EP2787178A1 - Ensemble d'aube directrice - Google Patents

Ensemble d'aube directrice Download PDF

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Publication number
EP2787178A1
EP2787178A1 EP13162067.6A EP13162067A EP2787178A1 EP 2787178 A1 EP2787178 A1 EP 2787178A1 EP 13162067 A EP13162067 A EP 13162067A EP 2787178 A1 EP2787178 A1 EP 2787178A1
Authority
EP
European Patent Office
Prior art keywords
radial flange
vane assembly
lateral surface
axial ribs
guide vane
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.)
Granted
Application number
EP13162067.6A
Other languages
German (de)
English (en)
Other versions
EP2787178B1 (fr
Inventor
Manuel Hein
Markus Uecker
Rudolf Stanka
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
Priority to EP13162067.6A priority Critical patent/EP2787178B1/fr
Priority to US14/244,668 priority patent/US10151208B2/en
Publication of EP2787178A1 publication Critical patent/EP2787178A1/fr
Application granted granted Critical
Publication of EP2787178B1 publication Critical patent/EP2787178B1/fr
Active 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
    • 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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material

Definitions

  • the present invention relates to a vane arrangement for a turbomachine, in particular a gas turbine, and a gas turbine, in particular an aircraft engine gas turbine, with one or more such vane arrangements.
  • vane arrangements in aircraft engine gas turbines are subject to high thermal loads. Therefore, it is known in the industry practice to provide, for example, in the rear radial flanges of outer shrouds of such vane assemblies, cooling slots through which cooling air flows during operation to cool the vane assemblies and / or a housing in which it is disposed.
  • Such cooling slots represent a possible root of component cracks.
  • An object of an embodiment of the present invention is to provide an improved turbomachine, in particular an improved gas turbine.
  • a gas turbine in particular an aircraft engine gas turbine, one or more compressor and / or turbine stages, in particular low-pressure compressor and / or turbine stages, each having a vane arrangement.
  • a vane assembly according to one aspect of the present invention comprises a single or multiple outer shroud from which two or more vanes project radially inward.
  • the vanes can be detachable or permanent with connected to the outer shroud, in particular integrally made with this or materially connected thereto, in particular welded, be.
  • the outer shroud has a, in particular at least substantially conical, circumferential surface and at least one radial flange, in particular a rear radial flange in the throughflow direction and / or a front radial flange in the flow direction.
  • the guide vane assembly by means of at least one radial flange, in particular by a rear radial flange and / or a front radial flange with a housing, in particular releasably and / or positively connected, preferably suspended in this.
  • a pair of axial ribs are arranged between at least one pair of adjacent guide vanes, which are spaced apart in the circumferential direction from the outer surface of the outer ⁇ -end band and radially outward or between each guide vanes of the stator vane assembly. protrude on the side facing away from the guide vane.
  • the axial ribs of at least one, in particular rear and / or front, radial flange of the outer cover strip can project axially toward the lateral surface.
  • a possible crack propagation in one embodiment can advantageously be guided between the axial ribs and thus limited to a specific area of the outer platform.
  • crack propagation into an area in which a guide blade is arranged is made more difficult, preferably blocked, and thus directed into areas which are more favorable for this purpose.
  • a, in particular rear, radial flange has one or more recesses.
  • these can be designed like slits and / or produced by electrical discharge machining (EDM) and / or extend, at least substantially, in the radial direction, in particular from a radially outer edge of the radial flange. They are provided or arranged in an embodiment for the flow through a cooling fluid, in particular cooling air.
  • the axial ribs of an axial rib pair of adjacent axial ribs extend radially outward at least to a radially inner end of a recess, in particular a cooling slot, in a, in particular rear, radial flange. They are in one embodiment, seen in the circumferential direction, arranged on either side of the recess or the recess is arranged in the circumferential direction between the two adjacent axial ribs of an axial rib pair. In this way, in one embodiment, a possible crack propagation is already limited at its root between the axial ribs and guided by them at the radial flange and further in the lateral surface.
  • the axial ribs are in one embodiment, at least substantially, L-shaped, wherein one leg is integrally connected to the radial flange and the other leg is materially connected to the lateral surface.
  • the leg connected to the radial flange may, as stated above, in particular extend radially outward at least up to a radially inner end of a recess.
  • Such an axial rib thus extends radially outward at least as far as the radially inner end only in the area of the radial flange.
