US3761200A - Bladed rotors - Google Patents

Bladed rotors Download PDF

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Publication number
US3761200A
US3761200A US00202395A US3761200DA US3761200A US 3761200 A US3761200 A US 3761200A US 00202395 A US00202395 A US 00202395A US 3761200D A US3761200D A US 3761200DA US 3761200 A US3761200 A US 3761200A
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US
United States
Prior art keywords
platform
slots
blade
ring
disc
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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.)
Expired - Lifetime
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US00202395A
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English (en)
Inventor
T Gardiner
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SECR DEFENCE
STATE FOR DEFENCE GB
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SECR DEFENCE
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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
    • 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
    • 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
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F01D5/3015—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type with side plates
    • 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 bladed rotors for gas turbine engines and has particular reference to a construction of platforms for the rotor blades of the rotor.
  • each stage comprises an individual bladed rotor disposed between stator blades. It is necessary to provide a seal between the bladed rotors and their adjacent stator blades.
  • the blades must be provided with platforms which make up the inner surface of the gas flow passage and the weight of these has been supported by the disc, adding to the disc weight.
  • the object of the present invention is to provide a bladed rotor in which said seals and platforms are provided in a manner which enables the weight of the rotor to be reduced.
  • a bladed rotor for a gas turbine engine comprises an annular rotor disc having a plurality of slots in its outer periphery, a plurality of aero-foil shaped blades having root portions engaged in the slots, a platform ring surrounding the periphery of the disc, said ring having slots therein into which the blades are fitted, and having at least at one axial end, a continuous circumferential sealing member, means being provided for removably attaching the platform ring to the disc.
  • the blades may have a vestigial platform portion on each aerofoil surface adjacent the rootof the blade, and the platform ring may be in the form of a cage hav: ing a plurality of axially extending bars defining the slots therebetween, the slots being dimensioned to receive said platform portions.
  • the vestigial blade platforms may be shaped to provide friction surfaces for damping vibration of the blades and for giving additional support to the portions of the platform ring which extend between the blades.
  • the radially outer part of the platform ring extends axially of the disc and has radially inwardly extending flanges at each end which extend at least partly over the radially outer portions of the axial end faces of the disc, the sealing members extending axially from the flanges.
  • At least one of the flanges may extend radially to a point radially inwardly of the bottoms of the slots so that the flange also acts as a sealing plate, to prevent fluid from the high pressure side of the rotor from leaking back to the low pressure side.
  • the sealing ring is preferably attached to the disc by means of bolts.
  • FIG. 1 is a diagrammatic representation of a gas turbine engine incorporating a bladed rotor of the present invention
  • FIG. 2 is a sectional view of the compressor of the engine of FIG. 1 which is constructed according to the present invention
  • FIG. 3 is a pictorial view of a portion of the compressor of FIG. 2,
  • FIG. 4 is a section of the first row of blades of FIG. 2 on the line A-A, and
  • FIG. 5 is an alternative blade construction to that shown in FIG. 4.
  • FIG. 1 a gas turbine jet propulsion engine, which may be of any desired form having compressor means 1, combustion equipment 2, turbine means 3 and a propulsion nozzle 4 all in flow series.
  • the engine may be a single or multi-shaft engine since the invention is concerned only with a bladed rotor of the engine.
  • the engine is a three shaft engine having a three stage front fan in which all the fan rotors are constructed according to the invention.
  • the frontengine casing is shown broken away in FIG. 1 to illustrate the front fan, and FIG. 2 shows the fan in greater detail.
  • the fan comprises three bladed rotors 10 each having a row of stator vanes 11 downstream thereof.
  • the rotors are drivingly interconnected by means of a shaft 12 to a turbine of the engine.
  • Each rotor comprises a disc 15 which is slotted at its outerperiphery to receive the roots 16 of a plurality of rotor blades 17.
  • the roots of the rotor blades are of substantially dove-tail cross-section having an included angle of but clearly any other suitable shapes may be used, for example a conventional fir tree root.
  • Each rotor blade has vestigial platform portions 18, and a platform ring 19 is provided in the form of a cage which surrounds the outer periphery of the rotor to provide a platform surface between the adjacent platform portions 18 of the blade, which surface forms the inner boundary of the fluid flow passage through the duct.
  • the radially outer portion of the cage comprises a plurality of bars 20, which extend axially of the rotor between the axial end faces of the disc, and between which are defined a plurality of slots for receiving the vestigial platform portions 18 of the blade.
  • the assembly is such that the blades can be inserted into the platform ring from the outside, and thus the slots in the platform ring and the vestigial blade platforms, must be slightly wider than the blade roots to enable the blade roots to pass through said slots. Once all the blades are in position in the platform ring the assembly is offered up to the rotor disc and the blade roots are inserted into the slots in the disc.
