Guide vane arrangement

Patent metadata

Field	Value
Publication number	US-10151208-B2
Application number	US-201414244668-A
Country	US
Kind code	B2
Filing date	Apr 3, 2014
Priority date	Apr 3, 2013
Publication date	Dec 11, 2018
Grant date	Dec 11, 2018

Abstract

A guide vane arrangement for a turbomachine, particularly a gas turbine, is disclosed. The guide vane arrangement includes an external shroud band, on which at least two guide vanes protrude radially inward. A pair of axial ribs spaced apart from each other in the peripheral direction and disposed between two adjoining guide vanes in the peripheral direction protrude radially outward from a circumferential surface of the external shroud band and/or axially from a radial flange of the external shroud band.

First claim

What is claimed is:

1. A guide vane arrangement for a turbomachine, comprising:
an external shroud band including a single, continuous circumferential surface and a radial flange;
at least two guide vanes that protrude radially inward from the external shroud band; and
a pair of axial ribs, each axial rib of the pair of axial ribs being attached to the single, continuous circumferential surface such that the one axial rib is spaced apart from the other axial rib along the single, continuous circumferential surface in a peripheral direction, the attached pair of axial ribs being disposed in the external shroud band between two adjoining guide vanes in the peripheral direction, wherein each axial rib of the pair of axial ribs protrudes radially outward from the circumferential surface and axially from the radial flange, wherein the pair of axial ribs are each formed in an L-shape in the axial direction, and wherein each of the axial ribs have a first arm that is bonded to the radial flange and a second arm that is bonded to the circumferential surface.

2. The guide vane arrangement according to claim 1 , wherein the pair of axial ribs extend, at least in a region of the radial flange, radially outward at least to a radial inner end of a recess in the radial flange.

3. The guide vane arrangement according to claim 2 , wherein the recess is a slit.

4. The guide vane arrangement according to claim 1 , wherein the pair of axial ribs end in front of an additional radial flange of the external shroud band.

5. The guide vane arrangement according to claim 1 , wherein the circumferential surface is sunken in radially inwards in the peripheral direction between the pair of axial ribs.

6. The guide vane arrangement according to claim 1 , wherein the pair of axial ribs are welded to the circumferential surface and/or the radial flange in a fillet or produced by deposition welding.

7. A gas turbine, comprising:
at least one compressor and/or turbine stage with the guide vane arrangement according to claim 1 , wherein the guide vane arrangement is connected by the radial flange to a housing of the gas turbine.

8. The gas turbine according to claim 7 , wherein the gas turbine is an aircraft engine gas turbine.

9. The gas turbine according to claim 7 , wherein the guide vane arrangement is detachably connected by the radial flange to the housing of the gas turbine.

10. A guide vane arrangement for a turbomachine, comprising:
an external shroud band including a single, continuous circumferential surface and a radial flange;
at least two guide vanes that protrude radially inward from the external shroud band; and
a pair of axial ribs, each axial rib of the pair of axial ribs being attached to the single, continuous circumferential surface such that the one axial rib is spaced apart from the other axial rib along the single, continuous circumferential surface in a peripheral direction, the attached pair of axial ribs being disposed in the external shroud band between two adjoining guide vanes in the peripheral direction, wherein each axial rib of the pair of axial ribs protrudes radially outward from the circumferential surface and axially from the radial flange and wherein the pair of axial ribs are welded to the circumferential surface and/or the radial flange in a fillet.

Assignees

• MTU Aero Engines AG

Inventors

• Hein Manuel
• Uecker Markus
• Stanka Rudolf

Classifications

• F01D9/042 Primary

  fixing blades to stators (fixing stator-rings in the casing or to each other F01D25/246) · CPC title

• F01D9/041 Primary

  using blades (F01D5/148 takes precedence) · CPC title

• F01D25/246

  Fastening of diaphragms or stator-rings · CPC title

• F05D2230/30

  with deposition of material · CPC title

• F05D2230/232

  by welding · CPC title