US3661475A - Turbomachinery rotors - Google Patents

Turbomachinery rotors Download PDF

Info

Publication number
US3661475A
US3661475A US33220A US3661475DA US3661475A US 3661475 A US3661475 A US 3661475A US 33220 A US33220 A US 33220A US 3661475D A US3661475D A US 3661475DA US 3661475 A US3661475 A US 3661475A
Authority
US
United States
Prior art keywords
rim
rotor
web
platforms
slots
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.)
Expired - Lifetime
Application number
US33220A
Inventor
Bernard J Anderson
William E Schoenborn
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.)
General Electric Co
Original Assignee
General Electric Co
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 Co filed Critical General Electric Co
Application granted granted Critical
Publication of US3661475A publication Critical patent/US3661475A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • 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 improvements in gas turbine rotors and more particularly to improvements in rotors .which carry blades of an axial flow compressor or turbine incorporated on such engines.
  • Bladed rotors of axial flow turbomachinery provide a mounting means for the radially projecting blades.
  • blade attachment is accomplished by tangs on the blades which are received in dovetail slots on the rotor.
  • the composite rotor must have, between the blades, a smooth surface of revolution defining the inner bounds of the gas stream.
  • oppositely projecting wings referenced as platforms, are fonned on the blades to provide the desired flow surface. It has been recognized that platforms formed on blades increase the dead weight loading of the tangs when the rotor is operating at high speeds. Platforms have also been formed integrally with rotors and as separate plates.
  • Platform weight increases the centrifugal forces on the rotor and results in increased rotor mass and weight to carry these forces without overstressing the rotor material.
  • weight is, of course, a critical factor.
  • Platform weight becomes most significant in rotors and comprising discs having dovetails spaced around their peripheries and a large change in platform diameter from the inlet side to the discharge side of the platforms.
  • the object of the invention is to provide an improved turbomachinery rotor of the type last referenced wherein platforms are provided with a minimum of weight and assured structural integrity.
  • a turbomachinery rotor comprising a disc portion having integral periphery rim.
  • Dovetail slots extend across the rim for the mounting of tanged blades thereon.
  • the periphery of the structural portion of the rim terminates adjacent to and slightly outwardly of the necks of the dovetail slots.
  • Relatively thin platforms are provided between each adjacent pair of slots. The platforms are angled outwardly from one side of the rim to the other and the inner portion of the platforms are integrally formed with one side of the rim.
  • Relatively thin web means extend outwardly from the rim and integrally connect the remaining portions of each platform therewith.
  • the web means comprise a radially projecting web extending outwardly from the other side of the rim and an axial web extending from the radial web and from the rim to the inner side of the platform.
  • the web means can take the form of a pair of angularly spaced, axially extending webs projecting outwardly from the neck portions of the slots to said platforms throughout the width of the rim.
  • FIG. 1 is a schematic showing of a gas turbine engine
  • FIG. 2 is a fragmentary longitudinal section of a rotor embodying the present invention
  • FIG. 3 is a perspective view, with portions broken away and in section, of a rotor disc seen in FIG. 2, looking at the downstream side thereof;
  • FIG. 4 is a perspective view, similar to FIG. 3, of an alternate embodiment of the invention.
  • the illustrated gas turbine engine comprises an axial flow compressor 10 which pressurizes air discharged onto a combustor 12. This air supports combustion of fuel in the generation of an annular hot gas stream.
  • the hot gas stream drives a turbine 14 and then may be discharged from a nozzle 16 to generate a propulsive force.
  • the turbine 14 comprises a rotor 18 which is connected by a shaft 20 to the compressor rotor 22 to power the latter.
  • a portion of the rotor 22 is shown in greater detail in FIG. 2.
  • a stub shaft 24 provides a mounting means for the forward end of the rotor.
  • the stub shaft is integral with a cone portion 26 and connected by bolts 29 to two spacers 28 and 30.
  • the spacer 28 is integral with a first stage rotor disc 32 while conical spacer 30 is integral with a second stage disc 34.
  • the remainder of the rotor may be similarly fabricated with blades b projecting radially therefrom in circumferential rows or stages.
  • the second stage disc 34 is shown in greater detail in FIG. 3.
  • the disc has a rim 36 across which are machined dovetail slots 38.
  • the slots 38 receive correspondingly tapered tangs on the blades b.
  • the high centrifugal, radial force loadings of the blades transition from the slots 38 into the rim 36 and then into the inner portions of the disc 34.
  • the rim 36 terminates above and closely adjacent to the necks n of the slots 38.
  • Between the slots 38 are relatively thin platform portions 40 which angle outwardly from the inlet side of the rim or disc to the discharge side thereof.
  • the inlet sides of the platforms 48 are integrally formed and connected to the rim 36.
  • the discharge sides of the platforms 40 are connected to the rim 36 by relatively thin radial flanges 42.
  • the platforms 40 are also connected to the rim 36 by integrally fonned webs 44 which extend in an axial direction centrally of the platforms.
  • the rim may be lightened by axially extending cavities 46 intermediate each pair of slots 38.
  • the platform portions 40 may be connected to the rim 36 again having their inlet sides integrally formed with the rim 36 and with a pair of thin, axial webs 48 extending outwardly from the rim at the neck portions n of the slots 38.
  • the desired fiowpath function is provided across the bladed portion of the rotor with a minimum of centrifugal force loading on the disc and thus with a minimum of weight.
  • a bladed turbomachinery rotor comprising:
  • dovetail slots extending across said rim with their necks spaced below and closely adjacent to the outer periphery of the rim
  • platforms between each pair of slots said platforms being relatively thin and angled outwardly from one side of the rim to the other, the inner portion of said platforms being integrally formed with one side of said rim, and

