EP2586969B1 - Rotor à disques pour un moteur à turbine à gaz - Google Patents

Rotor à disques pour un moteur à turbine à gaz Download PDF

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Publication number
EP2586969B1
EP2586969B1 EP12190261.3A EP12190261A EP2586969B1 EP 2586969 B1 EP2586969 B1 EP 2586969B1 EP 12190261 A EP12190261 A EP 12190261A EP 2586969 B1 EP2586969 B1 EP 2586969B1
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EP
European Patent Office
Prior art keywords
rotor
recited
blades
spoke
turbine
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.)
Active
Application number
EP12190261.3A
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German (de)
English (en)
Other versions
EP2586969A3 (fr
EP2586969A2 (fr
Inventor
Gabriel L. Suciu
Stephen P. Muron
Ioannis Alvanos
Christopher M. Dye
Brian D. Merry
Arthur M. Salve
James W. Norris
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.)
RTX Corp
Original Assignee
United Technologies Corp
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
Priority claimed from US13/283,710 external-priority patent/US8784062B2/en
Priority claimed from US13/283,733 external-priority patent/US8944762B2/en
Priority claimed from US13/283,689 external-priority patent/US9938831B2/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2586969A2 publication Critical patent/EP2586969A2/fr
Publication of EP2586969A3 publication Critical patent/EP2586969A3/fr
Application granted granted Critical
Publication of EP2586969B1 publication Critical patent/EP2586969B1/fr
Active 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
    • 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
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • F01D11/008—Sealing the gap between rotor blades or blades and rotor by spacer elements between the blades, e.g. independent interblade platforms
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3061—Fixing blades to rotors; Blade roots ; Blade spacers by welding, brazing

