US20120057988A1 - Rotor for a turbomachine - Google Patents

Rotor for a turbomachine Download PDF

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Publication number
US20120057988A1
US20120057988A1 US13/202,010 US201013202010A US2012057988A1 US 20120057988 A1 US20120057988 A1 US 20120057988A1 US 201013202010 A US201013202010 A US 201013202010A US 2012057988 A1 US2012057988 A1 US 2012057988A1
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US
United States
Prior art keywords
rotor
further characterized
rotor according
blades
blade
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.)
Abandoned
Application number
US13/202,010
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English (en)
Inventor
Frank Stiehler
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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 MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Assigned to MTU AERO ENGINES GMBH, A COMPANY OF GERMANY reassignment MTU AERO ENGINES GMBH, A COMPANY OF GERMANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: STIEHLER, FRANK
Publication of US20120057988A1 publication Critical patent/US20120057988A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/666—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps by means of rotor construction or layout, e.g. unequal distribution of blades or vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/668—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps damping or preventing mechanical vibrations
    • 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 invention relates to a rotor for a turbomachine, in particular for a gas turbine, having a basic rotor body and a plurality of blades.
  • rotors having integral blading depends on whether a rotor or rotor support (called a basic rotor body in the following) is present that is shaped like a disk in cross section (blisk) or is ring-shaped in cross section (bling).
  • Blisk is the abbreviated form of bladed disk and bling of bladed ring.
  • integrally bladed rotors are known, during the production of which, the blades will be joined together into a blade ring in a first step and subsequently will be fastened to a basic rotor body. Vibrations can be transferred from one blade to the adjacent blades via the structure of such a blade ring. In addition, a replacing of individual blades in order to repair the rotor is not provided and is difficult for such rotors.
  • the problem of the invention is to create a rotor for a turbomachine with fewer vibrations.
  • a rotor according to the invention for a turbomachine, in particular for a gas turbine, having a basic rotor body and a plurality of blades, wherein at least one damping element is provided in the circumferential direction between adjacent blades of the rotor. In this way, vibrations of a blade can be damped and in particular, the transfer of vibrations to adjacent blades can be reduced.
  • the rotor is an integrally bladed rotor, in particular a rotor in which the basic rotor body and blades are welded directly to one another or via a separate intermediate piece; in particular, they are friction-welded.
  • integrally bladed rotors in turbomachines makes possible a savings in weight when compared with multi-part rotors and is thus particularly of advantage for aircraft engines.
  • the blades have a blade neck, by means of which they are joined to the basic rotor body.
  • the neck region of the blade can have different structural and functional features that can be integrated into the blade as one piece in this way.
  • the blade necks have, e.g., extensions that together form a shroud that bounds by its radially outward-pointing surface an annular flow channel of the turbomachine.
  • a shroud can separate the flow channel from the basic rotor body, so that the basic rotor body is protected against loads, particularly by hot gases. It is possible that a hollow space between adjacent rotors in the turbomachine is formed by the shroud, and this space can be used, in particular, for cooling the rotor.
  • the at least one damping element is essentially disposed in the region of the blade neck. In this way, the damping element does not occupy any structural space in the flow channel of the turbomachine.
  • the blade necks of adjacent blades are distanced from one another in the circumferential direction, and a free space will be formed between the blade necks. Due to this free space, the free structural space around a blade is enlarged, whereby it is possible, for example, to fasten an individual blade to the basic rotor body or to separate this blade from it, independently from the adjacent blades.
  • the damping element preferably forms a seal of the free space between adjacent blade necks relative to a flow channel of the turbomachine. This prevents the free space between the blade necks from acting negatively on the flow ratios or aerodynamic conditions in the turbomachine.
  • the damping element can form a part of the shroud and can limit the annular flow channel by its surface that is directed radially outward.
  • the damping element thus makes possible a smoother transition between the extensions of adjacent blades.
  • the damping element is disposed as an inserted piece between the blades.
  • one inserted piece can be inserted or removed independently from other components of the rotor.
  • the damping element can be joined to the adjacent blades in a form-fitting manner, via a press connection and/or cohesively. This reinforces the mechanical stability of the rotor.
  • damping element is joined to the basic rotor body in a form-fitting manner, via a press connection and/or cohesively.
  • a channel is provided between the at least one damping element and the basic rotor body.
  • a channel for example, makes possible a connection between the front and back sides of the rotor.
  • the channel can be formed at least partially by a notching in the basic rotor body and/or in the blade neck.
