초록 ▼

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 fou

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.

대표청구항 ▼

Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims in which we claim: 1. A gas turbine engine comprising: a fan, compressor, combustor, high pressure turbine, and low pressure turbine disposed in s

Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims in which we claim: 1. 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. 2. An engine according to claim 1 wherein: said second bleed circuit is joined to said compressor to provide to said vanes said second coolant at a second pressure lower than a first pressure of said first coolant; and said fourth circuit is joined to said compressor to provide to said blades said fourth coolant at a fourth pressure lower than a third pressure of said third coolant. 3. An engine according to claim 2 wherein: said vanes and blades include opposite leading and trailing edges with corresponding cooling holes thereat, and are configured for dropping pressure in combustion gases in turn; and said fourth circuit is joined to said compressor at a stage preceding said second circuit to provide backflow margins at said trailing edge holes of both said vanes and blades. 4. An engine according to claim 3 wherein said third circuit is joined to said compressor at a stage preceding said first circuit to provide a backflow margin at said blade leading edge holes. 5. An engine according to claim 4 wherein said third circuit is configured to effect a static pressure in said third coolant greater than the total relative pressure in said combustion gases at said blade leading edge holes. 6. An engine according to claim 5 wherein: said compressor includes rows of compressor blades for pressurizing air sequentially in corresponding stages; said first circuit bypasses said combustor to reach said vane forward channels; said second and fourth circuits extend radially outwardly from said compressor outboard of said combustor to reach said aft channels of said vanes and blades; and said third circuit extends inboard of said combustor to reach said blade forward channels. 7. An engine according to claim 5 wherein: said LP turbine includes a row of first stage LP nozzle vanes directly following said HP blades; and said fourth circuit is routed radially inwardly through said LP vanes to reach said blade aft channels. 8. An engine according to claim 7 wherein: said blades are trapped in corresponding dovetail slots in the perimeter of said disk by forward and aft blade retainers; said third circuit joins said row of blades at said forward blade retainer; and said fourth circuit joins said row of blades at said aft blade retainer. 9. An engine according to claim 8 wherein: said blades are joined to said compressor blades by a drive shaft extending inboard of said combustor; said forward blade retainer extends inwardly along the forward face of said disk to define an annular manifold disposed in flow communication with said dovetail slots around said disk perimeter; and said third circuit is joined in flow communication with said manifold. 10. An engine according to claim 9 further comprising: an outer case surrounding said combustor and HP nozzle to define said first circuit from said compressor to said vane forward channels; a plurality of spoolies extending radially through said outer case to said HP nozzle to define an outlet end of said second circuit to said vane aft channels; said blade forward channels extending to said dovetail slots for receiving said third coolant from said third circuit; and said blade aft channels extending aft through the shanks of said blades to said aft blade retainer for receiving said fourth coolant from said fourth circuit. 11. 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. 12. An engine according to claim 11 wherein: said second bleed circuit is joined to said compressor to provide to said vanes said second coolant at a second pressure lower than a first pressure of said first coolant; and said fourth circuit is joined to said compressor to provide to said blades said fourth coolant at a fourth pressure lower than a third pressure of said third coolant. 13. An engine according to claim 12 wherein: said nozzle vanes are configured for dropping pressure in combustion gases across the leading and trailing edges thereof; said blades are configured for additionally dropping pressure in said combustion gases across the leading and trailing edges thereof; and said fourth circuit is joined to an earlier stage in said compressor than said second circuit to provide a backflow margin in said fourth coolant at said blade trailing edge holes and a backflow margin in said second coolant at said vane trailing edge holes. 14. An engine according to claim 13 wherein said third circuit is joined to an earlier stage in said compressor than said first circuit to provide a backflow margin in said third coolant at said blade leading edge holes. 15. An engine according to claim 13 wherein said third circuit is configured to effect a static pressure in said third coolant greater than the total relative pressure in said combustion gases at said blade leading edge holes. 16. An engine according to claim 13 wherein: said compressor includes rows of compressor blades for pressurizing air sequentially in corresponding stages; said first circuit bypasses said combustor to reach said vane forward channels; said second and fourth circuits extend radially outwardly from said compressor outboard of said combustor to reach said aft channels of said vanes and blades; and said third circuit extends inboard of said combustor to reach said blade forward channels. 17. An engine according to claim 16 further comprising: a low pressure turbine following said HP turbine in serial flow communication, and including a row of first stage LP nozzle vanes directly following said HP blades; and said fourth circuit is routed radially inwardly through said LP nozzle vanes to reach said aft channels of said HP blades. 18. An engine according to claim 17 wherein: said blades are trapped in corresponding dovetail slots in the perimeter of said disk by forward and aft blade retainers; said third circuit joins said row of blades at said forward blade retainer; and said fourth circuit joins said row of blades at said aft blade retainer. 19. An engine according to claim 18 wherein: said blades are joined to said compressor blades by a drive shaft extending inboard of said combustor; said forward blade retainer extends inwardly along the forward face of said disk to define an annular manifold disposed in flow communication with said dovetail slots around said disk perimeter; and said third circuit is joined in flow communication with said manifold. 20. An engine according to claim 19 wherein: said blade forward channels extend to said dovetail slots for receiving said third coolant from said third circuit; and said blade aft channels extend aft through the shanks of said blades to said aft blade retainer for receiving said fourth coolant from said fourth circuit. 21. An engine according to claim 20 further comprising: an outer case surrounding said combustor and HP nozzle to define said first circuit from said compressor to said vane forward channels; and a plurality of spoolies extending radially through said outer case to said HP nozzle to define an outlet end of said second circuit to said vane aft channels. 22. 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. 23. A method according to claim 22 further comprising: dropping pressure in combustion gases across said vanes and blades in turn; bleeding from said compressor said second coolant at a lower pressure than said first coolant, and said fourth coolant at a lower pressure than said third coolant; and bleeding said fourth coolant from said compressor earlier than said second coolant for providing backflow margins at said trailing edge holes of both said vanes and blades. 24. A method according to claim 23 further comprising bleeding said first, second, third, and fourth coolants from said compressor in independent flowpaths at correspondingly different pressures and temperatures.