Active cooling of grommet bosses for a combustor panel of a gas turbine engine

Patent metadata

Field	Value
Publication number	US-10088159-B2
Application number	US-201414773990-A
Country	US
Kind code	B2
Filing date	Mar 11, 2014
Priority date	Mar 12, 2013
Publication date	Oct 2, 2018
Grant date	Oct 2, 2018

Abstract

A liner assembly for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a grommet with a multiple of grommet cooling passages.

First claim

What is claimed is:

1. A liner assembly for a combustor of a gas turbine engine comprising:
a grommet with a plurality of grommet cooling passages,
wherein each of said plurality of grommet cooling passages includes an inlet in a sidewall of said grommet and an outlet in a surface of said grommet that interfaces to a combustion chamber of said combustor,
wherein said gas turbine engine is defined about a central longitudinal axis, and
wherein each grommet cooling passage of said plurality of grommet cooling passages has a grommet cooling passage longitudinal axis which is parallel to said central longitudinal axis.

2. The liner assembly as recited in claim 1 , wherein each of said plurality of grommet cooling passages define a respective angle with respect to a hot side of a heat shield.

3. The liner assembly as recited in claim 1 , wherein said plurality of grommet cooling passages include a plurality of leading edge grommet cooling passages directed downstream relative to a flow of combustion gases.

4. The liner assembly as recited in claim 1 , wherein said plurality of grommet cooling passages include a plurality of trailing edge grommet cooling passages directed upstream relative to a flow of combustion gases.

5. The liner assembly as recited in claim 1 , wherein said plurality of grommet cooling passages include a plurality of leading edge grommet cooling passages directed downstream relative to a flow of combustion gases and a plurality of trailing edge grommet cooling passages directed upstream relative to the flow of combustion gases.

6. The liner assembly as recited in claim 5 , wherein said plurality of leading edge grommet cooling passages are opposed to said plurality of trailing edge grommet cooling passages.

7. The liner assembly as recited in claim 1 , wherein said grommet surrounds a dilution hole.

8. The liner assembly as recited in claim 7 , wherein said grommet is integral with a heat shield.

9. The liner assembly as recited in claim 7 , wherein said grommet is mounted to a heat shield.

10. The liner assembly as recited in claim 9 , wherein each of said outlets of said plurality of grommet cooling passages is co-axial with a respective passage in said heat shield.

11. A method of increasing durability of a liner assembly in a combustor of a gas turbine engine, comprising:
directing an airflow through a plurality of leading edge grommet cooling passages directed downstream with respect to a flow of combustion gases and a plurality of trailing edge grommet cooling passages directed upstream with respect to the flow of combustion gases,
wherein each of said plurality of leading edge grommet cooling passages and plurality of trailing edge grommet cooling passages includes an inlet in a sidewall of a grommet and an outlet in a surface of said grommet that interfaces to a combustion chamber of said combustor,
wherein said gas turbine engine is defined about a central longitudinal axis, and wherein each leading edge grommet cooling passage of said plurality of leading edge grommet cooling passages and each trailing edge grommet cooling passage of said plurality of trailing edge grommet cooling passages has a grommet cooling passage longitudinal axis which is parallel to said central longitudinal axis.

12. The method as recited in claim 11 , further comprising:
directing the airflow at an angle with respect to a hot side of a heat shield.

13. The method as recited in claim 11 , further comprising:
opposing the plurality of leading edge grommet cooling passages with respect to the plurality of trailing edge grommet cooling passages.

14. The method as recited in claim 11 , further comprising:
locating the plurality of leading edge grommet cooling passages parallel to the plurality of trailing edge grommet cooling passages.

15. A liner assembly for a combustor of a gas turbine engine comprising:
a grommet with a plurality of grommet cooling passages,
wherein each of said plurality of grommet cooling passages includes an inlet in a sidewall of said grommet and an outlet in a surface of said grommet that interfaces to a combustion chamber of said combustor,
wherein said grommet surrounds a dilution hole,
wherein said grommet is mounted to a heat shield, and
wherein each of said outlets of said plurality of grommet cooling passages is co-axial with a respective passage in said heat shield.

Assignees

• United Technologies Corp

Inventors

• Erbas-Sen Nurhak
• Hoke James B
• Tu John S
• Pacheco-Tougas Monica

Classifications

• Y02T50/675

  Cross-Sectional Technologies · mapped topic

• F23R2900/03041

  Effusion cooled combustion chamber walls or domes · CPC title

• F23R2900/03044

  Impingement cooled combustion chamber walls or subassemblies · CPC title

• F23R2900/03042

  Film cooled combustion chamber walls or domes · CPC title

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title