Gas turbine engine wave geometry combustor liner panel

Patent metadata

Field	Value
Publication number	US-2016109128-A1
Application number	US-201414893781-A
Country	US
Kind code	A1
Filing date	Jun 12, 2014
Priority date	Jun 14, 2013
Publication date	Apr 21, 2016
Grant date	—

Abstract

A panel for a combustor of a gas turbine engine includes a cold side defining at least one convex portion and at least one concave portion. The concave portion is in communication with a passage. A method of operating a combustor section of a gas turbine engine includes: directing an impingement flow toward a multiple of peaks on a cold side of a panel; directing the impingement flow from the multiple of peaks toward a multiple of troughs with a multiple of entrances on the cold side of the panel; and directing the impingement flow through the multiple of entrances and a respective multiple of effusion passages through the panel.

First claim

What is claimed is:

1 . A panel for a combustor of a gas turbine engine, the panel comprising:
a cold side defining at least one convex portion and at least one concave portion in communication with a passage.

2 . A panel as recited in claim 1 , wherein the convex and the concave portions form a wave pattern.

3 . The panel as recited in claim 2 , wherein
the concave portion in communication with the passage defines an entrance to the passage; and the entrance is displaced from a peak of said wave pattern.

4 . The panel as recited in claim 2 , wherein said passage is an effusion flow passage.

5 . The panel as recited in claim 2 , wherein said wave pattern is skewed.

6 . The panel as recited in claim 2 , wherein said wave pattern includes transverse ridges.

7 . The panel as recited in claim 5 , wherein
at least one concave portion defines an entrance to the passage; and said at least one entrance is separated longitudinally by adjacent peaks and laterally by adjacent transverse ridges.

8 . The panel as recited in claim 2 , wherein the wave pattern is continuous.

9 . The panel as recited in claim 1 , wherein at least one concave portion and at least one convex portion are integrally formed adjacent one another from a common surface of the cold side.

10 . A combustor of a gas turbine engine, the combustor comprising:
a shell with a multiple of impingement flow passages; and a panel mounted to said shell, said panel defining a wave pattern on a cold side thereof, said wave pattern including a multiple of peaks and troughs, said troughs defining a multiple of corresponding entrances to a respective multiple of effusion passages through said panel.

11 . The combustor as recited in claim 10 , further comprising a plurality of studs which extend from said cold side of said panel.

12 . The combustor as recited in claim 10 , wherein at least one of said multiple of impingement flow passages is directed at one of said multiple of peaks.

13 . The combustor as recited in claim 10 , wherein at least one of said multiple of effusion flow passages defines an angle through the panel.

14 . The combustor as recited in claim 10 , further comprising a multiple of transverse ridges along said multiple of peaks and troughs.

15 . The combustor as recited in claim 10 , wherein said wave pattern is skewed toward downstream nozzle guide vanes.

16 . A method of operating a combustor section of a gas turbine engine, the method comprising:
directing an impingement flow toward a multiple of peaks on a cold side of a panel; directing the impingement flow from the multiple of peaks toward a multiple of troughs with a multiple of entrances on the cold side of the panel; and directing the impingement flow through the multiple of entrances and a respective multiple of effusion passages through the panel.

17 . The method as recited in claim 16 , further comprising segregating each of the entrances.

18 . The method as recited in claim 16 , further comprising separating each of the entrances from one another longitudinally by adjacent peaks and laterally by adjacent transverse ridges.

19 . The method as recited in claim 16 , further comprising locating each of the entrances in a lowest portion of the multiple of troughs.

20 . The method as recited in claim 16 , further comprising displacing each of the multiple of entrances from respective exits of the effusion passages such that the effusion passages define respective angles relative to a hot side of the liner panel.

Assignees

• United Technologies Corp

Inventors

• Drake Christopher

Classifications

• F05D2250/611

  undulated · CPC title

• F23R3/44

  Combustion chambers comprising a {single} tubular flame tube within a tubular casing (reverse-flow combustion chambers F23R3/54) · CPC title

• F23R3/002 Primary

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

• F23R3/50

  Combustion chambers comprising an annular flame tube within an annular casing (toroidal combustion chambers F23R3/52) · CPC title

• F02C7/18

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