Abstract

A compressor rotor is disclosed having relatively thin angled platforms connected to the structural portion of a rotor disc by web means.

Description

United States Patent Anderson et a1.
451 May 9,1972
[54] TURBOMACHINERY ROTORS [72] Inventors: Bernard J. Anderson, Danvers, Mass;
William E. Schoenborn, Cincinnati, Ohio [73] Assignee: General Electric Company [22] Filed: Apr. 30, 1970 [21] Appl. No.: 33,220
[52] U.S.C1 ..416/2l9,416/244 [51] Int. Cl ..FOld 5/32, FOld 5/02 [58] Field of Search ..416/219, 244, 220,221
[56] References Cited UNITED STATES PATENTS Anderson ..416/220 3,393,862 7/1968 Harrison ..416/220 2,452,782 11/1948 McLeod et a1... ..416/220 3,291,446 12/1966 Huebner ..416/244 FOREIGN PATENTS OR APPLICATIONS 590,294 7/1955 Canada ..416/220 Primary Examiner-Everette A. Powell, Jr. Assistant Examiner-Clemens Schimikowski Almrney-Derek P, Lawrence, Frank L. Neuhauser, Oscar B. Waddell, Joseph Bv Forman and Edward S. Roman [57] ABSTRACT A compressor rotor is disclosed having relatively thin angled platforms connected to the structural portion of a rotor disc by web means.
4 Claims, 4 Drawing Figures PATENTEDMAY 9 m2 SHEET 1 BF 2 INVENTORS. BERNARD J. ANDERSON Q MLUAM E. SCHOENBORN PATENTEDMAY 9 I972 SHEET 2 BF 2 INVENT R5. BERNARD J. ANDEQSON WILLflAM E. SCHOENBORN raezvzy- TURBOMACHINERY ROTORS The invention described and claimed in the United States patent application herein resulted from work done under United States Government contract FA-SS-66-6. The United States Government has an irrevocable, non-exclusive license under said application to practice and have practiced the invention claimed herein, including the unlimited right to sublicense others to practice and have practiced the claimed invention for any purpose whatsoever.
The present invention relates to improvements in gas turbine rotors and more particularly to improvements in rotors .which carry blades of an axial flow compressor or turbine incorporated on such engines.
Bladed rotors of axial flow turbomachinery provide a mounting means for the radially projecting blades. Generally, blade attachment is accomplished by tangs on the blades which are received in dovetail slots on the rotor. When the blades are assembled the composite rotor must have, between the blades, a smooth surface of revolution defining the inner bounds of the gas stream. Frequently oppositely projecting wings, referenced as platforms, are fonned on the blades to provide the desired flow surface. It has been recognized that platforms formed on blades increase the dead weight loading of the tangs when the rotor is operating at high speeds. Platforms have also been formed integrally with rotors and as separate plates.
Platform weight, whether on the blades or on the rotor itself, increases the centrifugal forces on the rotor and results in increased rotor mass and weight to carry these forces without overstressing the rotor material. In aircraft propulsion engines weight is, of course, a critical factor.
Platform weight becomes most significant in rotors and comprising discs having dovetails spaced around their peripheries and a large change in platform diameter from the inlet side to the discharge side of the platforms.
The object of the invention is to provide an improved turbomachinery rotor of the type last referenced wherein platforms are provided with a minimum of weight and assured structural integrity.
These ends are attained by a turbomachinery rotor comprising a disc portion having integral periphery rim. Dovetail slots extend across the rim for the mounting of tanged blades thereon. The periphery of the structural portion of the rim terminates adjacent to and slightly outwardly of the necks of the dovetail slots. Relatively thin platforms are provided between each adjacent pair of slots. The platforms are angled outwardly from one side of the rim to the other and the inner portion of the platforms are integrally formed with one side of the rim. Relatively thin web means extend outwardly from the rim and integrally connect the remaining portions of each platform therewith.
Preferably the web means comprise a radially projecting web extending outwardly from the other side of the rim and an axial web extending from the radial web and from the rim to the inner side of the platform. Alternatively the web means can take the form of a pair of angularly spaced, axially extending webs projecting outwardly from the neck portions of the slots to said platforms throughout the width of the rim.
