EP2294286B1 - Rotor avex aubes mobiles carenées d'une turbomachine - Google Patents

Rotor avex aubes mobiles carenées d'une turbomachine Download PDF

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Publication number
EP2294286B1
EP2294286B1 EP09745441.7A EP09745441A EP2294286B1 EP 2294286 B1 EP2294286 B1 EP 2294286B1 EP 09745441 A EP09745441 A EP 09745441A EP 2294286 B1 EP2294286 B1 EP 2294286B1
Authority
EP
European Patent Office
Prior art keywords
rotor
shroud
gaps
turbomachine
damping
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.)
Not-in-force
Application number
EP09745441.7A
Other languages
German (de)
English (en)
Other versions
EP2294286A2 (fr
Inventor
Hans-Peter Borufka
Hernan Victor Arrieta
Klemens Hain
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 AG
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 AG filed Critical MTU Aero Engines AG
Publication of EP2294286A2 publication Critical patent/EP2294286A2/fr
Application granted granted Critical
Publication of EP2294286B1 publication Critical patent/EP2294286B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/183—Two-dimensional patterned zigzag
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/96—Preventing, counteracting or reducing vibration or noise

Definitions

  • the present invention relates to a shroud for rotor blades of a turbomachine, in particular a gas turbine, wherein the shroud is arranged on the circumference of a blade row arranged on a rotor with a plurality of blades and at its periphery has at least one separating gap.
  • the invention further relates to a turbomachine, in particular gas turbine, comprising at least one rotor having at least one blade row with a plurality of blades.
  • the sealing gap between rotating blades and the stationary engine casing is an influencing variable, which is of considerable importance for the efficiency of the engine.
  • a shroud which is arranged on the blade tips of the blades.
  • From the DE 40 15 206 C1 is a shroud for an integral wheel with at least one circumferentially arranged Z-shaped separating gap with angularly arranged to the axial direction damping columns smallest gap width and adjoining open gap sections known.
  • the mutually parallel damping gaps form the two legs of the Z-shaped separating gap and are aligned at an angle of 70 ° to 90 ° to the axial direction of the integral wheel.
  • the damping gaps are close together, while the Z-web is formed as an open gap extending in the direction of the shroud edges.
  • the two damping gaps provide a friction surface which allows a corresponding friction damping of the vibrations during operation of the integral wheel.
  • Shrouds with serrated separating gaps are from the EP0536575A1 , the EP0528138A1 and the FR1519898A known.
  • a shroud for moving blades of a turbomachine, in particular a gas turbine, of a rotor according to the invention is arranged on the circumference of a rotor blade row arranged on a rotor with a plurality of rotor blades and has at least one separating gap on its circumference.
  • the separating gap is serrated and has at least three mutually spaced and angularly extending to an axis of rotation of the rotor damping gaps and adjoining the damping column respectively connecting or extending in the direction of the shroud edges connecting gaps, wherein the gap width of the damping column upon rotation of the rotor up to a juxtaposition the partition walls forming the attenuation gaps are reduced.
  • the total available contact surface for friction damping is significantly increased under operating conditions of the turbomachine. This ensures a predominantly resonance-free operation of bladed turbomachines and also an improvement in the coupling stiffness of adjacent blades to one another. There is a continuous distribution of the power flow along the damping column or the contact points of the corresponding gap walls.
  • the resonance-free main operating range results from the control of the fundamental natural frequencies of the bladed disc (overall system blade / rotor) by the possibility of an individual design of the support and stress kinematics of the formed separating gaps.
  • the damping gaps are aligned at an angle of 60 ° to 90 ° to the axis of rotation of the rotor.
  • this ensures that when the turbomachine is started up, the circumferential expansion of the shroud resulting from the centrifugal forces can take place unhindered.