  • each blade and its vestigial platform can be made more easily as a precision forging from the raw material, and there is much less machining to be done on the blade and hence much less waste of material. Because of this, much more of the material of the blade which has been worked during the forging process, i.e., the parts near the surfaces of the material, is retained in the finished article and the blades are therefore stronger. In addition the weight of the blade and its vestigial platform is less than if the full platform were formed on the blade, and the weight of disc required to support this reduced weight is therefore reduced.
  • each axial end of the platform ring 19 there are radially inwardly extending flanges 21 and 22 which carry axially extending continuous sealing rings 23 and 24.
  • the sealing rings 23 and 24 are disposed at such a radius from the centre of the engine, as to be capable of co-operating with corresponding sealing surfaces 25 on the radially inner surfaces of shrouds 26 ofthe adjacent stator vanes.
  • the upstream rotor blade row has only one sealing ring, which extends downstream therefrom.
  • each platform ring there is a further radially extending flange 27 which is locally enlarged at several places around the rotor to provide bosses for bolting the platform ring to the disc 15.
  • the advantage of making the platform in this way is that the cylindrical sealing rings 23 and 24 are selfsupporting, as long as they lie within the free hoop radius of the rotor, and can, with the aid of the radial flanges 21 and 22, support the bars of the cage.
  • the cylindrical sealing rings 23 and 24 are selfsupporting, as long as they lie within the free hoop radius of the rotor, and can, with the aid of the radial flanges 21 and 22, support the bars of the cage.
  • the radially extending flanges 21 and 22 act as locking means for preventing axial movement of the blade roots in the slots in the disc, and the flanges 21, being continuous, further act as sealing plates for preventing recirculation of high pressure fluid from downstream of the blade row back to the low pressure upstream side through the slots in the disc.
  • Separate sealing plates and locking tabs can therefore be dispensed with, thus saving time in the assembly of the rotor.
  • a still further advantage of this construction is provided by having the platform rings carry the upstream and downstream sealing rings 23 and 24.
  • This provides a single, easily removable component, which not only aids assembly of the rotor but allows the seals to be replaced, should they become worn or damaged, without having to disturb the disc. Such replacement can be done easily in the field, which is not the case where the seals are integral with or shrunk onto the disc rim.
  • the seal is carried at the greatest possible radius, and the volume 28 between the rotor blade row and the stator vanes between the radii of the platform surface and the seal, is minimized.
  • Nuts 30 are removed thus releasing the nose bullet from the first row of rotor blades.
  • the ring of bolts 31 are then removed thus allowing the platform ring and the blades to slide out forwardly, disengaging the blade roots from the disc slots.
  • the blades can then be removed from the platform ring.
  • the ring of stators are then removable axially forwardly by removing bolts (not shown) which connect the stators to the engine casing.
  • the second stage blades and platform ring can be removed.
  • the second stage stators and third stage blades are removed in exactly similar manner.
  • platform rings as described above combine into a single removable piece, the sealing rings, locking means and sealing plates, most of which on present engines are connected separately to the rotor.
  • FIGS. 4 and 5 illustrate how vibration damping features can also be added simply to the rotor.
  • the vestigial platforms can be shaped to provide a friction damping surface where they abut the platform rmg. I
  • the axial surfaces of the blade platforms are provided with oppositely radially outwardly inclined surfaces 35 and 36 adapted to engage with corresponding surfaces on the platform ring, and in FIG. 5 the axial surfaces have circumferentially extending lips 37 and 38 which are arranged to fit over a corresponding lip on the platform ring.
  • the contacting surfaces may aid in the assembly of the rotor by preventing the vestigial blade platforms from passing completely through slots in the platform ring in place of the axially extending lips at the bottom of the slots on which the ends of the blade platforms seat.
  • the bars 20 of the cage may be provided with radially extending ribs 40 to cut down outward deflection of the bars under centrifugal loads.
  • damping means which may be in the form of U-shaped spring members 42 to be fitted into the spaces between the ribs 40 and the roots 16 of the blades.
  • the blades may be made without any platform at all on their aerofoil surfaces, and the slots in the platform ring may be reduced to aerofoil shaped slots only.
  • the blades may be inserted into the slots from inside the platform ring, or may have separate roots brazed or welded on after the aerofoil part has been inserted.
  • a bladed rotor for a gas turbine engine comprising an annular rotor disc having means defining a plurality of slots in its outer periphery, and a plurality of aerofoil shaped blades having root portions engaged in the slots, a platform ring surrounding the periphery of the disc, said ring having means defining slots therein into which the blades are fitted, and having at least at one axial end a continuous circumferential sealing member,
  • each blade has a vestigial platform portion on each aerofoil surface adjacent the root of the blade, and the platform ring is in the form of a cage having a plurality of axially extending bars defining the slots therebetween, the slots being dimensioned to receive said plat form portions.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00202395A 1970-12-05 1971-11-26 Bladed rotors Expired - Lifetime US3761200A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB5790470 1970-12-05

