Load bearing feature for ceramic matrix composite turbine components

Patent metadata

Field	Value
Publication number	US-12428965-B2
Application number	US-202418428856-A
Country	US
Kind code	B2
Filing date	Jan 31, 2024
Priority date	Jan 31, 2024
Publication date	Sep 30, 2025
Grant date	Sep 30, 2025

Abstract

A gas turbine engine component includes an inner surface and an outer surface. The outer surface and the inner surface are formed of ceramic matrix composite layers, and mount member. The mount member is in contact with an insert that is inserted within outer layers of the outer platform to form a thickened portion having an end surface that will react against forces from the mount member. A gas turbine engine and a method are also disclosed.

First claim

What is claimed is:

1. A gas turbine engine component comprising:
an inner surface and an outer surface, with the outer surface and the inner surface being formed of ceramic matrix composite layers, and a mount member, the mount member being in contact with an insert inserted within outer ones of the ceramic matrix composite layers of the outer surface to form a thickened portion having an end surface that will react against forces from the mount member.

2. The gas turbine engine component as set forth in claim 1 , wherein the insert has an inner end remote from the end surface, and sides extending outwardly from the inner end such that the end surface is thicker in a radial direction and a lateral direction than at the inner end.

3. The gas turbine engine component as set forth in claim 1 , wherein the outer surface is a radially outer platform, the inner surface is a radially inner platform and an airfoil connecting the radially inner platform to the radially outer platform.

4. The gas turbine engine component as set forth in claim 3 , wherein the airfoil has a leading edge and a trailing edge, and the thickened portion is axially forward of the leading edge.

5. The gas turbine engine component as set forth in claim 4 , wherein the outer platform has a pressure side edge and a suction side edge and the thickened portion is adjacent the suction side edge.

6. The gas turbine engine component as set forth in claim 1 , wherein the insert is formed a plurality of layers of ceramic matrix composite layers.

7. The gas turbine engine component as set forth in claim 1 , wherein the insert is formed as a monolithic structure.

8. The gas turbine engine component as set forth in claim 1 , wherein the insert is a 3D woven insert.

9. A gas turbine engine comprising:
a compressor section, a combustor and a turbine section, the turbine section including rotating turbine blades and at least one stationary component mounted adjacent a rotating turbine blade; and
the stationary component having an inner surface and an outer surface, the outer surface and the inner surface being formed of ceramic matrix composite layers, and a mount member, the mount member being in contact with an insert inserted within outer one of the ceramic matrix composite layers of the outer surface to form a thickened portion having an end surface that will react against forces from the mount member.

10. The gas turbine engine as set forth in claim 9 , wherein the insert has with an inner end remote from the end surface, and sides extending outwardly from the inner end such that the end surface is thicker in a radial direction and a lateral direction than at the inner end.

11. The gas turbine engine as set forth in claim 9 , wherein the outer surface is a radially outer platform and the inner surface is a radially inner platform and an airfoil connecting the radially inner platform to the outer platform.

12. The gas turbine engine as set forth in claim 11 , wherein the airfoil has a leading edge and a trailing edge, and the thickened portion is axially forward of the leading edge.

13. The gas turbine engine as set forth in claim 12 , wherein the outer platform has a pressure side edge and a suction side edge and the thickened portion is adjacent the suction side edge.

14. The gas turbine engine as set forth in claim 9 , wherein the insert is formed of a plurality of layers of ceramic matrix composite layers.

15. The gas turbine engine as set forth in claim 9 , wherein the insert is formed as a monolithic structure.

16. The gas turbine engine as set forth in claim 9 , wherein the insert is a 3D woven insert.

17. A method of forming a gas turbine engine component comprising the steps of:
providing a base layer of a plurality of CMC plies;
attaching a rigid insert onto the base layer, the rigid insert has an inner end remote from an end surface, with sides extending outwardly from the inner end such that the insert surface is thicker in a radial direction and a lateral direction than at the inner end;
applying outer cover plies over the rigid insert, but not covering the end surface such that the end surface may contact a mount member when the gas turbine engine component is mounted within a gas turbine engine; and
then densifying the base layers and the cover plies.

18. The method as set forth in claim 17 , wherein the rigid insert is formed of a plurality of CMC plies, and is densified prior to being added to the base layer.

19. The method as set forth in claim 17 , wherein the rigid insert is formed as a monolithic material.

20. The method as set forth in claim 17 , wherein the base layer forms a portion of an outer platform of a vane having an airfoil.

Assignees

• Rtx Corp

Inventors

• Patel Rutva A
• Roach James T

Classifications

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F05D2230/20

  essentially without removing material · CPC title

• F02C7/00

  Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants (controlling F02C9/00) · CPC title

• F01D9/042

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

• F01D5/147

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title