The above and other related objects and features of the invention will be apparent from a reading of the following description of the disclosure found on the accompanying drawings and the novelty thereof pointed out in the appended claims.
IN'THE DRAWINGS FIG. 1 is a schematic showing of a gas turbine engine;
FIG. 2 is a fragmentary longitudinal section of a rotor embodying the present invention;
FIG. 3 is a perspective view, with portions broken away and in section, of a rotor disc seen in FIG. 2, looking at the downstream side thereof; and
FIG. 4 is a perspective view, similar to FIG. 3, of an alternate embodiment of the invention.
Referencing FIG. 1, the illustrated gas turbine engine comprises an axial flow compressor 10 which pressurizes air discharged onto a combustor 12. This air supports combustion of fuel in the generation of an annular hot gas stream. The hot gas stream drives a turbine 14 and then may be discharged from a nozzle 16 to generate a propulsive force. The turbine 14 comprises a rotor 18 which is connected by a shaft 20 to the compressor rotor 22 to power the latter.
A portion of the rotor 22 is shown in greater detail in FIG. 2. A stub shaft 24 provides a mounting means for the forward end of the rotor. The stub shaft is integral with a cone portion 26 and connected by bolts 29 to two spacers 28 and 30. The spacer 28 is integral with a first stage rotor disc 32 while conical spacer 30 is integral with a second stage disc 34. The remainder of the rotor may be similarly fabricated with blades b projecting radially therefrom in circumferential rows or stages.
The second stage disc 34 is shown in greater detail in FIG. 3. The disc has a rim 36 across which are machined dovetail slots 38. The slots 38 receive correspondingly tapered tangs on the blades b. The high centrifugal, radial force loadings of the blades transition from the slots 38 into the rim 36 and then into the inner portions of the disc 34. As a structural member carrying these radial loads, the rim 36 terminates above and closely adjacent to the necks n of the slots 38. Between the slots 38 are relatively thin platform portions 40 which angle outwardly from the inlet side of the rim or disc to the discharge side thereof. The inlet sides of the platforms 48 are integrally formed and connected to the rim 36. The discharge sides of the platforms 40 are connected to the rim 36 by relatively thin radial flanges 42. The platforms 40 are also connected to the rim 36 by integrally fonned webs 44 which extend in an axial direction centrally of the platforms.
Further, the rim may be lightened by axially extending cavities 46 intermediate each pair of slots 38.
Alternately as shown in FIG. 4 the platform portions 40 may be connected to the rim 36 again having their inlet sides integrally formed with the rim 36 and with a pair of thin, axial webs 48 extending outwardly from the rim at the neck portions n of the slots 38.
By minimizing the structural portions of the rim as described and by integrally connecting the platforms 40 with integral ribs, the desired fiowpath function is provided across the bladed portion of the rotor with a minimum of centrifugal force loading on the disc and thus with a minimum of weight.
Having thus described the invention what is claimed as novel and desired to be secured by Letters Patent of the United States is:
1. A bladed turbomachinery rotor comprising:
a disc portion,
an integral peripheral rim formed on said disc,
dovetail slots extending across said rim with their necks spaced below and closely adjacent to the outer periphery of the rim,
platforms between each pair of slots, said platforms being relatively thin and angled outwardly from one side of the rim to the other, the inner portion of said platforms being integrally formed with one side of said rim, and
relatively thin web means extending outwardly from said rim and integrally connecting the remaining portion of each platform therewith.
' 2. A rotor as in claim I wherein the web means take the form of a radially projecting web extending from between the neck portion of each slot outwardly from the other side of the rim to the platform and an integral axially extending web extending from said radial web and from the other side of the rim to the inner side of said platform, said axial web being disposed generally centrally of said platform.
3. A rotor as in claim 1 wherein the web means take the form of a pair of angularly spaced, axially extending webs projecting outwardly from the neck portion of said slots to said platforms throughout the width of said rim.
4. A rotor as in claim 1 wherein axially extending cavities are provided between each adjacent pair of slots.

