US3847506A - Turbomachine rotor - Google Patents

Turbomachine rotor Download PDF

Info

Publication number
US3847506A
US3847506A US00420204A US42020473A US3847506A US 3847506 A US3847506 A US 3847506A US 00420204 A US00420204 A US 00420204A US 42020473 A US42020473 A US 42020473A US 3847506 A US3847506 A US 3847506A
Authority
US
United States
Prior art keywords
hub
holes
blades
rotor
rim
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
US00420204A
Inventor
S Straniti
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.)
Avco Corp
Original Assignee
Avco 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
Application filed by Avco Corp filed Critical Avco Corp
Priority to US00420204A priority Critical patent/US3847506A/en
Application granted granted Critical
Publication of US3847506A publication Critical patent/US3847506A/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
    • 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
    • F01D5/10—Anti- vibration 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means

Definitions

  • turbomachine rotor assemblies generally comprise a machined or forged hub having a series of axial dovetailed slots around its periphery. Individual blades, generally cast, have root sections received in the dovetailed grooves. Suitable retaining devices lock the blades axially to make up a completed rotor assembly.
  • these rotors are quite effective from an aerodynamic and thermodynamic standpoint, they are extremely difficult to manufacture. The reason for this is that the unit has a larger number of individual-pieces that must be made to particular tolerances to permit final assembly. In addition, they must be assembled by hand using timeconsuming and expensive methods.
  • a turbomachine rotor comprising a central annular hub and a plurality of integral airfoil blades.
  • the blades extend outward from the rim of the hub at spaced circumferential positions around its periphery.
  • the annular hub had cutaway portions adjacent its rim for minimizing stresses caused by differential thermal expansion and for minimizing vibratory stresses.
  • FIG. l is a simplified longitudinal section view of a gas turbine engine incorporating a turbomachine rotor which embodies the present invention
  • FIG. 2 is a greatly enlarged fragmentary section view of the turbomachine rotor shown in FIG. 1;
  • FIG. 3 is an end view of the turbomachine rotor of FIG. 2, taken on line 3-3 of FIG. 2;
  • FIG. 4 is a plan view of the turbomachine rotor of FIG. 2, taken on line 4-4 of FIG. 2.
  • FIG. 1 illustrates a gas turbine'engine comprising an output gearbox which supports an annular compressor'inlet housing 12. Air passes inward through an annular compressor inlet passage 14 formed in housing 12 into a compressor 16 comprising an axial flow bladed hub 18, a fixed axial flow stator 20, and a centrifugal impeller 22. Hub l8 and impeller 22 are mounted on a shaft 24 journaled at its forward end by bearing assembly 26 received in housing 12. Air is discharged from the periphery of impeller 22 into a diffuser 28 which has expanding passages for increasing the static pressure of the air.
  • pressurized air passes through a turning vane assembly 30 and into a chamber 32 formed by an annular outer housing 34 and an inner cast strut assembly 36.
  • a perforated annular combustor 38 is positioned in chamber 32 and has a plurality of nozzles 40 (only one is shown) which inject metered fuel into combustor 38 for mixing with pressurized air passing inward through the perforations.
  • a suitable device ignites the fuel/air mixture to produce a hot gas stream.
  • the hot gas stream passes through a turbine inlet duct 42 and from a turbine inlet nozzle 44.
  • Turbine rotor 46 is also mounted on shaft 24 and spaced from centrifugal hub 22 by a conical element 48.
  • a bearing assembly 50 adjacent rotor 46 journals the aft end of shaft 24 for rotation.
  • Hub 52 has an integral output shaft 56 extending forward to a speed-reduction gearset (not shown) in output gearbox l0. 7
  • turbine rotor 46 is designed in such a way that it is extremely simple and economical to manufacture.
  • turbine rotor 46 comprises an annular hub 60 having a rim 62 and a central opening 64 telescoped over shaft 24 (see FIG. 1).
  • a plurality of airfoil-shaped blades 66 are integral with hub 60 and project radially outwardly from rim 62 of hub 60.
  • Hub 60 and blades 66 are preferably cast from a hightemperature alloy normally used to form cast turbine blades. Thesetypes of alloys are strong enough to be used for the hub 60 and withstand centrifugal forces during rotation.
  • Blades 66 have a root section stagger angle a which defines an acute angle relative to the axis A of the hub 60.
  • Hub 60 has a plurality of holes 68 extending from its forward face 70 to its aft face 72 adjacent its rim 62.
  • the holes 68 are positioned in between adjacent blades 66 and have their axis generally parallel to the line S defining the stagger angle a.
  • the upstream ends of the holes 68 are radially outward relative to the downstream ends.
  • Each hole 68 has a slot 74 extending radially outward from the hole 68 to the rim 62 of hub 60.
  • slot 74 is ex tremely thin and it is conveniently formed by EDM or other suitable production technique.
  • Each hole 68 has an annular shoulder '76 formed at its upstream end and a circumferential groove 78 formed at its downstream end.
  • a cup 80 is received in each hole 68 and is maintained against shoulder 76 by a tubular element 82.
  • Tubular element 82 is deformed radially outward at 84 to hold the assembly in each hole 68.
  • the series of holes and slots also define sections 86 in between adjacent holes that act to in effect elongate the blades and permit them to flex.
  • the position of the holes 68 relative to the rim 62 can be varied to achieve blades with predetermined natural frequencies of vibration. This frequency is selected so that it will lie outside of the normal operating range for the engine.
  • This frequency is selected so that it will lie outside of the normal operating range for the engine.
  • the walls of slots 74 frictionally engage one another to damp the vibrations.
  • the tube 82 is forced against the radially outward walls of holes 68 by centrifugal force to damp vibration.
  • the cup 80 prevents the passage of fluid through holes 68 to maintain high efficiency. While some flow could be expected across slot 74, it is so small in width (approximately .003 to .007 inches) that the leakage through this slot is negligible. Although the pressure differential acts to push the cup 80 and tube 82 out of holes 68, centrifugal force urges them toward the upstream end of the holes. ln addition, the flared-out section 84 adequately holds the tube 82 in place.
  • the above turbomachine rotor is extremely simplified and is capable of being manufactured using lowcost, high-volume production techniques. At the same time, however, it enables the high level of thermodynamic efficiency and minimizes problems associated with differential thermal expansion and vibratory effects.
  • a one-piece cast turbomachine rotor comprising:
  • said holes extending generally parallel to the stagger angle line of said blades, said holes each having an annular shoulder of reduced diameter formed adjacent one end thereof, and an annular recess of enlarged diameter adjacent the opposite end thereof;
  • each of said slots extending radially outward from a respective hole and through the rim of said hub, said slots also extending in a line parallel to said stagger angle line;
  • a turbomachine rotor as in claim 1 wherein said holes are positioned at a predetermined location relative to the periphery of said hub thereby producing a given natural frequency of vibration for said blades whereby said rotor may be tuned to a frequency lying outside its normal operating range. 2: 4:

