US3734646A - Blade fastening means - Google Patents

Blade fastening means Download PDF

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Publication number
US3734646A
US3734646A US00222847A US3734646DA US3734646A US 3734646 A US3734646 A US 3734646A US 00222847 A US00222847 A US 00222847A US 3734646D A US3734646D A US 3734646DA US 3734646 A US3734646 A US 3734646A
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United States
Prior art keywords
blade
grooves
root
radial
outer periphery
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Expired - Lifetime
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US00222847A
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English (en)
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A Perkins
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General Electric Co
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General Electric Co
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Publication date
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    • 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
    • F01D5/32—Locking, e.g. by final locking blades or keys
    • F01D5/326—Locking of axial insertion type blades by other means
    • 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

Definitions

  • Appl. No.: 222,847 Blade fastening means are provided for securing airfoil blades around the periphery of a rotating member wherein the blades include contiguous stiffening or [521 U.S. Cl ..4l6/220, 416/196 [51] Int.
  • this invention relates to a blade fastening means for use in an axial flow turbomachine and more particularly to a blade fastening means which facilitates uninhibited insertion of individual blade elements within a shrouded blade assembly.
  • turbomachinery airfoils for example, rotor blades and stator vanes used in steam or gas turbine components, including axial flow compressors and turbines, are subject to vibrations during operation. Frequently, with use of extremely long and thin blades or vanes, the vibrations can assume such proportions that dampening or stiffening of the airfoils is required. It is also known to use means projecting laterally of the airfoil portion of the blades or vanes and located intermediate or at the radial ends thereof for stiffening or dampening, these tongue-like projections conventionally interengaging like projections of adjacent airfoil members. Vibration dampening may thus be provided by frictional interengagement, as is disclosed in the U.S. Pat. No.
  • a spacer in order to assemble or remove each individual blade, a spacer must also be assembled or removed axially in order that the blade may be moved radially within the groove permitting axial translation of the blade along the bottom surface of the groove without interference between the lateral partitions.
  • blade spacer increases the weight of the turbine assembly and provides an additional part which adds to the overall complexity of the turbine assembly.
  • the blade fastening and locking means of this invention is provided for securing blades within an axial flow turbomachine assembly.
  • the turbomachine assembly is of the type including a rotating member having a plurality of axial grooves disposed about the outer periphery thereof. The grooves extend axially through the rotating member and include opposing longitudinal sidewalls which diverge radially inward.
  • a plurality of blade members having root and airfoil portions are provided for insertion within the grooves.
  • the root portions are shaped to abut the groove sidewalls for retention within the grooves, there being clearance between the root portions and grooves to allow radial movement of the blades within the grooves.
  • the blades further include stiffening means which project laterally from opposite surfaces of the airfoil portions. Opposing stiffening means contiguously form a shroud ring along lines of abutment directed at an angle to the axis of the grooves.
  • a retaining ring means is fixedly attached to one of the radial surfaces of the rotating member and overlaps the open ended grooves to provide restraint against axial translation of the blade root portions in one direction.
  • Flange means are provided which extend radially inward from the bottom surface of the blade root for engagement with the bottom wall of the groove so as to inhibit radial movement of the blade at one end of the root.
  • the blade further includes a lip longitudinally extending from the other end of the blade root for engagement with the outer periphery of the retaining ring means to inhibit radial movement of the blade root at this end. Additional locking means are provided to restrain axial translation of the blade root portions in the opposing direction.
  • the clearance between the root portions and grooves is substantially equal to the radial extent of the shroud ring.