Claims (4)

1. A bladed turbomachinery rotor comprising: a disc portion, an integral peripheral rim formed on said disc, dovetail slots extending across said rim with their necks spaced below and closely adjacent to the outer periphery of the rim, platforms between each pair of slots, said platforms being relatively thin and angled outwardly from one side of the rim to the other, the inner portion of said platforms bEing integrally formed with one side of said rim, and relatively thin web means extending outwardly from said rim and integrally connecting the remaining portion of each platform therewith.
2. A rotor as in claim 1 wherein the web means take the form of a radially projecting web extending from between the neck portion of each slot outwardly from the other side of the rim to the platform and an integral axially extending web extending from said radial web and from the other side of the rim to the inner side of said platform, said axial web being disposed generally centrally of said platform.
3. A rotor as in claim 1 wherein the web means take the form of a pair of angularly spaced, axially extending webs projecting outwardly from the neck portion of said slots to said platforms throughout the width of said rim.
4. A rotor as in claim 1 wherein axially extending cavities are provided between each adjacent pair of slots.
US33220A 1970-04-30 1970-04-30 Turbomachinery rotors Expired - Lifetime US3661475A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US3322070A 1970-04-30 1970-04-30

Publications (1)

Publication Number Publication Date
US3661475A true US3661475A (en) 1972-05-09

Family

ID=21869177

Family Applications (1)

Application Number Title Priority Date Filing Date
US33220A Expired - Lifetime US3661475A (en) 1970-04-30 1970-04-30 Turbomachinery rotors

Country Status (6)

Country Link
US (1) US3661475A (en)
BE (1) BE766401A (en)
CA (1) CA943467A (en)
DE (1) DE2120470A1 (en)
FR (1) FR2092118B1 (en)
GB (1) GB1340937A (en)

