "loosely" retained between the inner band 12 and the outer band 14 such that they are free to move to a predetermined, limited degree, for example about 0.25 mm (0.010 in.) to about 0.5 mm ( (0.020 in.).
  • gas pressure on the non-structural vanes 16B loads them against the pockets 26 and 34, preventing further movement in longitudinal (axial) and tangential directions, while permitting the inner and outer bands 12 and 14 to move radially relative to the non-structural vanes 16B.
  • FIGS. 5 and 6 illustrate a configuration in which a pin 33 passes transversely through the root end 22 of the nonstructural vane 16B and the root pocket 26, and a rib 35 formed as part of the tip pocket 34 engages a transverse slot 37 of the tip end 24 of the nonstructural vane 16B.
  • the effect is to make the nonstructural vane 16B "ride" radially with the inner band 12 while allowing free radial motion relative to the outer band 14, and simultaneously preventing axial motion.
  • FIGS. 3 and 4 illustrate a portion of a turbine nozzle 1 10 similar in construction to the turbine nozzle 10 described above, illustrating variations in how the vanes may be mounted.
  • the nozzle 1 10 includes an inner band 112, an outer band 114, and structural vanes 1 16A alternating with non-structural vanes 116B.
  • Each structural vane 116A has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 118 and a tip end 120.
  • the number and position of structural vanes 1 16A is selected as described above.
  • the structural vanes 1 16A are functionally integral with the inner and outer bands 112 and 114.
  • the structural vanes 116 A are constructed from a strong, ductile material such as a metal alloy.
  • a strong, ductile material such as a metal alloy.
  • a known type of nickel-, iron-, or cobalt-based "superalloy" may be used for this purpose.
  • Each non-structural vane 116B has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 122 and a tip end 124.
  • One or more non-structural vanes 116B are disposed circumferentially between each pair of structural vanes 1 16A.
  • Each of the non-structural vanes 1 16B is constructed from a low-ductility, high-temperature-capable material.
  • a suitable material for the nonstructural vanes 1 16B is a CMC material as described above.
  • the outer band 114 incorporates an array of blind tip pockets 126 formed therein.
  • Each tip pocket 126 receives the tip end 124 of one of the nonstructural vanes 1 16B.
  • the tip pockets 126 are sized and shaped so that each permits a small gap between the tip pocket 126 and the associated non-structural vane 116B.
  • the inner band 1 12 incorporates an array of split sections formed therein. Each split section is positioned between adjacent structural vanes 1 16A, and includes a fixed forward segment 128 that is functionally and structurally integral to the adjoining portions of the inner band 112, and a discrete aft segment 130.
  • the forward segment 128 and the aft segment 130 meet along a splitline "S" that lies in a radial-tangential plane.
  • the forward segment 128 defines forward sections 132 of a plurality of airfoil-shaped root pockets 134, and the aft segment 130 defines aft sections 136 of the root pockets 134.
  • the aft segments 130 are sized and shaped such that when installed against the forward segments 128, they form a continuous annular structure in cooperation with the forward segments 128.
  • This configuration of forward and aft segments may be referred herein to as a "split band" configuration.
  • Each root pocket 134 receives the root end 122 of one of the non-structural vanes 116B.
  • the root pockets 134 are sized and shaped such that each permits a small gap between itself and the associated nonstructural vane 116B.
  • the forward section 132 of the root pocket 134 may be made deeper in a radial direction than the aft section 136, to facilitate installation of the nonstructural vanes 1 16B.
  • the turbine nozzle 1 10 is assembled as follows. First, the non-structural vanes 1 16B are inserted from radially inside the outer band 114, until their tip ends 124 engage the tip pockets 126. The root ends 122 of the non-structural vanes 116B are pivoted into the forward sections 132 of the root pockets 134. The aft segments 130 are then installed with the aft sections 136 of the root pockets 134 receiving the root ends 122 of the nonstructural vanes 1 16B. [0041] Finally, the aft segments 130 are secured to the forward segments 128. This could be done, for example, using known brazing or welding techniques, or by using mechanical fasteners (not shown). After engine service, the aft segments 130 may optionally be removed, permitting the non-structural vanes 116B to be replaced as needed, without replacing the entire nozzle 110.
  • the configuration of the inner and outer bands may be varied as required to suit a particular application, so long as one of the two bands includes one of the features described above permitting assembly of the non-structural vanes into the turbine nozzle.
  • one of the two bands of a turbine nozzle would include apertures and associated covers, or a fore/aft split structure.
  • the other of the two bands could include only blind pockets, apertures and associated covers, or a fore/aft split structure.
  • the turbine nozzle described above has several advantages compared to the prior art.
  • the turbine nozzle of the present invention has a lower weight as compared to a completely-metallic turbine nozzle, by using a majority of CMC airfoils within a metallic frame. This turbine nozzle can also work to reduce cooling flow, because the majority of airfoils do not require air cooling.
  • This configuration allows the metal frame to dictate the thermal growth response of the nozzle, while the CMC airfoils are free thermally to grow and carry only aerodynamic pressure loading.
  • the CMC airfoils are seated to the inner and outer bands under running conditions by the aerodynamic loading, and the metallic bands and airfoil struts transfer the load to the outer case to allow conventional cantilevered nozzle configuration.
  • the invention maintains very similar thermal response of the nozzle assembly to the rest of the engine, compared to a completely-metallic nozzle.
  • Other features of a cantilevered nozzle e.g. seals and shields