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A turbine rotor assembly comprising a cast hub and a series of integral airfoil blades extending radially outward from its rim. A series of holes and thin slots extend through the hub in between adjacent blades to minimize stress caused by differential thermal expansion and to place the natural frequency of the blades outside the normal operating range of the rotor. End caps and tubular inserts are received in the holes and held in place by deforming the inserts into circumferential grooves.

Description

United, States Patent 11 1 Straniti Nov. 12, 1974 TURBOMACHINE ROTOR FOREIGN PATENTS OR APPLICATIONS 1751 lnvemori Salvatore Stranifi, Orange, Com 609.446 9/1948 ore-61 13min 4l6/244 A [73] Assignee: Avco CorporatiomStratford, Conn. P E E A P H J rimarv xaminerverette owe r. [.22] F'led: Attorney, Agent, or Firm-Charles M. Hogan; lrwin P. [2]] Appl. No.: 420,204 Garfinkle 52 us. 01. 416/244 [57] ABSTRACT v [51] Int. Cl......' .f. F0 l d 5(08 A turbine rotor assembly comprising a cast hub and a [58] Field Of Search 416/244 A, 221 series of integral airfoil blades extending radially utward from its rim. A series of holes and thin slots ex- [56] References Cited tend through the hub in between adjacent blades to UNITED STATES PATENTS minimize stress caused by differential thermal expan- 737,042 8/1903 Stumpf 416 244 A and to Place the natural frequemy of the blades 2,753,149 7/l956 Kurti 416 221 ux Outside the normal Operating range of the FOIOF- End 2,953,348 9/1960 Leland 416/221 p and tubular inserts are received in the holes and 3,255,994 6/1966 Dreimimis 416/244 A held in place by deforming the inserts into circumfer- 3,262,676 7/1966 Huebner et al. 416/244 A ential grooves. 3,291,446 l2/l966 Huebner 4l6/244 A X 3 2 5 3 1/1967 2 Claims, 4 Drawing Figures Nickles 416/221 PATENIEUHBVIZ i974 I 3; L sum 10? 2 847 506 TURBOMACHINE RoToR The present invention relates to turbomachine rotors and more specifically to simplified rotors of this type.
One of the biggest obstacles to the mass production of a truly economical gas turbine engine has been the rotor construction. Present day turbomachine rotor assemblies generally comprise a machined or forged hub having a series of axial dovetailed slots around its periphery. Individual blades, generally cast, have root sections received in the dovetailed grooves. Suitable retaining devices lock the blades axially to make up a completed rotor assembly. Although these rotors are quite effective from an aerodynamic and thermodynamic standpoint, they are extremely difficult to manufacture. The reason for this is that the unit has a larger number of individual-pieces that must be made to particular tolerances to permit final assembly. In addition, they must be assembled by hand using timeconsuming and expensive methods.
The above problems are solved in accordance with the present invention by a turbomachine rotor comprising a central annular hub and a plurality of integral airfoil blades. The blades extend outward from the rim of the hub at spaced circumferential positions around its periphery. The annular hub had cutaway portions adjacent its rim for minimizing stresses caused by differential thermal expansion and for minimizing vibratory stresses.
The above and other related features of the present invention will be apparent from a reading of the following description of the disclosure shown in the accompanying drawings and the novelty thereof pointed out in the appended claims.
In the drawings:
FIG. l is a simplified longitudinal section view of a gas turbine engine incorporating a turbomachine rotor which embodies the present invention;
FIG. 2 is a greatly enlarged fragmentary section view of the turbomachine rotor shown in FIG. 1;
FIG. 3 is an end view of the turbomachine rotor of FIG. 2, taken on line 3-3 of FIG. 2; and
FIG. 4 is a plan view of the turbomachine rotor of FIG. 2, taken on line 4-4 of FIG. 2.
FIG. 1 illustrates a gas turbine'engine comprising an output gearbox which supports an annular compressor'inlet housing 12. Air passes inward through an annular compressor inlet passage 14 formed in housing 12 into a compressor 16 comprising an axial flow bladed hub 18, a fixed axial flow stator 20, and a centrifugal impeller 22. Hub l8 and impeller 22 are mounted on a shaft 24 journaled at its forward end by bearing assembly 26 received in housing 12. Air is discharged from the periphery of impeller 22 into a diffuser 28 which has expanding passages for increasing the static pressure of the air.
From these the pressurized air passes through a turning vane assembly 30 and into a chamber 32 formed by an annular outer housing 34 and an inner cast strut assembly 36. A perforated annular combustor 38 is positioned in chamber 32 and has a plurality of nozzles 40 (only one is shown) which inject metered fuel into combustor 38 for mixing with pressurized air passing inward through the perforations. A suitable device ignites the fuel/air mixture to produce a hot gas stream.
The hot gas stream passes through a turbine inlet duct 42 and from a turbine inlet nozzle 44.
The hot gas stream then passes across a bladed turbine rotor 46, to be described in detail later. Turbine rotor 46 is also mounted on shaft 24 and spaced from centrifugal hub 22 by a conical element 48. A bearing assembly 50 adjacent rotor 46 journals the aft end of shaft 24 for rotation. After the hot gas stream leaves the turbine rotor 46 it passes across a bladed power turbine hub 52 which is journaled in strut assembly 36 by a bearing 54. Hub 52 has an integral output shaft 56 extending forward to a speed-reduction gearset (not shown) in output gearbox l0. 7