  • the channel is a cooling channel.
  • a cooling of the basic rotor body, the blades, and/or the damping element is made possible in this way.
  • a plurality of damping elements that are joined together are provided. For example, this makes possible the coupling of the plurality of damping elements and reduces the number of individual parts in the assembly of the rotor.
  • the plurality of damping elements is introduced on a support ring that can be axially attached.
  • the support ring can be designed self-supporting, so that its mass does not contribute to the load on the basic rotor body when it is rotating.
  • FIG. 1 shows a section of a rotor according to the invention with blade necks distanced from one another prior to the insertion of the damping elements
  • FIG. 2 shows the rotor from FIG. 1 with inserted damping elements
  • FIG. 3 shows another view of the rotor according to FIG. 2 ;
  • FIG. 4 shows a detail view of the region of the blade neck of a rotor according to the invention.
  • FIG. 1 shows a section of a rotor 10 having a basic rotor body 12 , on which are fastened several blades 14 .
  • the blades 14 have a blade surface 16 and a blade neck 18 .
  • the blade neck 18 On the side lying in the circumferential direction, the blade neck 18 has a recess 20 , which makes possible a lightweight construction of the rotor 10 by reducing the weight of the blades 14 .
  • the blade necks 18 In the axial direction, the blade necks 18 have extensions 22 that together form a shroud that bounds by its radially outward-pointing surface an annular gas flow channel of the turbomachine.
  • Rotor 10 is designed for a gas turbine, whereby it can be disposed in the turbine section or in the compressor section.
  • the invention can also be applied to rotors of other turbomachines.
  • the blades 14 are joined to the basic rotor body 12 via the blade necks 18 .
  • basic rotor body 12 and blades 14 are welded together directly.
  • blades 14 are manufactured particularly from materials that do not permit fusion welding methods, for example, monocrystalline materials. Blades 14 are thus joined to the basic rotor body 12 via friction-welding processes, in particular linear friction-welding processes, or inductive high-frequency press welding.
  • Blades 14 are distanced from one another in the circumferential direction, particularly in the region of the blade neck 18 , whereby a free space is formed between the blades 14 .
  • the free spaces on both sides of a blade 14 provide sufficient structural space to fasten an individual blade 14 to the basic rotor body 12 or to separate this blade from it, independently from the adjacent blades 14 . This makes possible a simplified production process and particularly a simplified repair process by being able to replace an individual blade 14 .
  • blades 14 are welded to the basic rotor body 12 via a separate (in particular, polycrystalline) intermediate piece. In this way, one need only pay attention to the particular joining method when the intermediate piece and blades 14 are joined.
  • blades 14 are not integrally joined to the basic rotor body 12 , but are anchored in the basic rotor body 12 via blade feet.
  • FIG. 2 and FIG. 3 show the section of rotor 10 from FIG. 1 , whereby damping elements 24 are inserted between the blade necks 18 .
  • FIG. 4 shows a detail view of the region of the blade neck with inserted damping elements 24 .
  • the damping elements 24 reduce the transfer of vibrations from one blade 14 to adjacent blades 14 and also may have a sealing function.
  • the damping elements 24 are joined to the adjacent blade necks 18 in a form-fitting manner and/or cohesively and/or via a press fitting. Of course, the damping elements 24 can also be joined correspondingly to the basic rotor body 12 .
  • the damping elements 24 fill the free space between the adjacent blade necks 18 and seal this free space relative to the flow channel of the turbomachine.
  • the shape of the damping elements 24 is adapted to the shape of the blade necks 18 , particularly in the region of extensions 22 .
  • the radially outwardly pointing surface of the damping element 24 thus bounds the annular flow channel of the turbomachine.
  • the damping elements 24 are disposed as individual inserted pieces, each between two adjacent blades 14 .
  • a plurality of damping elements 24 can also be joined together. In this way, on the one hand, the damping elements 24 can be coupled together, and, on the other hand, the number of individual parts in the production of rotor 10 is reduced. It is also possible that a damping element 24 is not introduced between each adjacent pair of blades.
  • a plurality of damping elements 24 can be introduced on a support ring that can be fastened axially and that is inserted into the rotor 10 in the axial direction during assembly.
  • a support ring is designed self-supporting, so that its mass does not contribute to the load of the rotor 10 or of the basic rotor body 12 during operation.
  • Channels 26 which are formed by a notches 28 in the basic rotor body 12 and in the blade neck 18 , are provided between the damping elements 24 and the basic rotor body 12 .
  • the channels 26 serve as cooling channels because a cooling fluid flows through them. Channels 26 thus make possible a flow of cooling fluid in the axial direction from the axially front side of the rotor 10 to the axially back side of rotor 10 .
  • Ni-based alloys are particularly considered as the material for the damping elements 24 .
  • the material should be softer than the adjacent blade necks.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/202,010 2009-03-05 2010-02-27 Rotor for a turbomachine Abandoned US20120057988A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
DE102009011964.7 2009-03-05
DE102009011964A DE102009011964A1 (de) 2009-03-05 2009-03-05 Rotor für eine Strömungsmaschine
PCT/DE2010/000220 WO2010099781A2 (fr) 2009-03-05 2010-02-27 Rotor pour turbomachine

