Method for treating a gas turbine blade and gas turbine having said blade

Patent metadata

Field	Value
Publication number	US-10995625-B2
Application number	US-201916249247-A
Country	US
Kind code	B2
Filing date	Jan 16, 2019
Priority date	Oct 5, 2012
Publication date	May 4, 2021
Grant date	May 4, 2021

Abstract

The use of different ceramic layers allows different configurations of gas turbines to be produced each of which is optimized for a respective use of base load operation or peak load operation.

First claim

What is claimed is:

1. A method for operating a gas turbine system, comprising:
altering turbine blades or vanes of a first gas turbine to produce a second gas turbine, the first turbine blades or vanes having first ceramic thermal barrier coatings configured for base load operation;
removing completely the first ceramic thermal barrier coating of a first turbine blade or vane of the first gas turbine;
applying a new, second ceramic thermal barrier coating configured for peak load operation, different from the base load operation, to the first turbine blade or vane from which the first ceramic thermal barrier coating has been completely removed to produce a second turbine blade or vane by coating with a ceramic and a polymer or with a ceramic with grains of at least 20% greater mean particle diameter to generate pores in the new, second ceramic thermal barrier coating; and
starting the second gas turbine daily instead of continuously operating the second gas turbine;
wherein the new, second ceramic thermal barrier coating differs significantly from the completely removed first ceramic thermal barrier coating, in that porosities of the first and the second ceramic thermal barrier coatings are different, and the absolute difference in the reduced or increased porosity is at least 2%, and difference in coating thicknesses of the first and the second ceramic thermal barrier coatings is at least 50 μm;
wherein the porosity of the completely removed first ceramic thermal barrier coating, and the porosity of the second ceramic thermal barrier coating are in the range 12%±4% to 25%±4%, and
wherein the second turbine blade or vane is incorporated in the second gas turbine.

2. The method as claimed in claim 1 , wherein the completely removed first ceramic thermal barrier coating is a two-layer ceramic thermal barrier coating; with a bottommost layer and an outer layer, the bottommost layer having a porosity in the range 25%±4% and the outer layer having a porosity in the range 12%±4%, and
the second ceramic thermal barrier coating is a single layer thermal barrier coating with a porosity in the range 18%±4%.

3. The method as claimed in claim 1 , wherein the porosity of the second ceramic thermal barrier coating is elevated compared to the porosity of the first ceramic thermal barrier coating.

4. The method as claimed in claim 1 , wherein the porosity of the second ceramic thermal barrier coating is lower than the porosity of the first ceramic thermal barrier coating.

5. The method as claimed in claim 1 , wherein the first ceramic thermal barrier coating is thinner than the second ceramic thermal barrier coating and the difference in the thicknesses is at least +50 μm.

6. The method as claimed in claim 1 , wherein the first ceramic thermal barrier coating is thicker than the second ceramic thermal barrier coating and the difference in the thickness being at least −50 μm.

7. The method as claimed in claim 1 , wherein the second ceramic thermal barrier coating is a single layer coating with a porosity in the range 18%±4%, and the first ceramic thermal barrier coating has a bottommost ceramic layer with a porosity of 12%±4% and an outer ceramic layer with a porosity of 25%±4%, wherein an absolute difference in the porosities of the ceramic layers is at least 2%.

8. The method as claimed in claim 7 , wherein the mean pore diameter of the outer ceramic layer is greater than the mean pore diameter of the bottommost ceramic layer.

9. The method as claimed in claim 1 , wherein the second turbine blade or vane has a ceramic thermal barrier coating consisting of a single layer with a uniform porosity of 18%±4%.

10. The method as claimed in claim 1 , wherein the first turbine blade or vane has a two-layer ceramic thermal barrier coating which consists of a bottommost ceramic layer having a porosity of 12%±4% and a top ceramic layer having a porosity of 25%±4%, and the absolute difference in the porosities of the ceramic layers is at least 2%.

11. The method as claimed in claim 10 , wherein the bottommost ceramic layer comprises partially stabilized zirconium oxide and the top ceramic layer comprises partially stabilized zirconium oxide.

12. The method as claimed in claim 10 , wherein the top ceramic layer has a perovskite or pyrochlore structure, and the bottommost ceramic layer comprises zirconium oxide.

13. The method as claimed in claim 10 , wherein the mean pore diameter of the top ceramic layer is greater than the mean pore diameter of the bottommost ceramic layer, by at least 20 μm.

Assignees

• Siemens Ag

Inventors

• Bullinger Patrick
• Grosshäuser Martin
• Pahl Andreas
• Reiermann Dietmar
• Rindler Michael
• Stamm Werner

Classifications

• B23P6/002

  Repairing turbine components, e.g. moving or stationary blades, rotors, (B23P6/045 takes precedence) · CPC title

• F01D5/288 Primary

  Protective coatings for blades · CPC title

• F01D5/005

  Repairing methods or devices · CPC title

• F05D2230/90

  Coating; Surface treatment · CPC title

• B05D7/50

  Multilayers · CPC title