Turbine vane assembly with ceramic matrix composite components

Patent metadata

Field	Value
Publication number	US-10612399-B2
Application number	US-201815995369-A
Country	US
Kind code	B2
Filing date	Jun 1, 2018
Priority date	Jun 1, 2018
Publication date	Apr 7, 2020
Grant date	Apr 7, 2020

Abstract

A turbine vane assembly adapted for use in a gas turbine engine includes an airfoil, an endwall, and a spar. The airfoil is shaped to interact with hot gases moving axially along a primary gas path of the gas turbine engine. The endwall is shaped to define a boundary of the primary gas path near a radial end of the airfoil. The spar is located in an interior region of the airfoil to carry loads that act on the airfoil during operation of the gas turbine engine.

First claim

What is claimed is:

1. A turbine vane assembly adapted for use in a gas turbine engine, the turbine vane assembly comprising
an airfoil comprising ceramic matrix composite materials and shaped to interact with hot gases moving axially along a primary gas path of the gas turbine engine relative to an axis, the airfoil formed to include a radial-inner wall and a sidewall that extends radially outward and away from a perimeter of the radial-inner wall to define an interior region of the airfoil, and the radial-inner wall formed to define an airfoil passageway that extends radially through the radial-inner wall and opens into the interior region,
an endwall comprising ceramic matrix composite materials and shaped to define a boundary of the primary gas path near a radial end of the airfoil and the endwall formed to define an end-wall passageway that extends radially through the endwall, and
a spar comprising metallic materials and located in the interior region of the airfoil to carry loads that act on the airfoil, the spar including a spar body that engages the radial-inner wall of the airfoil so that an interface between the spar and the airfoil is located radially inward toward the boundary of the primary gas path, a spar tail that extends radially inward away from the spar body through the airfoil passageway and the end-wall passageway, and a retainer coupled to the spar tail to block movement of the radial-inner wall and the endwall away from the spar body.

2. The turbine vane assembly of claim 1 , further comprising an inner panel arranged around the axis and adapted to block fluid communication between a pressurized region located axially upstream of the inner panel and a pressurized region located axially downstream of the inner panel, the inner panel is coupled with the spar tail, and the inner panel is spaced apart radially from the spar body to locate the radial-inner wall and the endwall radially between the spar body and the inner panel.

3. The turbine vane assembly of claim 2 , wherein the inner panel is formed to define a panel passageway that extends through the inner panel and the spar tail extends through the panel passageway such that a portion of the inner panel is located between the spar body and the retainer.

4. The turbine vane assembly of claim 2 , wherein the inner panel includes a body and a flange that extends radially outward away from the body, the flange is formed to define a panel passageway that extends axially through the flange, and the retainer extends through the panel passageway to couple the inner panel with the spar tail.

5. The turbine vane assembly of claim 2 , wherein the spar tail extends radially inward through and beyond the endwall and the inner panel.

6. The turbine vane assembly of claim 1 , wherein the spar tail has threads and the retainer is threaded and mates with the threads of the spar tail.

7. The turbine vane assembly of claim 6 , wherein the airfoil passageway is circular when viewed radially.

8. The turbine vane assembly of claim 6 , further comprising an outer-endwall spaced apart radially from the endwall and shaped to define another boundary of the primary gas path near another radial end of the airfoil.

9. The turbine vane assembly of claim 1 , wherein the spar tail is formed to define a spar passageway that extends at least partway into the spar tail and the retainer extends into the spar passageway.

10. A turbine vane assembly comprising
an airfoil that includes a radial-inner wall formed to define an airfoil passageway that extends radially through the radial-inner wall relative to an axis and a sidewall that extends radially away from the radial-inner wall to define an interior region of the airfoil,
an endwall formed to define an end-wall passageway that extends radially through the endwall, and
a spar that includes a spar body located in the interior region of the airfoil and a spar tail that extends radially inward away from the spar body toward the axis and through the airfoil passageway and the end-wall passageway,
wherein the airfoil passageway is formed in the radial-inner wall of the airfoil at a predetermined location, the sidewall is spaced apart from the spar body to define a gap therebetween, and the spar tail engages the radial-inner wall of the airfoil in the airfoil passageway such that at least a portion of aero loads acting on the airfoil are transmitted through the radial-inner wall at the predetermined location to the spar tail during use of the turbine vane assembly.

11. The turbine vane assembly of claim 10 , further comprising an inner panel arranged at least partway around the axis and coupled with the spar tail.

12. The turbine vane assembly of claim 11 , wherein the inner panel is spaced apart radially from the spar body to locate the radial-inner wall and the endwall radially between the spar body and the inner panel.

13. The turbine vane assembly of claim 11 , wherein the inner panel is formed to define a panel passageway that extends through the inner panel and the spar tail extends through and beyond the panel passageway.

14. The turbine vane assembly of claim 11 , wherein the inner panel includes a body arranged at least partway around the axis and a flange that extends radially outward away from the body, the flange is formed to define a panel passageway that extends axially through the flange, and the spar further includes a retainer that extends through the flange to couple the inner panel with the spar tail.

15. The turbine vane assembly of claim 10 , wherein the spar tail has threads.

16. The turbine vane assembly of claim 10 , wherein the spar tail is formed to define a spar passageway that extends at least partway into the spar tail and the spar includes a retainer that extends into the spar passageway.

17. A method comprising
providing an airfoil that comprises ceramic materials, a first endwall, and a spar that comprises metallic materials, the airfoil includes a radial-inner wall formed to define an airfoil passageway that extends radially through the radial-inner wall relative to an axis and a sidewall that extends radially away from the radial-inner wall to define an interior region of the airfoil, the first endwall formed to define an end-wall passageway that extends radially through the first endwall, and the spar includes a spar body and a spar tail that extends radially away from the spar body, and the spar tail is further formed to define a spar passageway that extends at least partway into the spar tail,
locating the spar body in the interior region of the airfoil such that the spar tail extends through the airfoil passageway,
moving the first endwall relative to the spar such that the spar tail extends through the end-wall passageway, and
inserting a retainer into the spar passageway to couple the airfoil, the end wall, and the spar together.

18. The method of claim 17 , further comprising providing an inner panel and coupling the inner panel to the spar tail to locate the radial-inner wall and the first endwall radially between the spar body and a portion of the inner panel.

19. The method of claim 18 , further comprising providing a second endwall and locating the second endwall around a portion of the airfoil and in radial spaced apart relation relative to the first endwall.

Assignees

• Rolls Royce Nam Tech Inc
• Rolls Royce Corp

Inventors

• Freeman Ted J
• Sippel Aaron D
• Thomas David J

Classifications

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F01D5/284 Primary

  Selection of ceramic materials · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F01D9/042 Primary

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