Publications (1)

Publication Number Publication Date
US20120057988A1 true US20120057988A1 (en) 2012-03-08

Family

ID=42538567

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/202,010 Abandoned US20120057988A1 (en) 2009-03-05 2010-02-27 Rotor for a turbomachine

Country Status (4)

Country Link
US (1) US20120057988A1 (fr)
EP (1) EP2350459A2 (fr)
DE (1) DE102009011964A1 (fr)
WO (1) WO2010099781A2 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20160032753A1 (en) * 2014-07-31 2016-02-04 United Technologies Corporation Gas turbine engine with axial compressor having improved air sealing
US9810075B2 (en) 2015-03-20 2017-11-07 United Technologies Corporation Faceted turbine blade damper-seal
US10196896B2 (en) 2015-04-13 2019-02-05 Rolls-Royce Plc Rotor damper
US10344601B2 (en) 2012-08-17 2019-07-09 United Technologies Corporation Contoured flowpath surface
US10443502B2 (en) 2015-04-13 2019-10-15 Rolls-Royce Plc Rotor damper

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120099995A1 (en) 2010-10-20 2012-04-26 General Electric Company Rotary machine having spacers for control of fluid dynamics
EP2644833A1 (fr) 2012-03-26 2013-10-02 Alstom Technology Ltd Anneau de support

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4175912A (en) * 1976-10-19 1979-11-27 Rolls-Royce Limited Axial flow gas turbine engine compressor
US4192633A (en) * 1977-12-28 1980-03-11 General Electric Company Counterweighted blade damper
US4457668A (en) * 1981-04-07 1984-07-03 S.N.E.C.M.A. Gas turbine stages of turbojets with devices for the air cooling of the turbine wheel disc
US5143517A (en) * 1990-08-08 1992-09-01 Societe Nationale D'etude Et De Construction De Moteurs D'aviation"S.N.E.M.C.A." Turbofan with dynamic vibration damping
US5222865A (en) * 1991-03-04 1993-06-29 General Electric Company Platform assembly for attaching rotor blades to a rotor disk
US5244345A (en) * 1991-01-15 1993-09-14 Rolls-Royce Plc Rotor
US5478207A (en) * 1994-09-19 1995-12-26 General Electric Company Stable blade vibration damper for gas turbine engine
US5791877A (en) * 1995-09-21 1998-08-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Damping disposition for rotor vanes
US7445433B2 (en) * 2004-02-24 2008-11-04 Rolls-Royce Plc Fan or compressor blisk
US7458769B2 (en) * 2005-07-21 2008-12-02 Snecma Device for damping vibration of a ring for axially retaining turbomachine fan blades
DE102007037208A1 (de) * 2007-08-07 2009-02-19 Mtu Aero Engines Gmbh Turbinenschaufel mit zumindest einer Einsatzhülse zum Kühlen der Turbinenschaufel
US7931442B1 (en) * 2007-05-31 2011-04-26 Florida Turbine Technologies, Inc. Rotor blade assembly with de-coupled composite platform
US8070448B2 (en) * 2008-10-30 2011-12-06 Honeywell International Inc. Spacers and turbines

Family Cites Families (10)

