US20110311355A1 - Axial turbo compressor for a gas turbine having low radial gap losses and diffuser losses - Google Patents

Axial turbo compressor for a gas turbine having low radial gap losses and diffuser losses Download PDF

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Publication number
US20110311355A1
US20110311355A1 US13/201,065 US201013201065A US2011311355A1 US 20110311355 A1 US20110311355 A1 US 20110311355A1 US 201013201065 A US201013201065 A US 201013201065A US 2011311355 A1 US2011311355 A1 US 2011311355A1
Authority
US
United States
Prior art keywords
vane
axial
shaft cover
turbocompressor
tips
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
US13/201,065
Other languages
English (en)
Inventor
Francois Benkler
Karl Klein
Torsten Matthias
Achim Schirrmacher
Oliver Schneider
Vadim Shevchenko
Ulrich Waltke
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BENKLER, FRANCOIS, KLEIN, KARL, MATTHIAS, TORSTEN, SCHEVCHENKO, VADIM, SCHIRRMACHER, ARCHIM, SCHNEIDER, OLIVER, WALTKE, ULRICH
Publication of US20110311355A1 publication Critical patent/US20110311355A1/en
Abandoned legal-status Critical Current

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Classifications

    • 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
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/545—Ducts
    • F04D29/547—Ducts having a special shape in order to influence fluid flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
    • 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
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • 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
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • 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
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers

