COMBUSTOR FOR USE IN A TURBINE ENGINE

• US 20180187888A1
• Filed: 01/04/2017
• Published: 07/05/2018
• Est. Priority Date: 01/04/2017
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Granted Patent: US 10,641,490 B2
• Time in Patent Office: Days
• CPC Class Codes: F23R 2900/03042 Film cooled combustion chamber walls or domes
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/02 characterised by the air-flow or gas-flow configuration reverse- flow combustion chambers F23R3/54; cyclone or vortex type combustion chambers F23R3/58
• CPC Class Codes: F23R 3/04 Air inlet arrangements
• CPC Class Codes: F23R 3/045 using pipes
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube
• CPC Class Codes: F23R 3/12 inducing a vortex
• CPC Class Codes: F23R 3/42 characterised by the arrangement or form of the flame tubes or combustion chambers
• CPC Class Codes: F23R 3/52 Toroidal combustion chambers
• CPC Class Codes: F23R 3/58 Cyclone or vortex type combustion chambers

Abstract

A combustor for use in a turbine engine that includes an inner combustion liner and an outer combustion liner. An interior is defined between the inner combustion liner and the outer combustion liner, and the interior includes a cavity portion and a main portion extending radially inward from the cavity portion. The cavity portion includes a flow inlet and the main portion includes a flow outlet. A plurality of film cooling holes are formed in at least one of the inner combustion liner and the outer combustion liner. The plurality of film cooling holes are configured such that cooling airflow discharged therefrom flows helically relative to a centerline of the turbine engine and towards the flow outlet.

Claims

1. A combustor for use in a turbine engine, said combustor comprising:

an inner combustion liner;

an outer combustion liner, wherein an interior is defined between said inner combustion liner and said outer combustion liner, said interior comprising a cavity portion and a main portion extending radially inward from said cavity portion, said cavity portion comprising a flow inlet and said main portion comprising a flow outlet; and

a plurality of film cooling holes formed in at least one of said inner combustion liner and said outer combustion liner, said plurality of film cooling holes configured such that cooling airflow discharged therefrom flows helically relative to a centerline of the turbine engine and towards said flow outlet.

2. The combustor in accordance with claim 1, wherein said inner combustion liner and said outer combustion liner are convex relative to the centerline of the turbine engine.

3. The combustor in accordance with claim 1, wherein each film cooling hole of said plurality of film cooling holes comprises a flow channel that extends through said at least one of said inner combustion liner and said outer combustion liner at an oblique angle relative to a radial axis of the turbine engine.

4. The combustor in accordance with claim 3, wherein said flow channel is angled in a circumferential direction relative to the centerline of the turbine engine.

5. The combustor in accordance with claim 3, wherein said flow channel is angled in an axial direction relative to the centerline of the turbine engine

6. The combustor in accordance with claim 3, wherein said plurality of film cooling holes are formed such that the oblique angle of said flow channel relative to the radial axis is greater than about 50 degrees.

7. The combustor in accordance with claim 1, wherein said flow inlet comprises a plurality of cavity inlet holes configured to discharge cavity airflow therefrom that flows from said flow inlet towards said flow outlet with a predetermined angular momentum.

8. The combustor in accordance with claim 7, wherein said plurality of film cooling holes are configured to discharge the cooling airflow therefrom such that the predetermined angular momentum of the cavity airflow is maintained when the cooling airflow mixes with the cavity airflow.

9. The combustor in accordance with claim 7 further comprising a plurality of dilution holes formed in said inner combustion liner, said plurality of dilution holes configured such that dilution airflow discharged therefrom flows helically relative to the centerline of the turbine engine.

10. The combustor in accordance with claim 9, wherein said plurality of dilution holes are configured to discharge the dilution airflow therefrom such that the predetermined angular momentum of the cavity airflow is maintained when the dilution airflow mixes with the cavity airflow.

11. A turbine engine comprising:

a compressor assembly configured to discharge compressed air therefrom; and

a combustor coupled in flow communication with said compressor assembly configured to receive the compressed air, said combustor comprising;

an inner combustion liner;

an outer combustion liner, wherein an interior is defined between said inner combustion liner and said outer combustion liner, said interior comprising a cavity portion and a main portion extending radially inward from said cavity portion, said cavity portion comprising a flow inlet and said main portion comprising a flow outlet; and

a plurality of film cooling holes formed in at least one of said inner combustion liner and said outer combustion liner, said plurality of film cooling holes configured such that cooling airflow discharged therefrom flows helically relative to a centerline of the turbine engine and towards said flow outlet.

12. The turbine engine in accordance with claim 11, wherein said inner combustion liner and said outer combustion liner are convex relative to the centerline of the turbine engine.

13. The turbine engine in accordance with claim 11, wherein each film cooling hole of said plurality of film cooling holes comprises a flow channel that extends through said at least one of said inner combustion liner and said outer combustion liner at an oblique angle relative to a radial axis of the turbine engine.

14. The turbine engine in accordance with claim 13, wherein said flow channel is angled in a circumferential direction relative to the centerline of the turbine engine.

15. The turbine engine in accordance with claim 13, wherein said flow channel is angled in an axial direction relative to the centerline of the turbine engine.

16. The turbine engine in accordance with claim 13, wherein said plurality of film cooling holes are formed such that the oblique angle of said flow channel relative to the radial axis is greater than about 50 degrees.

17. The turbine engine in accordance with claim 11, wherein said flow inlet comprises a plurality of cavity inlet holes configured to discharge cavity airflow therefrom that flows from said flow inlet towards said flow outlet with a predetermined angular momentum.

18. The turbine engine in accordance with claim 17, wherein said plurality of film cooling holes are configured to discharge the cooling airflow therefrom such that the predetermined angular momentum of the cavity airflow is maintained when the cooling airflow mixes with the cavity airflow.

19. The turbine engine in accordance with claim 17 further comprising a plurality of dilution holes formed in said inner combustion liner, said plurality of dilution holes configured such that dilution airflow discharged therefrom flows helically relative to the centerline of the turbine engine, the dilution airflow discharged at a greater flow rate than the cooling airflow.

20. The turbine engine in accordance with claim 19, wherein said plurality of dilution holes are configured to discharge the dilution airflow therefrom such that the predetermined angular momentum of the cavity airflow is maintained when the dilution airflow mixes with the cavity airflow.