  • an axial rib may also be triangular, rectangular, U-shaped or the like, for example.
  • the axial ribs extend or run in an embodiment, at least substantially, in the axial direction. Likewise, they may also include an angle with the axial direction or axis of rotation of the turbomachine in one embodiment corresponds to a stagger angle of the vanes. For a more compact representation such oblique ribs are referred to as axial ribs in the context of the present invention.
  • Axial ridges can be rectilinear or curved or extend axially.
  • the axial ribs may be equidistant from the adjacent vanes in the circumferential direction, at least substantially.
  • a minimum circumferential clearance between an axial rib of an axial rib pair and its adjacent vane may be, at least substantially, equal to the minimum circumferential clearance between the other axial rib of that axial rib pair and the adjacent vane.
  • the axial ribs may extend in one embodiment from a rear to a front radial flange.
  • the axial ribs in particular starting from a rear radial flange, end axially in front of a further, in particular front, radial flange. It has been found that this customary crack sufficient and yet the material cost of the axial ribs can be minimized.
  • the axial ribs are connected to the lateral surface and / or the radial flange in a rounding, in particular a so-called fillet. This can make it difficult to climb over the ribs by a crack.
  • the lateral surface is lowered in the circumferential direction between the axial ribs radially inward. In this way, a possible crack additionally be directed between the pair of axial ribs.
  • the lateral surface at least substantially, the same wall thickness as the radial flange.
  • the axial ribs can be made separately in one embodiment and then materially connected to the lateral surface and / or the radial flange, in particular be welded. Likewise, they can be produced by build-up welding or be urformed with the lateral surface and / or the radial flange.
  • Fig. 1, 2 show a portion of a vane assembly of a low pressure compressor stage of an aircraft engine gas turbine according to an embodiment of the present invention in perspective view and an axial section, respectively.
  • the vane assembly has an outer shroud from which a plurality of vanes 1 radially inward (vertically downward in Fig. 1, 2 ) stand out.
  • the foot contours of the guide vanes 1 are in Fig. 1 indicated by dashed lines.
  • the outer shroud has a substantially conical surface 2, one in the flow direction (from left to right in Fig. 1, 2 ) rear radial flange 3 and a front in the flow direction radial flange 4, in Fig. 1 is omitted.
  • the Au ⁇ endeckband in a housing (not shown) mounted, as for example in the EP 1 462 616 A2 is disclosed, to which reference is made.
  • Seen in the circumferential direction (perpendicular to the plane of the Fig. 2 ) is arranged between the two adjacent vanes 1, a pair of axial ribs 5, which are circumferentially spaced from each other and from the outer surface 2 of the Au ⁇ endeckbands radially outward (vertically upwards in FIG Fig. 1, 2 ) and from the rear Radial flange 3 axially towards the lateral surface 2 out (from right to left in Fig. 2 ) protrude and are welded to each.
  • a slot-like recess 6 is formed by spark erosion, which extends substantially in the radial direction from a radially outer edge of the rear radial flange 3 (top in FIG Fig. 1 ) and for the flow through a cooling fluid, in particular cooling air, is provided.
  • the axial ribs 5 are viewed in the circumferential direction (in the plan view of Fig. 2 ) L-shaped, with a leg (right in Fig. 2 ) is materially connected to the rear radial flange 3 and the other leg (bottom in Fig. 2 ) is materially connected to the lateral surface 2.
  • the leg connected to the rear radial flange 3 extends radially outward to the radially inner end of the recess 6, which is arranged in the circumferential direction between the two adjacent axial ribs 5. In this way, a propagation of the crack 7 can already be limited at its root between the axial ribs 5 and guided by these at the radial flange 3 and in the lateral surface 2.
  • the axial ribs 5 extend substantially in the axial direction or parallel to a stagger angle of the guide vanes 1 and are equidistant from the adjacent stator vanes 1 in the circumferential direction.
  • the axial ribs 5 end, starting from the rear radial flange 3, axially in front of the front radial flange 4 (see. Fig. 2 ). It has been found that this customary crack sufficient and yet the material cost of the axial ribs can be minimized.
  • the axial ribs are connected to the lateral surface 2 and the rear radial flange 3 in a fillet 8. As a result, an over-climbing of the ribs 5 are made difficult by the crack 7.
  • the lateral surface 2 is slightly lowered in the circumferential direction between the axial ribs 5 radially inward.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13162067.6A 2013-04-03 2013-04-03 Ensemble d'aube directrice Active EP2787178B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP13162067.6A EP2787178B1 (fr) 2013-04-03 2013-04-03 Ensemble d'aube directrice
US14/244,668 US10151208B2 (en) 2013-04-03 2014-04-03 Guide vane arrangement

