Thermal barrier coating of a combustor dilution hole
Aspects of the disclosure are directed to a liner associated with a combustor of an aircraft engine, comprising: a thermal barrier coating, and a base metal, wherein the thermal barrier coating comprises a contoured surface on a flowpath side proximate to an exit of a hole formed by the thermal barrier coating and the base metal.
1. A liner associated with a combustor of an aircraft engine, comprising:
a panel composed of a thermal barrier coating and a base metal; and
a shell coupled to the panel;
wherein the thermal barrier coating comprises a contoured surface on a combustion flowpath side proximate to an exit of a hole defined by the thermal barrier coating and the base metal;
wherein the contoured surface is defined by the base metal having a first material thickness adjacent to the hole and a second material thickness further from the hole;
wherein the contoured surface is defined by a ramp between the first thickness and the second thickness;
wherein the first thickness is greater than the second thickness; and
wherein the ramp projects away from the shell in transitioning from the second thickness to the first thickness.
2. The liner of claim 1 , wherein the first thickness defines a maximum thickness of the liner in a contour region and the second thickness defines a base thickness of the liner.
3. The liner of claim 1 , wherein the contoured surface is defined by the thermal barrier coating have a first thickness adjacent to the hole and a second thickness further from the hole.
4. The liner of claim 1 , wherein the thermal barrier coating comprises a top coat and a bond coat.
5. The liner of claim 4 , wherein the contoured surface is defined by the bond coat having a first thickness adjacent to the hole and a second thickness further from the hole.
6. The liner of claim 4 , wherein the contoured surface is defined by the top coat having a first thickness adjacent to the hole and a second thickness further from the hole.
7. The liner of claim 1 , wherein the hole is one of an air feed hole, a dilution hole, or a trim hole.
8. The liner of claim 1 , further comprising:
a grommet that is defined by a surface of the base metal and a shell.
9. The liner of claim 8 , wherein the grommet is integral with a panel.
10. The liner of claim 1 , wherein the contoured surface is formed during a casting or fabrication of a panel.
11. The liner of claim 1 , wherein the contoured surface includes a transition of 0.25 millimeters.
12. The liner of claim 1 , wherein the contoured surface is defined by a material that is present in substantially equal amounts relative to a first side and a second side of the hole.
13. The liner of claim 1 , wherein the contoured surface is defined by material that is included in a three-dimensional pattern about the hole to provide a three-dimensional contoured surface.
14. A method for forming a liner of a combustor of an aircraft engine, the method comprising:
providing a base metal;
coupling a thermal barrier coating to the base metal, wherein the base metal and the thermal barrier coating are included in a panel; and
coupling the panel to a shell;
wherein the thermal barrier coating comprises a contoured surface on a combustion flowpath side proximate to an exit of a hole formed by the thermal barrier coating and the base metal;
wherein the contoured surface is defined by the base metal having a first material thickness adjacent to the hole and a second material thickness further from the hole;
wherein the contoured surface is defined by a ramp between the first thickness and the second thickness;
wherein the first thickness is greater than the second thickness; and
wherein the ramp projects away from the shell in transitioning from the second thickness to the first thickness.
15. The method of claim 14 , wherein the hole is one of a dilution hole or an air feed hole.