Compressor rotor cooling system for a gas turbine

• US 5,685,158 A
• Filed: 03/31/1995
• Issued: 11/11/1997
• Est. Priority Date: 03/31/1995
• Status: Expired due to Term

• Current Assignee: GE Power Systems (General Electric Company)
• Original Assignee: General Electric Company
• Inventors: Plemmons, Larry Wayne, Stover, Curtis W., Glynn, Christopher Charles, Lenahan, Dean Thomas, Pedersen, Poul D., Miller, Frederick M.
• Primary Examiner(s): FREAY, CHARLES GRANT
• Application Number:
  US08/414,699
• Time in Patent Office: 956 Days
• Field of Search: 60/39.02, 60/726, 60/728, 60/730, 415/115, 416/95, 416/96 R
• US Class Current: 60/726
• CPC Class Codes: F01D 5/081 Cooling fluid being directed on the side of the rotor disc or at the roots of the blades F01D5/087 takes precedence
• CPC Class Codes: F02C 7/185 Cooling means for reducing the temperature of the cooling air or gas
• CPC Class Codes: F04D 29/321 for axial flow compressors
• CPC Class Codes: F04D 29/584 cooling or heating the machine F04D29/5846, F04D29/5853 take precedence

Abstract

In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor.

Claims

1. In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor, wherein said cooling air comprises compressor discharge air extracted from the combustion system downstream of the compressor and upstream of the plurality of combustors, and further including a heat exchanger arranged to cool said cooling air after extraction from the combustion system.

2. The improvement of claim 1 wherein said piping includes at least one insulated pipe extending through one or more radial struts which connect a casing of the combustion system to a stage one nozzle support structure.

3. The improvement of claim 1 wherein said means for dividing said cooling air includes an annular compressor discharge pressure seal which divides said annular compressor aft cavity into an upper cavity and a lower cavity, and wherein said cooling air is fed into said upper cavity.

4. The improvement of claim 1 and including means for dividing said cooling air into discrete cooling air components for

a) cooling said peripheral rim of the last stage of the compressor;

and

b) cooling said bore.

5. The improvement of claim 4 wherein said cooling air is introduced into an annular compressor aft cavity located adjacent said last stage, and wherein said means for dividing said cooling air provides an additional discrete cooling air component for preventing hot compressor discharge air from entering said annular compressor aft cavity.

6. The improvement of claim 5 wherein said means for dividing said cooling air includes an annular compressor discharge pressure seal which divides said annular compressor aft cavity into an upper cavity and a lower cavity, and wherein said means for supplying cooling air feeds said cooling air into said upper cavity.

7. The improvement of claim 6 wherein said annular compressor discharge pressure seal includes a discourager seal potion which permits a predetermined quantity of cooling air to pass from said upper cavity to said lower cavity.

8. The improvement of claim 7 wherein said annular compressor aft cavity is defined in part by an annular cooling plate, and wherein said discrete cooling air component for cooling said peripheral rim of the last stage of the compressor flows through a plurality of axial passages in said annular cooling plate from said upper cavity.

9. The improvement of claim 8 wherein said cooling plate includes a radially outer tip including means for permitting said additional discrete cooling air component to flow from said upper cavity into the hot compressor discharge air and thereby prevent said hot air from entering said upper cavity.

10. The improvement of claim 8 wherein said cooling plate includes at least one radial passage extending between said lower cavity and said bore and wherein said discrete cooling air component for cooling said bore flows through said at least one radial passage.

11. The improvement of claim 10 wherein annular cooling cavities are provided between rim portions of adjacent ones of said rotor disks and wherein slots are provided in said rotor disks, permitting cooling air in said bore to flow into said annular cooling cavities between said adjacent ones of said rotor disks.

12. The improvement of claim 11 wherein each of said blades has a blade shank seated within a corresponding dovetail slot in a respective disk and wherein at least one axial cooling air slot is formed by a gap between a bottom surface of the blade shank and the dovetail slot, and further wherein cooling air flowing through said axial passages in said cooling plate enters an annular cavity formed between said cooling plate and said peripheral rim of said last stage, and then through said cooling air axial slots and into said annular cooling cavities between said adjacent ones of said rotor disks.

13. The improvement of claim 11 wherein the peripheral rims of at least the rotor disks of the last stage and a next-to-last stage are constructed with passageways to permit some of the cooling air in the respective annular cooling cavities to escape into said compressor discharge air.

14. The improvement of claim 13 wherein said passageways include a pair of radial portions separated by an axially oriented, annular buffer cavity.

15. The improvement of claim 11 wherein said slots in said rotor disks have angled surfaces at entrances to said annular cavities to facilitate swirling of said cooling air within said cavities.

16. The improvement of claim 5 wherein said annular compressor aft cavity is defined in part by an annular cooling plate, and wherein said discrete cooling air component for cooling said peripheral rim of the last stage of the compressor flows through a plurality of axial passages in said annular cooling plate from said upper cavity.

