Gas turbine engine combustor with integrated combustor vane

Patent metadata

Field	Value
Publication number	US-9404654-B2
Application number	US-201213627697-A
Country	US
Kind code	B2
Filing date	Sep 26, 2012
Priority date	Sep 26, 2012
Publication date	Aug 2, 2016
Grant date	Aug 2, 2016

Abstract

A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber.

First claim

What is claimed is:

1. A combustor section for a gas turbine engine comprising:
a combustion chamber; and
a combustor vane that extends at least partially into said combustion chamber;
wherein said combustor vane includes an internal helical axial swirler configured with a passage that extends helically through said helical axial swirler and about and along an axis.

2. The combustor section as recited in claim 1 further comprising:
an outer liner; and
an inner liner disposed radially inward of the outer liner and the combustion chamber being defined between said outer and inner liners, said combustor vane located between said outer liner and said inner liner.

3. The combustor section as recited in claim 1 , wherein said combustor vane defines an axial length between 35%-65% of said combustion chamber, said axial length defined along an axis which extends from a fuel injector through said combustion chamber.

4. The combustor section as recited in claim 1 , wherein said combustor vane includes film cooling along a leading edge thereof.

5. The combustor section as recited in claim 1 , wherein said helical axial swirler is along a leading edge thereof.

6. The combustor section as recited in claim 1 , wherein said combustor vane is manufactured of a refractory metal core (RMC) material.

7. The combustor section as recited in claim 6 , wherein said combustor vane includes an RMC circuit along a main body thereof.

8. The combustor section as recited in claim 6 , wherein said combustor vane includes a centerline RMC microcircuit with pedestals along a trailing edge thereof.

9. The combustor section as recited in claim 1 , wherein said helical axial swirler is one of a multiple of helical axial swirlers along a leading edge of said combustor vane.

10. The combustor section as recited in claim 9 , wherein a first of said multiple of helical axial swirlers is axially offset with respect to a second of said multiple of helical axial swirlers.

11. The combustor section as recited in claim 9 , wherein said multiple of helical axial swirlers are stacked along said leading edge.

12. The combustor section as recited in claim 1 , wherein the combustor vane is configured with an axially extending passage operable to direct fluid axially into said helical axial swirler.

13. The combustor section as recited in claim 12 , wherein said helical axial swirler is disposed within said axially extending passage.

14. The combustor section as recited in claim 12 , wherein said helical axial swirler is disposed inline with said axially extending passage.

15. A combustor section vane for a combustor for a gas turbine engine comprising:
an outer airfoil wall surface between a leading edge and a trailing edge; and
at least one helical axial swirler within said leading edge;
wherein said helical axial swirler is configured with a passage that extends helically through said helical axial swirler and about and along an axis.

16. The combustor section vane as recited in claim 15 , further comprising a multiple of helical axial swirlers within said leading edge.

17. The combustor section vane as recited in claim 15 , wherein said helical axial swirler is configured within an axially extending passage in a wall of the combustor section vane.

18. The combustor section vane as recited in claim 17 , wherein said axially extending passage includes a convergent-divergent section downstream of said helical axial swirler.

19. A combustor section for a gas turbine engine comprising:
a combustion chamber; and
a combustor vane that extends at least partially into said combustion chamber;
wherein said combustor vane includes a dilution jet on a leading edge thereof, and said dilution jet includes an axially extending helical passage with a convergent-divergent section.

Assignees

• Cunha Frank J
• Erbas-Sen Nurhak
• United Technologies Corp

Inventors

• Cunha Frank J
• Erbas-Sen Nurhak

Classifications

• F23R2900/03041

  Effusion cooled combustion chamber walls or domes · CPC title

• F23R3/06

  Arrangement of apertures along the flame tube · CPC title

• F23R2900/03045

  Convection cooled combustion chamber walls provided with turbolators or means for creating turbulences to increase cooling · CPC title

• F23R2900/00018

  Manufacturing combustion chamber liners or subparts · CPC title

• F23R2900/03042

  Film cooled combustion chamber walls or domes · CPC title