GAS TURBINE ENGINE VANE END DEVICES

• US 20190330993A1
• Filed: 06/26/2019
• Published: 10/31/2019
• Est. Priority Date: 02/26/2013
• Status: Active Application

• Current Assignee: 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.
• Granted Patent: US 11,326,464 B2
• Time in Patent Office: Days
• 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 rotatable airfoil member for a gas turbine engine having an end structured to contact a wall of the gas turbine engine, the end having an abradable coating configured to wear according to a contact between the coating and the wall of the gas turbine engine.

2. The apparatus of claim 1, wherein the rotatable airfoil member is disposed adjacent a rotatable blade and is configured to turn a working fluid.

3. The apparatus of claim 2, wherein the end includes a squealer tip, and wherein the coating is applied to an end portion of the squealer tip.

4. The apparatus of claim 3, wherein the rotatable airfoil member is a turbine vane, wherein the vane includes a cooling hole disposed between opposing squealer tips, and wherein one of the opposing squealer tips is disposed toward a relatively low pressure side of the vane and the other of the opposing squealer tips is disposed toward a relatively high pressure side of the vane.

5. The apparatus of claim 1, wherein the abradable coating is made of a first material and the end of the vane is made of a second material, and which further includes a gas turbine engine having the rotatable airfoil member.

6. The apparatus of claim 5, wherein the first material is different than the second material.

7. The apparatus of claim 5, wherein a density of the first material is different than a density of the second material.

8. An apparatus comprising:

a rotatable airfoil member for a gas turbine engine, the airfoil member having a first end and a second end spaced apart radially from the first end relative to a central axis, the first end formed to include a first squealer member that extends radially outward away from the first end relative to the central axis, and the rotatable airfoil member configured to rotate about a rotation axis that extends perpendicular to the central axis; and

a wall arranged circumferentially around the first end of the rotatable airfoil member relative to the axis;

wherein the rotatable airfoil member further includes a first wear surface disposed on a radial terminating tip of the first squealer member and the first wear surface is configured to be abraded when relative movement between the rotatable airfoil member and the wall causes the rotatable airfoil member and the wall to contact each other.

9. The apparatus of claim 8, wherein the rotatable airfoil member includes a leading edge and a trailing edge spaced apart axially from the leading edge relative to the central axis and the first squealer member is located at the leading edge of the rotatable airfoil member.

10. The apparatus of claim 9, wherein the rotatable airfoil member is formed to include a cooling hole formed in the first end of the rotatable airfoil member.

11. The apparatus of claim 10, wherein the rotatable airfoil member includes a second squealer member and a second wear surface disposed on the second squealer member and the second squealer member is spaced apart from the first squealer member to locate the cooling hole between the first squealer member and the second squealer member.

12. The apparatus of claim 11, wherein the first wear surface has a first radial thickness and the second wear surface has a second radial thickness that is different than the first radial thickness.

13. The apparatus of claim 8, wherein the rotatable airfoil member includes a leading edge and a trailing edge spaced apart axially from the leading edge relative to the central axis and the first squealer member is located at the trailing edge of the rotatable airfoil member.

14. The apparatus of claim 8, wherein the rotatable airfoil member includes a leading edge and a trailing edge spaced apart axially from the leading edge relative to the central axis and a radial thickness of the first wear surface varies as a function of axial location of the first wear surface relative to the central axis.

15. The apparatus of claim 14, wherein the first wear surface is thicker near the trailing edge than that near the leading edge of the rotatable airfoil member.

16. The apparatus of claim 8, wherein the first squealer member is airfoil shaped and extends continuously around the first end of the rotatable airfoil member.

17. The apparatus of claim 8, wherein the rotatable airfoil member is made of a first material and the first wear layer is made of a second material that is different than the first material.

18. The apparatus of claim 8, wherein the first wear surface is formed from a material and an entirety of the first squealer member is formed form the material.

19. The apparatus of claim 8, wherein the rotatable airfoil member includes a second squealer member and a second wear surface disposed on the second squealer member, the first wear surface has a first radial thickness, and the second wear surface has a second radial thickness that is different than the first radial thickness.

20. The apparatus of claim 8, wherein the rotatable airfoil member includes a second squealer member spaced apart from the first squealer member, the rotatable airfoil member is a turbine vane, the turbine vane includes a cooling hole disposed between the first and second squealer members, and wherein the first squealer members is disposed toward a relatively low pressure side of the turbine vane and the second squealer member is disposed toward a relatively high pressure side of the turbine vane.