  • the shroud is divided into individual shroud segments, wherein each shroud segment associated with a blade and arranged thereon and form the individual shroud segments with the respective adjacent shroud segments in the circumferential direction of the separation column.
  • each shroud segment of a group of at least two blades is assigned and arranged thereon and form the individual shroud segments with the respective adjacent shroud segments in the circumferential direction of the separation column.
  • the blades are integrally formed with the shroud segments.
  • the shroud according to the invention may have different advantageous embodiments and arrangements.
  • the subdivision of the shroud into shroud segments increases the range of applications.
  • the blades with the shroud segments integral, that is to form in one piece. This leads to a simplified manufacturing process and thus to reduced manufacturing costs.
  • At least one sealing lip is arranged on the outer circumference of the shroud.
  • two spaced-apart and mutually parallel sealing lips may be integrally formed on the outer circumference.
  • the sealing lips may be interrupted in the region of the separating gaps. The arrangement of the sealing lips is a further advantageous reduction of the sealing gap between the rotating blades or the shroud and the stationary engine housing, the efficiency of the turbomachine, in particular the gas turbine is significantly improved.
  • the rotor according to the invention is used in a low-pressure turbine, in particular a low-pressure turbine of an aircraft engine.
  • a turbomachine according to the invention in particular gas turbine, comprises at least one rotor which has at least one blade row with a plurality of rotor blades, wherein a shroud according to the above-described embodiments is arranged on the circumference of the rotor blade row. Due to the design of the shroud, the turbomachine according to the invention ensures predominantly resonance-free operation and an improvement in the coupling rigidity of adjacent blades to one another. This results in a significant increase in the efficiency of the turbomachine.
  • the turbomachine may be a low-pressure turbine, in particular a low-pressure turbine of an aircraft engine.
  • the blades may include components of an integral rotor construction, i. BLISK or BLING, his.
  • FIG. 1 shows a schematic representation of a portion of a turbomachine consisting of a blade 14 with a blade root 52, wherein the blade root 52 is disposed on a rotor 12.
  • the rotor 12 is rotatable about an axis 18.
  • a shroud 10 is arranged on the rotor blade 14.
  • the shroud 10 is arranged on the circumference of a arranged on the rotor 12 blade row consisting of several blades.
  • FIG. 2 1 shows a schematic representation of a plan view of the shroud 10. It can be seen that the shroud 10 has a plurality of separating gaps 16, 16 ', 16 "on its circumference, wherein the separating gaps 16, 16", 16 "are formed in a serrated shape the separating gaps 16, 16 ', 16 "each consist of five spaced apart and at an angle to the axis of rotation 18 of the rotor 12 extending damping columns 20, 22, 24, 26, 28 and to this subsequent, the damping column 20, 22, 24, 26th , 28 in each case connecting or in the direction of the shroud edges 54, 56 extending connecting gaps 30, 32, 34, 36, 38, 40.
  • the individual shroud segments 46, 48 form with the respective adjacent shroud segments in the circumferential direction the separating gaps 16, 16 ', 16 " Separation gaps 16, 16 ', 16 "formed parallel to each other, that is, the individual damping and connecting gaps each extend parallel to each other.
  • damping gaps 20, 22, 24, 26, 28 are aligned at an angle of 60 ° to 90 ° to the axis of rotation 18 of the rotor 12.
  • a total of four damping gaps 20, 22, 24, 26 are aligned parallel to each other in the illustrated embodiment.
  • a further damping gap 28 extends at an acute angle to the previously described damping gaps 20, 22, 24, 26.
  • connection gaps 30, 32, 34, 36, 38, 40 have different angles in a range between 0 ° and 90 ° ° with respect to the axis of rotation 18 of the rotor can take.
  • two mutually parallel sealing lips 62, 64 are integrally formed on the outer circumference of the shroud 10.
  • the sealing lips are interrupted in the area of the separating gaps 16, 16 ', 16 ".
  • the sealing lips 62, 64 result in a further advantageous reduction of the sealing gap between the shroud 10 and a fixed housing of the turbomachine adjoining it, in particular a fixed engine housing (US Pat. not shown).