Definitions

  • the invention refers to an axial turbocompressor for a gas turbine, wherein the axial turbocompressor has low radial gap losses.
  • a gas turbine has a turbocompressor, for example in an axial type of construction.
  • the turbocompressor has a casing with a stator attached thereupon, and a rotor which is enclosed by the casing.
  • the rotor has a shaft on which the rotor can be rotationally driven. Provision is made for a shaft cover, encompassing the shaft, the outer contour of which together with the inner contour of the casing form a part of the flow passage through the turbocompressor.
  • the flow passage has a cross section which widens in the flow direction so that the flow passage is formed as a diffuser.
  • the rotor has a multiplicity of rotor stages which are formed in each case by a rotor blade row.
  • the stator has a multiplicity of stator blade rows which, as seen in the axial direction, are arranged in a manner in which they alternate with the rotor blade rows.
  • a guide vane row is customarily arranged after the last rotor blade row and a downstream guide vane row arranged after that.
  • the guide vane rows have a multiplicity of vanes which by their one end are fastened in each case on the casing and by their other end point towards the shaft.
  • a vane tip which faces the shaft cover and is arranged directly adjacent thereto, is formed on the other end of the guide vane.
  • the distance between the vane tips and the shaft cover is formed as a radial gap which is dimensioned in such a way that on the one hand the vane tips do not butt against the shaft cover during operation of the gas turbine and on the other hand the leakage flow through the radial gap which ensues during operation of the gas turbine is as low as possible.
  • This radial gap is therefore to be designed as small as possible so that high efficiency is achieved and the full blading potential of the compressor can be exploited.
  • the casing of the turbocompressor is solidly constructed in order to be able to withstand the pressure stresses and temperature stresses during operation of the gas turbine. Also, the casing is of a rigid construction so that load transfer onto the casing during operation of the gas turbine results in only minor deformation of the casing. In contrast to this, the shaft cover is subjected to lower mechanical stresses during operation of the gas turbine, as a result of which the shaft cover is of a thinner and less solid construction than the casing.
  • the shaft cover heats up more quickly than the casing with the guide vane rows fastened thereupon. This has the result that for starting and shutting down the gas turbine the shaft cover and the casing have a different rate of thermal expansion so that during starting and shutting down of the gas turbine the depth of the radial gap alters, wherein the radial gap is temporarily smaller during starting and larger during shutting down.
  • the radial gap is provided with a minimum depth which is dimensioned in such a way that in each operating state of the gas turbine—steady state as well as transient—the vane tips seldom if ever come into contact with the shaft cover. This has the result that a correspondingly dimensioned radial gap is provided at the vane tips which leads to a reduction of the efficiency of the gas turbine.
  • the blockage which is created by the radial gap leads to a reduction of the main flow components, as a result of which the pressure recovery in the diffuser is reduced and disadvantageous separation phenomena can occur.
  • the axial turbocompressor according to the invention for a gas turbine has a guide vane cascade, which is formed by guide vanes with vane tips which are free standing on the hub side, and a stationary shaft cover which is arranged directly adjacent to the vane tips on the hub side and delimits the flow passage of the axial compressor, wherein between the shaft cover and the vane tips a radial gap is formed and is minimally dimensioned in such a way that the assembly of the axial turbocompressor can only just be accomplished, and in the shaft cover provision is made for a multiplicity of blind hole-like recesses, wherein one of the recesses is associated with each vane tip and arranged directly adjacent to the vane tip which is associated therewith, and dimensioned in such a way that during operation of the axial turbocompressor each vane tip can sink into its associated recess without one of the vane tips coming into significant contact with the shaft cover.
  • the radial gap between the vane tip and the shaft cover is adjusted to the minimum required assembly gap so that the depth of the radial gap is reduced to the assembly-dependent minimum.
  • the depth of the minimum required assembly gap is selected in such a way that the rolling in of the guide vane cascade, especially of the rear guide vane cascade, can be accomplished.
  • the radial gap between the vane tips and the shaft cover is provided with a minimum required depth which is selected in such a way that in practically all conceivable operating states of the gas turbine the vane tips scarcely come into contact with, or do not make contact with, the shaft cover. Consequently, the radial gap is created with such depth that an appreciable mass flow of leakage flow flows through the radial gap, which leads to an undesirable lowering of efficiency of the gas turbine.
  • the radial gap is adjusted to the minimum possible radial gap, specifically to the minimum required assembly gap, so that the leakage flow through the radial gap is minimal.
  • the axial turbocompressor has a high pressure recovery in the diffuser section and therefore high efficiency.
  • the vane tips can sink into the recesses during operation of the axial turbocompressor so that although the radial gap is reduced to the minimum required assembly gap, a damaging contact of the vane tips with the shaft cover during operation of the axial turbocompressor is prevented.
  • the vane tip sinks into its associated recess during a specific operating state, then the flow around the vane tip decreases, as a result of which the leakage flow at the vane tip also decreases. Consequently, the efficiency of the guide vane cascade increases and losses and also separations in the diffuser which lies downstream of the axial turbocompressor are reduced. In all, a good overall machine performance and high overall machine efficiency of the gas turbine result from the improved radial gap behavior.
  • the divergence degree of the diffuser i.e. the diffuser angle of the diffuser, can be selected larger than would be the case with a conventional diffuser. A reduction of the overall length of the gas turbine compared with a conventional gas turbine is associated with this.
  • honeycomb-like and/or felt-like structure which can yield during contact by the vane tip, is applied to the base of the recess.
  • the honeycomb-like structure is preferably a honeycomb.
  • the vane tip can sink into the honeycomb-like and/or felt-like structure, wherein the vane tip is not damaged. Resulting from this is the advantage that the distance between the vane tip and the honeycomb-like and/or felt-like structure is designed to be small. Therefore, the flow around the vane tip decreases if the vane tip sinks into its associated recess during a specific operating state and digs into the honeycomb-like and/or felt-like structure. As a result, the leakage flow at the vane tip advantageously additionally decreases.
  • the recesses have an outline shape on the surface of the shaft cover which is adapted to the profile of the guide vanes associated therewith at the vane tip, and have a prespecified depth.
  • the material of the shaft cover is arranged in such a way around the vane tip which is sunk into the recess that on the one hand the vane tip does not butt against the shaft cover when sinking into the recess and on the other hand the flow around the vane tip decreases.
  • the depth of the recesses is determined in such a way that during operation of the axial turbocompressor the radial relative movements between the vane tips and the shaft cover can be compensated.
  • the outline shape of the recesses is determined in such a way that during operation of the axial turbocompressor the axial relative movements between the vane tips and the shaft cover can be compensated.
  • FIG. 1 shows a perspective view of a detail of the axial turbocompressor
  • FIG. 2 shows a view along the vane longitudinal axis of the detail from FIG. 1 .
  • an axial turbocompressor 1 has a guide vane cascade 2 which is formed by a multiplicity of guide vanes 3 .
  • the guide vanes 3 are arranged in a row in the circumferential direction of the axial turbocompressor 1 and have a longitudinal extent in the radial direction of the axial turbocompressor 1 .
  • the axial turbocompressor 1 has a casing 5 on which the guide vanes 3 are fastened on the inner side. Facing away from the casing 5 , the guide vanes 3 have a vane tip 4 which points inwardly into the casing 5 .
  • a shaft cover 6 which is designed as a circumferentially symmetrical ring, is arranged directly at the vane tips 3 .
  • a multiplicity of recesses 7 On the outer side of the shaft cover 6 which faces the vane tips 4 , provision is made for a multiplicity of recesses 7 .
  • Each recess 7 is associated with a different vane tip 4 , wherein the recess 7 is located directly adjacent to its associated guide vane tip 4 .
  • the recess 7 is formed like a blind hole and therefore terminates in a blind manner. That is to say, it is provided with a tightly sealed off base in order to avoid leakage losses.
  • Each recess 7 on the outer side of the shaft cover 6 facing the vane tips 4 , has a contour 8 which is adapted to the profile shape of the guide vane 3 at the guide vane tip 4 . Also, each recess 7 is provided with a depth 9 in the shaft cover 6 . The shape of the contour 8 and the depth 9 are determined in such a way that during operation of the axial turbocompressor each vane tip 4 can sink into its associated recess 7 , wherein during the sinking in the vane tip 4 does not come into contact, or barely comes into contact, with the shaft cover 6 .
  • each recess 7 Applied to the base of each recess 7 is a honeycomb structure 10 , as is shown in FIG. 1 by way of example for the middle recess 7 . If, during operation of the axial turbocompressor, the vane tips 4 come into contact with the honeycomb structure 10 , then the honeycomb structure 10 yields so that the vane tip 4 presses into the honeycomb structure 10 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US13/201,065 2009-02-13 2010-01-27 Axial turbo compressor for a gas turbine having low radial gap losses and diffuser losses Abandoned US20110311355A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP09002056A EP2218918A1 (de) 2009-02-13 2009-02-13 Axialturboverdichter für eine Gasturbine mit geringen Spaltverlusten und Diffusorverlusten
EP09002056.1 2009-02-13
PCT/EP2010/050933 WO2010091956A1 (de) 2009-02-13 2010-01-27 Axialturboverdichter für eine gasturbine mit geringen radialspaltverlusten und diffusorverlusten

