EP2787178B1 - Ensemble d'aube directrice - Google Patents
Ensemble d'aube directrice Download PDFInfo
- Publication number
- EP2787178B1 EP2787178B1 EP13162067.6A EP13162067A EP2787178B1 EP 2787178 B1 EP2787178 B1 EP 2787178B1 EP 13162067 A EP13162067 A EP 13162067A EP 2787178 B1 EP2787178 B1 EP 2787178B1
- Authority
- EP
- European Patent Office
- Prior art keywords
- radial flange
- axial ribs
- vane assembly
- lateral surface
- axial
- 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.)
- Active
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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
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- 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
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- 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
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- 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.
- a generic vane arrangement is for example from the EP 2 397 653 A1 and the US 2012/128472 A1 known. In addition, it was also on the US 5,248,240 pointed.
- vane arrangements in aircraft engine gas turbines are subject to high thermal loads. Therefore, it is known from the field of internal practice, for example, to provide cooling slots in rear radial flanges of outer shrouds of such vane assemblies 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.
- 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 region 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 as far as a radially inner end from 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 Umtangsraum between the two adjacent axial ribs of a Axialrippencoveres. 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 can also be triangular, rectangular, Ü-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 increases in the circumferential direction between the axial ribs radially outward, in particular thickened between the axial ribs, be to complicate the progression of a crack.
- 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 outer shroud is mounted in a housing (not shown), 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 spaced apart in the circumferential direction and from the outer surface 2 of the outer shroud 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)
Claims (7)
- Ensemble d'aubes directrices destiné à une turbomachine, notamment une turbine à gaz, pourvu d'un anneau de renforcement extérieur duquel au moins deux aubes directrices (1) font saillie radialement vers l'intérieur, ledit ensemble comprenant une paire de nervures axiales (5) qui sont disposées à distance l'une de l'autre dans la direction périphérique et qui font saillie, entre deux aubes directrices (1) adjacentes dans la direction périphérique, radialement vers l'extérieur depuis une surface d'enveloppe (2) de l'anneau de renforcement extérieur et/ou axialement en direction de la surface d'enveloppe depuis une bride radiale (3) de l'anneau de renforcement extérieur, caractérisé en ce que les nervures axiales s'étendent radialement vers l'extérieur, au moins dans la zone de la bride radiale, au moins jusqu'à une extrémité radialement intérieure d'un évidement (6), notamment en forme de fente, ménagé dans la bride radiale.
- Ensemble d'aubes directrices selon la revendication précédente, caractérisé en ce que les nervures axiales sont configurées, au moins sensiblement, en forme de L et comportent chacune une branche, qui est reliée à la bride radiale par liaison de matière, et une branche qui est reliée à la surface d'enveloppe par liaison de matière.
- Ensemble d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce que les nervures axiales se terminent en avant d'une autre bride radiale (4) de l'anneau de renforcement extérieur.
- Ensemble d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce que les nervures axiales sont reliées à la surface d'enveloppe et/ou à la bride radiale dans un arrondi (8).
- Ensemble d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce que la surface d'enveloppe est abaissée radialement vers l'intérieur entre les nervures axiales dans la direction périphérique.
- Ensemble d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce que les nervures axiales sont soudées à la surface d'enveloppe et/ou à la bride radiale, en particulier dans un arrondi (8), ou sont réalisées par soudage avec apport de matière.
- Turbine à gaz, notamment turbine à gaz de moteur d'avion, comprenant au moins un étage de compresseur et/ou de turbine équipé d'un ensemble d'aubes directrices selon l'une des revendications précédentes, caractérisée en ce que l'ensemble d'aubes directrices est relié, notamment de façon amovible, à un boîtier de la turbine à gaz au moyen de la bride radiale (3).
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 EP2787178A1 (fr) | 2014-10-08 |
EP2787178B1 true 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)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CA2951425C (fr) * | 2014-06-12 | 2019-12-24 | General Electric Company | Ensemble de suspension de carenage |
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 |
Family Cites Families (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5127793A (en) * | 1990-05-31 | 1992-07-07 | General Electric Company | Turbine shroud clearance control assembly |
US5248240A (en) * | 1993-02-08 | 1993-09-28 | General Electric Company | Turbine stator vane assembly |
DE10312956B4 (de) | 2003-03-22 | 2011-08-11 | MTU Aero Engines GmbH, 80995 | Anordnung für das axiale und radiale Festlegen einer Leitschaufelbaugruppe in dem Gehäuse eines Turbinentriebwerkes |
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 |
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 |
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2013
- 2013-04-03 EP EP13162067.6A patent/EP2787178B1/fr active Active
-
2014
- 2014-04-03 US US14/244,668 patent/US10151208B2/en active Active
Also Published As
Publication number | Publication date |
---|---|
US20140301840A1 (en) | 2014-10-09 |
EP2787178A1 (fr) | 2014-10-08 |
US10151208B2 (en) | 2018-12-11 |
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