Application of dense vertically cracked and porous thermal barrier coating to a gas turbine component
View Patent ↗A configuration for coating a turbine component such as a blade or vane with various forms of thermal barrier coating to provide enhanced temperature capability and increased strain tolerance is disclosed. A gas path surface of the platform, airfoil and airfoil fillet region are first coated with a bond coating. A dense vertically cracked (DVC) thermal barrier coating is then applied to at least the gas path surface of the platform and can be applied to the fillet region. A porous thermal barrier coating is then applied to at least the airfoil. The porous thermal barrier coating can also be applied over the DVC thermal barrier coating if desired.
1. A method of applying thermal barrier coatings to a turbine component having at least one platform and an airfoil, the method comprising:
identifying a first, second, and third areas of the turbine component requiring a bond coating and a thermal barrier coating;
applying the bond coating to the first, second, and third areas;
applying directly over the bond coating of the first area only a dense vertically cracked thermal barrier coating;
applying directly over the bond coating of the second area a dense vertically cracked thermal barrier coating; and,
applying directly over the bond coating of the third area only a porous thermal barrier coating to the third area.
2. The method of claim 1 , wherein a generally uniform transition occurs between the coatings applied to the first, second, and third areas.
3. The method of claim 1 , wherein the first area is a gas path surface of the platform.
4. The method of claim 3 , wherein the second area is a fillet region extending around a perimeter of the airfoil proximate an end of the airfoil at an interface of the airfoil and the gas path surface of the platform.
5. The method of claim 4 , wherein the third area is an outer surface of the airfoil.
6. The method of claim 1 , further comprising applying the porous thermal barrier coating to at least a portion of the second area such that the porous thermal barrier coating is oversprayed onto the dense vertically cracked thermal barrier coating.
7. The method of claim 1 , further comprising identifying a fourth area located along a leading edge of the airfoil and applying the dense vertically cracked thermal barrier coating to the fourth area in lieu of the porous thermal barrier coating.
8. A gas turbine component comprising:
a platform portion having a generally planar gas path surface;
an airfoil extending from the platform;
a fillet region extending around a perimeter of the airfoil at an interface of the airfoil and the platform portion;
a first coating applied to the airfoil, the fillet region, and the planar gas path surface of the platform portion;
only a second coating applied directly over the first coating on the planar gas path surface;
the second coating applied directly over the first coating on the fillet region; and,
only a third coating applied directly over the first coating on the airfoil;
wherein the third coating has a thermal conductivity lower than a thermal conductivity of the second coating.
9. The gas turbine component of claim 8 , wherein the airfoil further comprises a plurality of cooling holes.
10. The gas turbine component of claim 8 , wherein the first coating is a CoNiCrAlY bond coating applied to a metal substrate of the gas turbine component.
11. The gas turbine component of claim 8 , wherein the second coating is a thermal barrier coating applied to the bond coating such that dense vertically-oriented micro-cracks are formed in the coating to provide improved durability to high strain areas of the gas turbine component.
12. The gas turbine component of claim 11 , wherein the dense vertically-oriented micro-cracks have a density factor that is variable based on coating particle temperature, velocity, and temperature of the gas turbine component such that strain tolerance and cohesive strength of the second coating is variable.
13. The gas turbine component of claim 11 , wherein the second coating is a 7%-9% Yttria Stabilized Zirconia applied approximately 0.010″-0.025″ thick.
14. The gas turbine component of claim 8 , wherein the third coating is a porous thermal barrier coating.
15. The gas turbine component of claim 14 , wherein the third coating is a 7%-9% Yttria Stabilized Zirconia applied approximately 0.005″-0.019″ thick.
16. The gas turbine component of claim 8 , wherein the third coating has a higher porosity than the second coating.
17. A gas turbine component comprising:
one or more platforms having a generally planar gas path surface;
an airfoil extending from the platform;
a fillet region extending around a perimeter of the airfoil at an interface of the airfoil and the one or more platforms;
a first coating applied to the airfoil, the fillet region, and the planar gas path surface of the one or more platforms;
only a second coating applied directly over the first coating of the planar gas path surface;
a third coating applied directly over the first coating of the fillet region and,
only the third coating applied directly over the first coating of the airfoil;
wherein the third coating has a thermal conductivity lower than a thermal conductivity of the second coating.
18. The gas turbine component of claim 17 , wherein the airfoil further comprises a plurality of cooling holes.
19. The gas turbine component of claim 17 , wherein the first coating is a MCrAlXZ bond coating applied to a metal substrate of the gas turbine component, where M is Ni and/or Co, X is selected from the group comprising Y, Zr, Hf, and Si, and Z is one of Ta, Re, or Pt.
20. The gas turbine component of claim 17 , wherein the second coating is a thermal barrier coating applied to the bond coating such that dense vertically-oriented micro-cracks are formed in the coating to provide improved durability to high strain areas of the gas turbine component.
21. The gas turbine component of claim 17 , wherein the third coating is a porous thermal barrier coating.
22. A gas turbine component comprising:
a first platform and a second platform oriented generally parallel and spaced a radial distance apart, the first and second platforms having generally planar gas path surfaces;
one or more airfoils extending between the first and second platforms;
fillet regions extending around a perimeter of the one or more airfoils at interfaces of the one or more airfoils and the platforms;
a first coating applied to the one or more airfoils, the fillet regions, and the planar gas path surfaces of the platforms;
only a second coating applied directly over the first coating of the planar gas path surfaces of the platforms;
the second coating applied directly over the first coating of the fillet regions; and,
only a third coating applied directly over the first coating of the one or more airfoils;
wherein the third coating has a thermal conductivity lower than a thermal conductivity of the second coating.
23. The gas turbine component of claim 22 , wherein the one or more airfoils have a plurality of shaped cooling holes.
24. The gas turbine component of claim 23 , wherein the first coating is a CoNiCrAlY bond coating applied to a metal substrate of the gas turbine component.
25. The gas turbine component of claim 22 , wherein the second coating is a thermal barrier coating of 7%-9% Yttria Stabilized Zirconia approximately 0.010″-0.025″ thick applied to the bond coating such that dense vertically-oriented micro-cracks are formed in the coating to provide durability to high strain areas of the gas turbine component.
26. The gas turbine component of claim 22 , wherein the third coating is a porous thermal barrier coating of 7%-9% Yttria Stabilized Zirconia applied approximately 0.005″-0.019″ thick.