Machinable CMC insert

1. An assembly for use in a gas turbine engine, the assembly comprising

a ceramic matrix composite component comprising silicon carbide fibers in a silicon carbide matrix,

a ceramic insert adjacent to the ceramic matrix composite component, and

a ply at least partially covering the ceramic insert such that the ceramic insert is sandwiched between the ply and the ceramic matrix composite component,

wherein the ply extends beyond the ceramic insert in at least one direction so that the ply is joined to the ceramic matrix composite, and

wherein the ply comprises at least one layer of silicon carbide fibers or carbon fibers in a silicon carbide matrix.

2. The assembly of claim 1 , wherein the ceramic insert comprises a resin char comprising silicon ceramic material, chopped carbon fibers, chopped silicon carbide fibers, or a combination thereof.

3. The assembly of claim 1 , wherein the ceramic insert comprises a pressed powder comprising silicon carbide powder and a binding agent.

4. The assembly of claim 1 , wherein ceramic insert comprises a ceramic material comprising silicon carbide.

5. The assembly of claim 1 , wherein the ceramic insert comprises a reticulated foam having a continuous porosity.

6. The assembly of claim 1 , wherein the ceramic matrix composite component comprises a blade track, a vane, a blade, a combustor tile, an exhaust tail cone, an exhaust liner flap, or a combination thereof.

7. The assembly of claim 1 , wherein in the ceramic matrix composite component comprises a recess along an exterior surface of the ceramic matrix composite component, wherein the ceramic insert is positioned in the recess.

8. The assembly of claim 7 , wherein the ceramic matrix composite component comprises a blade track segment comprising the recess along the exterior surface.

9. The assembly of claim 8 , wherein the ceramic insert comprises a dovetail post arranged to mate with the recess of the blade track segment.

10. The assembly of claim 8 , wherein the ceramic insert comprises a positioning tab for mating with the recess of the blade track segment.

11. A method of joining an insert to a ceramic matrix composite component for use in a gas turbine engine, the method comprising

providing a ceramic preform comprising silicon carbide fibers,

positioning a ceramic insert adjacent to the ceramic preform,

covering at least a portion of the ceramic insert and at least a portion of the ceramic preform with a ply comprising silicon carbide fibers, and

co-infiltrating the ceramic preform and the ply with silicon metal or silicon alloy to form a silicon carbide matrix that extends through the ceramic preform and the ply.

12. The method of claim 11 , wherein co-infiltrating comprises chemical vapor infiltration, chemical vapor deposition, slurry infiltration, polymer impregnation and pyrolysis, or a combination thereof.

13. The method of claim 11 , further comprising depositing a pre-ceramic polymer along an exterior surface of the ceramic preform, wherein positioning the ceramic insert comprises sandwiching the pre-ceramic polymer between the ceramic insert and the ceramic preform.

14. The method of claim 13 , wherein the pre-ceramic polymer further comprises a ceramic phase, SiC, a transition metal, a transition metal boride, a transition metal silicide, or combinations thereof.

15. The method of claim 13 , further comprising heating the assembly to char the pre-ceramic polymer and produce a solid ceramic material.

16. The method of claim 15 , wherein heating is performed at a temperature between about 1300° C. and about 1500° C.

17. The method of claim 16 , further comprising infiltrating the ceramic preform and the ceramic insert with silicon metal, liquid silicon, silicon alloy, silicon carbide, or a combination thereof.

18. The method of claim 13 , further comprising machining the ceramic insert.