A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a com
A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. The liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end. The quench air jets include a plurality of primary stage air jets and a plurality of secondary stage air jets. The plurality of primary stage air jets are each spaced from the plurality of secondary stage air jets along the axial direction and together provide the combustion chamber with a quench airflow.
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1. A rich-quench-lean combustor assembly for a gas turbine engine defining an axial direction and a circumferential direction, the combustor assembly comprising: a fuel nozzle;a dome, the fuel nozzle attached to the dome; anda liner attached to or formed integrally with the dome, the liner and the d
1. A rich-quench-lean combustor assembly for a gas turbine engine defining an axial direction and a circumferential direction, the combustor assembly comprising: a fuel nozzle;a dome, the fuel nozzle attached to the dome; anda liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber, the liner extending between a first forward end and a first aft end, the liner comprising a first annular forward section disposed adjacent the first forward end, a first annular aft section disposed adjacent the first aft end and extending parallel to the first annular forward section, and a first annular convergence section interconnecting the first annular forward section and the first annular aft section, the first annular convergence section disposed at an oblique angle to the first annular forward section and the first annular aft section, the liner comprising a plurality of first quench air jets positioned between the first forward end and first aft end, the first quench air jets comprising a plurality of first primary stage air jets and a plurality of first secondary stage air jets, the plurality of first primary stage air jets each spaced from the plurality of first secondary stage air jets along the axial direction; andwherein the plurality of first primary stage air jets are positioned at a forward portion of the convergence section, and are oriented substantially perpendicularly to the axial direction, and the plurality of first secondary stage air jets are positioned at an aft portion of the first convergence section, and are oriented oblique relative to the axial direction such that air is injected in an upstream direction relative to a flow through the combustion chamber, wherein at least one of the plurality of first primary stage air jets defines a first cross-sectional area, and at least one of the plurality of first secondary stage air jets defines a second cross-sectional area less than the first cross-sectional area. 2. The combustor assembly of claim 1, wherein the liner is an outer liner, wherein the combustor assembly further comprises: an inner liner attached to or formed integrally with the dome, the inner liner defining at least in part the combustion chamber, the inner liner extending between a second forward end and a second aft end, the inner liner comprising a second annular forward section disposed adjacent the second forward end, a second annular aft section disposed adjacent the second aft end and extending parallel to the second annular forward section, and a second annular convergence section interconnecting the second annular forward section and the second annular aft section, the second annular convergence section disposed at an oblique angle to the second annular forward section and the second annular aft section, the inner liner comprising a plurality of second quench air jets positioned between the second forward end and the second aft end, the second quench air jets comprising a plurality of second primary stage air jets and a plurality of second secondary stage air jets, the plurality of second primary stage air jets each spaced from the plurality of second secondary stage air jets along the axial direction, wherein the plurality of second primary stage air jets are positioned at a forward portion of the second convergence section, and are oriented substantially perpendicularly to the axial direction, and the plurality of second secondary stage air jets are positioned at an aft portion of the second convergence section, and are oriented oblique relative to the axial direction, wherein at least one of the plurality of second primary stage air jets defines a third cross-sectional area, and at least one of the plurality of second secondary stage air jets defines a fourth cross-sectional area less than the third cross-sectional area, wherein the combustor assembly is configured to receive compressed air for combustion, wherein at least about sixty percent (60%) of the compressed air for combustion is introduced into the combustion chamber through the first and second quench air jets as a quench airflow. 3. The combustor assembly of claim 2, wherein the combustor assembly is configured to provide between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the first and second primary stage air jets and between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the first and second secondary stage air jets. 4. The combustor assembly of claim 1, wherein the combustor assembly defines a ratio of a number of the first secondary stage air jets to a number of the first primary stage air jets of at least about 2:1. 5. The combustor assembly of claim 1, wherein the plurality of first primary stage air jets are each spaced along the circumferential direction, and wherein the plurality of first secondary stage air jets are each spaced along the circumferential direction. 6. The combustor assembly of claim 5, wherein each of the plurality of first primary stage air jets and the plurality of first secondary stage air jets are evenly spaced along the circumferential direction. 7. The combustor assembly of claim 5, wherein at least one of the plurality of first primary stage air jets or the plurality of first secondary stage air jets are unevenly spaced along the circumferential direction. 8. The combustor assembly of claim 7, wherein the fuel nozzle comprises a plurality of fuel nozzles spaced evenly along the circumferential direction, and wherein the uneven spacing along the circumferential direction of the at least one of the plurality of first primary stage air jets or the plurality of first secondary stage air jets correlates to a position of the plurality of fuel nozzles. 9. The combustor assembly of claim 7, wherein the fuel nozzle comprises a plurality of fuel nozzles spaced evenly along the circumferential direction, and wherein at least one of the plurality of first primary stage air jets or the plurality of first secondary stage air jets vary in size along the circumferential direction, the variation in size correlating to a position of the plurality of fuel nozzles. 10. The combustor assembly of claim 1, wherein the combustion chamber defines a centerline, wherein the plurality of first secondary stage air jets are each configured as elongated slots having a widthwise direction defining an oblique angle relative to the centerline. 11. The combustor assembly of claim 1, wherein the liner includes an inlet transition immediately forward of each of the plurality of first primary stage air jets, wherein the inlet transition defines a radius of curvature of at least about 0.65 inches. 12. The combustor assembly of claim 1, wherein each of the plurality of first primary stage air jets defines an inlet having an elliptical shape, wherein the elliptical shape of the inlet includes a minor radius of curvature of at least about 0.25 inches and a major radius of curvature of at least about 0.4 inches. 