Film cooling hole arrangement for acoustic resonators in gas turbine engines

Patent metadata

Field	Value
Publication number	US-10359194-B2
Application number	US-201415502016-A
Country	US
Kind code	B2
Filing date	Aug 26, 2014
Priority date	Aug 26, 2014
Publication date	Jul 23, 2019
Grant date	Jul 23, 2019

Abstract

The present disclosure provides a gas turbine combustor liner ( 34 ) comprising an outer surface ( 38 ) and an inner surface ( 36 ), a plurality of film cooling holes ( 44 ) through a thickness of the gas turbine combustor liner ( 34 ), and a plurality of resonator boxes ( 32 ) affixed to the outer surface ( 38 ) of the gas turbine combustor liner ( 34 ). The film cooling holes ( 44 ) extend circumferentially around the gas turbine combustor liner ( 34 ) and comprise a first set of holes ( 56 ) having a first axial row spacing X and a second set of holes ( 58 ) having a second axial row spacing X′. The second set of holes ( 58 ) is formed in the gas turbine combustor liner ( 34 ) in a downstream direction relative to the first set of holes ( 56 ). The second axial row spacing X′ is greater than the first axial row spacing X.

First claim

What is claimed is:

1. A gas turbine combustor liner comprising:
an outer surface and an inner surface, the outer surface being exposed to a cooling airflow and the inner surface being exposed to hot combustion gases;
a plurality of film cooling holes through a thickness of the gas turbine combustor liner, the film cooling holes extending circumferentially around the gas turbine combustor liner, wherein the film cooling holes comprise:
a first set of holes having a first axial row spacing X, the first set of holes being defined by a first plurality of rows of holes extending in a circumferential direction; and
a second set of holes having a second axial row spacing X′, the second set of holes being defined by a second plurality of rows of holes extending in a circumferential direction, wherein the second set of holes is formed in the gas turbine combustor liner in a downstream direction relative to the first set of holes, the second axial row spacing X′ being greater than the first axial row spacing X; and
a plurality of resonator boxes affixed to the outer surface of the gas turbine combustor liner; and
wherein each of the resonator boxes extend axially over at least a portion of each of the first set of holes and the second set of holes.

2. The gas turbine combustor liner of claim 1 wherein an axis of the film cooling holes is substantially perpendicular to the outer surface and the inner surface of the gas turbine combustor liner.

3. The gas turbine combustor liner of claim 1 wherein a dimensionless first axial row spacing, X 0= X/d, of the first set of holes is greater than or equal to about 3 and less than 10, where d is the diameter of the holes, and wherein the second axial row spacing X 0 ′ = X′/d, of the second set of holes is between about 3 and 10.

4. The gas turbine combustor liner of claim 1 wherein the resonator boxes further comprise a plurality of impingement holes configured to introduce at least a portion of the cooling airflow into the resonator boxes.

5. The gas turbine combustor liner of claim 1 wherein the resonator boxes further comprise an upstream wall and a downstream wall, an upstream wall height being less than a downstream wall height.

6. The gas turbine combustor liner of claim 1 wherein the resonator boxes are affixed to a location of the gas turbine combustor liner wherein a flow temperature of the hot combustion gases is increasing in a downstream direction.

7. The gas turbine combustor liner of claim 1 wherein the first set of holes further comprises a first circumferential hole spacing and the second set of holes further comprises a second circumferential hole spacing, the first circumferential hole spacing Y being different than the second circumferential hole spacing.

8. A turbine engine assembly comprising:
a turbine engine having a compressor section, a combustor comprising a combustor liner, and a turbine section, wherein the combustor liner comprises:
a plurality of film cooling holes extending circumferentially around the combustor liner and extending through a thickness of the combustor liner, wherein the film cooling holes comprise a first set of holes having a first axial row spacing X and a second set of holes having a second axial row spacing X′, the first set of holes and the second set of holes each being defined by a plurality of rows of holes extending in a circumferential direction, wherein the second set of holes is located in a downstream direction relative to the first set of holes, the second axial row spacing X′ being greater than the first axial row spacing X; and

a plurality of resonator boxes affixed to and located circumferentially about an outer surface of the combustor liner, wherein each of the resonator boxes extend axially over at least a portion of each of the first set of holes and the second set of holes, the resonator boxes further comprising a plurality of impingement holes configured to introduce a cooling airflow into the resonator boxes.

9. The turbine engine assembly of claim 8 wherein the impingement holes are offset from the film cooling holes.

10. The turbine engine assembly of claim 8 wherein an interior of each resonator box is in fluid communication with an interior of the combustor.

11. The turbine engine assembly of claim 8 wherein the resonator boxes further comprise an upstream wall and a downstream wall, an upstream wall height being less than a downstream wall height.

12. A turbine engine assembly comprising:
a turbine engine having a compressor section, a combustor comprising a combustor liner, and a turbine section, wherein the combustor liner comprises:
a plurality of film cooling holes extending circumferentially around the combustor liner and extending through a thickness of the combustor liner, wherein the film cooling holes comprise a first set of holes having a first axial row spacing X and a second set of holes having a second axial row spacing X′, the first set of holes and the second set of holes each being defined by a plurality of rows of holes extending in a circumferential direction, wherein the second set of holes is located in a downstream direction relative to the first set of holes, the second axial row spacing X′ being greater than the first axial row spacing X; and

a plurality of resonator boxes affixed to and located circumferentially about an outer surface of the combustor liner, wherein each of the resonator boxes extend axially over at least a portion of each of the first set of holes and the second set of holes, the resonator boxes further comprising a plurality of impingement holes configured to introduce a cooling airflow into the resonator boxes; and
wherein the film cooling holes further comprise a first set of holes having a first axial row spacing X and a second set of holes having a second axial row spacing X′, each of the first set of holes and the second set of holes being defined by a plurality of rows of holes extending in a circumferential direction, wherein the second set of holes is formed in the gas turbine combustor liner in a downstream direction relative to the first set of holes, the second axial row spacing X′ being greater than the first axial row spacing X, wherein each of the resonator boxes extend axially over at least a portion of each of the first set of holes and the second set of holes.

13. The method of claim 12 further comprising providing a film cooling boundary layer of maximum thickness at the upstream end of the resonator boxes and maintaining the film cooling boundary layer at a substantially constant thickness in a direction downstream from the upstream end of the resonator boxes.

14. The method of claim 12 further comprising providing greater impingement cooling of the combustor liner at the upstream end of the resonator boxes as compared to the downstream end.

15. The method of claim 14 wherein the resonator boxes further comprise an upstream wall and a downstream wall and wherein providing greater impingement cooling of the combustor liner comprises forming the resonator boxes such that an upstream wall height is less than a downstream wall height.

16. The method of claim 12 further comprising locating the resonator boxes on the combustor liner such that a flow temperature of hot combustion gases in the interior of the combustor liner is increasing in an upstream to downstream direction along an axial length of the resonator boxes.

Assignees

• Siemens Energy Inc

Inventors

• Schilp Reinhard
• Fox Timothy A

Classifications

• F23R2900/03044

  Impingement cooled combustion chamber walls or subassemblies · CPC title

• F23R2900/00014

  Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators · CPC title

• F23R3/06

  Arrangement of apertures along the flame tube · CPC title

• F23R3/002 Primary

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