Combustion liner with bias effusion cooling

Patent metadata

Field	Value
Publication number	US-9429323-B2
Application number	US-201414278770-A
Country	US
Kind code	B2
Filing date	May 15, 2014
Priority date	May 15, 2014
Publication date	Aug 30, 2016
Grant date	Aug 30, 2016

Abstract

A system and method for improving the cooling to a portion of a combustion liner of a gas turbine combustor is disclosed. The combustion liner is cooled by supplying air through a plurality of cooling holes arranged in axially spaced rows in an upper and lower portion of the liner. The cooling holes are spaced accordingly so as to direct additional cooling flow to an area of the combustion liner not receiving sufficient flow due to maldistributions of air from the compressor discharge.

First claim

The invention claimed is:

1. A combustion liner for a gas turbine combustor comprising:
a generally annular body having an inner wall, an outer wall spaced a distance from the inner wall, an inlet end, and an opposing outlet end, the generally annular body also having a plane extending through
a centerline to bisect the generally annular body and define an upper portion and a lower portion of the generally annular body;
a seal positioned along the outer wall proximate the outlet end;
the upper portion having a plurality of cooling holes, the upper portion having a lower boundary defined by the plane extending through the centerline, a left boundary defined by a column of one or more cooling holes most near to the inlet end, and a right boundary defined by a column of one or more cooling holes most near to the outlet end;
the lower portion having a plurality of cooling holes, the lower portion having an upper boundary defined by the plane extending through the centerline, a left boundary defined by a column of one or more cooling holes most near to the inlet end, and a right boundary defined by a column of one or more cooling holes most near to the outlet end, wherein the lower portion comprises a section of the generally annular body that is positioned closer to an engine axis upon installation of the combustion liner in the gas turbine engine; and
wherein the entire lower portion of the generally annular body has a higher concentration of cooling holes than the entire upper portion of the generally annular body.

2. The combustion liner of claim 1 , wherein the plurality of cooling holes are arranged in a plurality of axially spaced rows, the plurality of axially spaced rows comprising:
a first set of axially spaced rows located in the upper portion of the generally annular body;
a second set of axially spaced rows located in the lower portion of the generally annular body; and
a third set of axially spaced rows located in the lower portion of the generally annular body;
wherein the first set of axially spaced rows and the second set of axially spaced rows each comprise at least three rows and the second set of axially spaced rows are in axial alignment with the first set of axially spaced rows.

3. The combustion liner of claim 2 , wherein the third set of axially spaced rows comprise at least two rows with each row positioned between one of the second set of axially spaced rows.

4. The combustion liner of claim 2 , wherein the third set of axially spaced rows extend approximately 120 degrees across the lower portion of the generally annular body.

5. The combustion liner of claim 1 further comprising a thermal barrier coating applied to the inner wall from proximate the plurality of openings to the outlet end.

6. The combustion liner of claim 1 , wherein the plurality of cooling holes are oriented at a surface angle relative to the plane extending through the centerline of the generally annular body.

7. The combustion liner of claim 6 , wherein the plurality of cooling holes are also oriented at a tangential angle relative to the plane extending through the centerline of the generally annular body.

8. A cooling pattern for a combustion liner having a generally annular body, the cooling pattern comprising:
a first set of axially spaced rows, the first set of axially spaced rows spaced a first distance apart;
a second set of axially spaced rows, the second set of axially spaced rows spaced a second distance apart;
a third set of axially spaced rows, the third set of axially spaced rows spaced a third distance apart;
an upper half having a lower boundary defined by a plane extending through a centerline, a left boundary defined by a left-most row in the first set of axially spaced rows, and a right boundary defined by a right-most row in the first set of axially spaced rows;
a lower half having an upper boundary defined by the plane extending through the centerline, a left boundary defined by a left-most row in the second set of axially spaced rows, and a right boundary defined by a right-most row in the second set of axially spaced rows; and
wherein the second set of axially spaced rows and the third set of axially spaced rows are arranged such that of the entire lower half of the combustion liner has a higher concentration of cooling holes than the entire upper half of the combustion liner.

9. The cooling pattern of claim 8 , wherein the third set of axially spaced rows extend approximately 120 degrees across the lower half of the combustion liner.

10. The cooling pattern of claim 8 further comprising a thermal barrier coating applied to an inner wall of the combustion liner.

11. The combustion liner of claim 8 , wherein the first, second and third plurality of rows of cooling holes are oriented at a surface angle relative to the annular body.

12. The combustion liner of claim 11 , wherein the first, second, and third plurality of cooling holes are also oriented at a tangential angle relative to the plane extending through the centerline.

13. The cooling pattern of claim 8 , wherein the cooling holes are equally spaced in each of the first set of axially spaced rows.

14. The cooling pattern of claim 8 , wherein there are an equal number of cooling holes in each of the second set and third set of axially spaced rows of cooling holes.

15. A method of providing increased cooling to a portion of a combustion liner comprising:
providing a generally annular body having a plane extending through a centerline, wherein the plane extending through the centerline bisects the generally annular body and defines an upper portion and a lower portion, where the lower portion is a section of the combustion liner positioned closest to the centerline of a gas turbine engine when the combustion liner is installed in the gas turbine engine;
providing a plurality of cooling holes in the upper portion of the combustion liner, the upper portion having a lower boundary defined by the plane extending through the centerline, a left boundary defined by a column of one or more cooling holes most near to the inlet end, and a right boundary defined by a column of one or more cooling holes most near to the outlet end:
providing a plurality of cooling holes in the lower portion of the combustion liner, the lower portion having an upper boundary defined by the plane extending through the centerline: a left boundary defined by a left-most row in the second set of axially spaced rows, and a right boundary defined by a right-most row in the second set of axially spaced rows; and
wherein the plurality of cooling holes are arranged such that the entire lower portion of the generally annular body has a higher concentration of cooling holes than the entire upper portion of the generally annular body.

16. The method of claim 15 , wherein the plurality of cooling holes are arranged in a first plurality of rows of cooling holes, a second plurality of rows of cooling holes, and a third plurality of rows of cooling holes, wherein the second and third plurality of rows of cooling holes extend approximately 120 degrees across the lower portion of the generally annular body.

17. The method of claim 16 , wherein the cooling holes of the first plurality of rows, second plurality of rows, and third plurality of rows are each oriented at a surface angle relative to an axis of the generally annular body as well as a tangential angle.

18. The method of claim 16 , wherein the cooling holes in the first plurality of rows are equally spaced at a first distances apart, the cooling holes in the second

Assignees

• Richardson Brian
• Oumejjoud Khalid
• Cutright John
• Mcmullen Nikki
• Stuttaford Peter John
• Parker Richard
• Soni Sumit
• General Electric Technology Gmbh

Inventors

• Richardson Brian
• Oumejjoud Khalid
• Cutright John
• Mcmullen Nikki
• Stuttaford Peter John
• Parker Richard
• Soni Sumit

Classifications

• F23R3/04 Primary

  Air inlet arrangements · CPC title

• F23R3/002 Primary

  Wall structures (F23R3/02 and F23R3/007 take precedence) · CPC title

• Y02T50/675

  Cross-Sectional Technologies · mapped topic

• F23R2900/03041

  Effusion cooled combustion chamber walls or domes · CPC title

• F23R3/06

  Arrangement of apertures along the flame tube · CPC title