Method for producing and restoring ceramic heat insulation coatings in gas turbines and associated gas turbine

• US 10,513,935 B2
• Filed: 12/18/2012
• Issued: 12/24/2019
• Est. Priority Date: 03/28/2012
• Status: Active Grant

• Current Assignee: Siemens Energy Global GmbH & Co. KG (Siemens AG)
• Original Assignee: Siemens AG
• Inventors: Bullinger, Patrick, Grosshauser, Martin, Pahl, Andreas, Reiermann, Dietmar, Rindler, Michael, Stamm, Werner
• Primary Examiner(s): Wilensky, Moshe
• Assistant Examiner(s): Prager, Jesse M
• Application Number:
  US14/388,888
• Publication Number:
  US 20150071772A1
• Time in Patent Office: 2,562 Days
• CPC Class Codes: B05D 3/002 Pretreatement
• CPC Class Codes: C23C 28/3215 at least one MCrAlX layer
• CPC Class Codes: C23C 28/345 with at least one oxide layer
• CPC Class Codes: C23C 28/3455 with a refractory ceramic layer, e.g. refractory metal oxide, ZrO2, rare earth oxides or a thermal barrier system comprising at least one refractory oxide layer
• CPC Class Codes: C23C 4/02 Pretreatment of the material to be coated, e.g. for coating on selected surface areas
• CPC Class Codes: C23C 4/11 Oxides
• CPC Class Codes: F01D 5/005 Repairing methods or devices
• CPC Class Codes: F01D 5/284 Selection of ceramic materials
• CPC Class Codes: F01D 5/288 Protective coatings for blades
• CPC Class Codes: F01D 9/02 Nozzles; Nozzle boxes; Stator blades; Guide conduits , e.g. individual nozzles nozzle boxes F01D9/047
• CPC Class Codes: F05D 2230/80 Repairing, retrofitting or upgrading methods
• CPC Class Codes: F05D 2230/90 Coating; Surface treatment manufacture with deposition of material F05D2230/30
• CPC Class Codes: F05D 2260/231 Preventing heat transfer
• CPC Class Codes: F05D 2300/2118 Zirconium oxides
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

Through the use of different ceramic heat insulation coatings (7, 13) on turbine blades, it is possible to produce or restore different configurations of gas turbines, which are then optimized for a respective baseload operation application or peak load operation application. Differences in the sprayed-on heat insulation coatings (7, 13) can be generated by means of a different material composition of the powder, for example zirconium oxide, by means of a different number of layers (10), by means of different coating thicknesses and/or by means of different porosities.

Claims

1. A process for operating a second gas turbine using first turbine blades or vanes of a first gas turbine comprising:

providing first turbine blades or vanes of the first gas turbine having either a single-layer first ceramic thermal barrier coating previously operated at peak load operations or a two-layer first ceramic thermal barrier coating previously operated at base load operations,

removing the first ceramic thermal barrier coating from the first turbine blades or vanes of the first gas turbine to produce uncoated first turbine blades or vanes;

applying a second ceramic thermal barrier coating to the uncoated first turbine blades or vanes to produce second turbine blades or vanes;

operating the second turbine blades or vanes at base load operations with a bottom layer having a porosity of 12% to 16% and a top layer having a porosity of 18%±

4%, or the bottom layer having a porosity of 18%±

4% and the top layer having a porosity of 25%±

4%, when the removed first ceramic thermal barrier coating was the single-layer ceramic barrier coating configured for peak load operations,

wherein, when applied, at least the top layer is comprised of partially stabilized zirconium oxide,

operating the second turbine blades or vanes at peak load operations with a uniform porosity of 18%±

4% when the removed first ceramic thermal barrier coating was a two-layer first ceramic thermal barrier coating configured for base load operations.

2. The process as claimed in claim 1, wherein the absolute difference in the porosity of the bottom and the top layers is at least 2%, when the two-layer second ceramic thermal barrier coating is applied,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.

3. The process as claimed in claim 1, wherein, when applied, the bottom layer of the applied two-layer second ceramic thermal barrier coating is thinner than the top layer thereof;

and

the bottom layer of the applied two-layer second ceramic thermal barrier coating has a thickness of 75 μ

m to 150 μ

m,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.

4. The process as claimed in claim 1, wherein the bottom layer is comprised of partially stabilized zirconium oxide.

5. The process as claimed in claim 1, wherein the zirconium oxide is sprayed with a powder having a monoclinic proportion of less than 3%.

6. The process as claimed in claim 4, wherein a tetragonal proportion has the greatest proportion of zirconium oxide.

7. The process as claimed in claim 5, further comprising applying a heat treatment to reduce the monoclinic proportion of the zirconium oxide powder to be sprayed by at least 50%.

8. The process as claimed in claim 1, wherein, when applied, the applied two-layer second ceramic thermal barrier coating is applied by spraying a bottom layer without a polymer and spraying a top layer with a polymer,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.

9. The process as claimed in claim 1, wherein, when applied, the top layer of the applied two-layer second ceramic thermal barrier coating has pores with a mean pore diameter and the bottom ceramic layer has pores with a mean pore diameter, and the mean pore diameter of the top layer is greater than the mean pore diameter of the bottom layer.

10. The process as claimed in claim 1, wherein, when applied, a same powder with a same composition and a same grain size distribution is used in applying the two-layer second ceramic thermal barrier coating,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.

11. The process as claimed in claim 1, wherein, when applied, the bottom ceramic layer and the top ceramic layer of the applied two-layer second ceramic thermal barrier coating are comprised of different materials,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.

12. The process as claimed in claim 2, wherein the absolute difference in the porosity of the top and the bottom layers is at most 4%,

wherein the first turbine blades or vanes have a single layer first ceramic thermal barrier coating.