  • the blade members may be individually translated a limited axial distance by slightly twisting the airfoil so as to allow engagement or disengagement of the radial flange and longitudinal lip with the bottom wall of the groove and the outer periphery of the retaining ringv DESCRIPTION OF THE DRAWINGS While the specification concludes with a series of claims which particularly point out and distinctly claim the invention described herein, it is believed that the invention will be more readily understood by reference to the discussion below and the accompanying drawings in which:
  • FIG. 1 is a cutaway perspective view of the blade fastening means of this invention
  • FIG. 2 is a top view of the blade fastening means of FIG. 1;
  • FIG. 3 is a partial cross-sectional view of the blade fastening means of FIG. 1;
  • FIG. 4 is a partial cutaway perspective view of the blade fastening means of FIG. 1;
  • FIG. 5 is a partial cross-sectional view of an alternate embodiment of the blade fastening means of HG. 1.
  • the blade root or base portion includes pairs of oppositely directed, longitudinally extending faces and 16, end faces 17 and 18, and a bottom portion 19.
  • the blade supporting structure comprises a rotor wheel disk, indicated generally at 22, having an enlarged rim portion 24 and a pair of oppositely directed (axially of the rotor) faces 26 and 28.
  • a rotor wheel disk Extending through the rim portion in a generally axial direction relative to the rotor are individual blade receiving grooves, indicated at 30, the best viewed in FIG. 4. These grooves comprise pairs of Iongitudinally extending angled sidewalls 32 and 33 and a wider base or bottom wall 34.
  • the general overall shape of the disk or wheel grooves generally corresponds to the shape of the blade roots with a clearance between the bottom portion 19 of the blade roots and the base 34 of the grooves as indicated at A.
  • longitudinal faces 15 and 16 of the blade roots are shaped to abut and be retained against action of centrifugal force by the angled walls 32 and 33 respectively, of the rotor wheel grooves 30.
  • the dampening or stiffening means may be integral with the blades, such as projections laterally of the blade airfoils and may be located intermediate or at the radially spaced airfoil ends.
  • the blade lashings or partitions may take the form of the contiguous, interengaging tongue-like elements, indicated generally at 40 and 42 in FIGS. 1 and 2. Each of these partition elements includes a bottom wall or surface 44, a top surface 46, and an end or lateral portion or wall 48. As shown, therefore, the partitions are generally flat, laterally extending members which, when in interengagement with adjacent members, form what may be termed a blade dampening ring or shroud.
  • the action of the motive fluid against the airfoils will cause the blades 10to untwist.
  • the ring or shroud is formed of the contiguous abutting partitions 40 and 42.
  • the line of abutment coincides with the lateral portions 48 of adjacent partition elements and is angled or slanted with respect to the rotor grooves. This improves frictional dampening by causing increased rubbing during untwisting motion of the blades about their longitudinal axis during rotor operation.
  • the plane of the abutting edges 48 of the adjacent partitions 40, 42 is in a generally radial direction with respect to the rotor axis.
  • FIG. 4 illustrates that the blade base groove 30 in the rotor disk extends between the opposed faces 26 28 of the disk in a generally axial direction, although they may be slightly angled to the axis of the turbomachine without departing from the intended scope of invention.
  • movement axially longitudinally of the grooves will not permit the bladed members to be removed individually because of interference between the adjacent laterally extending partitions 40 42 of each blade of the row.
  • a circumferential, integral flange 50 projects radially inward from the underside of the forward portion of the wheel rim 24 and the forward face 52 of the flange 50 is made flush to face 26 of the rim 24.
  • the root portion of each blade includes an integral lip 54, as best shown in FIGS. 3 and 4, which extends longitudinally outward from the outer radial portion of the end face 17 and wherein the forward edge of the lip may be radiused as shown in the drawings.
  • the rim of the rotor wheel disk may also include corresponding integral lip sections 58 which extend longitudinally outward from the outer radial portion of the rim face 26.
  • the longitudinally extending lips 54 of the blade elements register with the longitudinally extending lip sections 58 of the wheel rim to form a completed annular lip.
  • Each blade root further includes an integral flange 60 as shown in FIG. 3 which extends radially inward from the aft end of the bottom portion 19 of the blade root.
  • the length of the flange 60 is made equal to the clearance A between the bottom portion 19 of the blade root and the base 34 of the groove.
  • the width of the flange 60 is also of critical dimension as will be made obvious from the following discussion.
