Gas turbine engine component having an airfoil with internal cross-ribs

Patent metadata

Field	Value
Publication number	US-12196095-B2
Application number	US-202218071121-A
Country	US
Kind code	B2
Filing date	Nov 29, 2022
Priority date	Nov 29, 2022
Publication date	Jan 14, 2025
Grant date	Jan 14, 2025

Abstract

A gas turbine engine component includes an airfoil body extending between a leading edge and a trailing edge and having a suction wall and a pressure wall. An outer surface of the airfoil body is formed by an outer coat defining around the pressure and suction sides, a leading edge and a trailing edge. An internal cross-rib is formed within a cavity in the airfoil body. The internal cross-rib extends to be secured to the outer coat adjacent the leading edge and the trailing edge, and extending across the internal cavity to form a junction such that the cross-rib is x-shaped. The outer coat and the cross-rib are formed of ceramic matrix composites. A gas turbine engine is also disclosed.

First claim

What is claimed is:

1. A gas turbine engine component comprising:
an airfoil body extending between a leading edge and a trailing edge and having a suction wall and a pressure wall, an outer surface of the airfoil body formed by an outer coat defining around the pressure and suction sides, a leading edge and a trailing edge;
an internal cross-rib formed within a cavity in the airfoil body, said internal cross-rib extending to be secured to the outer coat adjacent the leading edge and the trailing edge, and extending across the internal cavity to form a junction such that the cross-rib is x-shaped;
said outer coat and said cross-rib formed of ceramic matrix composites; and
the internal cross rib has a portion that bends across the entire leading edge and another portion that extends across the entire trailing edge.

2. The component as set forth in claim 1 , wherein said component is a static vane for use in a turbine section of a gas turbine engine.

3. The component as set forth in claim 1 , wherein major chambers are defined between said junction and said leading edge and said trailing edge, and minor chambers are formed in the cavity between said junction and an inner surface of said pressure side and inner surface of said suction side, and there is a cooling air supply supplying cooling air into the major chambers and the minor chambers.

4. The component as set forth in claim 3 , wherein a pressure of the cooling air supplied into the major air chambers is equal to a pressure of the cooling air supplied into the minor chambers.

5. The component as set forth in claim 3 , wherein a pressure of the cooling air supplied into the minor chambers is less than a pressure of the cooling air supplied into the major chambers.

6. The component as set forth in claim 5 , wherein the pressure of the cooling air supplied into the minor chambers is between 10% and 90% of the pressure of the cooling air supplied into the major chambers.

7. The component as set forth in claim 3 , wherein there are at least two of said cross-ribs in the cavity.

8. The component as set forth in claim 7 , wherein the major chambers are defined between said junction on one of the cross-ribs and said leading edge, between said junction on said one of the cross-ribs and the junction on the other of said cross-ribs, and the junction on the other of said cross-ribs and said trailing edge, and the minor chambers are formed in the cavity between each said junction and an inner surface of said pressure side and inner surface of said suction side, and there is a cooling air supply supplying cooling air into the major chambers and the minor chambers.

9. The component as set forth in claim 8 , wherein a pressure of the cooling air supplied into the minor chambers is less than a pressure of the cooling air supplied into the major chambers.

10. The component as set forth in claim 1 , wherein said cross-rib is formed by woven plies which are woven together to form a shape, and then densified with an injection of a matrix.

11. A gas turbine engine comprising:
a compressor section for delivering air into a combustor, a turbine section positioned downstream of the combustor to receive products of combustion from the combustor;
the turbine section comprising rows of circumferentially spaced static vanes, axially spaced from rows of rotating turbine blades, said static vanes and said turbine blades both being formed with an airfoil;
the airfoils in at least one of said static vanes and said turbine blades having an airfoil body extending between a leading edge and a trailing edge and having a suction wall and a pressure wall, an outer surface of the airfoil body formed by an outer coat defining the pressure and suction sides, a leading edge and a trailing edge;
an internal cross-rib formed within a cavity in the airfoil body, said internal cross-rib extending to be secured to the outer coat adjacent the leading edge and the trailing edge, and extending across the internal cavity to form a junction such that the cross-rib is x-shaped; and
said outer coat and said cross-rib formed of ceramic matrix composites; and
the internal cross rib has a portion that bends across the entire leading edge and another portion that extends across the entire trailing edge.

12. The gas turbine engine as set forth in claim 11 , wherein said airfoil is part of said static vanes.

13. The gas turbine engine as set forth in claim 11 , wherein major chambers are defined between said junction and said leading edge and said trailing edge, and minor chambers are formed in the cavity between said junction and an inner surface of said pressure side and inner surface of said suction side, and there is a cooling air supply supplying cooling air into the major chambers and the minor chambers.

14. The gas turbine engine as set forth in claim 13 , wherein a pressure of the cooling air supplied into the major air chambers is equal to a pressure of the cooling air supplied into the minor chambers.

15. The gas turbine engine as set forth in claim 13 , wherein a pressure of the cooling air supplied into the minor chambers is less than a pressure of the cooling air supplied into the major chambers.

16. The gas turbine engine as set forth in claim 15 , wherein the pressure of the cooling air supplied into the minor chambers is between 10% and 90% of the pressure of the cooling air supplied into the major chambers.

17. The gas turbine engine as set forth in claim 13 , wherein there are at least two of said cross-ribs in the cavity.

18. The gas turbine engine as set forth in claim 17 , wherein the major chambers are defined between said junction on one of the cross-ribs and said leading edge, between said junction on said one of the cross-ribs and the junction on the other of said cross-ribs, and the junctions on the other of said cross-ribs and said trailing edge, and the minor chambers are formed in the cavity between said junction and an inner surface of said pressure side and inner surface of said suction side, and there is a cooling air supply supplying cooling air into the major chambers and the minor chambers.

19. The gas turbine engine as set forth in claim 18 , wherein a pressure of the cooling air supplied into the minor chambers is less than a pressure of the cooling air supplied into the major chambers.

20. The gas turbine engine as set forth in claim 11 , wherein said cross-rib is formed by woven plies which are woven together to form a shape, and then densified with an injection of a matrix.

Assignees

• Raytheon Tech Corp
• Rtx Corp

Inventors

• Roach James T
• Banhos Jonas
• Kim Russell

Classifications

• F05D2260/20

  Heat transfer, e.g. cooling · CPC title

• F05D2240/30

  Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor · CPC title

• F05D2240/126

  Baffles or ribs · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F01D9/02

  Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} · CPC title