MULTI-COMPONENT ASSEMBLY CASTING

• US 20110299999A1
• Filed: 06/07/2010
• Published: 12/08/2011
• Est. Priority Date: 06/07/2010
• Status: Active Grant

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Energy Incorporated (Siemens AG)
• Inventors: James, Allister W.
• Granted Patent: US 9,156,086 B2
• Time in Patent Office: Days
• US Class Current: 416/97.R
• CPC Class Codes: B22C 9/04 Use of lost patterns
• CPC Class Codes: B22D 19/00 Casting in, on, or around objects which form part of the product B22D23/04 takes precedence; alumino-thermic welding B23K23/00; coating by casting molten material on the substrate C23C6/00
• CPC Class Codes: C22C 19/03 based on nickel
• 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 5/187 Convection cooling
• CPC Class Codes: F01D 5/28 Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
• CPC Class Codes: F05D 2230/211 by precision casting, e.g. microfusing or investment casting
• CPC Class Codes: F05D 2300/607 Monocrystallinity
• CPC Class Codes: Y10T 29/4932 Turbomachine making
• CPC Class Codes: Y10T 29/49341 with cooling passage

Abstract

Multi-component vane segment and method for forming the same. Assembly includes: positioning a pre-formed airfoil component (12) and a preformed shroud heat resistant material (18) in a mold, wherein the airfoil component (12) and the shroud heat resistant material (18) each comprises an interlocking feature (24); preheating the mold; introducing molten structural material (46) into the mold; and solidifying the molten structural material such that it interlocks the pre-formed airfoil component (12) with respect to the preformed shroud heat resistant material (18) and is effective to provide structural support for the shroud heat resistant material (18). Surfaces between the airfoil component (12) and the structural material (46), between the airfoil component (12) and the shroud heat resistant material (18), and between the shroud heat resistant material (18) and the structural material (46) are free of metallurgical bonds.

Claims

1. A vane segment comprising;

an airfoil; and

a shroud segment comprising a shroud heat resistant material and a shroud structural material underlying and providing structural support for the shroud heat resistant material, wherein the shroud structural material fixes the shroud heat resistant material relative to the airfoil via geometric interlocking features there between,

wherein surfaces between the airfoil and the shroud structural material, between the airfoil and the shroud heat resistant material, and between the shroud heat resistant material and the shroud structural material are free of metallurgical bonds.

2. The vane segment of claim 1, wherein a heat exposed surface of the shroud heat resistant material intersects the airfoil without a fillet.

3. The vane segment of claim 1, wherein the shroud structural material has superior structural properties but inferior thermal properties compared to the shroud heat resistant material.

4. The vane segment of claim 1, wherein the airfoil comprises an airfoil spar underlying and providing structural support for an airfoil heat resistant material, wherein the airfoil spar fixes the airfoil heat resistant material relative to the shroud heat resistant material via geometric interlocking features.

5. The vane segment of claim 1, wherein at least one of the shroud heat resistant material and the airfoil comprises an oxide layer on a surface abutting a respective structural material effective to prevent a metallurgical bond between the abutting surfaces.

6. The vane segment of claim 4, comprising a cooling channel formed between surfaces of the airfoil spar and the airfoil heat resistant material.

7. The vane segment of claim 1, comprising a cooling channel formed between surfaces of the shroud structural material and the shroud heat resistant material.

8. The vane segment of claim 1, wherein the airfoil and the shroud heat resistant material are different.

9. The vane segment of claim 1, wherein the airfoil is a monolithic single crystal superalloy and the shroud heat resistant material is an oxide dispersion strengthened alloy.

10. The vane segment of claim 1, wherein the airfoil is a monolithic single crystal superalloy and the shroud heat resistant material is a single crystal superalloy.

11. The vane segment of claim 4, wherein the airfoil heat resistant material and the shroud heat resistant material are different.

12. The vane segment of claim 11, wherein the airfoil heat resistant material is a single crystal superalloy and the shroud heat resistant material is comprised of an oxide dispersion strengthened alloy.

13. The vane segment of claim 11, wherein the airfoil heat resistant material is a single crystal superalloy and the shroud heat resistant material is a single crystal superalloy.

14. The vane segment of claim 1, wherein the shroud heat resistant material is an oxide dispersion strengthened alloy, the vane segment comprising a thermal barrier coating applied to the oxide dispersion strengthened alloy.

15. The vane segment of claim 14, comprising a bonding layer between the thermal barrier coating and the oxide dispersion strengthened alloy.

16. A gas turbine engine comprising the vane segment of claim 1.

17. A method for forming a vane segment comprising:

positioning a pre-formed airfoil component and a preformed shroud heat resistant material in a mold, wherein the airfoil component and the shroud heat resistant material each comprises an interlocking feature;

preheating the mold;

introducing molten structural material into the mold; and

solidifying the molten structural material such that it interlocks the pre-formed airfoil component with respect to the preformed shroud heat resistant material and is effective to provide structural support for the shroud heat resistant material,

wherein surfaces between the airfoil component and the structural material, between the airfoil component and the shroud heat resistant material, and between the shroud heat resistant material and the structural material are free of metallurgical bonds, and

wherein a heat exposed surface of the shroud heat resistant material intersects the airfoil component without a fillet.

18. The method of claim 17, wherein solidifying the molten structural material occurs in a manner that prevents melting of exposed surfaces of the airfoil component and the shroud heat resistant material.

19. The method of claim 18, wherein preheating the mold occurs in a manner that forms an oxide layer on a surface of at least one of the airfoil component and the shroud heat resistant material, wherein the oxide layer is effective to prevent a metallurgical bond.

20. The method of claim 17, comprising placing fugitive material such that a void that remains once the fugitive material is removed forms a cooling channel in the vane segment, and the method comprises removing the fugitive material.

21. The method of claim 20, wherein the cooling channel is formed into the structural material and an abutting heat resistant material.

22. The method of claim 17, wherein the airfoil component and the shroud heat resistant material are different.

23. The method of claim 17, wherein the airfoil component is a monolithic airfoil.

24. The method of claim 23, wherein the airfoil component is a single crystal superalloy and the shroud heat resistant material is an oxide dispersion strengthened alloy.

25. The method of claim 23, wherein the airfoil component is a single crystal superalloy and the shroud heat resistant material is a single crystal superalloy.

26. The method of claim 17, wherein the airfoil component is an airfoil heat resistant material and the structural material serves as an airfoil spar and further fixes the airfoil heat resistant material with respect to the shroud heat resistant material.

27. The method of claim 26, wherein the airfoil heat resistant material is a single crystal structure and the shroud heat resistant material is an oxide dispersion strengthened alloy.

28. The method of claim 26, wherein the airfoil heat resistant material is a single crystal superalloy and the shroud heat resistant material is a single crystal superalloy.

29. The method of claim 17, wherein the shroud heat resistant material is an oxide dispersion strengthened alloy, the method comprising coating the shroud heat resistant material with a thermal barrier coating.

30. The method of claim 29, comprising applying a bonding layer between the shroud heat resistant material and the thermal barrier coating.

31. The method of claim 17, wherein the preformed components are metal alloys, and wherein the structural material is a metal alloy.

32. A method for making a gas turbine engine comprising the method of claim 17.