  • An annular retaining ring shown generally at 62 is provided for restraining forward longitudinal motion of the blade roots within their respective grooves.
  • One element of the retaining ring 62 includes a washer-type annulus 64 having opposed side faces 65 and 66 which extend in generally radial directions.
  • the retaining ring 62 is affixed to the wheel 10 so that face 66 of the washer-type annulus abuts face 26 of the wheel rim 24 and face 52 of the flange 50.
  • the retaining ring is maintained in affixed relationship to the wheel by a plurality of circumferentially spaced bolts 67 which extend through and clamp the washer-type annulus 64 to the radially inwardly directed flange 50. Attachment of the retaining ring could alternatively be achieved without the inwardly directed flange 50 by circumferentially disposing the bolt receiving holes intermediate the grooves 30.
  • Another element of the locking ring includes a flaired annulus 68 formed integral with the outer circumferential edge of the washer-type annulus 64. From the outer periphery of the flaired annulus there is provided an integral lip 70 extending in an outward radial direction. The outer periphery of the radial lip 70 is notched at 72 for engagement with the lips 54 of the blade roots and the lip sections 58 of the wheel rim.
  • the notch 72 may be radiused to conform to the shape of the engaging lips of the blade roots and as is readily apparent from FIG. 3 the notch includes a portion 74 which supports the root lips 54 in an outward radial direction and a portion 73 which restrains the root lips from forward longitudinal translation.
  • the portion 74 of the notch providing outward radial support to the root lips has an effective width designated generally at B wherein B is preferably equivalent to the width C of flange 60.
  • the steps of assembly may proceed as follows. Assuming the retaining ring 62 remains permanently affixed to the wheel rim after initial assembly, individual blades must be inserted into respective grooves from the aft side of the wheel. The individual blade members are slid into the respective grooves 30 with the bottom portion, or surfaces 19, of the blade roots resting on the bases or bottoms 34 of the grooves. Each blade is inserted all the way axially into the groove until the radial side of the flange 60 engages the face 28 of the wheel rim 24.
  • the clearance A between the bottom portion 19 of the blade roots and the base 34 of the grooves is made substantially equal to the thickness of the partitions 40, 42, thus insuring that the contiguous partitions of adjacent blades do not interfere with the inserted blade when it is slid along the bottom portion of the wheel groove.
  • the blade With the blade thus inserted, it must be raised and moved longitudinally forward only a limited distance equal to the width C of flange 60 or the width B of notch 72, whichever is greater, in order that the longitudinal faces and 16 of the blade root abut the oppositely directed sidewalls 32 and 33 of groove 30.
  • the limited longitudinal translation of the blade necessary in order to engage the blade lip 54 within the notch 72 and to seat the inside edge of the flange 60 on the bottom portion of the wheel groove is made possible by the blade flexibility which permits a limited amount of airfoil twisting. Since the thickness of the lateral partition is substantially equal to the clearance A, the laterally extending dampening or stiffening partition means 40 and 42 of the blades now form the contiguous ring or shroud. Final restraint of the blades against translation in an aft direction is accomplished by insertion of the snap ring 74 within the overlapping hooks 75.
  • the snap ring 74 is first removed, whereupon a blade is then translated a limited distance in the aft direction by twisting the airfoil. The blade can then be dropped to the bottom surface of the wheel groove from where it can be entirely withdrawn from the groove without interference from the contiguous partitions of adjacent blades. lt should be understood that the designations of forward and aft are only presented by way of illustration and may be entirely reversed.
  • FIG. 5 where like numerals refer to previously described elements, there is shown a simplified alternate embodiment for the blade locking device of FIGS. 1 4.
  • a simplified retaining ring 62' having opposed parallel sides 65', 66' has been substituted for the more complex retaining ring previously described.
  • Retaining ring 62' is affixed to the wheel rim in identical manner utilizing circumferentially spaced bolts 67.
  • the circumferential notch 72 has been eliminated and forward axial translation of the blades 10 is prevented by engagement of face 66 with faces 17 of the blade roots.
  • the longitudinal faces 15 and 16 of the blade roots are maintained in abutting relation with the oppositely directed sidewalls 32 and 33 of groove 30 by engagement of the outer periphery of the retaining ring 62' with the interior surfaces of the blade root lips 54 and by engagement of the bottom portions 34 of the wheel grooves 30 with the inside edges of the root flanges 60.