In accordance with the present invention turbine rotor 46 is designed in such a way that it is extremely simple and economical to manufacture. Referring particularly to FIGS. 2, 3 and 4, turbine rotor 46 comprises an annular hub 60 having a rim 62 and a central opening 64 telescoped over shaft 24 (see FIG. 1). A plurality of airfoil-shaped blades 66 are integral with hub 60 and project radially outwardly from rim 62 of hub 60. Hub 60 and blades 66 are preferably cast from a hightemperature alloy normally used to form cast turbine blades. Thesetypes of alloys are strong enough to be used for the hub 60 and withstand centrifugal forces during rotation. Blades 66 have a root section stagger angle a which defines an acute angle relative to the axis A of the hub 60. Hub 60 has a plurality of holes 68 extending from its forward face 70 to its aft face 72 adjacent its rim 62. As is particularly evident in FIG. 4, the holes 68 are positioned in between adjacent blades 66 and have their axis generally parallel to the line S defining the stagger angle a. In addition, as seen in FIG. 2, the upstream ends of the holes 68 are radially outward relative to the downstream ends. Each hole 68 has a slot 74 extending radially outward from the hole 68 to the rim 62 of hub 60. As described below, slot 74 is ex tremely thin and it is conveniently formed by EDM or other suitable production technique. Each hole 68 has an annular shoulder '76 formed at its upstream end and a circumferential groove 78 formed at its downstream end. A cup 80 is received in each hole 68 and is maintained against shoulder 76 by a tubular element 82. Tubular element 82 is deformed radially outward at 84 to hold the assembly in each hole 68.
When the engine is in operation the hot gas stream passing across blades 66 and rim 62 causes a very large temperature gradient relative to the inner sections of hub 60. The holes 68 and slots 74 permit the rim section of the hub 60 to freely expandin response to the high temperature without creating undue stresses.
The series of holes and slots also define sections 86 in between adjacent holes that act to in effect elongate the blades and permit them to flex. The position of the holes 68 relative to the rim 62 can be varied to achieve blades with predetermined natural frequencies of vibration. This frequency is selected so that it will lie outside of the normal operating range for the engine. When vibrating conditions are encountered the walls of slots 74 frictionally engage one another to damp the vibrations. Also, the tube 82 is forced against the radially outward walls of holes 68 by centrifugal force to damp vibration.
The cup 80 prevents the passage of fluid through holes 68 to maintain high efficiency. While some flow could be expected across slot 74, it is so small in width (approximately .003 to .007 inches) that the leakage through this slot is negligible. Although the pressure differential acts to push the cup 80 and tube 82 out of holes 68, centrifugal force urges them toward the upstream end of the holes. ln addition, the flared-out section 84 adequately holds the tube 82 in place.
The above turbomachine rotor is extremely simplified and is capable of being manufactured using lowcost, high-volume production techniques. At the same time, however, it enables the high level of thermodynamic efficiency and minimizes problems associated with differential thermal expansion and vibratory effects.
While a preferred embodiment of the present invention has been described, it should be apparent to those skilled in the art that it may be practiced in other forms without departing from its spirit and scope.
Having thus described the invention what is claimed as novel and desired to be secured by Letters Patent of the United States is:
l. A one-piece cast turbomachine rotor comprising:
a central cylindrical hub;
a plurality of airfoil blades integral with and extending outward from the periphery of said hub at spaced circumferential positions around the rim of said hub, said blades having an acute stagger angle relative to the axis of said hub;
a plurality of holes in said annular hub, said holes extending generally parallel to the stagger angle line of said blades, said holes each having an annular shoulder of reduced diameter formed adjacent one end thereof, and an annular recess of enlarged diameter adjacent the opposite end thereof;
a like plurality of continuous slots, each of said slots extending radially outward from a respective hole and through the rim of said hub, said slots also extending in a line parallel to said stagger angle line; and
a generally cylindrical cup compressed into each of said holes, said cup being closed at one end and having a configuration complementary to said holes whereby said cups are retained in said holes between said shoulders and said recesses.
2. A turbomachine rotor as in claim 1 wherein said holes are positioned at a predetermined location relative to the periphery of said hub thereby producing a given natural frequency of vibration for said blades whereby said rotor may be tuned to a frequency lying outside its normal operating range. =2: 4:

