EP2090748A2 - Deckbandanordnung für eine Turbomaschine - Google Patents

Deckbandanordnung für eine Turbomaschine Download PDF

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Publication number
EP2090748A2
EP2090748A2 EP09152448A EP09152448A EP2090748A2 EP 2090748 A2 EP2090748 A2 EP 2090748A2 EP 09152448 A EP09152448 A EP 09152448A EP 09152448 A EP09152448 A EP 09152448A EP 2090748 A2 EP2090748 A2 EP 2090748A2
Authority
EP
European Patent Office
Prior art keywords
shroud
interface member
hard face
face interface
recess
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.)
Withdrawn
Application number
EP09152448A
Other languages
English (en)
French (fr)
Other versions
EP2090748A3 (de
Inventor
Kevin Leon Bruce
Ronald Ralph Cairo
Richard Gordon Rollings
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
Publication of EP2090748A2 publication Critical patent/EP2090748A2/de
Publication of EP2090748A3 publication Critical patent/EP2090748A3/de
Withdrawn legal-status Critical Current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
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding

Definitions

  • the present invention relates to the art of turbomachines and, more particularly, to a rotating assembly for a turbomachine.
  • Turbomachines employ a number of rotating components or assemblies.
  • Turbines for example, employ compressor stages and turbine stages that rotate at high speed when the turbine is in operation.
  • a stage includes a plurality of free-floating blades that extend radially outward from a central hub.
  • Some blades include a shroud that limits vibration within a stage.
  • the shroud is typically positioned at a tip portion of the blade, a mid portion of the blade or at both the mid portion and the tip portion of the blade.
  • the shrouds are designed such that at high or operational speeds, the free-floating blades interlock to form an integral rotating member. At lower speeds, such as on turbine turning gear, the blades do not interlock and will often times impact one another. Impacts between the blades can cause damage that will shorten service life of the turbomachine.
  • a hard face coating is applied to potential contact points.
  • the hard face coating increases wear resistance of the potential contact points to increase both impact resistance and durability.
  • the hard face coating is metallurgically bonded to the blade through, for example, a welding or brazing process.
  • Using a welding process to bond the hard face interface to the blade inherently produces a great deal of localized heat which, if not properly controlled, can weaken the impact resistance and other metallurgical properties at the interface of the materials being joined. Excessive heat can also cause cracking in adjacent material during manufacture.
  • a Z-notch radius portion of the blade The z-notch radius portion is subjected to high stresses and therefore subject to cracking causing a release of material.
  • An additional failure mechanism can occur during high speed operation when tension and/or shear forces develop that over-stress the metallurgical bond. Over time, the metallurgical bond can fail and the hard face coating is released from the blade becoming foreign object debris (FOD) in the turbomachine. FOD flying around the turbomachine can damage the rotating components as well as inner surfaces of the turbine and lead to engine failure.
  • FOD foreign object debris
  • a rotating assembly for a turbomachine includes a central hub and a plurality of rotating members that extend radially outward from a central hub.
  • Each of the plurality of rotating members includes a base portion detachably mounted to the central hub, a tip portion, and a mid-span portion that extends between the base portion and the tip portion.
  • Each of the plurality of rotating members also includes a shroud portion that is positioned at one of the tip portion and the mid-span portion.
  • the shroud portion includes a pressure side and a suction side.
  • At least one hard face interface member is secured to at least one of the suction side and the pressure side of the shroud portion.
  • the at least one hard face interface member is both mechanically interlocked with, and metallurgically bonded to, the shroud portion.
  • a method of securing a hard face interface member to a shroud portion of a rotating member for a turbomachine includes forming a rotating member having a base portion, a mid-span portion and a tip portion.
  • a shroud portion is formed on at least one of the mid-span portion and the tip portion of the rotating member.
  • the shroud portion includes a pressure side and a suction side.
  • a hard face interface member is secured with both a mechanical interlock and a metallurgical bond to at least one of the pressure side and the suction side of the shroud portion.
  • a turbomachine shown in the form of a gas turbine engine, constructed in accordance with an exemplary embodiment of the present invention is indicated generally at 2.
  • Engine 2 includes a compressor 4 and a plurality of combustor assemblies arranged in a can annular array, one of which is indicated at 8.
  • combustor assembly 8 includes an endcover assembly 9 that seals, and at least partially defines, a combustion chamber 12.
  • a plurality of nozzles 14-16 are supported by endcover assembly 9 and extend into combustion chamber 12. Nozzles 14-16 receive fuel through a common fuel inlet (not shown) and compressed air from compressor 4.
  • Turbine 30 includes a plurality of rotating assemblies or stages 31-33 that are operationally connected to compressor 4 through a compressor/turbine shaft 34 (sometimes referred to as a rotor).
  • the high pressure gas is supplied to combustor assembly 8 and mixed with fuel, for example process gas and/or synthetic gas (syngas), in combustion chamber 12.
  • fuel for example process gas and/or synthetic gas (syngas)
  • the fuel/air or combustible mixture ignited to form a high pressure, high temperature combustion gas stream of approximately 538 degrees Celsius (°C) to 1593 °C (1000 degrees Fahrenheit (°F) to 2900 °F).
  • combustor assembly 8 can combust fuels that include, but are not limited, to natural gas and/or fuel oil.
  • combustor assembly 8 channels the combustion gas stream to turbine 30 which coverts thermal energy to mechanical, rotational energy.
  • stage 31 is shown to include a plurality of rotating members, one of which is indicated at 46, which extend radially outward from a central hub 47.
  • rotating member 46 includes a base portion 48, a tip portion 49 and a mid-span portion 50 that extends between base portion 48 and tip portion 49.
  • rotating member 46 includes a first or mid-span shroud 60 having a first or suction side 63 and a second or pressure side 64.
  • Rotating member 46 is also shown to include a second or tip shroud 70 having a first or suction side 72 and a second or pressure side 73.
  • tip shroud 70 includes first and second opposing wing members 78 and 79 and a third wing member 80.