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Publication number Priority date Publication date Assignee Title
BE791375A (fr) * 1971-12-02 1973-03-01 Gen Electric Deflecteur et amortisseur pour ailettes de turbomachines
GB1457417A (en) * 1973-06-30 1976-12-01 Dunlop Ltd Vibration damping means
GB1549152A (en) * 1977-01-11 1979-08-01 Rolls Royce Rotor stage for a gas trubine engine
US4422827A (en) * 1982-02-18 1983-12-27 United Technologies Corporation Blade root seal
US4580946A (en) * 1984-11-26 1986-04-08 General Electric Company Fan blade platform seal
US5562419A (en) * 1995-06-06 1996-10-08 General Electric Company Shrouded fan blisk
GB2344383B (en) * 1998-12-01 2002-06-26 Rolls Royce Plc A bladed rotor
DE10361882B4 (de) * 2003-12-19 2013-08-22 Rolls-Royce Deutschland Ltd & Co Kg Rotor für die Hochdruckturbine eines Flugtriebwerks
US7448844B1 (en) * 2005-08-16 2008-11-11 Florida Turbine Technologies, Inc. Blisk having partially cut blade attachment
US9133720B2 (en) * 2007-12-28 2015-09-15 United Technologies Corporation Integrally bladed rotor with slotted outer rim

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4175912A (en) * 1976-10-19 1979-11-27 Rolls-Royce Limited Axial flow gas turbine engine compressor
US4192633A (en) * 1977-12-28 1980-03-11 General Electric Company Counterweighted blade damper
US4457668A (en) * 1981-04-07 1984-07-03 S.N.E.C.M.A. Gas turbine stages of turbojets with devices for the air cooling of the turbine wheel disc
US5143517A (en) * 1990-08-08 1992-09-01 Societe Nationale D'etude Et De Construction De Moteurs D'aviation"S.N.E.M.C.A." Turbofan with dynamic vibration damping
US5244345A (en) * 1991-01-15 1993-09-14 Rolls-Royce Plc Rotor
US5222865A (en) * 1991-03-04 1993-06-29 General Electric Company Platform assembly for attaching rotor blades to a rotor disk
US5478207A (en) * 1994-09-19 1995-12-26 General Electric Company Stable blade vibration damper for gas turbine engine
US5791877A (en) * 1995-09-21 1998-08-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Damping disposition for rotor vanes
US7445433B2 (en) * 2004-02-24 2008-11-04 Rolls-Royce Plc Fan or compressor blisk
US7458769B2 (en) * 2005-07-21 2008-12-02 Snecma Device for damping vibration of a ring for axially retaining turbomachine fan blades
US7931442B1 (en) * 2007-05-31 2011-04-26 Florida Turbine Technologies, Inc. Rotor blade assembly with de-coupled composite platform
DE102007037208A1 (de) * 2007-08-07 2009-02-19 Mtu Aero Engines Gmbh Turbinenschaufel mit zumindest einer Einsatzhülse zum Kühlen der Turbinenschaufel
US8070448B2 (en) * 2008-10-30 2011-12-06 Honeywell International Inc. Spacers and turbines

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10344601B2 (en) 2012-08-17 2019-07-09 United Technologies Corporation Contoured flowpath surface
US20160032753A1 (en) * 2014-07-31 2016-02-04 United Technologies Corporation Gas turbine engine with axial compressor having improved air sealing
US10107127B2 (en) * 2014-07-31 2018-10-23 United Technologies Corporation Gas turbine engine with axial compressor having improved air sealing
US9810075B2 (en) 2015-03-20 2017-11-07 United Technologies Corporation Faceted turbine blade damper-seal
US10196896B2 (en) 2015-04-13 2019-02-05 Rolls-Royce Plc Rotor damper
US10385696B2 (en) 2015-04-13 2019-08-20 Rolls-Royce Plc Rotor damper
US10443502B2 (en) 2015-04-13 2019-10-15 Rolls-Royce Plc Rotor damper

Also Published As

Publication number Publication date
WO2010099781A2 (fr) 2010-09-10
EP2350459A2 (fr) 2011-08-03
DE102009011964A1 (de) 2010-09-09
WO2010099781A3 (fr) 2011-05-19

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Legal Events

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AS Assignment

Owner name: MTU AERO ENGINES GMBH, A COMPANY OF GERMANY, GERMA

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:STIEHLER, FRANK;REEL/FRAME:026772/0542

Effective date: 20110809

STCB Information on status: application discontinuation

Free format text: ABANDONED -- AFTER EXAMINER'S ANSWER OR BOARD OF APPEALS DECISION