Definitions

  • the present disclosure relates to a gas turbine engine, and more particularly to a rotor system therefor.
  • Gas turbine rotor systems include successive rows of blades, which extend from respective rotor disks that are arranged in an axially stacked configuration.
  • the rotor stack may be assembled through a multitude of systems such as fasteners, fusion, tie-shafts and combinations thereof.
  • TMF thermo-mechanical fatigue
  • DE 675 222 C discloses a turbine engine and a production method thereof.
  • US 2010/329849 A1 discloses a turbine rotor.
  • GB 802,871 A discloses a turbine rotor for gas or steam turbines.
  • DE 103 40 823 A1 discloses a blade for a compressor or a turbine disk of a gas turbine.
  • US 4 784 572 A discloses a circumferentially bonded rotor.
  • DE 10 2007 050142 A1 discloses a method for producing a blisk or a bling, the component produced therewith and a turbine blade.
  • DE 10 2008 057160 A1 discloses a method for exchanging an inner disk member of an integrally bladed disk.
  • the invention provides a rotor for a gas turbine engine as recited in claim 1.
  • the invention also provides a spool for a gas turbine engine as recited in claim 11.
  • the invention also provides a spool for a gas turbine engine as recited in claim 13.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 may be connected to the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30 which in one disclosed non-limiting embodiment includes a gear reduction ratio of, for example, at least 2.4:1.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor (HPC) 52 and high pressure turbine (HPT) 54.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the turbines 54, 46 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the gas turbine engine 20 is typically assembled in build groups or modules ( Figure 2 ).
  • the high pressure compressor 52 includes eight stages and the high pressure turbine 54 includes two stages in a stacked arrangement. It should be appreciated, however, that any number of stages will benefit hereform as well as other engine sections such as the low pressure compressor 44 and the low pressure turbine 46. Further, other gas turbine architectures such as a three-spool architecture with an intermediate spool will also benefit herefrom as well.
  • the high pressure compressor (HPC) 52 is assembled from a plurality of successive HPC rotors 60C which alternate with HPC spacers 62C arranged in a stacked configuration.
  • the rotor stack may be assembled in a compressed tie-shaft configuration, in which a central shaft (not shown) is assembled concentrically within the rotor stack and secured with a nut (not shown), to generate a preload that compresses and retains the HPC rotors 60C with the HPC spacers 62C together as a spool. Friction at the interfaces between the HPC rotor 60C and the HPC spacers 62C is solely responsible to prevent rotation between adjacent rotor hardware.
  • each HPC rotor 60C generally includes a plurality of blades 64 circumferentially disposed around a rotor disk 66.
  • the rotor disk 66 generally includes a hub 68, a rim 70, and a web 72 which extends therebetween.
  • Each blade 64 generally includes an attachment section 74, a platform section 76 and an airfoil section 78 ( Figure 5 ).
  • the HPC rotor 60C may be a hybrid dual alloy integrally bladed rotor (IBR) in which the blades 64 are manufactured of one type of material and the rotor disk 66 is manufactured of different material.
  • IBR integrally bladed rotor
  • Bi-metal construction provides material capability to separately address different temperature requirements.
  • the blades 64 are manufactured of a single crystal nickel alloy that are transient liquid phase bonded with the rotor disk 66 which is manufactured of a different material such as an extruded billet nickel alloy.
  • the blades 64 may be subject to a first type of heat treat and the rotor disk 66 to a different heat treat. That is, the Bi-metal construction as defined herein includes different chemical compositions as well as different treatments of the same chemical compositions such as that provided by differential heat treatment.
  • a spoke 80 is defined between the rim 70 and the attachment section 74.
  • the spoke 80 is a circumferentially reduced section defined by interruptions which produce axial or semi-axial slots which flank each spoke 80.
  • the spokes 80 may be machined, cut with a wire EDM or other processes to provide the desired shape.
  • An interface 80I that defines the transient liquid phase bond and or heat treat transition between the blades 64 and the rotor disk 66 is defined within the spoke 80. That is, the spoke 80 contains the interface 801.
  • Heat treat transition as defined herein is the transition between differential heat treatments.
  • the spoke 80 provides a reduced area subject to the thermo-mechanical fatigue (TMF) across the relatively high temperature gradient between the blades 64 which are within the relatively hot core gas path and the rotor disk 66 which is separated therefrom and is typically cooled with a secondary cooling airflow.
  • TMF thermo-mechanical fatigue
  • the HPC spacers 62C provide a similar architecture to the HPC rotor 60C in which a plurality of core gas path seals 82 are bonded or otherwise separated from a rotor ring 84 at an interface 86I defined along a spoke 86.
  • the seals 82 may be manufactured of the same material as the blades 64 and the rotor ring 84 may be manufactured of the same material as the rotor disk 66. That is, the HPC spacers 62C may be manufactured of a hybrid dual alloy which are transient liquid phase bonded at the spoke 86.
  • the HPC spacers 62C may be manufactured of a single material but subjected to the differential heat treat which transitions within the spoke 86.
  • a relatively low-temperature configuration will benefit from usage of a single material such that the spokes 86 facilitate a weight reduction.
  • low-temperature bi-metal designs may further benefit from dissimilar materials for weight reduction where, for example, low density materials may be utilized where load carrying capability is less critical.
  • the rotor geometry provided by the spokes 80, 86 reduces the transmission of core gas path temperature via conduction to the rotor disk 66 and the seal ring 84.