Landscapes

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

Claims (10)

  1. Rotor (12) destiné à une turbomachine, notamment une turbine à gaz, avec une rangée d'aubes mobiles disposée sur le rotor (12) avec plusieurs aubes mobiles (14, 50) et un carénage, le carénage (10) étant disposé sur le pourtour de la rangée d'aubes mobiles et présentant sur son pourtour au moins une fente de séparation (16, 16', 16"), la fente de séparation (16, 16', 16") est conçue en dents de scie et présente au moins trois fentes d'amortissement (20, 22, 24, 26, 28) espacées les unes des autres et s'étendant avec un angle par rapport à un axe de rotation (18) du rotor (12) et des fentes de connexion (30, 32, 34, 36, 38, 40) leur faisant suite, reliant respectivement les fentes d'amortissement (20, 22, 24, 26, 28) ou les prolongeant en direction des bords du carénage (54, 56), caractérisé en ce que la largeur de fente de la fente d'amortissement (20, 22, 24, 26, 28) est diminuée lors de la rotation du rotor (12) jusqu'à une position adjacente des parois de fente (42, 44) formant la fente d'amortissement (20, 22, 24, 26, 28) et les fentes d'amortissement (20, 22, 24, 26, 28) sont orientées avec un angle de 60° à 90° par rapport à l'axe de rotation (18) du rotor (12).
  2. Rotor (12) selon la revendication 1, caractérisé en ce qu'au moins deux des fentes d'amortissement (20, 22, 24, 26, 28) sont orientées parallèlement l'une à l'autre.
  3. Rotor (12) selon l'une des revendications précédentes, caractérisé en ce que le carénage (10) est subdivisé en segments de carénage (46, 48) individuels, où chaque segment de carénage (46, 48) est associé à une aube mobile (14, 50) et y est fixé et les segments de carénage (46, 48) individuels forment la fente de séparation (16, 16', 16") avec les segments de carénage respectivement voisins dans la direction circonférentielle.
  4. Rotor (12) selon la revendication 1 ou 2, caractérisé en ce que le carénage (10) est subdivisé en segments de carénage individuels, où chaque segment de carénage est associé à un groupe d'au moins deux aubes mobiles et y est fixé et les segments de carénage individuels forment la fente de séparation avec les segments de carénage respectivement voisins dans la direction circonférentielle.
  5. Rotor (12) selon la revendication 3 ou 4, caractérisé en ce que les aubes mobiles (14, 50) sont formées intégralement avec les segments de carénage (46, 48).
  6. Rotor (12) selon l'une des revendications précédentes, caractérisé en ce qu'au moins une lèvre d'étanchéité (62, 64) est disposée sur le pourtour extérieur du carénage (10).
  7. Rotor (12) selon l'une des revendications précédentes, caractérisé en ce que la turbomachine est une turbine basse pression, notamment une turbine basse pression d'un moteur d'avion.
  8. Turbomachine, notamment turbine à gaz, comprenant au moins un rotor (12) selon l'une des revendications 1 à 6.
  9. Turbomachine selon la revendication 8, caractérisée en ce que la turbomachine est une turbine basse pression, notamment une turbine basse pression d'un moteur d'avion.
  10. Turbomachine selon la revendication 8 ou 9, caractérisée en ce que les aubes mobiles (14, 50) sont des composantes d'une construction de rotor intégrale.
EP09745441.7A 2008-05-13 2009-04-30 Rotor avex aubes mobiles carenées d'une turbomachine Not-in-force EP2294286B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102008023326A DE102008023326A1 (de) 2008-05-13 2008-05-13 Deckband für Laufschaufeln einer Strömungsmaschine und Strömungsmaschine
PCT/DE2009/000630 WO2009138057A2 (fr) 2008-05-13 2009-04-30 Carénage pour aubes mobiles d'une turbomachine, et turbomachine correspondante

Publications (2)

Publication Number Publication Date
EP2294286A2 EP2294286A2 (fr) 2011-03-16
EP2294286B1 true EP2294286B1 (fr) 2019-02-27

