Gas turbine engine vane end devices

• US 9,938,845 B2
• Filed: 12/18/2013
• Issued: 04/10/2018
• Est. Priority Date: 02/26/2013
• Status: Active Grant

• Current Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc), Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc), Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Inventors: Taketa, Linnea, Munson, John, Freeman, Ted J.
• Primary Examiner(s): VERDIER, CHRISTOPHER M
• Application Number:
  US14/132,738
• Publication Number:
  US 20150003963A1
• Time in Patent Office: 1,574 Days
• Field of Search: 4151701, 4151733, 4151742, 4151743, 415230, 415231, 415148, 415150-151, 415159-166, 277355
• CPC Class Codes: F01D 11/00 Preventing or minimising internal leakage of working-fluid, e.g. between stages sealings in general F16J ; sealing arrangements for transition ducts of combustor cans F01D9/023
• CPC Class Codes: F01D 11/005 Sealing means between non relatively rotating elements
• CPC Class Codes: F01D 11/02 by non-contact sealings, e.g. of labyrinth type for sealing space between rotor blade tips and stator F01D11/08
• CPC Class Codes: F01D 11/025 Seal clearance control; Floating assembly; Adaptation means to differential thermal dilatations
• CPC Class Codes: F01D 11/122 with erodable or abradable material blades having cutting or grinding tips F01D5/20
• CPC Class Codes: F01D 17/148 by means of rotatable members, e.g. butterfly valves
• CPC Class Codes: F01D 17/162 for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line F01D17/167 takes precedence
• CPC Class Codes: F01D 17/165 for radial flow, i.e. the vanes turning around axes which are essentially parallel to the rotor centre line F01D17/167 takes precedence
• CPC Class Codes: F01D 5/28 Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F04D 19/02 Multi-stage pumps
• CPC Class Codes: F04D 27/002 by varying geometry within the pumps, e.g. by adjusting vanes
• CPC Class Codes: F04D 27/0246 by varying geometry within the pumps, e.g. by adjusting vanes
• CPC Class Codes: F04D 29/083 especially adapted for elastic fluid pumps
• CPC Class Codes: F04D 29/164 of an axial flow wheel
• CPC Class Codes: F05B 2240/571 Brush seals
• CPC Class Codes: F05D 2240/56 Brush seals
• CPC Class Codes: F16J 15/3288 Filamentary structures, e.g. brush seals

Abstract

A turbomachinery component of a gas turbine engine is disclosed having a number of techniques of reducing the effects of a gap flow between an airfoil member of the gas turbine engine and a wall of the gas turbine engine. The airfoil member can be variable and in one form is a variable turbine vane. In one embodiment a brush seal is included between the vane and the wall. In another form a wear surface is disposed between the vane and the wall. In yet another form a moveable member capable of being actuated to change position can be disposed between the vane and the wall to alter the size of a gap between the two.

Claims

1. An apparatus comprising:

a moveable airfoil member structured for use in a working fluid flow path of a gas turbine engine; and

a brush seal disposed at an end of the moveable airfoil member, the brush seal having a plurality of extensions projecting outwardly and configured to discourage a flow of working fluid through the extensions as the working fluid traverses the working fluid flow path,

wherein the plurality of extensions each have a first end and a second end both disposed toward a distal side of the plurality of extensions, the first end and second end connected via a body that is looped around a central member, and

which further includes a crimp to couple the plurality of extensions to the central member, the central member extending along a chord of the moveable airfoil member, and wherein the moveable airfoil member is a rotatable turbine vane,

wherein the moveable airfoil member includes a leading edge and a trailing edge spaced apart from the leading edge to define the chord of the moveable airfoil member, the brush seal extends at least partway along the chord, and a height of the plurality of extensions varies along the chord, the plurality of extensions have a first uniform height near the leading edge, the plurality of extensions have a second uniform height near the trailing edge, and the second height is greater than the first height.

