BACKGROUND

The present disclosure is directed to a geometrically segmented abradable ceramic (GSAC) coating with a multiple of toughened layers for spallation resistance.

Components that are exposed to high temperatures, such as gas turbine engine turbine blades, turbine vanes, blade outer air seals, and compressor components, typically include protective coatings. The protective coatings may include one or more layers that function to protect the component from erosion, oxidation, corrosion, or the like to enhance component durability and maintain efficient operation of the engine.

A geometrically segmented abradable ceramic (GSAC) coating has been demonstrated in high gradient thermal cycle testing to survive surface temperatures in the range of 3200 degrees F. (1760 degrees C.). Recent testing has shown significant gains in durability and temperature capability over conventional thermal barrier coatings, however, some GSAC coatings may still be subject to spallation at inter-segment edges due to crack formation from the stresses associated with sintering shrinkage.

SUMMARY

A turbine article according to one disclosed non-limiting embodiment includes a substrate with a geometric surface having a multiple of divots recessed into the substrate; and a ceramic topcoat disposed over the geometric surface, the ceramic topcoat comprising a first layer having a first hardness, a second layer on the first layer, the second layer having a second hardness, the first hardness different than the second hardness, a third layer on the second layer, the third layer having a hardness about equivalent to the first hardness, wherein the third layer is flush with a top surface of the geometric surface.

A further aspect includes that the first layer is located within a bottom of each of the multiple of divots and on the raised area surrounding each divot.

A further aspect includes a fourth layer on the third layer, the fourth layer having a hardness about equivalent to the second hardness.

A further aspect includes a fifth layer on the fourth layer, the fifth layer having a hardness about equivalent to the first hardness.

A further aspect includes that the coating is machined to remove the portion of the fourth and fifth layers that are raised above the fourth and fifth layers located in the divot locations to produce a smooth surface.

A further aspect includes that the fifth layer is about 5-20 mils thick prior to being machined.

A further aspect includes that the fifth layer is about 3-10 mils thick after being machined.

A further aspect includes that the first layer is 1-3 mils thick and the second layer is 6-50 mils thick.

A further aspect includes that a divot depth of each of the multiple of divots of the geometric surface is 10-50 mils, the first layer thickness spans from 0-15% of the divot depth (with 0% being the bottom of the divot) and the third layer thickness spans 90-115% the divot depth.

A further aspect includes that each of the multiple of divots of the geometric surface comprises a substantially uniform divot depth extending into the substrate and a divot width, wherein the ratio of the divot width to the divot depth is 1 to 10, with the divot depth being at least 0.01 inches (0.254 millimeters).

A further aspect includes a bond coat between the substrate and the first layer.

A further aspect includes that the substrate is at least one of metallic, monolithic ceramic, metal matrix composite and a ceramic matrix composite.

A further aspect includes that the article is a blade outer air seal.

A process for manufacturing a geometrically segmented abradable coating on a turbine engine component, the process according to one disclosed non-limiting embodiment includes depositing a first layer of a ceramic topcoat on a geometric surface and within a bottom of each of a multiple of divots, the first layer having a first hardness; depositing a second layer of the ceramic topcoat on the first layer, the second layer having a second hardness, the first hardness higher than the second hardness; and depositing a third layer on the second layer, the third layer having a hardness about equivalent to the first hardness, at the divot locations, the third layer corresponds with a top surface of the geometric surface and the first layer on the top surface of the geometric surface.

A further aspect includes depositing a fourth layer on the third layer, the fourth layer having a hardness about equivalent to the second hardness; depositing a fifth layer on the fourth layer, the fifth layer having a hardness about equivalent to the first hardness; and machining the ceramic topcoat.

A further aspect includes that machining the ceramic topcoat comprises removing raised material over a substrate feature.

A further aspect includes redepositing at least a portion of the fifth layer.

A further aspect includes that the third layer corresponds with the top surface of the geometric surface and the first layer on the op surface of the geometric surface.

A further aspect includes that the third layer thickness spans the top surface of the geometric surface.

A further aspect includes a divot depth of the geometric surface is 10-50 mils, the first layer thickness spans from 0-15% of the divot depth (with 0% being the bottom of the divot) and the third layer thickness spans 90-115% the divot depth.

The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be appreciated that the following description and drawings are intended to be exemplary in nature and non-limiting.

