Compound turbine cooled engine
AssigneeGeneral Electric Company
InventorsChing-Pang Lee, Thomas Ory Moniz, Robert Joseph Orlando
Life of this patent
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Abstract
A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. The turbine includes a row of nozzle vanes followed by a row of rotor blades. The vanes and blades have corresponding forward and aft internal cooling channels. First, second, third, and fourth bleed circuits are joined in flow communication with different stages of the compressor for bleeding pressurized air therefrom at different pressures to provide coolant to the forward and aft channels of the turbine vanes and blades.
Claims24
A gas turbine engine comprising:
- a fan, compressor, combustor, high pressure turbine, and low pressure turbine disposed in serial flow communication;
- said compressor including rows of compressor blades for pressurizing air sequentially in corresponding stages;
- said HP turbine including a row of hollow nozzle vanes having forward and aft internal cooling channels followed by a row of hollow rotor blades having forward and aft internal cooling channels;
- and
- first, second, third, and fourth bleed circuits joined in flow communication with different stages of said compressor for bleeding pressurized air therefrom at different pressures to provide first and second coolants to said vane forward and aft channels from correspondingly different stages of said compressor, and third and fourth coolants to said blade forward and aft channels, respectively.
A gas turbine engine comprising:
- a multistage compressor disposed in serial flow communication with a combustor and high pressure turbine;
- said turbine including a nozzle having a row of stator vanes followed by a row of rotor blades extending from a rotor disk;
- each of said vanes including a forward cooling channel feeding leading edge cooling holes, and an aft cooling channel feeding trailing edge cooling holes;
- each of said blades including a forward cooling channel feeding leading edge cooling holes, and an aft cooling channel feeding trailing edge cooling holes;
- a first bleed circuit joining a last stage of said compressor with said vane forward channels to provide thereto a first coolant;
- a second bleed circuit joining an intermediate stage of said compressor and said vane aft channels to provide thereto a second coolant;
- a third bleed circuit joining said compressor and said blade forward channels to provide thereto a third coolant;
- and
- a fourth bleed circuit joining an intermediate stage of said compressor and said blade aft channels to provide thereto a fourth coolant.
A method of cooling nozzle vanes and rotor blades in a high pressure turbine comprising:
- bleeding a first coolant between the last stage of a multistage compressor and a forward cooling channel in said vanes for discharge from leading edge cooling holes therein;
- bleeding a second coolant between an intermediate stage of said compressor and an aft cooling channel in said vanes for discharge from trailing edge cooling holes therein;
- bleeding a third coolant between said compressor and a forward cooling channel in said blades for discharge from leading edge cooling holes therein;
- and
- bleeding a fourth coolant between an intermediate stage of said compressor and an aft cooling channel in said blades for discharge from trailing edge cooling holes therein.
Dependencies are inferred from the claim text.