  • Third wing member 80 extends perpendicularly relative to first and second wing members 78 and 79.
  • Tip shroud 70 covers a bucket or throat portion (not separately labeled) of rotating member 46.
  • Tip shroud 70 is designed to receive, or nest with, tip shrouds on adjacent rotating members in order to form a continuous ring 81 that extends circumferentially about stage 31.
  • Continuous ring 81 creates an outer flow path boundary that reduces gas path air leakage over top portions (not separately labeled) of stage 31 so as to increase stage efficiency and overall turbine performance.
  • mid-span shroud 60 is configured to receive, or nest within, mid-span shrouds on adjacent rotating members to form an inner ring indicated generally at 82 that further increases stage efficiency.
  • each rotating member 46 could alternatively be provided with a single shroud positioned either at tip portion 49 or mid-span portion 50.
  • adjacent rotating members 46 interlock through mid-span shroud 60 and tip-shroud 70 by virtue of centrifugal forces created by the operation of turbine 30.
  • the rotational force is not sufficient to establish the interlock and thus, often times adjacent rotating members impact one another. The impacts can create wear on the rotating members thereby lowering an overall service life of turbine 30.
  • each mid-span shroud 60 and tip-shroud 70 is provided with a wear resistant/impact resistant member in a manner that will be described more fully below.
  • suction side 72 includes a first surface 90 and an opposing second surface 91.
  • Second surface 91 includes a recess or cavity 94 having a base portion 96 and a peripheral wall portion 98.
  • a hard face interface member 100 that in accordance with an exemplary embodiment of the present invention is formed from a pre-sintered, preformed (PSP) material, is positioned within cavity 94.
  • PSP pre-sintered, preformed
  • Hard face interface member 100 is mechanically interlocked with, and metallurgically bonded to, suction side 72 to provide an impact and wear resistant surface for tip shroud 70.
  • hard face interface member 100 is positioned within cavity 94 and brazed to tip shroud 70.
  • cavity 94 provides a mechanical interlock, and brazing provides a metallurgical bond to tip shroud 70.
  • the mechanical interlock at an outer radial position or z-notch radius portion of hard face interface member 100 establishes a secondary bond that resists compressive or tensile (depending of the relative position on tip-shroud 70) forces developed during the operation of stage 31.
  • cavity 94 is provided with a plurality of surface elements, which, in the exemplary embodiment shown, take the form of dimples or indentations 105 formed on a base portion 96. Indentations 105 increase an overall surface area of cavity 94 so as to establish a more robust mechanical interlock for hard face interface member 100.
  • hard face interface member 100 is machined so as to be substantially flush with second surface 91 to provide an overall finish for tip-shroud 70 that reduces localized airflow turbulences.
  • pressure side 73 includes a first surface 110 and an opposing second surface 111.
  • Second surface 111 includes a recess or cavity 114 having a base portion 116 and a peripheral wall portion 118.
  • a hard face interface member 120 is positioned within cavity 114.
  • Hard face interface member 120 in accordance with an exemplary embodiment of the present invention, is formed from a PSP material and is mechanically interlocked with, and metallurgically bonded to, pressure side 73 to provide another impact and wear resistant surface for tip shroud 70.
  • hard face interface member 120 is positioned within cavity 114 and brazed to tip shroud 70 to establish a metallurgical bond.
  • cavity 114 provides a mechanical interlock
  • brazing provides a metallurgical bond to tip shroud 70.
  • Recess 114 also includes a plurality of surface elements, which in the exemplary embodiment shown, take the form of dimples or indentions 124 formed on base portion 116. Indentations 124 increase an overall surface area of base portion 116 to increase the overall strength of the mechanical interlock.
  • hard face interface member 120 is machined so as to the substantially flush with second surface 111 to provide an overall finish for tip-shroud 70 that reduces localized airflow turbulences.
  • mid-span shroud 60 includes a first hard face interface member 130 positioned within a recess (not separately labeled) formed on suction side 63 and a second hard face interface member 140 provided within a recess (not separately labeled) provided on pressure side 64.
  • hard face interface members 100, 120, 130 and 140 provide localized wear/impact resistant surface enhancements that remain secured to rotating member 46 during the operation of stage 31. That is, the combination of the mechanical link or interlock and the metallurgical bond provides a fail-safe multiple-load path, attachment/retention mechanism that is resistant to multiple directional loading for hard face interface member 100.
  • FIG. 8 illustrates a plurality of continuous protuberances 200 arranged on base portion 96.
  • FIG. 9 illustrates a plurality of continuous grooves 210 formed in base portion 96.
  • FIG. 10 depicts base portion 96 with a plurality of discontinuous grooves 220 while FIG.11 shows base portion 96 with a plurality of discontinuous or segmented protuberances 230.
  • various elements can be employed to enhance the mechanical bond between the hard face interface member and the blade.
  • Tip and/or mid span shroud 70 and 60 can include a machined cavity or recess 300 having a radius portion 310 and a plurality of continuous protuberances 320 such as shown in FIG. 12 .
  • Recess 300 is formed by removing for example, approximately 1/10 " (2.54 mm) of a face portion of suction side 72.
  • FIG. 13 illustrates a cavity or recess 360 having a radius portion 370 formed on a face portion of suction side 72.
  • Recess 360 includes a plurality of discontinuous protuberances 380 for mechanically interlocking a hard face interface member.
  • FIG. 14 illustrates a cavity or recess 400 having a radius portion 410 formed in suction side 72.
  • Recess 400 includes a plurality of continuous grooves 420 for mechanically interlocking the hard face interface member.
  • FIG. 15 illustrates a cavity or recess 460 having a radius 470 formed in suction side 72.
  • Recess 460 includes a plurality of discontinuous grooves 480 for mechanically interlocking the hard face interface member.
  • the particular hard face interface member would include structure corresponding to the protuberances or grooves to facilitate the mechanical interlock.
EP09152448A 2008-02-13 2009-02-10 Deckbandanordnung für eine Turbomaschine Withdrawn EP2090748A3 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/030,317 US20090202344A1 (en) 2008-02-13 2008-02-13 Rotating assembly for a turbomachine

