Compound nozzle cooled engine

• US 20080112794A1
• Filed: 11/10/2006
• Published: 05/15/2008
• Est. Priority Date: 11/10/2006
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Moniz, Thomas Ory, Orlando, Robert Joseph, Lee, Ching-Pang
• Granted Patent: US 7,870,743 B2
• Time in Patent Office: Days
• US Class Current: 415/115
• CPC Class Codes: F01D 5/189 the insert having a tubular cross-section, e.g. airfoil shape
• CPC Class Codes: F02C 7/12 Cooling of plants of component parts, see the relevant subclasses, e.g. F01D; cooling of engines in general F01P
• CPC Class Codes: F05D 2220/3212 the first stage of a turbine
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. The turbine includes a nozzle followed by a row of rotor blades. A first bleed circuit is joined in flow communication between the last stage of the compressor and a forward cooling channel in vanes of the nozzle for feeding first cooling holes therein with pressurized primary air at a first pressure. A second bleed circuit is joined in flow communication between an intermediate stage of the compressor and aft cooling channels in the nozzle vanes to feed second cooling holes with pressurized secondary air at a second pressure less than the first pressure.

Claims

1. A gas turbine engine comprising:

a fan, compressor, combustor, high pressure (HP) turbine, and low pressure turbine in serial flow communication;

said compressor including rows of compressor blades for pressurizing air sequentially in corresponding stages;

said HP turbine including a nozzle having a row of stator vanes followed by a row of rotor blades;

each of said vanes having circumferentially opposite pressure and suction sides extending in span radially and in chord axially between opposite leading and trailing edges, and being hollow with an imperforate septum bridging said opposite sides to define forward and aft cooling channels feeding corresponding first cooling holes along said leading edge and second cooling holes along said trailing edge;

a first bleed circuit joined in flow communication between a last stage of said compressor and said vane forward channels to provide thereto pressurized primary air at a first pressure; and

a second bleed circuit joined in flow communication between an intermediate stage of said compressor and said vane aft channels to provide thereto pressurized secondary air at a second pressure less than said first pressure.

2. An engine according to claim 1 wherein said aft channels are imperforate along both said pressure and suction sides, and said second holes extend aft therefrom toward said trailing edges.

3. An engine according to claim 2 wherein said intermediate stage is configured to effect a static pressure in said secondary air greater than static pressure in combustion gases at said trailing edges to provide a backflow margin at said second holes.

4. An engine according to claim 3 wherein said first holes include a row of aft holes extending aft from said forward channel along both said pressure side and said aft channel.

5. An engine according to claim 4 further comprising:

an outer case surrounding said combustor and nozzle to define said first circuit from said compressor to said nozzle; and

a plurality of spoolies extending radially through said outer case to said nozzle to define an outlet end of said second circuit.

6. An engine according to claim 5 wherein said second holes extend aft through said pressure side short of said trailing edge.

7. An engine according to claim 5 wherein said second holes extend aft through said suction side short of said trailing edge.

8. An engine according to claim 5 wherein said second holes extend aft through said trailing edge between said pressure and suction sides which are imperforate thereat.

9. An engine according to claim 5 further comprising:

an inner case disposed radially inwardly of said combustor to further define said first circuit; and

said outer and inner bands have corresponding outer and inner inlets corresponding with each of said vanes to define a discharge end of said first circuit.

10. A method of cooling said nozzle vanes according to claim 5 comprising:

bleeding primary air from the last stage of said compressor to said forward cooling channels in said vanes for discharge from said first holes; and

bleeding secondary air from said intermediate stage of said compressor to said aft cooling channels in said vanes for discharge from said second holes.

11. A gas turbine engine comprising:

a multistage compressor disposed in serial flow communication with a combustor and high pressure (HP) turbine;

said turbine including a nozzle having a row of stator vanes followed by a row of rotor blades;

each of said vanes including a forward cooling channel feeding first cooling holes, and an aft cooling channel feeding second cooling holes;

a first bleed circuit joined in flow communication between a last stage of said compressor and said vane forward channels to provide thereto pressurized primary air at a first pressure; and

a second bleed circuit joined in flow communication between an intermediate stage of said compressor and said vane aft channels to provide thereto pressurized secondary air at a second pressure less than said first pressure.

12. An engine according to claim 11 wherein:

said nozzle further comprises radially outer and inner bands integrally joined to corresponding end of said vanes, and each of said vanes further includes chordally opposite leading and trailing edges and a septum extending radially between said bands to separate said forward and aft channels;

said first holes are disposed near said leading edges, and said second holes are disposed near said trailing edges; and

said intermediate stage is configured to effect a static pressure in said secondary air greater than static pressure in combustion gases at said trailing edges to provide a backflow margin at said second holes.

13. An engine according to claim 12 wherein:

each of said vanes further includes a concave pressure side and an opposite, convex suction side, and said septum bridges said opposite sides to isolate said forward and aft channels;

said aft channel is imperforate along said pressure side; and

said second holes extend aft from said aft channel.

14. An engine according to claim 13 wherein said aft channel is imperforate along said suction side.

15. An engine according to claim 13 wherein said first holes include a row of aft holes extending aft from said forward channel along both said pressure side and said aft channel.

16. An engine according to claim 13 wherein said second holes extend aft through said pressure side short of said trailing edge.

17. An engine according to claim 13 wherein said second holes extend aft through said suction side short of said trailing edge.

18. An engine according to claim 13 wherein said second holes extend aft through said trailing edge between said pressure and suction sides which are imperforate thereat.

19. An engine according to claim 13 further comprising:

an outer case surrounding said combustor and nozzle to define said first circuit from said compressor to said nozzle; and

a plurality of spoolies extending radially through said outer case to said nozzle to define an outlet end of said second circuit.

20. An engine according to claim 19 further comprising:

an inner case disposed radially inwardly of said combustor to further define said first circuit; and

said outer and inner bands have corresponding outer and inner inlets corresponding with each of said vanes to define a discharge end of said first circuit.

21. A method of cooling first stage nozzle vanes following in turn a combustor and compressor in a gas turbine engine, comprising:

bleeding primary air from the last stage of said compressor to a forward cooling channel in said vanes for discharge from first cooling holes therein; and

bleeding secondary air from an intermediate stage of said compressor to an aft cooling channel in said vanes for discharge from second cooling holes therein.

22. A method according to claim 21 further comprising:

dropping pressure in combustion gases across leading and trailing edges of said vanes; and

bleeding said secondary air from said intermediate stage at a static pressure greater than the static pressure of said combustion gases along said vane trailing edges.