Publications (1)

Publication Number Publication Date
US3761200A true US3761200A (en) 1973-09-25

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ID=10480326

Family Applications (1)

Application Number Title Priority Date Filing Date
US00202395A Expired - Lifetime US3761200A (en) 1970-12-05 1971-11-26 Bladed rotors

Country Status (5)

Country Link
US (1) US3761200A (OSRAM)
DE (1) DE2166499A1 (OSRAM)
FR (1) FR2116522B1 (OSRAM)
GB (1) GB1318654A (OSRAM)
IT (1) IT945233B (OSRAM)

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3841792A (en) * 1973-03-09 1974-10-15 Westinghouse Electric Corp Turbomachine blade lock and seal device
US3867069A (en) * 1973-05-04 1975-02-18 Westinghouse Electric Corp Alternate root turbine blading
US3923420A (en) * 1973-04-30 1975-12-02 Gen Electric Blade platform with friction damping interlock
US4265594A (en) * 1978-03-02 1981-05-05 Bbc Brown Boveri & Company Limited Turbine blade having heat localization segments
US4668167A (en) * 1985-08-08 1987-05-26 Societe National D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Multifunction labyrinth seal support disk for a turbojet engine rotor
US4685863A (en) * 1979-06-27 1987-08-11 United Technologies Corporation Turbine rotor assembly
US4767247A (en) * 1987-02-24 1988-08-30 Westinghouse Electric Corp. Apparatus and method for preventing relative blade motion in steam turbine
US5049035A (en) * 1988-11-23 1991-09-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Bladed disc for a turbomachine rotor
US5244345A (en) * 1991-01-15 1993-09-14 Rolls-Royce Plc Rotor
US5277548A (en) * 1991-12-31 1994-01-11 United Technologies Corporation Non-integral rotor blade platform
US5601404A (en) * 1994-11-05 1997-02-11 Rolls-Royce Plc Integral disc seal
US6189891B1 (en) * 1997-03-12 2001-02-20 Mitsubishi Heavy Industries, Ltd. Gas turbine seal apparatus
US20060248929A1 (en) * 2005-05-06 2006-11-09 I.L.S.A. Spa Fabric articles dry cleaning machine by solvent nebulization
US20060275107A1 (en) * 2005-06-07 2006-12-07 Ioannis Alvanos Combined blade attachment and disk lug fluid seal
US20060275108A1 (en) * 2005-06-07 2006-12-07 Memmen Robert L Hammerhead fluid seal
US20060275106A1 (en) * 2005-06-07 2006-12-07 Ioannis Alvanos Blade neck fluid seal
US20070098545A1 (en) * 2005-10-27 2007-05-03 Ioannis Alvanos Integrated bladed fluid seal
FR2895021A1 (fr) * 2005-12-16 2007-06-22 Snecma Sa Systeme d'etancheite inter-etages dans une turbomachine
US20070189901A1 (en) * 2003-03-22 2007-08-16 Dundas Jason E Separable blade platform
