Combustor for use in a turbine engine

Patent metadata

Field	Value
Publication number	US-10641490-B2
Application number	US-201715398496-A
Country	US
Kind code	B2
Filing date	Jan 4, 2017
Priority date	Jan 4, 2017
Publication date	May 5, 2020
Grant date	May 5, 2020

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.

First claim

What is claimed is:

1. A combustor for use in a turbine engine, said combustor comprising:
an inner combustion hner comprising a forward inlet end and an aftward outlet end;
an outer combustion liner comprising a forward inlet end and an aftward outlet end, said inner and outer combustion liners circumscribing a centerline extending through the combustor, 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 defined at said forward inlet ends and said main portion comprising a flow outlet defined at said aftward outlet ends, said flow outlet positioned axially aftward from said flow inlet relative to the centerline, said flow inlet comprising a plurality of cavity inlet holes configured to discharge cavity airflow therefrom in an axially aftward and radially inward direction, relative to the centerline, to induce bulk swirl in the cavity airflow, wherein the cavity airflow is channeled from said flow inlet to said flow outlet with a predetermined angular momentum defined by the bulk swirl; and
a plurality of film cooling holes formed in at least one of said inner combustion liner or said outer combustion liner, said plurality of film cooling holes configured such that cooling airflow discharged therefrom flows helically relative to the centerline and towards said flow outlet;
wherein said inner combustion liner comprises a converging cross-sectional shape from the forward inlet end of the inner combustion liner to the aftward outlet end of the inner combustion liner such that the converging cross-sectional shape of the inner combustion liner is convex, relative to the centerline, from the forward inlet end of the inner combustion liner to the aftward outlet end of the inner combustion liner,
wherein said outer combustion lineer comprises a converging cross-sectional shape from the forward inlet end of the outer combustion liner to the aftward outlet end of the outer combustion liner such that the converging cross-sectional shape of the outer combustion liner is convex, relative to the centerline, from the forward inlet end of the outer combustion liner to the aftward outlet end of the outer combustion liner;
wherein the cavity portion is defined at a radially outermost region of the combustor and wherein the flow inlet is defined at the respective forward inlet end of the inner combustion liner and the forward inlet end of the outer combustion liner.

2. 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 or said outer combustion liner at an oblique angle relative to a radial axis oriented perpendicularly relative to the centerline.

3. The combustor in accordance with claim 2 , wherein said flow channel is angled in a circumferential direction relative to the centerline.

4. The combustor in accordance with claim 2 , wherein said flow channel is angled in an axial direction relative to the centerline.

5. The combustor in accordance with claim 2 , 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 50 degrees.

6. The combustor in accordance with claim 1 , 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, and wherein the cooling airflow is discharged so as to not disrupt the predetermined angular momentum of the cavity airflow.

7. The combustor in accordance with claim 1 , 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, wherein each dilution hole of the plurality of dilution holes comprises a chute coupled to the inner combustor liner.

8. The combustor in accordance with claim 7 , 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, and
wherein each chute facilitates channeling airflow from at least one airflow source through each dilution hole.

9. A turbine engine comprising:
a compressor assembly configured to discharge cornpressed 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 comprising a forward inlet end and an aftward outlet end:
an outer combustion liner comprising a forward inlet end and an aftward outlet end, 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, said flow inlet and said flow outlet positioned at opposing ends of said combustor, and said flow outlet positioned axially aftward from said flow inlet relative to a centerline of the turbine engine, said flow inlet comprising a plurality of cavity inlet holes configured to discharge cavity airflow therefrom in an axially aftward and radially inward direction, relative to the centerline, to induce bulk swirl in the cavity airflow, wherein the cavity airflow is channeled from said flow inlet to said flow outlet with a predetermined angular momentum defined by the bulk swirl; and
a plurality of film cooling holes formed in at least one of said inner combustion liner or said outer combustion liner, said plurality of film cooling holes configured such that cooling airflow discharged therefrom flows helically relative to the centerline and towards said flow outlet;

wherein said inner combustion liner comprises a converging cross-sectional shape from the forward inlet end of the inner combustion liner to the aftward outlet end of the inner combustion liner such that the converging cross-sectional shape of the inner combustion liner is convex, relative to the centerline, from the forward inlet end of the inner combustion liner to the aftward outlet end of the inner combustion liner,
wherein said outer combustion liner comprises a converging cross-sectional shape from the forward inlet end of the outer combustion liner to the aftward outlet end of the outer combustion liner such that the converging cross-sectional shape of the outer combustion liner is convex, relative to the centerline, from the forward inlet end of the outer combustion liner to the aftward outlet end of the outer combustion liner;
wherein the cavity portion is defined at a radially outermost region of the combustor and wherein the flow inlet is defined at the respective forward inlet end of the inner combustion liner and the forward inlet end of the outer combustion liner.

10. The turbine engine in accordance with claim 9 , 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 or said outer combustion liner at an oblique angle relative to a radial axis of the turbine engine, and wherein each flow channel is angled to channel the cooling airflow in an aftward axial direction relative to the centerline.

Assignees

• Gen Electric

Inventors

• Tangirala Venkat Eswarlu
• Joshi Narendra Digamber
• Monahan Sarah Marie

Classifications

• F23R3/42

  characterised by the arrangement or form of the flame tubes or combustion chambers · CPC title

• F23R3/58

  Cyclone or vortex type combustion chambers · CPC title

• F23R3/12

  inducing a vortex · CPC title

• F23R3/04

  Air inlet arrangements · CPC title

• F23R3/045

  using pipes · CPC title