Publications (1)

Publication Number Publication Date
US20110311355A1 true US20110311355A1 (en) 2011-12-22

Family

ID=40469901

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/201,065 Abandoned US20110311355A1 (en) 2009-02-13 2010-01-27 Axial turbo compressor for a gas turbine having low radial gap losses and diffuser losses

Country Status (5)

Country Link
US (1) US20110311355A1 (ja)
EP (2) EP2218918A1 (ja)
JP (1) JP5567036B2 (ja)
CN (1) CN102317634B (ja)
WO (1) WO2010091956A1 (ja)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20160319840A1 (en) * 2015-05-01 2016-11-03 General Electric Company Compressor system and airfoil assembly
US9822645B2 (en) 2014-02-27 2017-11-21 Rolls-Royce Deutschland Ltd & Co Kg Group of blade rows

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2961564B1 (fr) * 2010-06-17 2016-03-04 Snecma Compresseur et turbomachine a rendement optimise
EP2538031A1 (de) * 2011-06-22 2012-12-26 Siemens Aktiengesellschaft Rotor mit Dichtelement für eine stationäre Gasturbine
CN104074799B (zh) * 2013-11-17 2017-01-18 成都中科航空发动机有限公司 一种具有扩张型子午流道的轴流压气机及其设计方法
EP2977559B1 (fr) * 2014-07-25 2017-06-07 Safran Aero Boosters SA Stator de turbomachine axiale et turbomachine associée
FR3133886B1 (fr) * 2022-03-24 2024-03-01 Safran Helicopter Engines Module pour turbomachine d’aéronef