  • the spokes 80, 86 enable an IBR rotor to withstand increased T3 levels with currently available materials. Rim cooling may also be reduced from conventional allocations.
  • the overall configuration provides weight reduction at similar stress levels to current configurations.
  • the spokes 80, 86 in the disclosed non-limiting embodiment are oriented at a slash angle with respect to the engine axis A to minimize windage and the associated thermal effects. That is, the spokes are non-parallel to the engine axis A.
  • the passages which flank the spokes 80, 86 may also be utilized to define airflow paths to receive an airflow from an inlet HPC spacer 62CA.
  • the inlet HPC spacer 62CA includes a plurality of inlets 88 which may include a ramped flow duct 90 to communicate an airflow into the passages defined between the spokes 80, 86.
  • the airflow may be core gas path flow which is communicated from an upstream, higher pressure stage for use in a later section within the engine such as the turbine section 28.
  • various flow paths may be defined through combinations of the inlet HPC spacers 62CA to include but not limited to, core gas path flow communication, secondary cooling flow, or combinations thereof.
  • the airflow may be communicated not only forward to aft toward the turbine section, but also aft to forward within the engine 20. Further, the airflow may be drawn from adjacent static structure such as vanes to effect boundary flow turbulence as well as other flow conditions. That is, the HPC spacers 62C and the inlet HPC spacer 62CA facilitate through-flow for use in rim cooling, purge air for use downstream in the compressor, turbine, or bearing compartment operation.
  • the inlets 88' may be located through the inner diameter of an inlet HPC spacer 62CA' ( Figure 8 ).
  • the inlet HPC spacer 62CA' may be utilized to, for example, communicate a secondary cooling flow along the spokes 80, 86 to cool the spokes 80, 86 as well as communicate secondary cooling flow to other sections of the engine 20.
  • the inlets 88, 88' may be arranged with respect to rotation to essentially "scoop" and further pressurize the flow. That is, the inlets 88, 88' include a circumferential directional component.
  • each rotor ring 84 defines a forward circumferential flange 92 and an aft circumferential flange 94 which is captured radially inboard of the associated adjacent rotor rim 70. That is, each rotor ring 84 is captured therebetween in the stacked configuration.
  • the stacked configuration is arranged to accommodate the relatively lower-load capability alloys on the core gas path side of the rotor hardware, yet maintain the load-carrying capability between the seal rings 84 and the rims 70 to transmit rotor torque.
  • the alternating rotor rim 70 to seal ring 84 configuration carries the rotor stack preload - which may be upward of 150,000 lbs (66.7 kN) - through the high load capability material of the rotor rim 70 to seal ring 84 interface, yet permits the usage of a high temperature resistant, yet lower load capability materials in the blades 64 and the seal surface 82 which are within the high temperature core gas path. Divorce of the sealing area from the axial rotor stack load path facilitates the use of a disk-specific alloy to carry the stack load and allows for the high-temp material to only seal the rotor from the flow path.
  • the inner diameter loading and outer diameter sealing permits a segmented airfoil and seal platform design which facilitates relatively inexpensive manufacture and highly contoured airfoils.
  • the disclosed rotor arrangement facilitates a compressor inner diameter bore architectures in which the reduced blade/platform pull may be taken advantage of in ways that produce a larger bore inner diameter to thereby increase shaft clearance.
  • the HPC spacers 62C and HPC rotors 60C of the IBR may also be axially asymmetric to facilitate a relatively smooth axial rotor stack load path ( Figure 10 ).
  • the asymmetry may be located within particular rotor rims 70A and/or seal rings 84A.
  • the seal ring 84A includes a thinner forward circumferential flange 92 compared to a thicker aft circumferential flange 94 with a ramped interface 84Ai.
  • the ramped interface 84Ai provides a smooth rotor stack load path.
  • the load path along the spool may be designed in a more efficient manner as compared to the heretofore rather torturous conventional rotor stack load path ( Figure 11 ; RELATED ART).
  • the blades 64 and seal surface 82 are formed as segments that include tangential wire seals 96 between each pair of the multiple of seal surfaces 82 and each pair of the multiple of blades 64 as well as axial wire seals 98 between the adjacent HPC spacers 62C and HPC rotors 60C.
  • the tangential wire seals 96 and the axial wire seals 98 are located within teardrop shaped cavities 100 ( Figure 13 ) such that centrifugal forces increase the seal interface forces.
  • the high pressure compressor (HPC) 52 is discussed in detail above, it should be appreciated that the high pressure turbine (HPT) 54 ( Figure 14 ) is similarly assembled from a plurality of successive respective HPT rotor disks 60T which alternate with HPT spacers 62T ( Figure 15 ) arranged in a stacked configuration and the disclosure with respect to the high pressure compressor (HPC) 52 is similarly applicable to the high pressure turbine (HPT) 54 as well as other spools of the gas turbine engine 20 such as a low spool and an intermediate spool of a three-spool engine architecture. That is, it should be appreciated that other sections of a gas turbine engine may alternatively or additionally benefit herefrom.
  • each HPT rotor 60T generally includes a plurality of blades 102 circumferentially disposed around a rotor disk 124.
  • the rotor disk 124 generally includes a hub 126, a rim 128, and a web 130 which extends therebetween.
  • Each blade 102 generally includes an attachment section 132, a platform section 134, and an airfoil section 136 ( Figure 16 ).
  • the blades 102 may be bonded to the rim 128 along a spoke 136 at an interface 1361 as with the high pressure compressor (HPC) 52.
  • Each spoke 136 also includes a cooling passage 138 generally aligned with each turbine blade 102.
  • the cooling passage 138 communicates a cooling airflow into internal passages (not shown) of each turbine blade 102.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Claims (15)