Publications (2)

Publication Number Publication Date
EP2090748A2 true EP2090748A2 (de) 2009-08-19
EP2090748A3 EP2090748A3 (de) 2010-05-19

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP09152448A Withdrawn EP2090748A3 (de) 2008-02-13 2009-02-10 Deckbandanordnung für eine Turbomaschine

Country Status (2)

Country Link
US (1) US20090202344A1 (de)
EP (1) EP2090748A3 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3085419A1 (fr) * 2018-09-05 2020-03-06 Safran Aircraft Engines Aube mobile

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2985759B1 (fr) 2012-01-17 2014-03-07 Snecma Aube mobile de turbomachine
US20130202439A1 (en) * 2012-02-08 2013-08-08 General Electric Company Rotating assembly for a turbine assembly
US10215032B2 (en) 2012-10-29 2019-02-26 General Electric Company Blade having a hollow part span shroud
US9328619B2 (en) * 2012-10-29 2016-05-03 General Electric Company Blade having a hollow part span shroud
FR3001758B1 (fr) * 2013-02-01 2016-07-15 Snecma Aube de rotor de turbomachine
DE102016214234A1 (de) 2016-08-02 2018-02-08 MTU Aero Engines AG Laufschaufel mit Impulskörper
JP7434199B2 (ja) * 2021-03-08 2024-02-20 株式会社東芝 タービン動翼

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5083903A (en) 1990-07-31 1992-01-28 General Electric Company Shroud insert for turbomachinery blade
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

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4986737A (en) * 1988-12-29 1991-01-22 General Electric Company Damped gas turbine engine airfoil row
US6733907B2 (en) * 1998-03-27 2004-05-11 Siemens Westinghouse Power Corporation Hybrid ceramic material composed of insulating and structural ceramic layers
US6568908B2 (en) * 2000-02-11 2003-05-27 Hitachi, Ltd. Steam turbine
DE10342207A1 (de) * 2003-09-12 2005-04-07 Alstom Technology Ltd Laufschaufelbindung einer Turbomaschine
US7360991B2 (en) * 2004-06-09 2008-04-22 General Electric Company Methods and apparatus for fabricating gas turbine engines
US7771171B2 (en) * 2006-12-14 2010-08-10 General Electric Company Systems for preventing wear on turbine blade tip shrouds

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5083903A (en) 1990-07-31 1992-01-28 General Electric Company Shroud insert for turbomachinery blade
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

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3085419A1 (fr) * 2018-09-05 2020-03-06 Safran Aircraft Engines Aube mobile
WO2020049252A1 (fr) * 2018-09-05 2020-03-12 Safran Aircraft Engines Aube mobile

Also Published As

Publication number Publication date
US20090202344A1 (en) 2009-08-13
EP2090748A3 (de) 2010-05-19

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