17. The improvement of claim 16 wherein said cooling plate includes at least one radial passage extending between said lower cavity and said bore and wherein said discrete cooling air component for cooling said bore flows through said at least one radial passage.

18. The improvement of claim 17 wherein annular cooling cavities are provided between rim portions of adjacent ones of said rotor disks and wherein slots are provided in said rotor disks, permitting cooling air in said bore to flow into said annular cooling cavities between said adjacent ones of said rotor disks.

19. The improvement of claim 18 wherein each of said blades has a blade shank seated within a corresponding dovetail slot in a respective disk and wherein at least one axial cooling air slot is formed by a gap between a bottom surface of the blade shank and the dovetail slot, and further wherein cooling air flowing through said axial passages in said cooling plate enters an annular cavity formed between said cooling plate and said peripheral rim of said last stage, and then through said cooling air axial slots and into said annular cooling cavities between said adjacent ones of said rotor disks.

20. The improvement of claim 18 wherein the peripheral rims of at least the rotor disks of the last stage and a next-to-last stage are constructed with passageways to permit some of the cooling air in the respective annular cooling cavities to escape into said compressor discharge air.

21. The improvement of claim 20 wherein said passageways include a pair of radial portions separated by an axially oriented, annular buffer cavity.

22. The improvement of claim 18 wherein said slots in said rotor disks have angled surfaces at entrances to said annular cavities to facilitate swirling of said cooling air within said cavities.

23. In a gas turbine comprising a compressor including a rotor and a plurality of bladed disks and associated rims extending from a forward end to an aft end of the compressor;

a plurality of combustors, and a casing surrounding said combustors arranged to direct air from the compressor to the combustors; and

a turbine including multiple stages driven by combustion gases from the plurality of combustors, and wherein cooling air is utilized to cool the compressor rotor, the improvement comprising;

at least one extraction port in said casing and associated piping for supplying compressor discharge air as cooling air to at least some of said plurality of bladed disks and associated rims at the aft end of the compressor.

24. The gas turbine of claim 23 wherein said associated piping communicates with passageways in said compressor for supplying said cooling air to said compressor rotor and to said at least some of said bladed disks and associated rims.

25. The gas turbine of claim 24 wherein a heat exchanger is located within said piping between said extraction port and said compressor rotor to cool said compressor discharge air before reaching said rotor.

26. The gas turbine of claim 25 and including a flow control valve between said heat exchanger and said rotor.

27. The gas turbine of claim 26 including a compressor part divided into an upper part and a lower part by an annular compressor discharge seal, and wherein said seal incorporates a discourager segment which allows some of said cooling air to flow into said lower part.

28. The gas turbine of claim 27 wherein said compressor cavity is defined in part by an annular cooling plate, and wherein a part of said cooling air for cooling a last of the associated rims in the aft end of the compressor flows through a plurality of axial passages in said annular cooling plate from said upper part.

29. The gas turbine of claim 28 wherein said cooling plate includes at least one radial passage extending between said lower part and a compressor bore, and wherein another part of said cooling air for cooling said bore flows through said at least one radial passage.

30. The gas turbine of claim 29 wherein said cooling plate includes a radially outer tip including means for permitting a portion of said cooling air to flow from said upper part into the hot compressor discharge air and thereby prevent said hot air from entering said upper part.

31. The gas turbine of claim 27 and including temperature indicator means for monitoring air temperature in said upper part of said cavity.

32. In a land-based gas turbine comprising a compressor including a rotor and a plurality of bladed disks and associated rims extending from a forward end to an aft end of the compressor;

a plurality of combustors surrounded by a casing arranged to receive combustion air from the compressor; and

a turbine including multiple stages driven by combustion gases from the plurality of combustors and wherein cooling air is utilized to cool the compressor rotor, a method of maintaining at least the last of said disks and its associated rim at a predetermined temperature comprising the steps of;

a) extracting cooling air from said casing;

b) introducing said cooling air into said compressor in a direction counter to a flow direction of the combustion air; and

c) cooling said last disk and its associated rim to a predetermined temperature.

33. The method of claim 32 wherein said predetermined temperature is about 825°

F.

34. The method of claim 32 wherein each of said plurality of bladed disks includes a plurality of blades having corresponding blade shanks seated within dovetail slots provided in a corresponding one of said disks, and wherein at least a portion of said cooling air is flowed axially toward a forward end of said compressor by means of axial slots formed between bottom surfaces of said blade shanks and bottom surfaces of said dovetail slots.

35. The method of claim 32 wherein step c) includes flowing at least a portion of said cooling air radially into an annular cooling cavity located between rims of said last disk and a disk adjacent said last disk.

36. The method of claim 35 wherein at least a portion of said cooling air is permitted to escape into the combustion air by flowing radially between said last disk and an adjacent disk to thereby further cool the rim associated with said last disk.

37. The method of claim 36 wherein said portion of said cooling air permitted to escape radially between said last disk and an adjacent disk upstream of said last disk passes through a buffer cavity which prevents the combustion air from flowing radially inwardly into said annular cooling cavity.