Claims (2)

1. A one-piece cast turbomachine rotor comprising: a central cylindrical hub; a plurality of airfoil blades integral with and extending outward from the periphery of said hub at spaced circumferential positions around the rim of said hub, said blades having an acute stagger angle relative to the axis of said hub; a plurality of holes in said annular hub, said holes extending generally parallel to the stagger angle line of said blades, said holes each having an annular shoulder of reduced diameter formed adjacent one end thereof, and an annular recess of enlarged diameter adjacent the opposite end thereof; a like plurality of continuous slots, each of said slots extending radially outward from a respective hole and through the rim of said hub, said slots also extending in a line parallel to said stagger angle line; and a generally cylindrical cup compressed into each of said holes, said cup being closed at one end and having a configuration complementary to said holes whereby said cups are retained in said holes between said shoulders and said recesses.
2. A turbomachine rotor as in claim 1 wherein said holes are positioned at a predetermined location relative to the periphery of said hub thereby producing a given natural frequency of vibration for said blades whereby said rotor may be tuned to a frequency lying outside its normal operating range.
US00420204A 1973-11-29 1973-11-29 Turbomachine rotor Expired - Lifetime US3847506A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US00420204A US3847506A (en) 1973-11-29 1973-11-29 Turbomachine rotor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US00420204A US3847506A (en) 1973-11-29 1973-11-29 Turbomachine rotor

Publications (1)

Publication Number Publication Date
US3847506A true US3847506A (en) 1974-11-12

Family

ID=23665506

Family Applications (1)

Application Number Title Priority Date Filing Date
US00420204A Expired - Lifetime US3847506A (en) 1973-11-29 1973-11-29 Turbomachine rotor

Country Status (1)

Country Link
US (1) US3847506A (en)