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3873234A (en) * 1971-11-10 1975-03-25 Robert Noel Penny Turbine rotor
FR2456835A1 (en) * 1979-05-17 1980-12-12 United Technologies Corp ROTOR ASSEMBLY HAVING MULTI-STAGE DISC
US4536129A (en) * 1984-06-15 1985-08-20 United Technologies Corporation Turbine blade with disk rim shield
FR2568308A1 (en) * 1984-07-30 1986-01-31 Gen Electric AUBE DE ROTOR
US5193982A (en) * 1991-07-17 1993-03-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Separate inter-blade platform for a bladed rotor disk
US5271718A (en) * 1992-08-11 1993-12-21 General Electric Company Lightweight platform blade
FR2697051A1 (en) * 1992-10-21 1994-04-22 Snecma Turbomachine rotor comprising a disc whose periphery is occupied by oblique cavities which alternate with teeth of variable cross section.
FR2708036A1 (en) * 1993-07-21 1995-01-27 Gen Electric Gas turbine engine rotor disk and fin with which this disk is provided.
US6471474B1 (en) 2000-10-20 2002-10-29 General Electric Company Method and apparatus for reducing rotor assembly circumferential rim stress
US6511294B1 (en) 1999-09-23 2003-01-28 General Electric Company Reduced-stress compressor blisk flowpath
US6524070B1 (en) 2000-08-21 2003-02-25 General Electric Company Method and apparatus for reducing rotor assembly circumferential rim stress
US20040007830A1 (en) * 2001-10-10 2004-01-15 Kazuo Uematsu Sealing structure of spindle bolt, and gas turbine
US20040062651A1 (en) * 2001-11-14 2004-04-01 Suciu Gabriel L. Blade for turbine engine
US20050246889A1 (en) * 2004-04-09 2005-11-10 Snecma Moteurs Device for assembling annular flanges together, in particular in a turbomachine
US20060127214A1 (en) * 2004-12-10 2006-06-15 David Glasspoole Gas turbine gas path contour
US20080206485A1 (en) * 2005-06-28 2008-08-28 General Electric Company Devices for evaluating material properties, and related processes
US20140174098A1 (en) * 2012-12-20 2014-06-26 United Technologies Corporation Turbine disc with reduced neck stress concentration

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB9405473D0 (en) * 1994-03-19 1994-05-04 Rolls Royce Plc A gas turbine engine fan blade assembly

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2452782A (en) * 1945-01-16 1948-11-02 Power Jets Res & Dev Ltd Construction of rotors for compressors and like machines
CA590294A (en) * 1960-01-05 Rolls-Royce Limited Axial-flow compressors and turbines
US3291446A (en) * 1965-04-13 1966-12-13 Chrysler Corp Turbine wheel
US3383095A (en) * 1967-09-12 1968-05-14 Gen Electric Lock for turbomachinery blades
US3393862A (en) * 1965-11-23 1968-07-23 Rolls Royce Bladed rotors

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3556675A (en) * 1969-01-29 1971-01-19 Gen Electric Turbomachinery rotor with integral shroud

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA590294A (en) * 1960-01-05 Rolls-Royce Limited Axial-flow compressors and turbines
US2452782A (en) * 1945-01-16 1948-11-02 Power Jets Res & Dev Ltd Construction of rotors for compressors and like machines
US3291446A (en) * 1965-04-13 1966-12-13 Chrysler Corp Turbine wheel
US3393862A (en) * 1965-11-23 1968-07-23 Rolls Royce Bladed rotors
US3383095A (en) * 1967-09-12 1968-05-14 Gen Electric Lock for turbomachinery blades

