Divot pattern for thermal barrier coating

1. A component for a gas turbine engine, comprising:

a surface adjacent a flow of hot gases;

a plurality of cavities in a portion of the surface, the plurality of cavities having a first group of cavities with a first cross-section and a second group of cavities with a second cross-section different from the first cross-section, wherein the first and second groups of cavities are arranged such that there is no straight line across the portion of the surface that does not intersect one of the plurality of cavities, wherein the plurality of cavities span at least 30% of the surface, wherein the first cross-section is a slot shape arranged in a first direction and the second cross-section is a slot shape arranged in a second direction that is perpendicular to the first direction, and the first and second cavities alternate such that the first and second cavities form a herringbone pattern; and

a thermal barrier coating over the surface and filling each of the plurality of cavities.

2. The component of claim 1 , wherein the surface is on a blade outer air seal.

3. The component of claim 2 , wherein the blade outer air seal is a ceramic material.

4. The component of claim 1 , wherein the thermal barrier coating is an abradable coating that fills each of the plurality of cavities.

5. The component of claim 4 , wherein the abradable coating is a geometrically segmented abradable coating formed from a ceramic material.

6. The component of claim 1 , wherein the first cavities and the second cavities are the same size.

7. The component of claim 1 , wherein the first cavities have a length and a width and the length is more than two times the width.

8. A component for a gas turbine engine, comprising:

a surface adjacent a flow of hot gases;

a plurality of cavities in a portion of the surface, the plurality of cavities having a first group of cavities with a first cross-section and a second group of cavities with a second cross-section different from the first cross-section, wherein the first and second groups of cavities are arranged such that there is no straight line across the portion of the surface that does not intersect one of the plurality of cavities, wherein the first cross-section is a circle and the second cross-section is a slot shape.

9. The component of claim 8 , wherein a center of each cavity in the first group is arranged at a corner of a square.

10. The component of claim 9 , wherein a ratio of an area of the plurality of cavities within the square to an area of the square is greater than 0.53.

11. The component of claim 10 , wherein the ratio is greater than 0.56.

12. The component of claim 10 , wherein the ratio is between 0.56 and 0.58.

13. A turbine section for a gas turbine engine, comprising:

a turbine blade extending radially outwardly to a radially outer tip and for rotation about an axis of rotation; and

a component having a first side and a second side, wherein the first side has a cavity layer having a plurality of cavities on a portion of the first side, the plurality of cavities having a first group of cavities with a first cross-section and a second group of cavities with a second cross-section different from the first cross-section and a thermal barrier coating over the cavity layer, wherein a center of each cavity in the first group is arranged at a corner of a square and wherein the first cross-section is a circle and the second cross section is a slot shape, and each cavity in the second group is arranged along an edge of the square.

14. The turbine section of claim 13 , wherein a ratio of an area of the plurality of cavities in the square to an area of the square is greater than about 0.56.

15. The turbine section of claim 13 , wherein the thermal barrier coating is an abradable coating.

16. The turbine section of claim 13 , wherein the component is one of a ceramic and a metallic material.