Combustor heat panel arrangement having holes offset from seams of a radially opposing heat panel

Patent metadata

Field	Value
Publication number	US-9416970-B2
Application number	US-62752809-A
Country	US
Kind code	B2
Filing date	Nov 30, 2009
Priority date	Nov 30, 2009
Publication date	Aug 16, 2016
Grant date	Aug 16, 2016

Abstract

A combustor module for a gas turbine engine is provided that includes a first annular liner assembly extending along a longitudinal axis of the engine. The first annular liner assembly includes a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell. The first heat shield panels form a segmented ring defining a plurality of first axial seams therebetween. The combustor module further includes a bulkhead coupled to the first annular liner assembly. The bulkhead provides a plurality of fuel nozzles for passing a first mass flow comprising fuel and air. The combustor module further includes a second annular liner assembly coupled to the bulkhead. The second annular liner assembly is in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H. The second annular liner assembly includes an air admittance hole having a mean diameter D extending along a hole axis. The hole axis is offset from the first axial seam defined by the first heat shield panels.

First claim

We claim:

1. A combustor module for a gas turbine engine, comprising:
a first annular liner assembly extending along a longitudinal axis, the first annular liner assembly comprising a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell, the first heat shield panels forming a segmented ring defining a plurality of first axial seams therebetween;
a bulkhead coupled to the first annular liner assembly, the bulkhead providing a plurality of fuel nozzles for passing a first mass flow comprising fuel and air;
a second annular liner assembly coupled to the bulkhead, the second annular liner assembly in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H, the second annular liner assembly comprising a plurality of air admittance holes each having a mean diameter D, each air admittance hole extending along a hole axis, and each hole axis being offset from one of the first axial seams by an equal distance S.

2. The combustor module of claim 1 wherein the mean diameter D of the air admittance hole is greater than about 1.27 centimeters.

3. The combustor module of claim 1 wherein the mean diameter D, a gas flow g through the combustor, a jet flow j through the air admittance hole, and a momentum flux ratio J of the combustor flow and the jet flow are sufficient to provide a jet penetration distance Y along the hole axis, wherein Y is greater than or equal to H.

4. The combustor module of claim 3 , wherein the jet penetration distance Y is defined by the equation Y=D j [g/(g+j)]√{square root over (J)}.

5. The combustor module of claim 1 wherein the first annular support shell is an inner support shell.

6. The combustor module of claim 1 wherein the first heat shield panels forming a segmented ring further comprise a forward heat shield panel and an aft heat shield panel defining a circumferential seam therebetween.

7. The combustor module of claim 1 wherein the second annular liner assembly comprises a second annular support shell and a plurality of second heat shield panels coupled to the second annular support shell, the second heat shield panels forming a segmented ring having second axial seams therebetween.

8. The combustor module of claim 7 , wherein the second heat shield panels further comprise forward heat shield panels and aft heat shield panels defining a circumferential seam therebetween.

9. The combustor module of claim 1 , wherein the first heat shield panels are thermally decoupled from the first annular support shell.

10. The combustor module of claim 1 , wherein the offset between the first axial seam and the axis of the air admittance hole is sufficient to assure the outer diameter of the air admittance hole is aligned with a pattern of film cooling holes on the first heat shield panel.

11. A method for mitigating effects of maldistributed thermal loads in a combustor module, the method comprising the steps of:
selecting a first annular liner assembly comprising a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell, the first heat shield panels forming a segmented ring defining a plurality of first axial seams therebetween;
coupling a bulkhead to the first annular liner assembly;
coupling a second annular liner assembly to the bulkhead, the second annular liner assembly in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H;
selecting an arrangement of air admittance holes penetrating through the second annular liner assembly, the air admittance holes having a mean diameter D and extending along a hole axis; and
offsetting each hole axis of the air admittance holes in the second annular liner assembly from one of the first axial seams by an equal distance S.

12. The method of claim 11 , wherein the step of selecting an arrangement of air admittance holes comprises establishing fully penetrating flow through the air admittance holes.

13. The method according to claim 12 , wherein the step of establishing fully penetrating flow comprises determining a jet penetration distance Y along the hole axis.

14. The method according to claim 13 , wherein the jet penetration distance Y is determined using a computational fluid dynamics model.

15. The method according to claim 13 wherein the jet penetration distance Y is determined using a computational fluid dynamics model.

16. The method according to claim 15 , wherein the empirical determination of the jet penetration distance Y correlates the mean diameter D, a gas flow g through the combustor, a jet flow j through the air admittance hole, and a momentum flux ratio J of the combustor flow and the jet flow.

17. The method according to claim 16 , wherein the empirical determination of the jet penetration distance Y is defined by the equation Y=D j [g/(g+j)]√{square root over (J)}.

18. The method according to claim 11 , wherein the step of offsetting the air admittance holes from the first axial seam comprises selecting the offset such that the outer diameter of the air admittance hole is aligned with a pattern of film cooling holes on the first heat shield panel.

19. The method according to claim 11 , wherein the step of selecting a plurality of first heat shield panels comprises selecting a forward heat shield panel and an aft heat shield panel.

20. The method according to claim 11 , wherein the second annular liner assembly comprises a second annular support shell and a plurality of second heat shield panels coupled to the second annular support shell, the second heat shield panels forming a segmented ring having second axial seams therebetween, the method further comprising the step of selecting an arrangement of air admittance holes penetrating through the first annular liner assembly.

21. The method according to claim 20 , further including the step of offsetting the air admittance holes penetrating through the first annular liner assembly from the second axial seam defined by the second heat shield panels.

Assignees

• Kirsopp Philip J
• Dornback Aaron J
• Hoke James B
• United Technologies Corp

Inventors

• Kirsopp Philip J
• Dornback Aaron J
• Hoke James B

Classifications

• F23R2900/03041

  Effusion cooled combustion chamber walls or domes · CPC title

• F23R3/002 Primary

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

• Y02T50/675

  Cross-Sectional Technologies · mapped topic

• Y02T50/60

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