Abrasive tips for ceramic matrix composite blades and methods for making the same
A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.
1. A blade for a gas turbine engine, the blade comprising
an airfoil comprising a ceramic material defining continuous porosity, the airfoil having a distal end;
a tip comprising a porous tip preform, the tip extending beyond the distal end of the of the airfoil; and
infiltrated material that fills the continuous porosity of the ceramic material of the airfoil and extends through the airfoil and the porous tip preform to join the airfoil and the tip, wherein a perimeter of the tip conforms to a perimeter of the distal end of the airfoil when viewed from beyond the distal end of the airfoil.
2. The blade of claim 1 , wherein the infiltrated material of the airfoil is the same as the infiltrated material of the tip.
3. The blade of claim 1 , wherein the ceramic material of the airfoil comprises silicon carbide, and the tip comprises silicon carbide.
4. The blade of claim 1 , wherein the airfoil includes a tip-receiving space that extends into the airfoil and a portion of the tip is arranged in the tip-receiving space of the airfoil.
5. The blade of claim 4 , wherein the airfoil includes a hollow core in fluid communication with the tip-receiving space.
6. The blade of claim 1 , wherein the tip further comprises tip particles suspended in a substrate.
7. The blade of claim 6 , wherein the tip particles include silicon carbide.
8. The blade of claim 6 , wherein the tip particles are selected from cubic boron nitride, silicon nitride, or a combination thereof.
9. The blade of claim 1 , wherein a porosity of the tip is between about 1 percent and about 10 percent by volume.
10. The blade of claim 1 , wherein the tip defines at least one cooling hole, wherein the airfoil defines a core, and wherein the core is in fluid communication with the at least one cooling hole.
11. A blade for a gas turbine engine, the blade comprising
an airfoil comprising a ceramic reinforcement material defining spaces between the ceramic reinforcement material, the airfoil having a distal end defining a tip receiving space;
a tip comprising a porous structure defining spaces between material of the porous structure, the tip being received in the tip receiving space and extending beyond the distal end of the of the airfoil; and
infiltrated material that extends from the spaces of the ceramic reinforcement material to the spaces of porous structure to join the airfoil and the tip, wherein a perimeter of the tip conforms to a perimeter of the distal end of the airfoil when viewed from beyond the distal end of the airfoil.
12. The blade of claim 11 , wherein the infiltrated material of the airfoil is the same as the infiltrated material of the tip.
13. The blade of claim 11 , wherein the ceramic material of the airfoil comprises silicon carbide, and the tip comprises silicon carbide.
14. The blade of claim 11 , wherein the tip further comprises tip particles suspended in the porous structure.
15. The blade of claim 14 , wherein the tip particles include silicon carbide.
16. The blade of claim 14 , wherein the tip particles are selected from cubic boron nitride, silicon nitride, or a combination thereof.
17. The blade of claim 11 , wherein the tip defines at least one cooling hole, wherein the airfoil defines a core, and wherein the core is in fluid communication with the at least one cooling hole.