Turbine component thermal barrier coating with depth-varying material properties

Patent metadata

Field	Value
Publication number	US-10323533-B2
Application number	US-201515121196-A
Country	US
Kind code	B2
Filing date	Feb 18, 2015
Priority date	Feb 25, 2014
Publication date	Jun 18, 2019
Grant date	Jun 18, 2019

Abstract

A thermal barrier coating (TBC) with depth-varying material properties is formed on a turbine component. Exemplary depth-varying material properties include physical ductility, strength and thermal resistivity that vary from the TBC layer inner to outer surface. Exemplary ways to modify physical properties include application of plural separate overlying layers of different material composition or by varying the applied material composition during the application of the TBC layer. Various embodiment described herein also apply a calcium-magnesium-aluminum-silicon (CMAS)-retardant material over the TBC layer to retard reaction with or adhesion of CMAS containing combustion particulates to the TBC layer. In other embodiments the CMAS retardant material is also applied within engineered groove features (EGFs) that are formed in the TBC surface.

First claim

What is claimed is:

1. A combustion turbine component having a heat insulating outer surface for exposure to combustion gas, comprising:
a metallic substrate having a substrate surface;
an anchoring layer built upon the substrate surface;
a thermal barrier coat (TBC) layer having a TBC total thickness, a TBC inner surface coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas, the TBC layer comprising a progressively decreasing fracture toughness, elastic modulus, and thermal conductivity properties from the TBC inner surface to the TBC outer surface and an increasing porosity from the TBC inner surface to the TBC outer surface;
a planform pattern of engineered surface features (ESFs) projecting from the anchoring layer having projection height between approximately 2-75 percent of the TBC layer total thickness; and
a planform pattern of engineered groove features (EGFs) formed into and penetrating the previously applied TBC layer through the TBC outer surface, having a groove depth.

2. The component of claim 1 , the anchoring layer further comprising a lower thermal barrier coat (LTBC) layer portion defining the planform pattern of ESFs and the thermal barrier coat (TBC) layer further comprising an outer thermal barrier coat (OTBC) layer portion separately applied over the LTBC, having an OTBC inner surface coupled to the LTBC and an OTBC outer surface for exposure to combustion gas;
the LTBC layer portion having greater fracture toughness and elastic modulus than the OTBC layer portion; and
the OTBC layer portion having greater porosity and lower thermal conductivity than the LTBC layer portion.

3. The component of claim 2 , further comprising a calcium magnesium-aluminum-silicon (CMAS)-retardant layer applied over the OTBC outer surface and into the EGFs.

4. The component of claim 1 , the anchoring layer further comprising a bond coat layer coupled to the substrate and the ESFs formed in the bond coat layer.

5. The component of claim 1 , the anchoring layer further comprising a bond coat (BC) layer coupled to a featureless substrate and the ESFs formed in the BC layer.

6. The component of claim 5 , the anchoring layer further comprising a rough bond coat layer applied over the BC layer.

7. A combustion turbine engine comprising the component of claim 1 , the component TBC outer surface in communication with a combustion path of the engine for exposure to combustion gas.

8. The combustion turbine engine of claim 7 , the component ESFs defining an aggregate surface area at least 20 percent greater than an equivalent flat surface.

9. A method for making a combustion turbine component having a heat insulating outer surface for exposure to combustion gas, comprising:
providing a metallic substrate having a substrate surface;
building an anchoring layer upon the substrate surface;
forming a thermal barrier coat (TBC) having a TBC layer thickness, an inner surface coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas; and
varying composition of the TBC layer material progressively as the TBC layer is being continuously applied over the anchoring layer by progressively decreasing fracture toughness, elastic modulus and thermal conductivity and progressively increasing porosity as the TBC layer is being applied over the anchoring layer.

10. The method of claim 9 , further comprising forming a planform pattern of engineered groove features (EGFs) penetrating the previously applied TBC layer through the TBC outer surface, having a groove depth.

11. The method of claim 10 , further comprising thermally spraying a calcium magnesium-aluminum-silicon (CMAS)-retardant layer over the TBC outer surface and into the EGFs.

12. The method of claim 9 , further comprising forming a planform pattern of engineered groove features (EGFs) penetrating the previously applied TBC layer through the TBC outer surface, having a groove depth.

13. The method of claim 12 , further comprising thermally spraying a calcium magnesium-aluminum-silicon (CMAS)-retardant layer over the TBC outer surface and into the EGFs.

14. A method for making a combustion turbine component having a heat insulating outer surface for exposure to combustion gas, comprising:
providing a metallic substrate having a substrate surface;
building an anchoring layer upon a substrate surface of a metallic substrate, the substrate surface including a planform pattern of engineered surface features (ESFs) projecting from the anchoring layer;
forming a thermal barrier coat (TBC) having a TBC layer thickness, an inner surface coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas, by progressively decreasing fracture toughness, elastic modulus and thermal conductivity and progressively increasing porosity as the TBC layer is being applied over the anchoring layer; and
forming a planform pattern of engineered groove features (EGFs) penetrating the previously applied TBC layer through the TBC outer surface, having a groove depth.

15. The method of claim 14 , the anchoring layer forming further comprising:
applying a lower thermal barrier coat (LTBC) layer portion defining the planform pattern of ESFs; and
the thermal barrier coat (TBC) layer further comprising an outer thermal barrier coat (OTBC) layer portion separately applied over the LTBC, having an OTBC inner surface coupled to the LTBC and an OTBC outer surface for exposure to combustion gas;
the LTBC layer portion having greater fracture toughness and elastic modulus than the OTBC layer portion; and
the OTBC layer portion having greater porosity and lower thermal conductivity than the LTBC layer portion.

Assignees

• Siemens Ag

Inventors

• Hitchman Neil
• Subramanian Ramesh
• Schillig Cora

Classifications

• F01D5/288

  Protective coatings for blades · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2230/90

  Coating; Surface treatment · CPC title

• F05D2230/311

  by torch or flame spraying · CPC title

• C23C4/12

  characterised by the method of spraying · CPC title