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US 4,820,116 AGranted
Turbine cooling for gas turbine engine
AssigneeUnited Technologies Corporation
InventorsEdward J. Hovan, Guilford E. Stephens, Samuel C. Peele
Life of this patent
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Abstract
A gas turbine engine cooling system has cooling air cooling in order, the first stage disk (50) and blades (22), the second stage vanes (24), and the second stage air seal (82). Air passes to the second stage vane from the rotor through a reaction nozzle (62) effecting a reaction stage adding energy to the rotor and cooling the air.
Claims2
A gas turbine engine having:
- a compressor;
- a first stage rotor blade set having a cooling flowpath through the blades;
- a first stage rim structure supporting said first stage rotor blades;
- a first stage disk supporting said rim structure;
- a second stage stator vane set having cooling flowpaths through the vanes;
- a second stage interstage labyrinth seal sealing around said second stage stator vane set;
- a second stage rotor blade set;
- stator nozzles mounted in said stator adjacent to said first stage disk and angled to direct airflow passing therethrough in the direction of rotation of said disk;
- means for conveying cooling air from substantially a compressor outlet to said stator nozzles;
- said rim structure having cooling airflow openings therethrough from an upstream side to a downstream side of said disk;
- means for conveying airflow through said rotor from said stator nozzles to said cooling airflow openings in said rim structure;
- nozzles secured to said rotor, angled to direct flow therethrough opposite the direction of rotation of said rotor and located to discharge air into a plenum between seal sections of said second stage interstage labyrinth seal;
- and
- the cooling flowpath through said second stage vanes being in fluid communication with said seal plenum.
11 more elements
Dependencies are inferred from the claim text.