  • insertion and withdrawal of the blades can be accomplished in the same manner as previously described.
  • Blade fastening and locking means for a turbomachine comprising:
  • a rotating member having a plurality of axial grooves disposed about the outer periphery thereof wherein the grooves include opposing sidewalls diverging radially inward;
  • a plurality of blade members having root and airfoil portions, the root portions being received in the grooves and shaped to abut the groove sidewalls for retention in the grooves, there being clearance between the root portions and grooves to allow radial movement therein;
  • a blade member stiffening means projecting laterally from the opposite surfaces of the airfoil portions, with opposing stiffening means contiguously forming a shroud ring along lines of abutment directed at an angle to the axis of the grooves;
  • a retaining ring means with an outer peripheral edge wherein the ring means is stationed in fixed relation relative to a radial surface of the rotating member so as to overlap one end of the grooves to provide restraint against axial translation of the blade root portions in one direction;
  • flange means extending radially inward from the bottom surface of the blade root for engagement with rotating member with a plurality of axial grooves disposed about the outer periphery thereof wherein the grooves include opposing sidewalls diverging radially inward; a plurality of blade members having root and airfoil portions, the root portions being received in the grooves and shaped to abut the groove sidewalls for retention in the grooves with clearance between the root portions and grooves to allow radial movement therein; and blade stiffening means projecting laterally from opthe bottom wall of the groove so as to inhibit radial movement of one end of the blade root;
  • any one of the blade members may be translated a limited axial distance by slightly twisting the airfoil so as to allow engagement or disengagement of the radial flange and longitudinal lip with the bottom wall of the groove and the outer periphery of the retaining ring respectively, further permitting the blade members to be moved radially and removed from or inserted into their associated grooves individually and without interference from adjacent member stiffening means.
  • the rotating member includes a rotor wheel disk having an enlarged rim portion around the outer periphery thereof, through which the blade receiving grooves extend, and the outer periphery of the retaining ring is circumferentially notched for axial and radial engagement with the edge of the longitudinally extending lip means of each blade.
  • one side of the wheel rim includes an inwardly directed radial flange to which the retaining ring means is fixedly connected by means of a plurality of circumferentially spaced bolts extending therethrough and the locking means includes a snap ring for insertion within a plurality of circumferentially spaced hooks which extend from the radial surface of the wheel rim.
  • blade fastening and locking means comprise:
  • a retaining ring means with an outside peripheral edge wherein the ring is stationed in fixed relation relative to a radial surface of the rotating member so as to overlap one end of the grooves to provide restraint against axial translation of the blade root portions in one direction;
  • flange means extending radially inward from the bottom surface of each blade root for engagement with the bottom wall of the groove so as to inhibit radial movement of one end of the blade root
  • lip means longitudinally extending from the other end of the blade root for engagement with the outer periphery of the retaining ring to inhibit radial movement of the blade root at this end wherein any one of the blade members may be translated a limited axial distance by slightly twisting the airfoil so as to allow engagement or disengagement of the radial flange and longitudinal lip with the bottom wall of the groove and the outer periphery of the retaining ring means respectively further permitting the blade members to be moved radially and removed from or inserted into their associated grooves individually and without interference from adjacent member stiffening means.
  • the rotating member includes a rotor wheel disk having an enlarged rim portion around the outer periphery thereof, through which the blade receiving grooves extend, and the outer periphery of the retaining ring is circumferentially notched for axial and radial engagement with the edge of the longitudinally extending lip means of each blade.
  • one side of the wheel rim includes an inwardly directed radial flange to which the retaining ring means is fixedly connected by means of a plurality of circumferentially spaced bolts extending therethrough, and
  • the other side of the wheel rim includes a locking means to provide restraint against axial translation of the blade root portions in the opposing directlon.