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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 distributeur de turbine pour moteur à turbine à gaz comprenant: une bande interne annulaire (12); une bande externe annulaire (14) entourant la bande interne; une pluralité d'aubes structurelles (16A) en forme de surface portante s'étendant entre et reliant la bande interne et la bande externe; et une pluralité d'aubes non structurelles (16A) en forme de surface portante s'étendant entre la bande interne et la bande externe, chaque aube non structurelle présentant une extrémité d'emplanture (18) reçue par la bande interne, et une extrémité de pointe (20) reçue par la bande externe, de telle sorte que chaque aube non structurelle est libre de se déplacer d'un degré limité par rapport aux bandes interne et externe.
PCT/US2013/065598 2012-12-21 2013-10-18 Distributeur de turbine hybride WO2014107217A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
CN201380065235.2A CN104870754A (zh) 2012-12-21 2013-10-18 混合涡轮机喷嘴
BR112015014809A BR112015014809A2 (pt) 2012-12-21 2013-10-18 aparelho de bocal de turbina para um motor de turbina a gás e método para montar um bocal de turbina para um motor de turbina a gás
EP13850000.4A EP2935798A1 (fr) 2012-12-21 2013-10-18 Distributeur de turbine hybride
CA2894854A CA2894854A1 (fr) 2012-12-21 2013-10-18 Distributeur de turbine hybride
JP2015549376A JP2016505103A (ja) 2012-12-21 2013-10-18 ハイブリッドタービンノズル

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/725,150 2012-12-21
US13/725,150 US20140212284A1 (en) 2012-12-21 2012-12-21 Hybrid turbine nozzle

Publications (1)

Publication Number Publication Date
WO2014107217A1 true WO2014107217A1 (fr) 2014-07-10

Family

ID=50588790

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2013/065598 WO2014107217A1 (fr) 2012-12-21 2013-10-18 Distributeur de turbine hybride

Country Status (7)

Country Link
US (1) US20140212284A1 (fr)
EP (1) EP2935798A1 (fr)
JP (1) JP2016505103A (fr)
CN (1) CN104870754A (fr)
BR (1) BR112015014809A2 (fr)
CA (1) CA2894854A1 (fr)
WO (1) WO2014107217A1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2853688A3 (fr) * 2013-09-30 2015-07-22 MTU Aero Engines GmbH Aube pour une turbine à gaz
JP2016114050A (ja) * 2014-12-15 2016-06-23 ゼネラル・エレクトリック・カンパニイ セラミックマトリックス複合材取付のための装置及びシステム
EP3244022A1 (fr) * 2016-05-10 2017-11-15 General Electric Company Ensemble de turbine, ensemble de paroi interne de turbine et procédé d'ensemble de turbine
US10982564B2 (en) 2014-12-15 2021-04-20 General Electric Company Apparatus and system for ceramic matrix composite attachment

