US-12618562-B1 - Optimized front end aerodynamics for advanced RQL combustor

Abstract

An injector for a combustor and a gas turbine engine with the same includes a stem extending from a mount to a nozzle and a swirler. The nozzle extends from the stem along an axis. The nozzle includes a center body and a fuel passage. The swirler circumscribes the nozzle and includes an inner air passage and an outer air passage. An array of circumferentially-spaced inner vanes within the inner air passage each form an inner acute angle with the axis. An array of circumferentially-spaced outer vanes within the outer air passage each form an outer acute angle less than the first acute angle.

Inventors

• Timothy Snyder

Assignees

• RTX CORPORATION

Dates

Publication Date

20260505

Application Date

20250630

Claims (17)

1. 1 . An injector for a combustor of a gas turbine engine, the injector comprising: a mount; a stem extending from a proximal end joined with the mount to a distal end; a nozzle extending from the distal end along an axis, the nozzle comprising: a center body coaxial with the axis; and a fuel passage extending through the center body; and a swirler circumscribing the center body, the swirler comprising: an inner air passage extending through the swirler; and an outer air passage extending through the swirler radially outward from the inner air passage; wherein the inner air passage includes a plurality of inner vanes spaced circumferentially about the axis, each inner vane of the plurality of inner vanes forming a first acute angle with a radial datum relative to the axis to generate an inner recirculation zone in a chamber of the combustor, and wherein the outer air passage includes a plurality of outer vanes spaced circumferentially about the axis, each outer vane of the plurality of outer vanes forming a second acute angle with the radial datum that is less than the first acute angle, and the inner acute angles of the plurality of inner vanes are operatively associated with an inner swirl number to generate a diverging airflow cone outward of the recirculation zone, and wherein the outer acute angles of the plurality of outer vanes are operatively associated with an outer swirl number, and wherein the outer swirl number is less than the inner swirl number.
2. 2 . The injector of claim 1 , wherein the inner air passage and the outer air passage have a co-swirl orientation.
3. 3 . The injector of claim 1 , wherein the inner air passage and the outer air passage have a counter swirl orientation.
4. 4 . The injector of claim 1 , wherein the fuel passage includes a primary fuel passage coaxial with the axis and a plurality of secondary fuel passages radially outward from the primary fuel passage.
5. 5 . The gas turbine engine of claim 1 , wherein the outer swirl number is is greater than zero and less than or equal to 2.5.
6. 6 . The injector of claim 1 , wherein the inner swirl number is greater than or equal to 0.5 and less than or equal to 0.7.
7. 7 . The injector of claim 1 , wherein the outer swirl number is greater than 0.0 and less than or equal to 0.5.
8. 8 . The injector of claim 7 , wherein the inner swirl number is greater than or equal to 0.5 and less than or equal to 0.7.
9. 9 . A gas turbine engine comprising: a combustor shell that bounds a combustion chamber; a casing surrounding the combustor shell to form a plenum; a plurality of quench openings extending through the combustor shell that divides the combustion chamber into a primary zone upstream from the plurality of quench openings and a secondary zone downstream from the plurality of quench openings; and an injector extending through the combustor shell into the primary zone, the injector comprising: a mount supported by the casing; a stem extending from a proximal end joined with the mount to a distal end; a nozzle extending from the distal end along an axis, the nozzle comprising: a center body coaxial with the axis; and a fuel passage extending through the center body; and a swirler circumscribing the center body, the swirler comprising: an inner air passage extending through the swirler; and an outer air passage extending through the swirler radially outward from the inner air passage; wherein the inner air passage includes a plurality of inner vanes spaced circumferentially about the axis, each inner vane of the plurality of inner vanes forming a first acute angle with a radial datum relative to the axis to generate an inner recirculation zone in the combustion chamber, and wherein the outer air passage includes a plurality of outer vanes spaced circumferentially about the axis, each outer vane of the plurality of outer vanes forming a second acute angle with the radial datum that is less than the first acute angle to generate a diverging airflow cone outward of the recirculation zone, and wherein the second acute angles of the plurality of outer vanes are operatively associated with an outer divergence angle, and wherein a projection of the outer divergence angle intersects the combustor shell downstream from the primary zone.
10. 10 . The gas turbine engine of claim 9 , wherein the inner air passage and the outer air passage have a co-swirl orientation.
11. 11 . The gas turbine engine of claim 9 , wherein the fuel passage is an annular fuel passage.
12. 12 . The gas turbine engine of claim 9 , wherein the inner acute angles of the plurality of inner vanes are operatively associated with an inner swirl number, and wherein the outer acute angles of the plurality of outer vanes are operatively associated with an outer swirl number, and wherein the outer swirl number is less than the inner swirl number.
13. 13 . The gas turbine engine of claim 12 , wherein the outer swirl number is is greater than zero and less than or equal to 2.0.
14. 14 . The gas turbine engine of claim 12 , wherein the outer swirl number greater than zero and less than or equal to 2.5.
15. 15 . The gas turbine engine of claim 14 , wherein the inner swirl number is greater than or equal to 0.5 and less than or equal to 0.7.
16. 16 . The gas turbine engine of claim 12 , wherein the outer swirl number is greater than 0.0 and less than or equal to 0.5.
17. 17 . The gas turbine engine of claim 16 , wherein the inner swirl number is greater than or equal to 0.5 and less than or equal to 0.7.