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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)
US00222847A 1972-02-02 1972-02-02 Blade fastening means Expired - Lifetime US3734646A (en)

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Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3957393A (en) * 1974-10-29 1976-05-18 United Technologies Corporation Turbine disk and sideplate construction
FR2440466A1 (fr) * 1978-11-02 1980-05-30 Gen Electric Aube de turbomachine perfectionnee
FR2472077A1 (fr) * 1979-12-13 1981-06-26 United Technologies Corp Bandage pour aubes de rotor
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
DE3221925A1 (de) * 1981-06-18 1983-01-20 General Electric Co., Schenectady, N.Y. Laufschaufel und verfahren zu deren herstellung und verstaerkung
US4480958A (en) * 1983-02-09 1984-11-06 The United States Of America As Represented By The Secretary Of The Air Force High pressure turbine rotor two-piece blade retainer
US4483661A (en) * 1983-05-02 1984-11-20 General Electric Company Blade assembly for a turbomachine
US4502841A (en) * 1982-11-08 1985-03-05 S.N.E.C.M.A. Fan blade axial locking device
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
US5030063A (en) * 1990-02-08 1991-07-09 General Motors Corporation Turbomachine rotor
US5067877A (en) * 1990-09-11 1991-11-26 United Technologies Corporation Fan blade axial retention device
EP0470763A1 (en) * 1990-08-06 1992-02-12 General Electric Company Protective coating for rotor blades
US5205714A (en) * 1990-07-30 1993-04-27 General Electric Company Aircraft fan blade damping apparatus
US5302086A (en) * 1992-08-18 1994-04-12 General Electric Company Apparatus for retaining rotor blades
JPH0886202A (ja) * 1994-09-14 1996-04-02 Kawasaki Heavy Ind Ltd セラミックブレードの取付構造
US20040131471A1 (en) * 2002-12-20 2004-07-08 Martyn Richards Blade arrangement for gas turbine engine
US6951448B2 (en) 2002-04-16 2005-10-04 United Technologies Corporation Axial retention system and components thereof for a bladed rotor
US20100166563A1 (en) * 2007-08-08 2010-07-01 Alstom Technology Ltd Method for improving the sealing on rotor arrangements
US20100322772A1 (en) * 2009-06-23 2010-12-23 Rolls-Royce Plc Annulus filler for a gas turbine engine
US20110038731A1 (en) * 2009-08-12 2011-02-17 Rolls-Royce Plc Rotor assembly for a gas turbine
US20110236185A1 (en) * 2010-03-23 2011-09-29 Rolls-Royce Plc Interstage seal
US20130156590A1 (en) * 2010-06-25 2013-06-20 Snecma Gas turbine engine rotor wheel having composite material blades with blade-root to disk connection being obtained by clamping
CN101382149B (zh) * 2007-08-16 2015-07-01 通用电气公司 一种制造叶片的方法
US20160237840A1 (en) * 2013-09-25 2016-08-18 Snecma Rotary assembly for a turbomachine
US20190277192A1 (en) * 2018-03-09 2019-09-12 General Electric Company Compressor rotor cooling apparatus
US10927683B2 (en) * 2017-12-14 2021-02-23 Safran Aircraft Engines Damping device