Family

ID=41180361

Family Applications (1)

Application Number Title Priority Date Filing Date
EP09745441.7A Not-in-force EP2294286B1 (fr) 2008-05-13 2009-04-30 Rotor avex aubes mobiles carenées d'une turbomachine

Country Status (5)

Country Link
US (1) US8573939B2 (fr)
EP (1) EP2294286B1 (fr)
DE (1) DE102008023326A1 (fr)
ES (1) ES2715798T3 (fr)
WO (1) WO2009138057A2 (fr)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2551460A1 (fr) * 2011-07-29 2013-01-30 Siemens Aktiengesellschaft Groupe d'aubes
US9840917B2 (en) * 2011-12-13 2017-12-12 United Technologies Corporation Stator vane shroud having an offset
US8894368B2 (en) * 2012-01-04 2014-11-25 General Electric Company Device and method for aligning tip shrouds
US20150354374A1 (en) * 2014-06-09 2015-12-10 General Electric Company Turbine blisk and method of manufacturing thereof
DE102016222720A1 (de) * 2016-11-18 2018-05-24 MTU Aero Engines AG Dichtungssystem für eine axiale Strömungsmaschine und axiale Strömungsmaschine
US10519783B2 (en) 2016-12-22 2019-12-31 General Electric Company Method for modifying a shroud and blade
KR101874243B1 (ko) * 2017-03-31 2018-07-03 두산중공업 주식회사 버킷의 진동감쇠구조와 이를 포함하는 버킷 및 터보머신
FR3084399B1 (fr) * 2018-07-24 2021-05-14 Safran Aircraft Engines Aube mobile pour une roue d'une turbomachine
CN110701102A (zh) * 2019-09-29 2020-01-17 中国航发沈阳发动机研究所 一种风扇转子叶片结构极具有其的发动机风扇转子

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Publication number Priority date Publication date Assignee Title
FR1519898A (fr) * 1967-02-24 1968-04-05 Creusot Forges Ateliers Perfectionnements aux aubages torses mobiles de turbo-machines
US4710102A (en) * 1984-11-05 1987-12-01 Ortolano Ralph J Connected turbine shrouding
GB2223276B (en) * 1988-09-30 1992-09-02 Rolls Royce Plc Turbine aerofoil blade
DE4015206C1 (fr) 1990-05-11 1991-10-17 Mtu Muenchen Gmbh
DE59202211D1 (de) * 1991-08-08 1995-06-22 Asea Brown Boveri Deckblatt für axialdurchströmte Turbine.
DE59201833D1 (de) * 1991-10-08 1995-05-11 Asea Brown Boveri Deckband für axialdurchströmte Turbine.
JPH07332003A (ja) * 1994-06-13 1995-12-19 Hitachi Ltd タービン動翼
US6164916A (en) * 1998-11-02 2000-12-26 General Electric Company Method of applying wear-resistant materials to turbine blades, and turbine blades having wear-resistant materials
GB0218060D0 (en) * 2002-08-03 2002-09-11 Alstom Switzerland Ltd Sealing arrangements
US7001152B2 (en) * 2003-10-09 2006-02-21 Pratt & Wiley Canada Corp. Shrouded turbine blades with locally increased contact faces
JP5124276B2 (ja) * 2004-10-07 2013-01-23 ボルボ エアロ コーポレイション ガスタービン中間構造および該中間構造を含むガスタービンエンジン
US7762779B2 (en) * 2006-08-03 2010-07-27 General Electric Company Turbine blade tip shroud

Non-Patent Citations (1)

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Title
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Also Published As

Publication number Publication date
US8573939B2 (en) 2013-11-05
DE102008023326A1 (de) 2009-11-19
WO2009138057A3 (fr) 2010-08-26
WO2009138057A2 (fr) 2009-11-19
ES2715798T3 (es) 2019-06-06
US20110103956A1 (en) 2011-05-05
EP2294286A2 (fr) 2011-03-16

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