Hybrid through holes and angled holes for combustor grommet cooling

Patent metadata

Field	Value
Publication number	US-2016201913-A1
Application number	US-201514886973-A
Country	US
Kind code	A1
Filing date	Oct 19, 2015
Priority date	Oct 20, 2014
Publication date	Jul 14, 2016
Grant date	—

Abstract

Aspects of the disclosure are directed to a cooling design feature for inclusion in a liner of an aircraft, comprising: a plurality of angled holes, and at least one through hole separating all combinations of any two of the angled holes, wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

First claim

What is claimed is:

1 . A cooling design feature for inclusion in a liner of an aircraft, comprising:
a plurality of angled holes; and at least one through hole separating all combinations of any two of the angled holes, wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

2 . The cooling design feature of claim 1 , wherein the cooling design feature includes a grommet, and wherein the entirety of the grommet is configured to include an alternating sequence of through holes and angled holes.

3 . The cooling design feature of claim 1 , wherein the at least one through hole includes at least two through holes that are directly adjacent to one another.

4 . The cooling design feature of claim 3 , wherein the at least two through holes that are directly adjacent to one another are proximate to a rib.

5 . The cooling design feature of claim 1 , wherein the cooling design feature includes a grommet, and wherein the grommet is configured to be included in a gas turbine engine liner of the aircraft.

6 . A gas turbine engine liner for an aircraft, comprising:
a shell; a panel coupled to the shell and comprising at least one cooling design feature, wherein the at least one cooling design feature includes:
a plurality of angled holes; and
at least one through hole separating all combinations of any two of the angled holes,
wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and
wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

7 . The liner of claim 6 , wherein the panel includes threaded studs configured to allow the panel to be bolted to the shell.

8 . The liner of claim 6 , wherein the shell includes a plurality of impingement holes and wherein the panel includes a plurality of effusion holes.

9 . The liner of claim 6 , wherein the cooling design feature includes a grommet, and wherein the entirety of the grommet is configured to include an alternating sequence of through holes and angled holes.

10 . The liner of claim 6 , wherein the at least one through hole includes at least two through holes that are directly adjacent to one another.

11 . The liner of claim 10 , wherein the at least two through holes that are directly adjacent to one another are proximate to a rib.

12 . The liner of claim 6 , further comprising:
a groove configured to provide a passage for cooling air with respect to the at least one through hole.

Assignees

• United Technologies Corp

Inventors

• Tu John S
• Hoke James B
• Pacheco-Tougas Monica
• Snyder Timothy S
• Moura Dennis M
• Van Anthony

Classifications

• F23R2900/03044

  Impingement cooled combustion chamber walls or subassemblies · CPC title

• 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

• F23R2900/03042

  Film cooled combustion chamber walls or domes · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title