Cited By (33)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4497611A (en) * 1982-03-25 1985-02-05 Kraftwerk Union Aktiengesellschaft Device for vibration damping in a guide vane ring
US4776763A (en) * 1987-12-02 1988-10-11 Sundstrand Corporation Mechanical damping of turbine wheel blades
US4787821A (en) * 1987-04-10 1988-11-29 Allied Signal Inc. Dual alloy rotor
US4813848A (en) * 1987-10-14 1989-03-21 United Technologies Corporation Turbine rotor disk and blade assembly
US5215442A (en) * 1991-10-04 1993-06-01 General Electric Company Turbine blade platform damper
US20040064945A1 (en) * 2001-12-27 2004-04-08 Todd Howley Method of forming turbine blade root
US20050254958A1 (en) * 2004-05-14 2005-11-17 Paul Stone Natural frequency tuning of gas turbine engine blades
WO2008041889A1 (en) * 2006-10-05 2008-04-10 Volvo Aero Corporation Rotor element and method for producing the rotor element
US20080304972A1 (en) * 2007-06-07 2008-12-11 Honeywell International, Inc. Rotary body for turbo machinery with mistuned blades
US20090155082A1 (en) * 2007-12-18 2009-06-18 Loc Duong Method to maximize resonance-free running range for a turbine blade
EP2075411A1 (en) * 2007-12-28 2009-07-01 United Technologies Corporation Integrally bladed rotor with slotted outer rim and gas turbine engine comprising such a rotor
DE102009007468A1 (en) * 2009-02-04 2010-08-19 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
US20100239422A1 (en) * 2009-03-19 2010-09-23 Honeywell International Inc. Components for gas turbine engines
US20100325852A1 (en) * 2009-06-29 2010-12-30 Frederick Michel Method and apparatus for providing rotor discs
EP2453108A1 (en) * 2010-11-15 2012-05-16 MTU Aero Engines GmbH Rotor for a turbomachine
US20120282109A1 (en) * 2011-05-02 2012-11-08 Mtu Aero Engines Gmbh Blade, Integrally Bladed Rotor Base Body and Turbomachine
DE102011100221A1 (en) * 2011-05-02 2012-11-08 Mtu Aero Engines Gmbh Covering device, integrally bladed rotor body, method and turbomachine
US20150118048A1 (en) * 2013-10-24 2015-04-30 Honeywell International Inc. Gas turbine engine rotors including intra-hub stress relief features and methods for the manufacture thereof
US9273563B2 (en) 2007-12-28 2016-03-01 United Technologies Corporation Integrally bladed rotor with slotted outer rim
US20160069203A1 (en) * 2013-04-12 2016-03-10 United Technologies Corporation Integrally bladed rotor
US10040122B2 (en) 2014-09-22 2018-08-07 Honeywell International Inc. Methods for producing gas turbine engine rotors and other powdered metal articles having shaped internal cavities
US10189100B2 (en) 2008-07-29 2019-01-29 Pratt & Whitney Canada Corp. Method for wire electro-discharge machining a part
DE102018200832A1 (en) * 2018-01-19 2019-07-25 MTU Aero Engines AG Rotor, in particular blisk of a gas turbine, with dissolved rim and method for producing the same
US20200056506A1 (en) * 2018-08-17 2020-02-20 United Technologies Corporation Gas turbine engine seal ring assembly
US10648354B2 (en) 2016-12-02 2020-05-12 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US10760429B1 (en) * 2017-01-17 2020-09-01 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US10760592B1 (en) * 2017-01-17 2020-09-01 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US10788049B1 (en) * 2017-01-17 2020-09-29 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US10982551B1 (en) 2012-09-14 2021-04-20 Raytheon Technologies Corporation Turbomachine blade
US11149651B2 (en) 2019-08-07 2021-10-19 Raytheon Technologies Corporation Seal ring assembly for a gas turbine engine
US11199096B1 (en) 2017-01-17 2021-12-14 Raytheon Technologies Corporation Turbomachine blade
US11698002B1 (en) * 2017-01-17 2023-07-11 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US11767763B1 (en) * 2017-01-17 2023-09-26 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US737042A (en) * 1903-04-03 1903-08-25 Johann Stumpf Wheel or disk for steam-engines.
GB609446A (en) * 1946-03-14 1948-09-30 Parsons C A & Co Ltd Improvements in or relating to the rotors of gas turbines or the like
US2753149A (en) * 1951-03-30 1956-07-03 United Aircraft Corp Blade lock
US2953348A (en) * 1952-12-30 1960-09-20 Gen Motors Corp Blade fastenings
US3255994A (en) * 1963-09-03 1966-06-14 Chrysler Corp Turbine wheel
US3262676A (en) * 1964-05-27 1966-07-26 Chrysler Corp Turbine wheel
US3291446A (en) * 1965-04-13 1966-12-13 Chrysler Corp Turbine wheel
US3295826A (en) * 1966-04-08 1967-01-03 Gen Motors Corp Blade lock

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US737042A (en) * 1903-04-03 1903-08-25 Johann Stumpf Wheel or disk for steam-engines.
GB609446A (en) * 1946-03-14 1948-09-30 Parsons C A & Co Ltd Improvements in or relating to the rotors of gas turbines or the like
US2753149A (en) * 1951-03-30 1956-07-03 United Aircraft Corp Blade lock
US2953348A (en) * 1952-12-30 1960-09-20 Gen Motors Corp Blade fastenings
US3255994A (en) * 1963-09-03 1966-06-14 Chrysler Corp Turbine wheel
US3262676A (en) * 1964-05-27 1966-07-26 Chrysler Corp Turbine wheel
US3291446A (en) * 1965-04-13 1966-12-13 Chrysler Corp Turbine wheel
US3295826A (en) * 1966-04-08 1967-01-03 Gen Motors Corp Blade lock