2. The apparatus of claim 1, which further includes the gas turbine engine, the engine including a plurality of the moveable airfoil members, wherein each moveable airfoil member includes a range of travel, and wherein the extensions contact a wall arranged around the moveable airfoil members in the range of travel.

3. The apparatus of claim 1, wherein the plurality of extensions cover a portion of the chord between the leading edge and trailing edge.

4. The apparatus of claim 1, wherein the plurality of extensions include flexible extensions capable of flexing when placed in contact with a wall of the gas turbine engine.

5. The apparatus of claim 4, wherein an amount of flexing is dependent upon a position of the vane, and wherein the plurality of extensions creates a plurality of tortuous flow paths between an upstream side and a downstream side of the plurality of extensions.

6. A gas turbine engine assembly comprising

a turbomachinery component having a wall that defines a flow path of the gas turbine engine assembly,

a rotatable vane positioned adjacent the wall and configured to move relative to the wall, and

a brush seal coupled to the rotatable vane for movement therewith, the brush seal being located between the rotatable vane and the wall, and the brush seal including a base member, a plurality of bristles that extend outwardly away from the base member toward the wall, and a clamp arranged around a portion of the base member and the bristles to couple the bristles with the base member,

wherein the rotatable vane includes a leading edge and a trailing edge spaced apart from the leading edge to define a chord of the rotatable vane, the brush seal extends at least partway along the chord, and a height of the plurality of bristles varies along the chord,

wherein the rotatable vane is moveable between a first position and a second position and at least one portion of the bristles does not contact the wall in at least one of the first position and the second position.

7. The gas turbine engine assembly of claim 6, wherein the plurality of bristles have a first height near the leading edge and the plurality of bristles have a second height near the trailing edge and the second height is greater than the first height.

8. The gas turbine engine assembly of claim 6, wherein the base member has a circular cross section, each of the plurality of bristles extend between a first end and second end of the bristle, and each bristle is wound about the base member to cause the first end and the second end of the bristle to extend toward the wall.

9. The gas turbine engine assembly of claim 6, wherein the plurality of bristles comprise ceramic fiber and are configured to flex in response to contacting the wall.

10. A gas turbine engine assembly comprising

a turbomachinery component having a wall that defines a flow path of the gas turbine engine assembly,

a rotatable vane positioned adjacent the wall to define a gap between the rotatable vane and the wall, the rotatable vane having a spindle and the rotatable vane configured to rotate about the spindle relative to the wall between a first position and a second position to vary a height of the gap formed between the rotatable vane and the wall, and

a brush seal coupled to the rotatable vane for movement therewith, the brush seal including a plurality of bristles located in the gap, and the brush seal arranged to cause the bristles to contact the wall when the rotatable vane is in at least one of the first position and the second position,

wherein the brush seal further includes a base member coupled with the rotatable vane and the plurality of bristles are coupled with the base member,

wherein the base member is angled relative to one of the rotatable vane and the wall,

wherein the rotatable vane includes a leading edge and a trailing edge spaced apart from the leading edge to define a chord of the rotatable vane, the brush seal extends at least partway along the chord, and a height of the plurality of bristles varies along the chord, the plurality of bristles have a first uniform height near the leading edge, the plurality of bristles have a second uniform height near the trailing edge, and the second height is greater than the first height.

11. The gas turbine engine assembly of claim 10, wherein the plurality of bristles are bonded with the base member.

12. The gas turbine engine assembly of claim 10, wherein the brush seal extends along the entirety of the chord.

13. The gas turbine engine assembly of claim 10, wherein the rotatable vane includes a pressure side and a suction side spaced apart from the pressure side and the plurality of bristles includes a first group of bristles disposed toward the pressure side of the vane and a second group of bristles disposed toward the suction side of the vane.

14. The gas turbine engine assembly of claim 10, wherein a portion of the brush seal is received within an opening formed in the rotatable vane.