13. The combustor assembly of claim 1, wherein the liner is an outer liner, wherein the combustor assembly further comprises: an inner liner attached to or formed integrally with the dome, the inner liner defining at least in part the combustion chamber, the inner liner extending between a second forward end and a second aft end, the inner liner comprising a second annular forward section disposed adjacent the second forward end, a second annular aft section disposed adjacent the second aft end and extending parallel to the second annular forward section, and a second annular convergence section interconnecting the second annular forward section and the second annular aft section, the second annular convergence section disposed at an oblique angle to the second annular forward section and the second annular aft section, the inner liner comprising a plurality of second quench air jets positioned between the second forward end and the second aft end, the second quench air jets comprising a plurality of second primary stage air jets and a plurality of second secondary stage air jets, the plurality of second primary stage air jets each spaced from the plurality of second secondary stage air jets along the axial direction, wherein the plurality of second primary stage air jets are positioned at a forward portion of the second convergence section, and are oriented substantially perpendicularly to the axial direction, and the plurality of second secondary stage air jets are positioned at an aft portion of the second convergence section, and are oriented oblique relative to the axial direction, wherein at least one of the plurality of second primary stage air jets defines a third cross-sectional area, and at least one of the plurality of second secondary stage air jets defines a fourth cross-sectional area less than the third cross-sectional area. 14. The combustor assembly of claim 13, wherein the combustor assembly defines a forward height within the combustion chamber between the outer liner and the inner liner at a location forward of the plurality of first and second quench air jets, wherein the combustor assembly defines an aft height within the combustion chamber between the outer liner and the inner liner at a location aft of the plurality of first and second quench air jets, and wherein a ratio of the forward height to the aft height is at least about 1.75:1. 15. The combustor assembly of claim 1, wherein the dome and the liner are each formed of a ceramic matrix composite material. 16. A gas turbine engine defining an axial direction and a circumferential direction, the gas turbine engine comprising: a compressor section and a turbine section arranged in serial flow order; anda rich-quench-lean combustor assembly positioned between the compressor section and the turbine section, the combustor assembly comprising:a fuel nozzle;a dome, the fuel nozzle attached to the dome; anda liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber, the liner extending between a first forward end and a first aft end, the liner comprising a first annular forward section disposed adjacent the first forward end, a first annular aft section disposed adjacent the first aft end and extending parallel to the first annular forward section, and a first annular convergence section interconnecting the first annular forward section and the first annular aft section, the first annular convergence section disposed at an oblique angle to the first annular forward section and the first annular aft section, the liner comprising a plurality of first quench air jets positioned between the first forward end and first aft end, the first quench air jets comprising a plurality of first primary stage air jets and a plurality of first secondary stage air jets, the plurality of first primary stage air jets each spaced from the plurality of first secondary stage air jets along the axial direction; andwherein the plurality of first primary stage air jets are positioned at a forward portion of the convergence section, and are oriented substantially perpendicularly to the axial direction, and the plurality of first secondary stage air jets are positioned at an aft portion of the first convergence section, and are oriented oblique relative to the axial direction such that air is injected in an upstream direction relative to a flow through the combustion chamber, wherein at least one of the plurality of first primary stage air jets defines a first cross-sectional area, and at least one of the plurality of first secondary stage air jets defines a second cross-sectional area less than the first cross-sectional area. 17. The gas turbine engine of claim 16, wherein the fuel nozzle is configured to provide a mixture of fuel and air to the combustion chamber having at least about a seventy percent (70%) mixedness. 18. The gas turbine engine of claim 16, wherein the liner is an outer liner, wherein the combustor assembly further comprises: an inner liner attached to or formed integrally with the dome, the inner liner defining at least in part the combustion chamber, the inner liner extending between a second forward end and a second aft end, the inner liner comprising a second annular forward section disposed adjacent the second forward end, a second annular aft section disposed adjacent the second aft end and extending parallel to the second annular forward section, and a second annular convergence section interconnecting the second annular forward section and the second annular aft section, the second annular convergence section disposed at an oblique angle to the second annular forward section and the second annular aft section, the inner liner comprising a plurality of second quench air jets positioned between the second forward end and the second aft end, the second quench air jets comprising a plurality of second primary stage air jets and a plurality of second secondary stage air jets, the plurality of second primary stage air jets each spaced from the plurality of second secondary stage air jets along the axial direction, wherein the plurality of second primary stage air jets are positioned at a forward portion of the second convergence section, and are oriented substantially perpendicularly to the axial direction, and the plurality of second secondary stage air jets are positioned at an aft portion of the second convergence section, and are oriented oblique relative to the axial direction, wherein at least one of the plurality of second primary stage air jets defines a third cross-sectional area, and at least one of the plurality of second secondary stage air jets defines a fourth cross-sectional area less than the third cross-sectional area, wherein the combustor assembly is configured to receive compressed air for combustion, wherein at least about sixty percent (60%) of the compressed air for combustion is introduced into the combustion chamber through the first and second quench air jets as a quench airflow. 19. The gas turbine engine of claim 18, wherein the combustor assembly is configured to provide between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the first and second primary stage air jets and between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the first and second secondary stage air jets.
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이 특허에 인용된 특허 (22)
Kim, Won-Wook; Snyder, Timothy S., Advanced quench pattern combustor.
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