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1479332A (en) * 1974-11-06 1977-07-13 Rolls Royce Means for retaining blades to a disc or like structure
US8206119B2 (en) * 2009-02-05 2012-06-26 General Electric Company Turbine coverplate systems
FR3109604B1 (fr) * 2020-04-27 2022-08-26 Safran Aircraft Engines Roue aubagee a performances d’etancheite et de retention des aubes ameliorees

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US3008689A (en) * 1954-08-12 1961-11-14 Rolls Royce Axial-flow compressors and turbines
US3047268A (en) * 1960-03-14 1962-07-31 Stanley L Leavitt Blade retention device
US3216699A (en) * 1963-10-24 1965-11-09 Gen Electric Airfoil member assembly
US3395891A (en) * 1967-09-21 1968-08-06 Gen Electric Lock for turbomachinery blades
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 (6)

* 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
US3047268A (en) * 1960-03-14 1962-07-31 Stanley L Leavitt Blade retention device
US3216699A (en) * 1963-10-24 1965-11-09 Gen Electric Airfoil member assembly
US3395891A (en) * 1967-09-21 1968-08-06 Gen Electric Lock for turbomachinery blades
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 (38)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3957393A (en) * 1974-10-29 1976-05-18 United Technologies Corporation Turbine disk and sideplate construction
FR2440466A1 (fr) * 1978-11-02 1980-05-30 Gen Electric Aube de turbomachine perfectionnee
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
FR2472077A1 (fr) * 1979-12-13 1981-06-26 United Technologies Corp Bandage pour aubes de rotor
DE3221925A1 (de) * 1981-06-18 1983-01-20 General Electric Co., Schenectady, N.Y. Laufschaufel und verfahren zu deren herstellung und verstaerkung
US4453890A (en) * 1981-06-18 1984-06-12 General Electric Company Blading system for a gas turbine engine
DE3221925C2 (xx) * 1981-06-18 1993-04-01 General Electric Co., Schenectady, N.Y., Us
US4502841A (en) * 1982-11-08 1985-03-05 S.N.E.C.M.A. Fan blade axial locking device
US4480958A (en) * 1983-02-09 1984-11-06 The United States Of America As Represented By The Secretary Of The Air Force High pressure turbine rotor two-piece blade retainer
US4483661A (en) * 1983-05-02 1984-11-20 General Electric Company Blade assembly for a turbomachine
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
US5030063A (en) * 1990-02-08 1991-07-09 General Motors Corporation Turbomachine rotor
US5205714A (en) * 1990-07-30 1993-04-27 General Electric Company Aircraft fan blade damping apparatus
EP0470763A1 (en) * 1990-08-06 1992-02-12 General Electric Company Protective coating for rotor blades
US5137426A (en) * 1990-08-06 1992-08-11 General Electric Company Blade shroud deformable protective coating
US5067877A (en) * 1990-09-11 1991-11-26 United Technologies Corporation Fan blade axial retention device
US5302086A (en) * 1992-08-18 1994-04-12 General Electric Company Apparatus for retaining rotor blades
JP2726895B2 (ja) 1994-09-14 1998-03-11 川崎重工業株式会社 セラミックブレードの取付構造
JPH0886202A (ja) * 1994-09-14 1996-04-02 Kawasaki Heavy Ind Ltd セラミックブレードの取付構造
US6951448B2 (en) 2002-04-16 2005-10-04 United Technologies Corporation Axial retention system and components thereof for a bladed rotor
US20040131471A1 (en) * 2002-12-20 2004-07-08 Martyn Richards Blade arrangement for gas turbine engine
US6971855B2 (en) * 2002-12-20 2005-12-06 Rolls-Royce Plc Blade arrangement for gas turbine engine
US20100166563A1 (en) * 2007-08-08 2010-07-01 Alstom Technology Ltd Method for improving the sealing on rotor arrangements
US9435213B2 (en) * 2007-08-08 2016-09-06 General Electric Technology Gmbh Method for improving the sealing on rotor arrangements
CN101382149B (zh) * 2007-08-16 2015-07-01 通用电气公司 一种制造叶片的方法
US20100322772A1 (en) * 2009-06-23 2010-12-23 Rolls-Royce Plc Annulus filler for a gas turbine engine
US8596981B2 (en) 2009-06-23 2013-12-03 Rolls-Royce Plc Annulus filler for a gas turbine engine
US8636474B2 (en) * 2009-08-12 2014-01-28 Rolls-Royce Plc Rotor assembly for a gas turbine
US20110038731A1 (en) * 2009-08-12 2011-02-17 Rolls-Royce Plc Rotor assembly for a gas turbine
US8864451B2 (en) 2010-03-23 2014-10-21 Rolls-Royce Plc Interstage seal
US20110236185A1 (en) * 2010-03-23 2011-09-29 Rolls-Royce Plc Interstage seal
US20130156590A1 (en) * 2010-06-25 2013-06-20 Snecma Gas turbine engine rotor wheel having composite material blades with blade-root to disk connection being obtained by clamping
US9422818B2 (en) * 2010-06-25 2016-08-23 Snecma Gas turbine engine rotor wheel having composite material blades with blade-root to disk connection being obtained by clamping
US20160237840A1 (en) * 2013-09-25 2016-08-18 Snecma Rotary assembly for a turbomachine
US10662795B2 (en) * 2013-09-25 2020-05-26 Snecma Rotary assembly for a turbomachine
US10927683B2 (en) * 2017-12-14 2021-02-23 Safran Aircraft Engines Damping device
US20190277192A1 (en) * 2018-03-09 2019-09-12 General Electric Company Compressor rotor cooling apparatus
US10746098B2 (en) * 2018-03-09 2020-08-18 General Electric Company Compressor rotor cooling apparatus

Also Published As

Publication number Publication date
BE794573A (fr) 1973-05-16
FR2170692A5 (xx) 1973-09-14
DE2304291A1 (de) 1973-08-09

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