  1. Rotor (60C ; 60T) pour un moteur à turbine à gaz comprenant :
    un disque de rotor (66 ; 124) défini le long d'un axe de rotation (A), ledit disque de rotor (66 ; 124) comportant un rebord (70 ; 128) ; et
    une pluralité d'aubes (64 ; 102) qui s'étend depuis ledit disque de rotor (66 ; 124), dans lequel chaque aube (64 ; 124) comporte une section de fixation (74 ; 132), une section de plate-forme (76 ; 134) et une section de surface portante (78 ; 136) s'étendant hors de ladite section de plate-forme (76 ; 134), et dans lequel chacune de ladite pluralité d'aubes (64 ; 102) s'étend depuis ledit disque de rotor (66 ; 124) au niveau d'une interface (80I ; 136I) entre ledit rebord (70 ; 128) et ladite section de fixation (74 ; 132), ladite interface (80I ; 136I) étant définie le long d'un disque (80 ; 136) qui comprend une section réduite de manière circonférentielle définie par des interruptions qui produisent des fentes axiales ou semi-axiales qui bordent chaque disque (80 ; 136), ladite section de plate-forme (76 ; 134) comprend au moins un évidement s'étendant le long d'un bord entier de la section de plate-forme (76 ; 134) pour définir avec un évidement adjacent une cavité en forme de goutte (100) qui reçoit un premier joint de fil (96) qui s'étend axialement, caractérisé en ce qu'un second joint de fil (98) s'étend de manière circonférentielle et est reçu dans une autre cavité en forme de goutte (100), dans lequel le second joint de fil (98) est configuré pour être situé entre le rotor (60C ; 60T) et une entretoise adjacente (62C ; 62T).
  2. Rotor selon la revendication 1, dans lequel ladite interface (80I ; 136I) comporte une transition de traitement thermique.
  3. Rotor selon la revendication 1, dans lequel ladite interface (80I ; 136I) comporte une liaison.
  4. Rotor selon une quelconque revendication précédente, dans lequel ledit disque de rotor (66 ; 124) est fabriqué en un premier matériau et ladite pluralité d'aubes (64 ; 102) est fabriquée en un second matériau, ledit premier matériau étant différent dudit second matériau.
  5. Rotor selon une quelconque revendication précédente, dans lequel chaque disque (80 ; 136) est parallèle audit axe de rotation.
  6. Rotor selon l'une quelconque des revendications 1 à 4, dans lequel chaque disque (80 ; 136) est incliné par rapport audit axe de rotation.
  7. Rotor selon une quelconque revendication précédente, dans lequel chaque disque (80 ; 136) est radialement à l'intérieur de l'une de ladite pluralité d'aubes (64 ; 102).
  8. Rotor selon une quelconque revendication précédente, dans lequel ladite pluralité d'aubes sont des aubes de compresseur haute pression (64).
  9. Rotor selon l'une quelconque des revendications 1 à 7, dans lequel ladite pluralité d'aubes sont des aubes de turbine haute pression (102).
  10. Rotor selon la revendication 9, dans lequel chacune de ladite pluralité d'aubes (102) comporte un passage de refroidissement (138) à l'intérieur dudit disque (136).
  11. Bobine pour un moteur à turbine à gaz comprenant :
    un rotor selon une quelconque revendication précédente, ledit disque de rotor étant un disque de rotor (66) de compresseur.
  12. Bobine selon la revendication 11, comprenant en outre :
    un disque de rotor (124) de turbine défini le long dudit axe de rotation et comportant un rebord (128) de disque de turbine ; et
    une pluralité d'aubes (102) de turbine qui s'étend depuis ledit disque de rotor (124) de turbine, chacune de ladite pluralité d'aubes (102) de turbine s'étendant depuis ledit disque de rotor (124) de turbine au niveau d'une interface entre ledit rebord (128) de disque de turbine et une plate-forme (134) d'aube de turbine, ladite interface (136I) étant définie le long d'un disque (136) qui comprend une section réduite de manière circonférentielle, chacune de ladite pluralité d'aubes (102) de turbine comportant éventuellement un passage de refroidissement (138) à l'intérieur dudit disque (136).
  13. Bobine pour un moteur à turbine à gaz comprenant :
    un rotor selon l'une quelconque des revendications 1 à 10 dans laquelle ladite interface d'aube (80I) est définie le long d'un disque (80) radialement à l'intérieur de la plate-forme (76) ; et comprenant en outre :
    une bague de rotor (84) qui est définie autour dudit axe de rotation, ladite bague de rotor (84) étant axialement adjacente audit disque de rotor (66) ;
    une pluralité de joints de trajet de gaz central (82) qui s'étend depuis ladite bague de rotor (84), chacun de ladite pluralité de joints de trajet de gaz central (82) s'étendant depuis ladite bague de rotor (84) au niveau d'une interface d'étanchéité (86I) s'étendant entre ladite bague de rotor (86I) et ledit joint (82), ladite interface d'étanchéité (86I) étant définie le long d'un disque (86) qui comprend une section réduite de manière circonférentielle, ladite pluralité de joints de trajet de gaz central (82) étant axialement adjacente à ladite plate-forme (76) d'aube.
  14. Bobine selon la revendication 13, dans laquelle ladite bague de rotor (84) et ledit disque de rotor (66) reçoivent une précharge de bloc rotor, dans laquelle, éventuellement, ladite précharge de bloc rotor définit un trajet de charge de bloc rotor axial radialement à l'intérieur de ladite interface d'aube (80I) et de ladite interface d'étanchéité (861).
  15. Bobine selon l'une quelconque des revendications 11 à 14, dans laquelle ladite bobine est une bobine haute pression.
EP12190261.3A 2011-10-28 2012-10-26 Rotor à disques pour un moteur à turbine à gaz Active EP2586969B1 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US13/283,710 US8784062B2 (en) 2011-10-28 2011-10-28 Asymmetrically slotted rotor for a gas turbine engine
US13/283,733 US8944762B2 (en) 2011-10-28 2011-10-28 Spoked spacer for a gas turbine engine
US13/283,689 US9938831B2 (en) 2011-10-28 2011-10-28 Spoked rotor for a gas turbine engine