US20070224035A1 (en) * 2005-09-16 2007-09-27 General Electric Company Angel wing seals for turbine blades and methods for selecting stator, rotor and wing seal profiles
US20080044284A1 (en) * 2006-08-16 2008-02-21 United Technologies Corporation Segmented fluid seal assembly
US20080112811A1 (en) * 2006-11-13 2008-05-15 United Technologies Corporation Hoop seal with partial slot geometry
EP1992787A1 (en) * 2007-05-15 2008-11-19 General Electric Company Turbine rotor blade assembly comprising a removable platform
US20090110548A1 (en) * 2007-10-30 2009-04-30 Pratt & Whitney Canada Corp. Abradable rim seal for low pressure turbine stage
US20090208326A1 (en) * 2006-09-08 2009-08-20 Eric Durocher Rim seal for a gas turbine engine
US20100166551A1 (en) * 2008-12-29 2010-07-01 Morrison Adam J Hybrid turbomachinery component for a gas turbine engine
US20100189556A1 (en) * 2009-01-28 2010-07-29 United Technologies Corporation Segmented ceramic matrix composite turbine airfoil component
US20100290914A1 (en) * 2009-05-15 2010-11-18 Souers Philip F Blade Closing Key System for a Turbine Engine
CN102733863A (zh) * 2011-03-31 2012-10-17 通用电气公司 具有气体流的增强抑制特征的定子-转子组件及相关方法
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
US20130071248A1 (en) * 2011-09-19 2013-03-21 General Electric Company Compressive stress system for a gas turbine engine
CN101315032B (zh) * 2007-05-30 2013-06-19 通用电气公司 具有增强的气流封闭表面特征的定子-转子组件和方法
US20180094638A1 (en) * 2016-10-05 2018-04-05 Pratt & Whitney Canada Corp. Integrally bladed fan rotor
US20180112543A1 (en) * 2016-10-26 2018-04-26 Rolls-Royce Corporation Turbine wheel assembly with ceramic matrix composite components
US10570767B2 (en) 2016-02-05 2020-02-25 General Electric Company Gas turbine engine with a cooling fluid path
US10633992B2 (en) 2017-03-08 2020-04-28 Pratt & Whitney Canada Corp. Rim seal

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4084919A (en) * 1976-06-29 1978-04-18 United Technologies Corporation Means of attaching a seal to a disk
GB1549152A (en) * 1977-01-11 1979-08-01 Rolls Royce Rotor stage for a gas trubine engine
US5156528A (en) * 1991-04-19 1992-10-20 General Electric Company Vibration damping of gas turbine engine buckets

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2931624A (en) * 1957-05-08 1960-04-05 Orenda Engines Ltd Gas turbine blade
US3008689A (en) * 1954-08-12 1961-11-14 Rolls Royce Axial-flow compressors and turbines
US3266770A (en) * 1961-12-22 1966-08-16 Gen Electric Turbomachine rotor assembly
US3572970A (en) * 1969-01-23 1971-03-30 Gen Electric Turbomachinery blade spacer
US3644058A (en) * 1970-05-18 1972-02-22 Westinghouse Electric Corp Axial positioner and seal for turbine blades