Description

BACKGROUND The present disclosure relates generally to combustion for gas turbine engines, and more particularly, to features of injectors and combustors that reduce front end combustor liner and shell temperatures during rich-quench-lean (“RQL”) combustion. The aerodynamic design of a combustor front end is affected by the shape of the combustor liner and the swirl flow strength. Injectors utilize swirling flow to aid atomization of fuel and mix fuel with air to achieve a rich mixture within the combustor front end. Swirling flow with relatively high tangential velocity can cause recirculation within corner regions of the front end, which can interfere with cooling flows discharged along the combustor liners and increase temperatures of the combustor liner and shell. While past aerodynamic designs for combustor front ends are considered satisfactory for the intended purpose, further improvements to the aerodynamic design for reducing temperatures of the combustor liner and shell can further improve operational life of the combustor for RQL combustion. SUMMARY An injector according to an example embodiment of this disclosure includes a mount, a stem, a nozzle, and a swirler. The stem extends from a proximal end joined with the mount to a distal end. The nozzle extends from the distal end of the stem along an axis. The nozzle includes a center body coaxial with the axis and a fuel passage, which extends from the stem through the center body. The swirler circumscribes the center body and includes an inner air passage and an outer air passage. The inner air passage and the outer air passage extend through the swirler. The outer air passage is radially outward from the inner air passage relative to the axis. The inner air passage includes a plurality of inner vanes spaced circumferentially about the axis, each inner vane forming a first acute angle with a radial datum relative to the axis. The outer air passage includes a plurality of outer vanes spaced circumferentially about the axis, each outer vane forming a second acute angle with the radial datum that is less than the first acute angle. A gas turbine engine according to a further example embodiment of this disclosure includes a combustor shell, a casing, a plurality of quench openings, and an injector. The combustor shell bounds a combustion chamber. The casing surrounds the combustor shell to form a plenum. The plurality of quench openings extends through the combustor shell that divides the combustion chamber into a primary zone upstream from the quench openings and a secondary zone downstream from the quench openings. The injector extends through the combustor shell into the primary zone. The injector includes a mount, a stem, a nozzle, and a swirler. The stem extends from a proximal end joined with the mount to a distal end. The nozzle extends from the distal end of the stem along an axis. The nozzle includes a center body coaxial with the axis and a fuel passage, which extends from the stem through the center body. The swirler circumscribes the center body and includes an inner air passage and an outer air passage. The inner air passage and the outer air passage extend through the swirler. The outer air passage is radially outward from the inner air passage relative to the axis. The inner air passage includes a plurality of inner vanes spaced circumferentially about the axis, each inner vane forming a first acute angle with a radial datum relative to the axis. The outer air passage includes a plurality of outer vanes spaced circumferentially about the axis, each outer vane forming a second acute angle with the radial datum that is less than the first acute angle. BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic cross-sectional view of an example gas turbine engine. FIG. 2 is a schematic cross-sectional view of an example combustor of the gas turbine engine that includes injectors with low outer swirl flow. FIG. 3 is a schematic cross-sectional view of another example combustor of the gas turbine engine that includes injectors with low outer swirl flow. FIG. 4 is a schematic cross-sectional view of an injector of the combustor depicted by FIG. 2, or the combustor depicted by FIG. 3. FIG. 5 is a partial developed view of inner air passage vanes from the injector depicted in FIG. 4. FIG. 6 is a partial developed view of outer air passage vanes from the injector depicted in FIG. 4. FIG. 7 is a partial developed view of inner air passage vanes from the injector depicted in FIG. 4 illustrating a counter-swirl inner air passage configuration. DETAILED DESCRIPTION As disclosed herein, injectors combine an inner air flow and an outer air flow to centralize a rich air-fuel mixture within the combustor front end. Swirl (i.e., tangential velocity) imposed on the inner air flow atomizes and mixes the inner air flow with fuel discharged into the combustor front end to produce a rich air-fuel mixture within a front end of the combustor. An inner recirculatio