Method for making gas turbine engine ceramic matrix composite airfoil

Patent metadata

Field	Value
Publication number	US-9683443-B2
Application number	US-201314141395-A
Country	US
Kind code	B2
Filing date	Dec 26, 2013
Priority date	Mar 4, 2013
Publication date	Jun 20, 2017
Grant date	Jun 20, 2017

Abstract

A method for making a gas turbine engine ceramic matrix composite airfoil is disclosed. The method includes fabricating an airfoil preform that has a slotted forward end and a continuous trailing end. The slotted forward end of the airfoil preform is coupled to an airfoil core insert. A ceramic matrix composite covering is applied to cover the slots of the airfoil perform. The continuous trailing end of the airfoil preform is removed to expose the slots. A gas turbine engine airfoil is also disclosed.

First claim

What is claimed is:

1. A method comprising:
fabricating an airfoil preform having a slotted forward end and a continuous trailing end;
coupling the slotted forward end of the airfoil preform to an airfoil core insert;
applying a ceramic matrix composite covering to cover the slots of the airfoil preform; and
removing the continuous trailing end of the airfoil preform to expose the slots;
wherein:
the fabricating step comprises providing a cooling passage preform having a spanwise extending trailing end portion and a plurality of flow dividing members projecting from the spanwise extending trailing end portion that define cooling passages therebetween;
the coupling step comprises coupling a cooling delivery core to a forward end of the cooling passage preform to close a forward end of the cooling passages;
the applying step comprises covering the cooling passages with a ceramic matrix material; and
the removing step comprises trimming the spanwise extending trailing end portion of the cooling passage preform to expose the cooling passages.

2. The method of claim 1 in which the airfoil preform comprises a monolithic ceramic.

3. The method of claim 1 in which the slots are substantially perpendicular to the continuous trailing end of the airfoil preform.

4. The method of claim 1 in which the slots extend through the thickness of the airfoil preform.

5. The method of claim 1 in which the slots are formed by machining material from the airfoil preform.

6. The method of claim 1 in which the coupling comprises capturing the slotted forward end of the airfoil preform in a spanwise groove in the trailing end of the airfoil core insert.

7. The method of claim 1 in which adjacent ones of the plurality of flow dividing members define cooling passages therebetween that have a substantially linear configuration in the chordwise direction.

8. The method of claim 1 in which the cooling passages are substantially the same size.

9. The method of claim 1 in which the cooling passages are equally spaced apart in the spanwise direction.

10. The method of claim 1 in which the cooling passages have a substantially uniform width.

11. A method comprising:
fabricating an airfoil preform as a cooling passage preform having a spanwise extending trailing end portion and a plurality of flow dividing members projecting from the spanwise extending trailing end portion towards a forward end and defining cooling passages therebetween;
coupling a cooling delivery core to the forward end of the cooling passage preform to close a forward end of the cooling passages;
covering the cooling passages with a ceramic matrix material; and
trimming the spanwise extending trailing end portion of the cooling passage preform to expose the cooling passages.

12. The method of claim 11 , wherein the airfoil preform comprises a monolithic ceramic.

13. The method of claim 11 , wherein the cooling passages are substantially perpendicular to the trailing end portion of the airfoil preform.

14. The method of claim 11 , wherein the cooling passages extend as slots through the thickness of the airfoil preform.

15. The method of claim 11 , wherein the cooling passages are formed by machining material from the airfoil preform.

16. The method of claim 11 , wherein the coupling comprises capturing the forward end of the airfoil preform in a spanwise groove of the cooling delivery core.

17. The method of claim 11 , wherein the plurality of flow dividing members are substantially linear in the chordwise direction.

18. The method of claim 11 , wherein the cooling passages are substantially the same size.

19. The method of claim 11 , wherein the cooling passages are equally spaced apart.

20. The method of claim 11 , wherein the cooling delivery core is formed as at least a portion of an airfoil core insert.

Assignees

• Rolls Royce Nam Tech Inc

Inventors

• Freeman Ted J
• Uskert Richard C

Classifications

• F01D5/187 Primary

  Convection cooling · CPC title

• Y10T29/49337

  Composite blade · CPC title

• F01D5/186 Primary

  Film cooling (F01D5/187 takes precedence) · CPC title

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title