Sealed combustor liner panel for a gas turbine engine

Patent metadata

Field	Value
Publication number	US-10816201-B2
Application number	US-201414913165-A
Country	US
Kind code	B2
Filing date	Jul 11, 2014
Priority date	Sep 13, 2013
Publication date	Oct 27, 2020
Grant date	Oct 27, 2020

Abstract

A liner panel for a combustor of a gas turbine engine includes a rail which at least partially defines an impingement cavity. The rail includes a notch which faces toward the impingement cavity. A method of cooling a wall assembly within a combustor for of a gas turbine engine includes directing air through a support shell and a liner panel to form a pressure drop across the support shell that is less than about 80% of a pressure drop across the combustor and to also form a pressure drop across the liner panel greater than about 20% of the pressure drop across the combustor.

First claim

What is claimed is:

1. A combustor of a gas turbine engine, the combustor comprising:
a shell comprising a multiple of cooling impingement passages;
a liner panel mounted to the shell, the liner panel comprising:
a cold side facing the shell and a hot side opposite the cold side and facing away from the shell;
a rail extending around a periphery of the cold side of the liner panel and in contact with the shell along a rail surface of the rail so as to define an impingement cavity between the liner panel and the shell, the rail comprising a notch which faces toward the impingement cavity; and
a multiple of effusion passages extending through the liner panel between the cold side and the hot side; and

a seal disposed within the notch and surrounding the impingement cavity, wherein the seal is a C-seal;
wherein an opening of the C-seal is directed toward the impingement cavity; and
wherein the multiple of effusion passages are more numerous than the multiple of cooling impingement passages.

2. The combustor as recited in claim 1 , further comprising a multiple of studs extending from the cold side of the liner panel.

3. The combustor as recited in claim 1 , wherein the combustor is configured so that air directed through the shell and the liner panel forms a pressure drop across the shell that is less than about 80% and greater than about 50% of a pressure drop across the combustor and a pressure drop across the liner panel that is greater than about 20% and less than about 50% of the pressure drop across the combustor.

4. A combustor of a gas turbine engine, the combustor comprising:
a shell comprising a multiple of impingement flow passages;
a liner panel mounted to the shell, the liner panel comprising:
a cold side facing the shell and a hot side opposite the cold side and facing away from the shell;
a rail extending around a periphery of the cold side of the liner panel and in contact with the shell along a rail surface of the rail which interfaces with the shell to define an impingement cavity between the liner panel and the shell, and the rail including a notch which faces toward the impingement cavity; and
a multiple of effusion passages extending through the liner panel between the cold side and the hot side; and

a seal disposed within the notch and surrounding the impingement cavity, wherein the seal is a C-seal;
wherein an opening of the C-seal is directed toward the impingement cavity;
wherein a dilution passages penetrates through both the shell and the liner panel along a common axis; and
wherein the multiple of effusion passages are more numerous than the multiple of cooling impingement passages.

5. The combustor as recited in claim 4 , further comprising a plurality of studs extending from the cold side of the liner panel, the studs extending through the shell.

6. A method of cooling a wall assembly within a combustor of a gas turbine engine, the method comprising:
directing aft through a support shell and a liner panel to form a pressure drop across the support shell that is less than about 80% of a pressure drop across the combustor and to further form a pressure drop across the liner panel greater than about 20% of the pressure drop across the combustor;
the combustor comprising a shell comprising a multiple of cooling impingement passages; the liner panel mounted to the shell, the liner panel comprising:
a cold side facing the shell and a hot side opposite the cold side and facing away from the shell;
a rail extending around a periphery of the cold side of the liner panel and in contact with the shell along a rail surface of the rail so as to define an impingement cavity between the liner panel and the shell, the rail comprising a notch which faces toward the impingement cavity; and
a multiple of effusion passages extending through the liner panel between the cold side and the hot side; and

a seal disposed within the notch and surrounding the impingement cavity, wherein the seal is a C-seal;
wherein an opening of the C-seal is directed toward the impingement cavity; and wherein the multiple of effusion passages are more numerous than the multiple of cooling impingement passages.

7. The method as recited in claim 6 , further comprising sealing an interface between the support shell and at least the liner panel of a multiple of liner panels.

8. The method as recited in claim 7 , further comprising compressing the C-seal between the support shell and the liner panel of the multiple of liner panels.

9. The method as recited in claim 7 , further comprising pressurizing the C-seal.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Drake Christopher

Classifications

• F02C7/24 Primary

  Heat or noise insulation (air intakes having provisions for noise suppression F02C7/045; turbine exhaust heads, chambers, or the like F01D25/30; silencing nozzles of jet-propulsion plants F02K1/00) · CPC title

• F23R2900/03041

  Effusion cooled combustion chamber walls or domes · CPC title

• F23R3/60

  Support structures; Attaching or mounting means · CPC title

• F23R3/002

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

• F23R2900/03042

  Film cooled combustion chamber walls or domes · CPC title