Combustor assemblies for use in turbine engines and methods of assembling same

Patent metadata

Field	Value
Publication number	US-10690345-B2
Application number	US-201615203110-A
Country	US
Kind code	B2
Filing date	Jul 6, 2016
Priority date	Jul 6, 2016
Publication date	Jun 23, 2020
Grant date	Jun 23, 2020

Abstract

A combustor assembly for use in a gas turbine engine includes a combustor liner that defines a combustion chamber and includes an axial combustion portion and a curved transition portion. The combustion liner also includes an inner surface and an outer surface and a first plurality of cooling channels defined between the inner and outer surfaces. The combustor assembly also includes a sleeve substantially circumscribing the combustor liner such that an annular cavity is defined between the combustor liner and the sleeve. The sleeve includes a second plurality of cooling channels defined therethrough that are configured to channel a fluid against the combustor liner outer surface.

First claim

What is claimed is:

1. A reverse flow combustor assembly for use in a gas turbine engine, the combustor assembly comprising:
a combustor liner defining a combustion chamber and comprising an axial combustion portion and a curved transition portion, the curved transition portion positioned between an inlet and outlet of the combustor liner such that the inlet and outlet are spaced-apart along a curved centerline, wherein the combustor liner comprises an inner surface and an outer surface and a first plurality of cooling channels defined between the inner and outer surfaces; and
a sleeve substantially circumscribing the combustor liner such that an annular cavity is defined between the combustor liner and the sleeve, wherein the sleeve comprises a second plurality of cooling channels defined therethrough, the second plurality of cooling channels configured to channel a fluid against the combustor liner outer surface;
a single solid partition positioned within the cavity between the combustor liner and the sleeve, the partition being positioned axially between the axial combustion portion and the transition portion such that the partition extends circumferentially about an entirety of the combustor liner within the cavity to define a first cooling zone within the cavity upstream of the partition and a second cooling zone within the cavity downstream of the partition, the partition causing a differential pressure between the first and the second cooling zones.

2. The combustor assembly in accordance with claim 1 , wherein the inlet and an outlet of the combustor define a first length therebetween, and wherein the sleeve comprises an inlet and an outlet that define a second length therebetween that is substantially equal to the first length.

3. The combustor assembly in accordance with claim 1 , wherein the axial combustion portion comprises a plurality of dilution openings defined therethrough.

4. The combustor assembly in accordance with claim 1 , wherein the first plurality of cooling channels are oriented obliquely with respect to the combustor liner inner and outer surfaces.

5. The combustor assembly in accordance with claim 4 , wherein the second plurality of cooling channels are oriented perpendicular to the sleeve.

6. The combustor assembly in accordance with claim 1 , wherein the partition is integral with the sleeve.

7. A turbine engine comprising:
a compressor; and
a reverse flow combustor coupled in flow communication with the compressor, the combustor comprising at least one combustor assembly comprising:
a combustor liner defining a combustion chamber and comprising an axial combustion portion and a curved transition portion, the curved transition portion positioned between an inlet and outlet of the combustor liner such that the inlet and outlet are spaced-apart along a curved centerline, wherein the combustor liner comprises an inner surface and an outer surface and a first plurality of cooling channels defined between the inner and outer surfaces; and
a sleeve substantially circumscribing the combustor liner such that an annular cavity is defined between the combustor liner and the sleeve, wherein the sleeve comprises a second plurality of cooling channels defined therethrough, the second plurality of cooling channels configured to channel a fluid against the combustor liner outer surface;
a single solid partition positioned within the cavity between the combustor liner and the sleeve, the partition being positioned axially between the axial combustion portion and the transition portion such that the partition extends circumferentially about an entirety of the combustor liner within the cavity to define a first cooling zone within the cavity upstream of the partition and a second cooling zone within the cavity downstream of the partition, wherein the partition prevents cooling air within the first cooling zone from traveling downstream to the second cooling zone thereby causing a differential pressure between the first and the second cooling zones.

8. The turbine engine in accordance with claim 7 , wherein the inlet and an outlet of the combustor define a first length therebetween, and wherein the sleeve comprises an inlet and an outlet that define a second length therebetween that is substantially equal to the first length.

9. The turbine engine in accordance with claim 7 , wherein the axial combustion portion comprises a radially inner portion having a first plurality of circumferentially-spaced dilution openings defined therethrough and a radially outer portion comprising a second plurality of circumferentially-spaced dilution openings defined therethrough.

10. The turbine engine in accordance with claim 9 , wherein the first plurality of dilution openings are circumferentially offset with and are axially aligned with the second plurality of dilution openings.

11. The turbine engine in accordance with claim 7 , wherein the first plurality of cooling channels are oriented obliquely with respect to the combustor liner inner and outer surfaces, and wherein the second plurality of cooling channels are oriented perpendicular to the sleeve.

12. A method of manufacturing a reverse flow combustor assembly for use in a gas turbine engine, the method comprising:
forming a first plurality of cooling channels between an inner surface and an outer surface of a combustor liner, wherein the combustor liner defines a combustion chamber and includes an axial combustion portion and a curved transition portion, the curved transition portion positioned between an inlet and outlet of the combustor liner such that the inlet and outlet are spaced-apart along a curved centerline;
forming a second plurality of cooling channels through a sleeve; and
coupling the sleeve to the combustor liner such that the sleeve substantially circumscribes the combustor liner to define an annular cavity between the combustor liner and the sleeve, wherein the second plurality of cooling channels configured to channel a fluid against the combustor liner outer surface;
coupling a single solid partition within the cavity between the combustor liner and the sleeve axially between the axial combustion portion and the transition portion such that the partition extends circumferentially about an entirety of the combustor liner within the cavity to define a first cooling zone within the cavity upstream of the partition and a second cooling zone within the cavity downstream of the partition, wherein the partition causes a differential pressure between the first and the second cooling zones.

13. The method in accordance with claim 12 , wherein coupling the sleeve to the combustor liner comprises coupling the sleeve to the combustor liner such that the sleeve extends an entire length of the combustor liner.

14. The combustor assembly in accordance with claim 1 , wherein the fluid is deflected and dispersed within the annular cavity in both a first direction and a different second direction along the outer surface of the combustor liner.

Assignees

• Gen Electric

Inventors

• Monty Joseph Douglas
• Jacobson John Carl
• Howell Stephen John

Classifications

• Y02T50/60

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

• F23R3/54

  Reverse-flow combustion chambers · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F02C7/18

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

• F23R2900/03044

  Impingement cooled combustion chamber walls or subassemblies · CPC title