Cited By (56)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4497611A (en) * 1982-03-25 1985-02-05 Kraftwerk Union Aktiengesellschaft Device for vibration damping in a guide vane ring
US4787821A (en) * 1987-04-10 1988-11-29 Allied Signal Inc. Dual alloy rotor
US4813848A (en) * 1987-10-14 1989-03-21 United Technologies Corporation Turbine rotor disk and blade assembly
US4776763A (en) * 1987-12-02 1988-10-11 Sundstrand Corporation Mechanical damping of turbine wheel blades
US5215442A (en) * 1991-10-04 1993-06-01 General Electric Company Turbine blade platform damper
US20040064945A1 (en) * 2001-12-27 2004-04-08 Todd Howley Method of forming turbine blade root
US20050254958A1 (en) * 2004-05-14 2005-11-17 Paul Stone Natural frequency tuning of gas turbine engine blades
WO2005111377A1 (en) * 2004-05-14 2005-11-24 Pratt & Whitney Canada Corp. Natural frequency tuning of gas turbine engine blades
US7252481B2 (en) 2004-05-14 2007-08-07 Pratt & Whitney Canada Corp. Natural frequency tuning of gas turbine engine blades
WO2008041889A1 (en) * 2006-10-05 2008-04-10 Volvo Aero Corporation Rotor element and method for producing the rotor element
US8757979B2 (en) 2006-10-05 2014-06-24 Volvo Aero Corporation Rotor element and method for producing the rotor element
US20100021305A1 (en) * 2006-10-05 2010-01-28 Hans Martensson Rotor element and method for producing the rotor element
US20080304972A1 (en) * 2007-06-07 2008-12-11 Honeywell International, Inc. Rotary body for turbo machinery with mistuned blades
US7887299B2 (en) * 2007-06-07 2011-02-15 Honeywell International Inc. Rotary body for turbo machinery with mistuned blades
US20090155082A1 (en) * 2007-12-18 2009-06-18 Loc Duong Method to maximize resonance-free running range for a turbine blade
EP2075411A1 (en) * 2007-12-28 2009-07-01 United Technologies Corporation Integrally bladed rotor with slotted outer rim and gas turbine engine comprising such a rotor
US20110182745A1 (en) * 2007-12-28 2011-07-28 Suciu Gabriel L Integrally bladed rotor with slotted outer rim
US9273563B2 (en) 2007-12-28 2016-03-01 United Technologies Corporation Integrally bladed rotor with slotted outer rim
US9133720B2 (en) 2007-12-28 2015-09-15 United Technologies Corporation Integrally bladed rotor with slotted outer rim
US11583947B2 (en) 2008-07-29 2023-02-21 Pratt & Whitney Canada Corp. Method for wire electro-discharge machining a part
US10189100B2 (en) 2008-07-29 2019-01-29 Pratt & Whitney Canada Corp. Method for wire electro-discharge machining a part
DE102009007468A1 (en) * 2009-02-04 2010-08-19 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
US8821122B2 (en) 2009-02-04 2014-09-02 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
US20100239422A1 (en) * 2009-03-19 2010-09-23 Honeywell International Inc. Components for gas turbine engines
US8157514B2 (en) 2009-03-19 2012-04-17 Honeywell International Inc. Components for gas turbine engines
EP2230382A3 (en) * 2009-03-19 2014-03-12 Honeywell International Inc. Gas turbine rotor stage
US8925201B2 (en) 2009-06-29 2015-01-06 Pratt & Whitney Canada Corp. Method and apparatus for providing rotor discs
US20100325852A1 (en) * 2009-06-29 2010-12-30 Frederick Michel Method and apparatus for providing rotor discs
EP2453108A1 (en) * 2010-11-15 2012-05-16 MTU Aero Engines GmbH Rotor for a turbomachine
