Turbine Vane With Non-Uniform Wall Thickness

• US 20160061046A1
• Filed: 05/20/2014
• Published: 03/03/2016
• Est. Priority Date: 06/10/2013
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Allen, Richard N., Porter, Steven D., Ols, John T.
• Granted Patent: US 10,641,114 B2
• Time in Patent Office: Days
• US Class Current: 1/1
• CPC Class Codes: F01D 25/24 Casings modified for heating or cooling F01D25/14; Casing parts, e.g. diaphragms, casing fastenings casings for rotary machines or engines in general F16M ; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045
• CPC Class Codes: F01D 5/143 Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F01D 9/042 fixing blades to stators fixing stator-rings in the casing or to each other F01D25/246
• CPC Class Codes: F01D 9/047 Nozzle boxes
• CPC Class Codes: F05D 2240/121 related to the leading edge of a stator vane
• CPC Class Codes: F05D 2240/122 related to the trailing edge of a stator vane
• CPC Class Codes: F05D 2240/80 Platforms for stationary or moving blades
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A vane for use in a gas turbine engine has an airfoil extending between a leading edge and a trailing edge, and is generally hollow. The vane has a radially outer platform and a radially inner platform. At least one of the radially inner and radially outer platforms has thicker portions circumferentially located beyond one of the leading edge and the trailing edge, and thinner portions circumferentially beyond circumferential edges of the airfoil. A mid-turbine frame and a gas turbine engine are also disclosed.

Claims

1. A vane for use in a gas turbine engine comprising:

an airfoil extending between a leading edge and a trailing edge, and being generally hollow;

a radially outer platform and a radially inner platform; and

with at least one of said radially inner and radially outer platforms having thicker portions circumferentially located beyond one of said leading edge and said trailing edge, and thinner portions circumferentially beyond circumferential edges of said airfoil.

2. The vane as set forth in claim 1, wherein said at least one platform is said radially inner platform.

3. The vane as set forth in claim 1, wherein said thicker portions vary in thickness, and extend circumferentially beyond circumferential edges of said airfoil.

4. The vane as set forth in claim 3, wherein a radial thickness of said thicker portion at a thickest portion compared to a radial thickness of thinner portions circumferentially spaced from said thicker portions is equal to or between about 1.1 and 5.0.

5. The vane as set forth in claim 1, wherein said thicker portions are provided by at least one thicker bump positioned axially beyond said leading edge, but circumferentially aligned with said airfoil.

6. The vane as set forth in claim 5, wherein there are at least a plurality of said bumps.

7. The vane as set forth in claim 6, wherein a radial thickness of a thickest portion of said bump is defined, and a ratio of said radial thickness of said thickest portion of said bump to a radially thickness of portions circumferentially spaced from said bump is equal to or between about 1.1 and 5.0.

8. The vane as set forth in claim 1, wherein said one of said leading edge and said trailing edge is said leading edge.

9. A mid-turbine frame comprising:

a plurality of vanes having airfoils extending between a radially inner and a radially outer platform;

said airfoil shape extending between a leading edge and a trailing edge, and being generally hollow; and

with at least one of said radially inner and radially outer platforms having thicker portions circumferentially located beyond said leading edge, and thinner portions circumferentially beyond circumferential edges of said airfoil.

10. The mid-turbine frame as set forth in claim 9, wherein said thicker portions vary in thickness, and extend circumferentially beyond circumferential edges of said airfoil.

11. The mid-turbine frame as set forth in claim 10, wherein a radial thickness of said thicker portion at a thickest portion compared to a radial thickness of thinner portions circumferentially spaced from said thicker portions is equal to or between about 1.1 and 5.0.

12. The mid-turbine frame as set forth in claim 9, wherein said thicker portions are provided by at least one thicker bump positioned axially beyond said leading edge, but circumferentially aligned with said airfoil.

13. The mid-turbine frame vane as set forth in claim 12, wherein a radial thickness of a thickest portion of said bump is defined, and a ratio of said radial thickness of said thickest portion of said bump to a radially thickness of portions circumferentially spaced from said bump is equal to or between about 1.1 and 5.0.

14. The mid-turbine frame as set forth in claim 12, wherein there are at least a plurality of said bumps.

15. A gas turbine engine comprising:

at least a first upstream turbine rotor and a downstream turbine rotor, with a mid-turbine frame located intermediate said upstream and downstream turbine rotors;

the mid-turbine frame including a plurality of vanes having a radially outer platform and a radially inner platform, and an airfoil extending between a leading edge and a trailing edge, and being generally hollow; and

at least one of said radially inner and radially outer platforms having thicker portions circumferentially located beyond said leading edge, and thinner portions circumferentially beyond circumferential edges of said airfoil.

16. The gas turbine engine as set forth in claim 15, wherein said at least one platform is said radially inner platform.

17. The gas turbine engine as set forth in claim 15, wherein said thicker portions vary in thickness, and extend circumferentially across circumferential edges of said airfoil.

18. The gas turbine engine as set forth in claim 17, wherein a radial thickness of said thicker portion at a thickest portion compared to a radial thickness of thinner portions circumferentially spaced from said thicker portions is equal to or between about 1.1 and 5.0.

19. The gas turbine engine as set forth in claim 15, wherein said thicker portions are provided by at least one thicker bump positioned axially beyond said leading edge, but circumferentially aligned with said airfoil.

20. The gas turbine engine as set forth in claim 19, wherein a radial thickness of a thickest portion of said bump is defined, and a ratio of said radial thickness of said thickest portion of said bump to a radially thickness of portions circumferentially spaced from said bump is equal to or between about 1.1 and 5.0.