Publications (3)

Publication Number Publication Date
EP2586969A2 EP2586969A2 (fr) 2013-05-01
EP2586969A3 EP2586969A3 (fr) 2017-05-03
EP2586969B1 true EP2586969B1 (fr) 2020-03-25

Family

ID=47148609

Family Applications (3)

Application Number Title Priority Date Filing Date
EP12190276.1A Active EP2586971B1 (fr) 2011-10-28 2012-10-26 Joints d'étanchéité, rotor, axe et procédé d'orientation des flux de force lors de l'assemblage du rotor
EP12190261.3A Active EP2586969B1 (fr) 2011-10-28 2012-10-26 Rotor à disques pour un moteur à turbine à gaz
EP12190264.7A Active EP2586970B1 (fr) 2011-10-28 2012-10-26 Espaceur de disque échelé pour un moteur à turbine à gaz

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EP12190276.1A Active EP2586971B1 (fr) 2011-10-28 2012-10-26 Joints d'étanchéité, rotor, axe et procédé d'orientation des flux de force lors de l'assemblage du rotor

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EP12190264.7A Active EP2586970B1 (fr) 2011-10-28 2012-10-26 Espaceur de disque échelé pour un moteur à turbine à gaz

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Publication number Priority date Publication date Assignee Title
EP3047103B1 (fr) 2013-09-17 2020-09-30 United Technologies Corporation Pales de turbine et procédés de fabrication
US10006364B2 (en) * 2014-08-20 2018-06-26 United Technologies Corporation Gas turbine rotors
US10837288B2 (en) 2014-09-17 2020-11-17 Raytheon Technologies Corporation Secondary flowpath system for a gas turbine engine
US9664058B2 (en) * 2014-12-31 2017-05-30 General Electric Company Flowpath boundary and rotor assemblies in gas turbines
EP3438410B1 (fr) 2017-08-01 2021-09-29 General Electric Company Système d'étanchéité pour machine rotative

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Also Published As

Publication number Publication date
EP2586969A3 (fr) 2017-05-03
EP2586971A3 (fr) 2017-05-24
EP2586970B1 (fr) 2019-04-24
EP2586971A2 (fr) 2013-05-01
EP2586971B1 (fr) 2019-06-12
EP2586970A2 (fr) 2013-05-01
EP2586969A2 (fr) 2013-05-01
EP2586970A3 (fr) 2017-05-24

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