Controlled variation of pressure drop through effusion cooling in a double walled combustor of a gas turbine engine

Patent metadata

Field	Value
Publication number	US-2016356500-A1
Application number	US-201414913795-A
Country	US
Kind code	A1
Filing date	Sep 16, 2014
Priority date	Sep 16, 2013
Publication date	Dec 8, 2016
Grant date	—

Abstract

A combustor of a gas turbine engine includes a multiple of liner panels mounted to the support shell, at least one of the multiple of liner panels includes a first impingement cavity that operates at a first pressure and a second impingement cavity that operates at a second pressure different than the first pressure. A method of cooling a wall assembly within a combustor of a gas turbine engine includes directing air through a support shell and a liner panel that defines a first impingement cavity and a second impingement cavity. The first impingement cavity operates at a first pressure and the second impingement cavity operates at a second pressure that is different than the first pressure.

First claim

What is claimed is:

1 . A combustor of a gas turbine engine, comprising:
a support shell defining a multiple of impingement flow passages; and a multiple of liner panels mounted to the support shell, at least one of the multiple of liner panels defines a first impingement cavity with the support shell operable at a first pressure and a second impingement cavity operable at a second pressure different than the first pressure.

2 . The combustor as recited in claim 1 , wherein one of the support shell and the multiple of liner panels includes a rail configured to segregate the first impingement cavity and the second impingement cavity.

3 . The combustor as recited in claim 2 , further comprising a reduced height rail adjacent to the support shell.

4 . The combustor as recited in claim 3 , wherein the reduced height rail defines a trailing edge of the at least one of the multiple of liner panels.

5 . The combustor as recited in claim 1 , wherein the support shell defines a first multiple of impingement flow passages in communication with the first impingement cavity and a second multiple of impingement flow passages in communication with the second impingement cavity.

6 . The combustor as recited in claim 5 , wherein at least one of the first multiple of impingement flow passages defines a diameter different than at least one of the second multiple of impingement flow passages.

7 . The combustor as recited in claim 5 , wherein each of the first multiple of impingement flow passages defines a diameter different than each of the second multiple of impingement flow passages.

8 . The combustor as recited in claim 5 , wherein the first multiple of impingement flow passages are more numerous than the second multiple of impingement flow passages.

9 . The combustor as recited in claim 5 , further comprising
a first multiple of effusion flow passages though the at least one of the multiple of liner panels in communication with the first impingement cavity; and a second multiple of effusion flow passages though the at least one of the multiple of liner panels in communication with the second impingement cavity.

10 . The combustor as recited in claim 9 , wherein
the first multiple of impingement flow passages and the first multiple of effusion flow passages define a first ratio and the second multiple of impingement flow passages; and the second multiple of effusion flow passages define a second ratio, the first ratio different than the second ratio.

11 . The combustor as recited in claim 1 , further comprising an annular grommet between the support shell and the at least one of the multiple of liner panels, the annular grommet and the support shell defines an annular impingement cavity.

12 . The combustor as recited in claim 11 , further comprising at least one effusion flow passage through the annular grommet and the at least one of the multiple of liner panels.

13 . A combustor of a gas turbine engine, comprising:
a support shell with a multiple of impingement flow passages; and a multiple of liner panels mounted to the support shell; a first liner panel of the multiple of liner panels includes a first impingement cavity with respect to the support shell that operates at a first pressure; and a second liner panel with respect to the support shell of the multiple of liner panels includes a second impingement cavity that operates at a second pressure different than the first pressure.

14 . The combustor as recited in claim 13 , further comprising a plurality of studs which extend from a cold side of each of the multiple of liner panels, the studs extending through the support shell.

15 . The combustor as recited in claim 14 , further comprising:
a first rail around a periphery of the first liner panel; and a second rail around a periphery of the second liner panel; the first rail and the second rail in contact with the support shell to segregate the first liner panel and the second liner panel.

16 . A method of cooling a wall assembly within a combustor of a gas turbine engine, comprising:
directing air through a support shell into a first impingement cavity and a second impingement cavity such that the first impingement cavity operates at a first pressure and the second impingement cavity operates at a second pressure different than the pressure.

17 . The method as recited in claim 16 , further comprising directing air from the second impingement cavity through a trailing edge of a liner panel mounted to the support shell.

18 . The method as recited in claim 16 , further comprising directing air from an annular impingement cavity at least partially defined by an annular grommet between the support shell and the liner panel.

19 . The method as recited in claim 16 , further comprising forming a first impingement pressure drop across the support shell in the first impingement cavity and a second impingement pressure drop across the support shell in the second impingement cavity, the first impingement pressure drop different than the second impingement pressure drop.

20 . The method as recited in claim 16 , further comprising forming a first effusion pressure drop across the liner panel with respect to the first impingement cavity and a second effusion pressure drop across the liner panel with respect to the second impingement cavity, the first effusion pressure drop different than the second effusion pressure drop.

Assignees

• United Technologies Corp

Inventors

• Bouldin Lee E
• Eastwood Jonathan J
• Moura Dennis M
• Pacheco-Tougas Monica

Classifications

• F23R3/06 Primary

  Arrangement of apertures along the flame tube · CPC title

• F23R3/002

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

• F23R3/60

  Support structures; Attaching or mounting means · CPC title

• F23R2900/03042

  Film cooled combustion chamber walls or domes · CPC title

• F23R2900/03044

  Impingement cooled combustion chamber walls or subassemblies · CPC title