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3008689A (en) * 1954-08-12 1961-11-14 Rolls Royce Axial-flow compressors and turbines
US2931624A (en) * 1957-05-08 1960-04-05 Orenda Engines Ltd Gas turbine blade
US3266770A (en) * 1961-12-22 1966-08-16 Gen Electric Turbomachine rotor assembly
US3572970A (en) * 1969-01-23 1971-03-30 Gen Electric Turbomachinery blade spacer
US3644058A (en) * 1970-05-18 1972-02-22 Westinghouse Electric Corp Axial positioner and seal for turbine blades

Cited By (53)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3841792A (en) * 1973-03-09 1974-10-15 Westinghouse Electric Corp Turbomachine blade lock and seal device
US3923420A (en) * 1973-04-30 1975-12-02 Gen Electric Blade platform with friction damping interlock
US3867069A (en) * 1973-05-04 1975-02-18 Westinghouse Electric Corp Alternate root turbine blading
US4265594A (en) * 1978-03-02 1981-05-05 Bbc Brown Boveri & Company Limited Turbine blade having heat localization segments
US4685863A (en) * 1979-06-27 1987-08-11 United Technologies Corporation Turbine rotor assembly
US4668167A (en) * 1985-08-08 1987-05-26 Societe National D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Multifunction labyrinth seal support disk for a turbojet engine rotor
US4767247A (en) * 1987-02-24 1988-08-30 Westinghouse Electric Corp. Apparatus and method for preventing relative blade motion in steam turbine
US5049035A (en) * 1988-11-23 1991-09-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Bladed disc for a turbomachine rotor
US5244345A (en) * 1991-01-15 1993-09-14 Rolls-Royce Plc Rotor
US5277548A (en) * 1991-12-31 1994-01-11 United Technologies Corporation Non-integral rotor blade platform
US5601404A (en) * 1994-11-05 1997-02-11 Rolls-Royce Plc Integral disc seal
US6189891B1 (en) * 1997-03-12 2001-02-20 Mitsubishi Heavy Industries, Ltd. Gas turbine seal apparatus
US20070189901A1 (en) * 2003-03-22 2007-08-16 Dundas Jason E Separable blade platform
US7284958B2 (en) 2003-03-22 2007-10-23 Allison Advanced Development Company Separable blade platform
US20060248929A1 (en) * 2005-05-06 2006-11-09 I.L.S.A. Spa Fabric articles dry cleaning machine by solvent nebulization
US7610780B2 (en) 2005-05-06 2009-11-03 I.L. S.A. SpA Fabric articles dry cleaning machine by solvent nebulization
US20060275108A1 (en) * 2005-06-07 2006-12-07 Memmen Robert L Hammerhead fluid seal
US20060275106A1 (en) * 2005-06-07 2006-12-07 Ioannis Alvanos Blade neck fluid seal
US20060275107A1 (en) * 2005-06-07 2006-12-07 Ioannis Alvanos Combined blade attachment and disk lug fluid seal
US20070224035A1 (en) * 2005-09-16 2007-09-27 General Electric Company Angel wing seals for turbine blades and methods for selecting stator, rotor and wing seal profiles
US7465152B2 (en) * 2005-09-16 2008-12-16 General Electric Company Angel wing seals for turbine blades and methods for selecting stator, rotor and wing seal profiles
US20070098545A1 (en) * 2005-10-27 2007-05-03 Ioannis Alvanos Integrated bladed fluid seal
US7334983B2 (en) * 2005-10-27 2008-02-26 United Technologies Corporation Integrated bladed fluid seal
FR2895021A1 (fr) * 2005-12-16 2007-06-22 Snecma Sa Systeme d'etancheite inter-etages dans une turbomachine
US20080044284A1 (en) * 2006-08-16 2008-02-21 United Technologies Corporation Segmented fluid seal assembly