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3873234A (en) * 1971-11-10 1975-03-25 Robert Noel Penny Turbine rotor
FR2456835A1 (en) * 1979-05-17 1980-12-12 United Technologies Corp ROTOR ASSEMBLY HAVING MULTI-STAGE DISC
US4536129A (en) * 1984-06-15 1985-08-20 United Technologies Corporation Turbine blade with disk rim shield
FR2568308A1 (en) * 1984-07-30 1986-01-31 Gen Electric AUBE DE ROTOR
US5193982A (en) * 1991-07-17 1993-03-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Separate inter-blade platform for a bladed rotor disk
US5271718A (en) * 1992-08-11 1993-12-21 General Electric Company Lightweight platform blade
FR2697051A1 (en) * 1992-10-21 1994-04-22 Snecma Turbomachine rotor comprising a disc whose periphery is occupied by oblique cavities which alternate with teeth of variable cross section.
US5395213A (en) * 1992-10-21 1995-03-07 Societe Nationale D'etude Et De Construction De Motors D'aviation "Snecma" Turbojet engine rotor
FR2708036A1 (en) * 1993-07-21 1995-01-27 Gen Electric Gas turbine engine rotor disk and fin with which this disk is provided.
US6511294B1 (en) 1999-09-23 2003-01-28 General Electric Company Reduced-stress compressor blisk flowpath
US6524070B1 (en) 2000-08-21 2003-02-25 General Electric Company Method and apparatus for reducing rotor assembly circumferential rim stress
US6471474B1 (en) 2000-10-20 2002-10-29 General Electric Company Method and apparatus for reducing rotor assembly circumferential rim stress
US20070234703A1 (en) * 2001-10-10 2007-10-11 Mitsubishi Heavy Industries Ltd. Sealing structure of spindle bolt, and gas turbine
US6991429B2 (en) * 2001-10-10 2006-01-31 Mitsubishi Heavy Industries, Ltd. Sealing structure of spindle bolt, and gas turbine
US20040007830A1 (en) * 2001-10-10 2004-01-15 Kazuo Uematsu Sealing structure of spindle bolt, and gas turbine
US20040062651A1 (en) * 2001-11-14 2004-04-01 Suciu Gabriel L. Blade for turbine engine
US6764282B2 (en) * 2001-11-14 2004-07-20 United Technologies Corporation Blade for turbine engine
US20050246889A1 (en) * 2004-04-09 2005-11-10 Snecma Moteurs Device for assembling annular flanges together, in particular in a turbomachine
US7390170B2 (en) * 2004-04-09 2008-06-24 Snecma Device for assembling annular flanges together, in particular in a turbomachine
US20060127214A1 (en) * 2004-12-10 2006-06-15 David Glasspoole Gas turbine gas path contour
US7179049B2 (en) 2004-12-10 2007-02-20 Pratt & Whitney Canada Corp. Gas turbine gas path contour
US20080206485A1 (en) * 2005-06-28 2008-08-28 General Electric Company Devices for evaluating material properties, and related processes
US20140174098A1 (en) * 2012-12-20 2014-06-26 United Technologies Corporation Turbine disc with reduced neck stress concentration

Also Published As

Publication number Publication date
CA943467A (en) 1974-03-12
BE766401A (en) 1971-09-16
DE2120470A1 (en) 1971-11-11
FR2092118B1 (en) 1974-10-11
FR2092118A1 (en) 1972-01-21
GB1340937A (en) 1973-12-19

Similar Documents

Publication Publication Date Title
US3661475A (en) Turbomachinery rotors
US3037742A (en) Compressor turbine
US4863352A (en) Blade carrying means
US3294366A (en) Blades for gas turbine engines
US5466123A (en) Gas turbine engine turbine
EP0643199B1 (en) Rotor blade
EP0900920B1 (en) One-piece blisk of a gas turbine engine
CN109538352B (en) Outer drum rotor assembly and gas turbine engine
US4907944A (en) Turbomachinery blade mounting arrangement
GB2207191A (en) Gas turbine engine
CA1312828C (en) Damped gas turbine engine airfoil row
US6079200A (en) Ducted fan gas turbine engine with fan shaft frangible connection
JPH04228836A (en) Inter-stage seal structure used for air wheel stage of turbine engine double reversing rotor
GB2159887A (en) A stator vane mounting arrangement
JPH079194B2 (en) Gas turbine engine cooling air transfer means
WO2014105102A1 (en) Platform with curved edges adjacent suction side of airfoil
GB2221259A (en) Turbines pumps & compressors
GB1217275A (en) Gas turbine engine axial flow multi-stage compressor
US3606574A (en) Cooled shrouded turbine blade
US3765795A (en) Compositely formed rotors and their manufacture
US3688371A (en) The method of manufacturing compositely formed rotors
US20140127007A1 (en) Gas turbine engine rotor seal
GB1047281A (en)
US3588277A (en) Bladed rotors
CA2845615C (en) Tip-controlled integrally bladed rotor for gas turbine engine