ABRASIVE TIPS FOR CERAMIC MATRIX COMPOSITE BLADES AND METHODS FOR MAKING THE SAME

• US 20160177745A1
• Filed: 12/02/2015
• Published: 06/23/2016
• Est. Priority Date: 12/18/2014
• Status: Active Grant

• Current Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc), Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc), Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Inventors: Uskert, Richard C., Lee, Kang N., Freeman, Ted J.
• Granted Patent: US 10,202,854 B2
• Time in Patent Office: Days
• US Class Current: 1/1
• CPC Class Codes: C23C 16/045 Coating cavities or hollow spaces, e.g. interior of tubes; Infiltration of porous substrates
• CPC Class Codes: F01D 11/12 using a rubstrip, e.g. erodible. deformable or resiliently-biased part
• CPC Class Codes: F01D 5/14 Form or construction selecting particular materials, measures against erosion or corrosion F01D5/28
• CPC Class Codes: F01D 5/147 Construction, i.e. structural features, e.g. of weight-saving hollow blades F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18
• CPC Class Codes: F01D 5/282 Selecting composite materials, e.g. blades with reinforcing filaments
• CPC Class Codes: F01D 5/284 Selection of ceramic materials
• CPC Class Codes: F01D 5/288 Protective coatings for blades
• CPC Class Codes: F05D 2230/314 by chemical vapour deposition
• CPC Class Codes: F05D 2300/2261 of silicon
• CPC Class Codes: F05D 2300/2282 of boron
• CPC Class Codes: F05D 2300/43 Synthetic polymers, e.g. plastics; Rubber

Abstract

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.

Claims

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 from 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.

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 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.

12. The blade of claim 1 wherein the infiltrated material of the airfoil is substantially identical to the infiltrated material of the tip.

11. A blade for a gas turbine engine, the blade comprising

an airfoil comprising a ceramic material, the airfoil having a distal end,

a tip comprising a porous structure, the tip extending beyond the distal end of the of the airfoil, and

infiltrated material that extends from the airfoil to the tip to join the airfoil and 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 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.

15. The blade of claim 14, wherein the airfoil includes a hollow core in fluid communication with the tip-receiving space.

16. The blade of claim 11, wherein the tip further comprises tip particles suspended in a substrate.

17. The blade of claim 16, wherein the tip particles include silicon carbide.

18. The blade of claim 16, wherein the tip particles are selected from cubic boron nitride, silicon nitride, or a combination thereof.

19. The blade of claim 11, wherein a porosity of the tip is between about 1 percent and about 10 percent by volume.

20. The blade of claim 11, wherein the tip has a shape that generally conforms to the shape of the airfoil when viewed from beyond the distal end of the airfoil.