US8172514B2 (en) 2006-09-08 2012-05-08 Pratt & Whitney Canada Corp. Rim seal for a gas turbine engine
US20090208326A1 (en) * 2006-09-08 2009-08-20 Eric Durocher Rim seal for a gas turbine engine
US20080112811A1 (en) * 2006-11-13 2008-05-15 United Technologies Corporation Hoop seal with partial slot geometry
US7927069B2 (en) 2006-11-13 2011-04-19 United Technologies Corporation Hoop seal with partial slot geometry
US20080286109A1 (en) * 2007-05-15 2008-11-20 Sean Robert Keith Turbine rotor blade and method of fabricating the same
US7976281B2 (en) 2007-05-15 2011-07-12 General Electric Company Turbine rotor blade and method of assembling the same
EP1992787A1 (en) * 2007-05-15 2008-11-19 General Electric Company Turbine rotor blade assembly comprising a removable platform
CN101315032B (zh) * 2007-05-30 2013-06-19 通用电气公司 具有增强的气流封闭表面特征的定子-转子组件和方法
US20090110548A1 (en) * 2007-10-30 2009-04-30 Pratt & Whitney Canada Corp. Abradable rim seal for low pressure turbine stage
US20100166551A1 (en) * 2008-12-29 2010-07-01 Morrison Adam J Hybrid turbomachinery component for a gas turbine engine
US8435007B2 (en) * 2008-12-29 2013-05-07 Rolls-Royce Corporation Hybrid turbomachinery component for a gas turbine engine
US20100189556A1 (en) * 2009-01-28 2010-07-29 United Technologies Corporation Segmented ceramic matrix composite turbine airfoil component
US8511980B2 (en) 2009-01-28 2013-08-20 United Technologies Corporation Segmented ceramic matrix composite turbine airfoil component
US8251651B2 (en) 2009-01-28 2012-08-28 United Technologies Corporation Segmented ceramic matrix composite turbine airfoil component
US8215915B2 (en) 2009-05-15 2012-07-10 Siemens Energy, Inc. Blade closing key system for a turbine engine
US20100290914A1 (en) * 2009-05-15 2010-11-18 Souers Philip F Blade Closing Key System for a Turbine Engine
CN102733863A (zh) * 2011-03-31 2012-10-17 通用电气公司 具有气体流的增强抑制特征的定子-转子组件及相关方法
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
EP2562355A3 (en) * 2011-08-23 2018-04-11 General Electric Company Coupled blade platforms and methods of sealing
US8888459B2 (en) * 2011-08-23 2014-11-18 General Electric Company Coupled blade platforms and methods of sealing
US20130071248A1 (en) * 2011-09-19 2013-03-21 General Electric Company Compressive stress system for a gas turbine engine
US8985956B2 (en) * 2011-09-19 2015-03-24 General Electric Company Compressive stress system for a gas turbine engine
US10570767B2 (en) 2016-02-05 2020-02-25 General Electric Company Gas turbine engine with a cooling fluid path
US20180094638A1 (en) * 2016-10-05 2018-04-05 Pratt & Whitney Canada Corp. Integrally bladed fan rotor
US10371162B2 (en) * 2016-10-05 2019-08-06 Pratt & Whitney Canada Corp. Integrally bladed fan rotor
US20180112543A1 (en) * 2016-10-26 2018-04-26 Rolls-Royce Corporation Turbine wheel assembly with ceramic matrix composite components
US10428661B2 (en) * 2016-10-26 2019-10-01 Roll-Royce North American Technologies Inc. Turbine wheel assembly with ceramic matrix composite components
US10633992B2 (en) 2017-03-08 2020-04-28 Pratt & Whitney Canada Corp. Rim seal

Also Published As

Publication number Publication date
IT945233B (it) 1973-05-10
GB1318654A (en) 1973-05-31
DE2159857A1 (de) 1972-06-29
DE2166499A1 (de) 1974-07-04
FR2116522A1 (OSRAM) 1972-07-13
DE2159857B2 (de) 1977-07-14
FR2116522B1 (OSRAM) 1975-08-29

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