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 method of making a blade for a gas turbine engine, the method comprising
forming a tip preform having continuous porosity substantially throughout the tip preform,
positioning the tip preform relative to an airfoil preform so that the tip preform extends beyond a distal end of the airfoil preform, the airfoil preform including ceramic material and having continuous porosity substantially throughout the airfoil preform, and
co-infiltrating, after the positioning the tip preform, the tip preform and the airfoil preform to densify the tip preform and the airfoil preform thereby joining the tip preform and the airfoil preform to form the blade.
2. The method of claim 1 , wherein the co-infiltrating comprises densifying the tip preform and the airfoil preform via slurry infiltration and melt infiltration.
3. The method of claim 2 , wherein the co-infiltrating further comprises densifying the tip preform and the airfoil preform via chemical vapor infiltration or chemical vapor deposition prior to slurry infiltration and melt infiltration.
4. The method of claim 1 , wherein the positioning the tip preform relative to the airfoil preform comprises inserting the tip preform into a tip-receiving space that extends into the airfoil preform from the distal end of the airfoil preform.
5. The method of claim 1 , wherein an infiltrated material extends from the airfoil preform to the tip preform.
6. The method of claim 1 , wherein the forming the tip preform comprises mixing tip particles with a substrate of the tip preform, and wherein, after co-infiltrating the tip preform and the airfoil preform, the tip preform defines an abrasive surface of an abrasive tip of the blade.
7. The method of claim 6 , wherein the tip particles comprise silicon carbide.
8. The method of claim 6 , wherein the tip particles comprise cubic boron nitride, silicon nitride or a combination thereof, and the forming the tip preform comprises encapsulating the tip particles and the substrate in a polymer.
9. The method of claim 8 , wherein the co-infiltrating the substrate of the tip preform and the airfoil preform includes burning off the polymer.
10. The method of claim 1 , wherein the airfoil preform is substantially silicon carbide based fibers.
11. The method of claim 1 , further comprising positioning the blade adjacent to an abradable coating such that the tip rubs the abradable coating under some operating conditions of the gas turbine engine.
12. A method comprising:
forming a tip preform having continuous porosity substantially throughout the tip preform by at least:
suspending tip particles in a substrate or material comprising at least one of chopped fibers, continuous fibers, woven fabrics, or ceramic foam;
positioning the tip preform relative to an airfoil preform so that the tip preform extends beyond a distal end of the airfoil preform, the airfoil preform including ceramic material and having continuous porosity substantially throughout the airfoil preform, and
co-infiltrating, after the positioning the tip preform, the tip preform and the airfoil preform to densify the tip preform and the airfoil preform thereby joining the tip preform and the airfoil preform to form the blade, wherein the infiltrated tip preform defines a distal-most surface of the blade, and wherein the distal-most surface of the blade comprises the tip particles.
13. The method of claim 12 , further comprising positioning the blade adjacent to an abradable coating such that the tip rubs the abradable coating under some operating conditions of the gas turbine engine.
14. The method of claim 12 , wherein positioning the tip preform relative to the airfoil preform comprises inserting the tip preform into a tip-receiving space that extends into the airfoil preform from the distal end of the airfoil preform, wherein the tip preform extends from inside the tip receiving space to beyond the distal end of the airfoil.