Families Citing this family (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2972380A1 (fr) * 2011-03-11 2012-09-14 Alstom Technology Ltd Procede de fabrication d'un diaphragme de turbine a vapeur
WO2014165518A1 (fr) * 2013-04-01 2014-10-09 United Technologies Corporation Dispositif d'aube de stator pour un moteur à turbine
BE1022361B1 (fr) * 2014-11-06 2016-03-17 Techspace Aero Sa Stator mixte de compresseur de turbomachine axiale.
US9863260B2 (en) 2015-03-30 2018-01-09 General Electric Company Hybrid nozzle segment assemblies for a gas turbine engine
US9845692B2 (en) * 2015-05-05 2017-12-19 General Electric Company Turbine component connection with thermally stress-free fastener
USD777212S1 (en) * 2015-06-20 2017-01-24 General Electric Company Nozzle ring
US9951632B2 (en) 2015-07-23 2018-04-24 Honeywell International Inc. Hybrid bonded turbine rotors and methods for manufacturing the same
US11067277B2 (en) * 2016-10-07 2021-07-20 General Electric Company Component assembly for a gas turbine engine
FR3072607B1 (fr) * 2017-10-23 2019-12-20 Safran Aircraft Engines Turbomachine comprenant un ensemble de redressement
US10927677B2 (en) 2018-03-15 2021-02-23 General Electric Company Composite airfoil assembly with separate airfoil, inner band, and outer band
US11466580B2 (en) * 2018-05-02 2022-10-11 General Electric Company CMC nozzle with interlocking mechanical joint and fabrication
US11454128B2 (en) * 2018-08-06 2022-09-27 General Electric Company Fairing assembly
US10767493B2 (en) 2019-02-01 2020-09-08 Rolls-Royce Plc Turbine vane assembly with ceramic matrix composite vanes
PL431184A1 (pl) * 2019-09-17 2021-03-22 General Electric Company Polska Spółka Z Ograniczoną Odpowiedzialnością Zespół silnika turbinowego
US11242770B2 (en) 2020-04-02 2022-02-08 General Electric Company Turbine center frame and method
US11319822B2 (en) 2020-05-06 2022-05-03 Rolls-Royce North American Technologies Inc. Hybrid vane segment with ceramic matrix composite airfoils
FR3111163B1 (fr) * 2020-06-04 2022-06-10 Safran Aircraft Engines Distributeur de turbine pour une turbomachine
US11773735B2 (en) 2021-12-22 2023-10-03 Rolls-Royce Plc Vane ring assembly with ceramic matrix composite airfoils
US12017782B2 (en) 2022-05-30 2024-06-25 Pratt & Whitney Canada Corp. Aircraft engine with stator having varying pitch
US11939886B2 (en) * 2022-05-30 2024-03-26 Pratt & Whitney Canada Corp. Aircraft engine having stator vanes made of different materials
US11952917B2 (en) * 2022-08-05 2024-04-09 Rtx Corporation Vane multiplet with conjoined singlet vanes

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS60209604A (ja) * 1984-04-04 1985-10-22 Mitsubishi Heavy Ind Ltd ガスタ−ビン静翼
GB2250782A (en) * 1990-12-11 1992-06-17 Rolls Royce Plc Stator vane assembly
EP1087103A2 (fr) * 1999-09-24 2001-03-28 General Electric Company Anneau de guidage de turbine en céramique
EP2213840A2 (fr) * 2009-01-30 2010-08-04 General Electric Company Support d'aubes statoriques pour une turbomachine et procédé pour minimiser le poids associé

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2654463A1 (fr) * 1989-11-15 1991-05-17 Snecma Element de stator de turbomachine.
FR2697285B1 (fr) * 1992-10-28 1994-11-25 Snecma Système de verrouillage d'extrémités d'aubes.
CN2155825Y (zh) * 1993-03-17 1994-02-16 陈森荣 膝关节
FR2738283B1 (fr) * 1995-08-30 1997-09-26 Snecma Agencement de turbomachine comprenant une grille d'aubes et un carter intermediaire
US8500394B2 (en) * 2008-02-20 2013-08-06 United Technologies Corporation Single channel inner diameter shroud with lightweight inner core
US8177513B2 (en) * 2009-02-18 2012-05-15 General Electric Company Method and apparatus for a structural outlet guide vane
EP2339120B1 (fr) * 2009-12-22 2015-07-08 Techspace Aero S.A. Étage redresseur de turbomachine et compresseur associé
US8454303B2 (en) * 2010-01-14 2013-06-04 General Electric Company Turbine nozzle assembly
CN102155825A (zh) * 2011-03-17 2011-08-17 杭州杰虹养殖有限公司 用于农业养殖地源热泵的蒸发器

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS60209604A (ja) * 1984-04-04 1985-10-22 Mitsubishi Heavy Ind Ltd ガスタ−ビン静翼
GB2250782A (en) * 1990-12-11 1992-06-17 Rolls Royce Plc Stator vane assembly
EP1087103A2 (fr) * 1999-09-24 2001-03-28 General Electric Company Anneau de guidage de turbine en céramique
EP2213840A2 (fr) * 2009-01-30 2010-08-04 General Electric Company Support d'aubes statoriques pour une turbomachine et procédé pour minimiser le poids associé

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2853688A3 (fr) * 2013-09-30 2015-07-22 MTU Aero Engines GmbH Aube pour une turbine à gaz
JP2016114050A (ja) * 2014-12-15 2016-06-23 ゼネラル・エレクトリック・カンパニイ セラミックマトリックス複合材取付のための装置及びシステム
EP3045685A1 (fr) * 2014-12-15 2016-07-20 General Electric Company Fixation mécanique et agencement associé de fixation d'aube de redresseur
US10982564B2 (en) 2014-12-15 2021-04-20 General Electric Company Apparatus and system for ceramic matrix composite attachment
EP3244022A1 (fr) * 2016-05-10 2017-11-15 General Electric Company Ensemble de turbine, ensemble de paroi interne de turbine et procédé d'ensemble de turbine

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BR112015014809A2 (pt) 2017-07-11
JP2016505103A (ja) 2016-02-18
CA2894854A1 (fr) 2014-07-10
US20140212284A1 (en) 2014-07-31
EP2935798A1 (fr) 2015-10-28
CN104870754A (zh) 2015-08-26

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