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1328426A (en) * 1971-08-28 1973-08-30 British Leyland Truck & Bus Gas turbine engines
US5226789A (en) * 1991-05-13 1993-07-13 General Electric Company Composite fan stator assembly
US6431830B1 (en) * 1998-03-28 2002-08-13 MTU Motoren-und Turbinen München GmbH Nozzle ring for a gas turbine
US6619917B2 (en) * 2000-12-19 2003-09-16 United Technologies Corporation Machined fan exit guide vane attachment pockets for use in a gas turbine
US8636466B2 (en) * 2008-12-11 2014-01-28 Techspace Aero S.A. Segmented composite inner ferrule and segment of diffuser of axial compressor

Family Cites Families (7)

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EP0536575B1 (de) * 1991-10-08 1995-04-05 Asea Brown Boveri Ag Deckband für axialdurchströmte Turbine
US5494404A (en) * 1993-12-22 1996-02-27 Alliedsignal Inc. Insertable stator vane assembly
US6409472B1 (en) * 1999-08-09 2002-06-25 United Technologies Corporation Stator assembly for a rotary machine and clip member for a stator assembly
US6450766B1 (en) * 1999-08-09 2002-09-17 United Technologies Corporation Stator vane blank and method of forming the vane blank
US6543995B1 (en) * 1999-08-09 2003-04-08 United Technologies Corporation Stator vane and stator assembly for a rotary machine
US6425736B1 (en) * 1999-08-09 2002-07-30 United Technologies Corporation Stator assembly for a rotary machine and method for making the stator assembly
US20060198726A1 (en) * 2005-03-07 2006-09-07 General Electric Company Apparatus for eliminating compressor stator vibration induced by tip leakage vortex bursting

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1328426A (en) * 1971-08-28 1973-08-30 British Leyland Truck & Bus Gas turbine engines
US5226789A (en) * 1991-05-13 1993-07-13 General Electric Company Composite fan stator assembly
US6431830B1 (en) * 1998-03-28 2002-08-13 MTU Motoren-und Turbinen München GmbH Nozzle ring for a gas turbine
US6619917B2 (en) * 2000-12-19 2003-09-16 United Technologies Corporation Machined fan exit guide vane attachment pockets for use in a gas turbine
US6910860B2 (en) * 2000-12-19 2005-06-28 United Technologies Corporation Machined fan exit guide vane attachment pockets for use in a gas turbine
US8636466B2 (en) * 2008-12-11 2014-01-28 Techspace Aero S.A. Segmented composite inner ferrule and segment of diffuser of axial compressor

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9822645B2 (en) 2014-02-27 2017-11-21 Rolls-Royce Deutschland Ltd & Co Kg Group of blade rows
US20160319840A1 (en) * 2015-05-01 2016-11-03 General Electric Company Compressor system and airfoil assembly
CN106194276A (zh) * 2015-05-01 2016-12-07 通用电气公司 压缩机系统和翼型件组件
US9988918B2 (en) * 2015-05-01 2018-06-05 General Electric Company Compressor system and airfoil assembly

Also Published As

Publication number Publication date
CN102317634B (zh) 2014-06-25
WO2010091956A1 (de) 2010-08-19
JP2012518109A (ja) 2012-08-09
CN102317634A (zh) 2012-01-11
EP2396555A1 (de) 2011-12-21
EP2218918A1 (de) 2010-08-18
JP5567036B2 (ja) 2014-08-06

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AS Assignment

Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:BENKLER, FRANCOIS;KLEIN, KARL;MATTHIAS, TORSTEN;AND OTHERS;REEL/FRAME:026850/0113

Effective date: 20110714

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION