Gas turbine engine including a compressor bleed air system
A gas turbine includes a compressor bleed air system that includes a compressor shroud, a compressor bleed air plenum, and a plurality of compressor bleed air passages arranged through the compressor shroud that provide fluid communication between the compressor airflow passage and the compressor bleed air plenum. The plurality of compressor bleed air passages are arranged circumferentially spaced apart from one another, where a first compressor bleed air passage and a second compressor bleed air passage circumferentially adjacent to one another are spaced apart a first angular spacing, and the respective inlets of at least a portion of a remainder of the plurality of compressor bleed air passages are circumferentially spaced apart from one another a second angular spacing, the second angular spacing being less than the first angular spacing.
1. A gas turbine engine comprising:
a compressor; and
a compressor bleed air system including:
a compressor shroud defining a compressor airflow passage therewithin;
a compressor bleed air plenum;
a plurality of compressor stator vanes connected with the compressor shroud and extending into the compressor airflow passage, the plurality of compressor stator vanes being arranged circumferentially spaced apart from each other within a stator vane stage of the compressor; and
a plurality of compressor bleed air passages arranged through the compressor shroud within the stator vane stage and providing fluid communication between the compressor airflow passage and the compressor bleed air plenum, each of the plurality of compressor bleed air passages including an inlet at the compressor airflow passage and an outlet at the compressor bleed air plenum,
wherein respective inlets of respective ones of the plurality of compressor bleed air passages are arranged circumferentially spaced apart from one another about a centerline axis of the compressor shroud, a first compressor bleed air passage of the plurality of compressor bleed air passages and a second compressor bleed air passage of the plurality of compressor bleed air passages are arranged circumferentially adjacent to one another and an inlet of the first compressor bleed air passage and an inlet of the second compressor bleed air passage are circumferentially spaced apart from one another by a first angular spacing, and respective inlets of at least a portion of a remainder of the plurality of compressor bleed air passages circumferentially sequentially arranged adjacent to each other are circumferentially spaced apart from one another by a second angular spacing, the second angular spacing being less than the first angular spacing,
wherein each one of the plurality of compressor stator vanes includes a leading edge and a trailing edge extending along an axial direction along the centerline axis, and a base connecting the compressor stator vane to the compressor shroud, the stator vane stage being defined axially between the leading edge and the trailing edge, and
wherein each one of the plurality of compressor bleed air passages is arranged circumferentially within the stator vane stage between a respective pair of compressor stator vanes among the plurality of compressor stator vanes.
2. The gas turbine engine according to claim 1 , wherein at least one of the plurality of compressor bleed air passages defines a constant width throat at the inlet, and defines a diverging portion from a downstream end of the throat to the outlet.
3. The gas turbine engine according to claim 1 , wherein the respective inlets of the portion of the remainder of the plurality of compressor bleed air passages are circumferentially spaced apart from one another by an equal angular spacing.
4. The gas turbine engine according to claim 3 , wherein the respective inlets of the portion of the remainder of the plurality of compressor bleed air passages are arranged circumferentially between the first compressor bleed air passage and the second compressor bleed air passage.
5. The gas turbine engine according to claim 4 , wherein the inlet of a first one of the portion of the remainder of the plurality of compressor bleed air passages is arranged adjacent to the first compressor bleed air passage and is arranged the second angular spacing from the inlet of the first compressor bleed air passage, and the inlet of a second one of the portion of the remainder of the plurality of compressor bleed air passages is arranged adjacent to the second compressor bleed air passage and is arranged the second angular spacing from the inlet of the second compressor bleed air passage.
6. The gas turbine engine according to claim 1 , wherein at least one of the plurality of compressor bleed air passages defines a diffuser flow passage in which a size of the diffuser flow passage increases from the inlet to the outlet.
7. The gas turbine engine according to claim 6 , wherein the inlet and the outlet are circular openings.
8. The gas turbine engine according to claim 7 , wherein the diffuser flow passage defines a centerline extending from a center of the inlet to a center of the outlet.
9. The gas turbine engine according to claim 8 , wherein the centerline is curvilinear extending from the center of the inlet to the center of the outlet.
10. A gas turbine engine comprising:
a compressor; and
a compressor bleed air system including:
a compressor shroud defining a compressor airflow passage therewithin;
a compressor bleed air plenum;
a plurality of compressor stator vanes connected with the compressor shroud and extending into the compressor airflow passage, the plurality of compressor stator vanes being arranged circumferentially spaced apart from each other within a stator vane stage of the compressor; and
a plurality of compressor bleed air passages arranged through the compressor shroud within the stator vane stage and providing fluid communication between the compressor airflow passage and the compressor bleed air plenum, each of the plurality of compressor bleed air passages including an inlet at the compressor airflow passage and an outlet at the compressor bleed air plenum,
wherein (a) the plurality of compressor bleed air passages are arranged in a first group of the plurality of the compressor bleed air passages and a second group of the plurality of the compressor bleed air passages, (b) respective inlets of respective ones of each of the plurality of compressor bleed air passages in the first group and respective inlets of respective ones of each of the plurality of compressor bleed air passages in the second group are arranged equally spaced apart a first angular spacing from one another about a centerline axis of the compressor shroud, and (c) a first end of the first group and a second end of the second group are arranged spaced apart from one another a second angular spacing greater than the first angular spacing,
wherein each one of the plurality of compressor stator vanes includes a leading edge and a trailing edge extending along an axial direction along the centerline axis, and a base connecting the compressor stator vane to the compressor shroud, the stator vane stage being defined axially between the leading edge and the trailing edge, and
wherein each one of the plurality of compressor bleed air passages is arranged circumferentially within the stator vane stage between a respective pair of compressor stator vanes among the plurality of compressor stator vanes.
11. The gas turbine engine according to claim 10 , wherein a second end of the first group and a first end of the second group are arranged spaced apart from one another a third angular spacing greater than the first angular spacing.
12. The gas turbine engine according to claim 11 , wherein the second angular spacing is the same as the third angular spacing.
13. The gas turbine engine according to claim 11 , wherein the second angular spacing is different from the third angular spacing.
14. The gas turbine engine according to claim 10 , wherein at least one of the plurality of compressor bleed air passages defines a constant width throat at the inlet, and defines a diverging portion from a downstream end of the throat to the outlet.
15. The gas turbine engine according to claim 10 , wherein at least one of the plurality of compressor bleed air passages defines a diffuser flow passage in which a size of the diffuser flow passage increases from the inlet to the outlet.
16. The gas turbine engine according to claim 15 , wherein the inlet and the outlet are circular openings.
17. The gas turbine engine according to claim 16 , wherein the diffuser flow passage defines a centerline extending from a center of the inlet to a center of the outlet.
18. The gas turbine engine according to claim 17 , wherein the centerline is curvilinear extending from the center of the inlet to the center of the outlet.