Coated gas turbine components

Patent metadata

Field	Value
Publication number	US-10113435-B2
Application number	US-201113184136-A
Country	US
Kind code	B2
Filing date	Jul 15, 2011
Priority date	Jul 15, 2011
Publication date	Oct 30, 2018
Grant date	Oct 30, 2018

Abstract

A gas turbine component subject to extreme temperatures and pressures includes a wall defined by opposite first and second surfaces. An airflow aperture through the wall is defined by an aperture wall surface which extends from a first opening in the first surface to a second opening in the second surface. The aperture wall surface is flared at a juncture with the first surface, such that the first opening has a greater cross-sectional flow area than the second opening. A high-pressure, high-temperature coating is adhered to the first surface, and adhered to at least a portion of the aperture wall surface.

First claim

The invention claimed is:

1. A method of forming a gas turbine engine component subject to extreme temperatures and pressures, the method comprising:
fabricating a wall having a first surface and a second surface which define opposite sides of the wall;
creating an airflow aperture that extends through the wall in a direction substantially perpendicular to the first surface, the airflow aperture defined by an aperture wall surface which extends from a first opening in the first surface to a second opening in the second surface, and which is flared at a juncture with the first surface such that the first opening has a greater cross-sectional flow area than the second opening; and
depositing a high-pressure, high-temperature resistant coating on the first surface, adhered to a portion of the aperture wall surface adjacent the first opening, such that a minimum flow width w of the airflow aperture is reduced and defined by the high-pressure, high-temperature resistant coating, where

w
=

W
major

-

W
minor

2

-

2
⁢

⁢
t
⁢

⁢
sin
⁢

⁢
Θ

,

W major is a maximum uncoated width of the airflow aperture, W minor is a minimum uncoated width of the airflow aperture, t is a thickness of the high-pressure, high-temperature resistant coating, and Θ is a surface angle between the aperture wall surface and a line normal to the first surface.

2. The method of claim 1 , wherein the gas turbine engine component is a gas turbine combustor liner or afterburner liner.

3. The method of claim 1 , wherein the aperture wall surface is substantially perpendicular to the first and second surfaces where adjacent the second surface.

4. The method of claim 1 , wherein the high pressure, high temperature resistant coating is adhered in a uniform thickness.

5. The method of claim 4 , wherein the portion of the aperture wall surface adjacent the first surface has cross-sectional profile with a radius of curvature greater than or equal to the uniform thickness of the high pressure, high temperature resistant coating.

6. The method of claim 1 , wherein the portion of the aperture wall surface adjacent the first surface has a substantially frusto-conical cross-sectional profile.

7. The method of claim 6 , wherein the aperture wall surface has a frusto-conical cross-sectional profile from the first surface to the second surface.

8. The method of claim 1 , wherein the high pressure, high temperature resistant coating is a ceramic-based protective coating.

9. The method of claim 1 , wherein the first and second openings are substantially circular.

10. The method of claim 1 , wherein at least one of the first or second openings is elliptical.

11. A gas turbine engine component subject to extreme temperatures and pressures, the gas turbine engine component comprising:
a wall having a first surface and a second surface which define opposite sides of the wall, and an airflow aperture that extends entirely through the wall, the airflow aperture defined by an aperture wall surface which meets the first surface in a hole perimeter, such that the aperture wall surface is angled at a uniform obtuse angle relative to the first surface at this hole perimeter; and
a high-pressure, high-temperature resistant coating adhered to the first surface, and adhered to a portion of the aperture wall surface adjacent the first opening, such that a minimum flow width w of the airflow aperture is reduced and defined by the high-pressure, high-temperature resistant coating, such that

w
=

W
major

-

W
minor

2

-

2
⁢

⁢
t
⁢

⁢
sin
⁢

⁢
Θ

,

where W major is a maximum uncoated width of the airflow aperture, W minor is a minimum uncoated width of the airflow aperture, t is a thickness of the high-pressure, high-temperature resistant coating, and Θ is a surface angle between the aperture wall surface and a line normal to the first surface.

12. The gas turbine engine component of claim 11 , wherein the wall is a gas turbine engine combustor liner or afterburner liner.

13. The gas turbine engine component of claim 11 , wherein the wall is an airfoil blade or vane surface.

14. The gas turbine engine component of claim 11 , wherein the high-pressure, high-temperature resistant coating comprises a ceramic-based plasma spray coating.

15. The gas turbine engine component of claim 14 , wherein the ceramic-based coating is a thermal barrier coating.

16. The gas turbine engine component of claim 11 , wherein the aperture wall surface has a substantially frusto-conical cross-section at the hole perimeter.

17. The gas turbine engine component of claim 11 , wherein the aperture wall surface is curved continuously with the first surface at the hole perimeter.

18. The gas turbine engine component of claim 11 , wherein the hole perimeter is elliptical.

19. The gas turbine engine component of claim 11 , wherein the aperture wall surface is substantially perpendicular to the first and second surfaces where adjacent the second surface.

Assignees

• Pater Christopher M
• United Technologies Corp

Inventors

• Pater Christopher M

Classifications

• F23R3/06

  Arrangement of apertures along the flame tube · CPC title

• F23R2900/00018

  Manufacturing combustion chamber liners or subparts · CPC title

• F05D2260/202

  by film cooling · CPC title

• F01D25/08

  Cooling (of machines or engines in general F01P); Heating; Heat-insulation (of blade-carrying members, of blades F01D5/00) · CPC title

• F05D2300/20

  Oxide or non-oxide ceramics · CPC title