DE102011100221A1 (en) * 2011-05-02 2012-11-08 Mtu Aero Engines Gmbh Covering device, integrally bladed rotor body, method and turbomachine
US20120282109A1 (en) * 2011-05-02 2012-11-08 Mtu Aero Engines Gmbh Blade, Integrally Bladed Rotor Base Body and Turbomachine
DE102011100221B4 (en) 2011-05-02 2017-03-09 MTU Aero Engines AG Integrally bladed rotor body, process and turbomachine
WO2012149925A3 (en) * 2011-05-02 2013-02-28 Mtu Aero Engines Gmbh Cover device, integrally bladed main rotor body, method and turbomachine
US9739151B2 (en) * 2011-05-02 2017-08-22 Mtu Aero Engines Gmbh Blade, integrally bladed rotor base body and turbomachine
US10982551B1 (en) 2012-09-14 2021-04-20 Raytheon Technologies Corporation Turbomachine blade
US10458265B2 (en) 2013-04-12 2019-10-29 United Technologies Corporation Integrally bladed rotor
US20160069203A1 (en) * 2013-04-12 2016-03-10 United Technologies Corporation Integrally bladed rotor
EP2984290A4 (en) * 2013-04-12 2016-06-22 United Technologies Corp Integrally bladed rotor
US9714577B2 (en) * 2013-10-24 2017-07-25 Honeywell International Inc. Gas turbine engine rotors including intra-hub stress relief features and methods for the manufacture thereof
US20150118048A1 (en) * 2013-10-24 2015-04-30 Honeywell International Inc. Gas turbine engine rotors including intra-hub stress relief features and methods for the manufacture thereof
US11305348B2 (en) 2014-09-22 2022-04-19 Honeywell International Inc. Methods for producing gas turbine engine rotors and other powdered metal articles having shaped internal cavities
US10040122B2 (en) 2014-09-22 2018-08-07 Honeywell International Inc. Methods for producing gas turbine engine rotors and other powdered metal articles having shaped internal cavities
US10807166B2 (en) 2014-09-22 2020-10-20 Honeywell International Inc. Methods for producing gas turbine engine rotors and other powdered metal articles having shaped internal cavities
US10851660B2 (en) 2016-12-02 2020-12-01 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US10648354B2 (en) 2016-12-02 2020-05-12 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US11015472B2 (en) 2016-12-02 2021-05-25 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US11199096B1 (en) 2017-01-17 2021-12-14 Raytheon Technologies Corporation Turbomachine blade
US10788049B1 (en) * 2017-01-17 2020-09-29 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US10760592B1 (en) * 2017-01-17 2020-09-01 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US10760429B1 (en) * 2017-01-17 2020-09-01 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US11698002B1 (en) * 2017-01-17 2023-07-11 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
US11767763B1 (en) * 2017-01-17 2023-09-26 Raytheon Technologies Corporation Gas turbine engine airfoil frequency design
DE102018200832A1 (en) * 2018-01-19 2019-07-25 MTU Aero Engines AG Rotor, in particular blisk of a gas turbine, with dissolved rim and method for producing the same
US10920617B2 (en) * 2018-08-17 2021-02-16 Raytheon Technologies Corporation Gas turbine engine seal ring assembly
US20200056506A1 (en) * 2018-08-17 2020-02-20 United Technologies Corporation Gas turbine engine seal ring assembly
US11149651B2 (en) 2019-08-07 2021-10-19 Raytheon Technologies Corporation Seal ring assembly for a gas turbine engine