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP13162067.6A EP2787178B1 (fr) 2013-04-03 2013-04-03 Ensemble d'aube directrice

Publications (2)

Publication Number Publication Date
EP2787178A1 true EP2787178A1 (fr) 2014-10-08
EP2787178B1 EP2787178B1 (fr) 2016-03-02

Family

ID=48013866

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13162067.6A Active EP2787178B1 (fr) 2013-04-03 2013-04-03 Ensemble d'aube directrice

Country Status (2)

Country Link
US (1) US10151208B2 (fr)
EP (1) EP2787178B1 (fr)

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6363232B2 (ja) * 2014-06-12 2018-07-25 ゼネラル・エレクトリック・カンパニイ シュラウドハンガーアセンブリ
US9863265B2 (en) 2015-04-15 2018-01-09 General Electric Company Shroud assembly and shroud for gas turbine engine
DE102015222834A1 (de) 2015-11-19 2017-05-24 MTU Aero Engines AG Schaufelcluster mit Umfangssicherung
US11168566B2 (en) 2016-12-05 2021-11-09 MTU Aero Engines AG Turbine blade comprising a cavity with wall surface discontinuities and process for the production thereof
US11286797B2 (en) * 2018-06-06 2022-03-29 Raytheon Technologies Corporation Gas turbine engine stator vane base shape

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5248240A (en) * 1993-02-08 1993-09-28 General Electric Company Turbine stator vane assembly
EP1462616A2 (fr) 2003-03-22 2004-09-29 MTU Aero Engines GmbH Ensemble de fixation axiale et radiale d'une aube statorique dans un boítier d'une turbine
EP2397653A1 (fr) * 2010-06-17 2011-12-21 Siemens Aktiengesellschaft Segment de plateforme pour porter une aube de guidage pour turbine à gaz et procédé de refroidissement de ce segment
US20120128472A1 (en) * 2010-11-23 2012-05-24 General Electric Company Turbomachine nozzle segment having an integrated diaphragm

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5127793A (en) * 1990-05-31 1992-07-07 General Electric Company Turbine shroud clearance control assembly
US8257028B2 (en) * 2007-12-29 2012-09-04 General Electric Company Turbine nozzle segment
FR2928963B1 (fr) * 2008-03-19 2017-12-08 Snecma Distributeur de turbine pour une turbomachine.
DE102009051552A1 (de) * 2009-10-31 2011-05-05 Mtu Aero Engines Gmbh Verfahren und Vorrichtung zur Herstellung eines Bauteils

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5248240A (en) * 1993-02-08 1993-09-28 General Electric Company Turbine stator vane assembly
EP1462616A2 (fr) 2003-03-22 2004-09-29 MTU Aero Engines GmbH Ensemble de fixation axiale et radiale d'une aube statorique dans un boítier d'une turbine
EP2397653A1 (fr) * 2010-06-17 2011-12-21 Siemens Aktiengesellschaft Segment de plateforme pour porter une aube de guidage pour turbine à gaz et procédé de refroidissement de ce segment
US20120128472A1 (en) * 2010-11-23 2012-05-24 General Electric Company Turbomachine nozzle segment having an integrated diaphragm

Also Published As

Publication number Publication date
US20140301840A1 (en) 2014-10-09
EP2787178B1 (fr) 2016-03-02
US10151208B2 (en) 2018-12-11

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