Similar Documents

Publication Publication Date Title
US3847506A (en) Turbomachine rotor
US3533712A (en) Cooled vane structure for high temperature turbines
US5224339A (en) Counterflow single rotor turbojet and method
US3505819A (en) Gas turbine power plant
US3703081A (en) Gas turbine engine
EP2169181B1 (en) Gas turbine engine rotor and balance weight therefor
KR960034693A (en) Compressor rotor cooling system for gas turbines
EP3088670B1 (en) Film cooling system
EP2365235A1 (en) Cooled turbine rim seal
US3203180A (en) Turbo-jet powerplant
JP2005155623A (en) Triple circuit turbine cooling
ES410317A1 (en) Centrifugal flow gas turbine engine with annular combustor
US2501614A (en) Compressor construction
EP3597875B1 (en) Debris separator for a gas turbine engine and corresponding installing method
US3844678A (en) Cooled high strength turbine bucket
US5201845A (en) Low pressure drop radial inflow air-oil separating arrangement and separator employed therein
EP3663022A1 (en) A method of manufacturing a planet carrier of a gearbox
US20230258134A1 (en) High fan tip speed engine
CN109083847A (en) Compressor bleed air equipment and its bleed method for turbogenerator
EP3388623B1 (en) Compressor having reinforcing disk, and gas turbine having same
US3873232A (en) Two-piece channel diffuser
US20180128119A1 (en) Minidisk balance flange
EP3246517B1 (en) Fastener openings for stress distribution
US20230243268A1 (en) Airfoils for gas turbine engines
JPH0347401U (en)