Rotor tip bleed in gas turbine engines

• US 6,574,965 B1
• Filed: 12/23/1998
• Issued: 06/10/2003
• Est. Priority Date: 12/23/1998
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Feulner, Matthew R.
• Primary Examiner(s): FREAY, CHARLES GRANT
• Application Number:
  US09/220,468
• Time in Patent Office: 1,630 Days
• Field of Search: 607/85, 602/62, 608/05, 60/397.5, 415/144, 415/168.1, 415/168.2, 415/168.4, 415/169.1, 415/914, 415/173.1, 415/57.1, 415/57.3, 415/57.4, 415/58.4
• US Class Current: 60/785
• CPC Class Codes: F01D 11/10 using sealing fluid, e.g. steam
• CPC Class Codes: F04D 27/023 Details or means for fluid extraction
• CPC Class Codes: F04D 29/526 Details of the casing section radially opposing blade tips ducts F04D29/545
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine includes a compressor, a combustor, and turbine enclosed in an engine case with the compressor having a plurality of alternating rows of rotating blades and stationary vanes. The gas turbine engine also includes a substantially circumferential groove formed in the engine case of the compressor section. The groove is substantially adjacent to a row of rotating blades to extract a portion of tip leakage flow from that row of blades and to route the extracted tip leakage flow to the turbine section for cooling turbine components. In the preferred embodiment of the present invention, the groove communicates with a plenum which communicates with the turbine section of the gas turbine engine via channeling pipes formed within the engine case.

Claims

1. A gas turbine engine having a compressor section, a combustor section, and a turbine section enclosed in an engine case with an annular flow path extending axially therethrough, said compressor section having a plurality of rows of rotating blades alternating with a plurality of rows of stationary vanes, said gas turbine engine comprising:

a groove formed in said engine case of said compressor section and extending substantially circumferentially therethrough, said groove being substantially adjacent to one row of said plurality of rows of rotating blades, said groove being in fluid communication with said turbine section for channeling cooling air from said compressor section to said turbine section to increase stability of said gas turbine engine.

2. The gas turbine engine according to claim 1 further comprising:

a plenum formed in said engine case and being in communication with said groove and said turbine section.

3. The gas turbine engine according to claim 2 further comprising:

a channeling pipe formed in said engine case and being in communication with said plenum and routing said cooling air from said plenum to said turbine section of said gas turbine engine.

4. The gas turbine engine according to claim 1 further comprising:

a channeling pipe formed in said engine case and being in communication with said groove and routing said cooling air from said groove to said turbine section of said gas turbine engine.

5. The gas turbine engine according to claim 4 further comprising:

a valve cooperating with said channeling pipe to modulate amount of cooling air being diverted to said turbine section of said gas turbine engine.

6. The gas turbine engine according to claim 1 further comprising:

a second groove formed in said engine case of said compressor section and extending substantially circumferentially therethrough, said second groove being substantially adjacent to another row of said plurality of rows of rotating blades, said second groove being in fluid communication with said turbine section for channeling cooling air from said compressor section to said turbine section.

7. The gas turbine engine according to claim 1 wherein said groove is substantially annular.

8. The gas turbine engine according to claim 1 wherein said groove is circumferentially interrupted.

9. The gas turbine engine according to claim 1 wherein said groove includes a plurality of circumferentially interrupted grooves.

10. The gas turbine engine according to claim 1 wherein said groove is disposed between a leading edge and a mid-cord section of said rotating blades.

11. A gas turbine engine having a compressor section, a combustor section, and a turbine section with an annular flow path extending axially therethrough, said compressor section having a plurality of rows of rotating blades alternating with a plurality of rows of stationary vanes, each said row of rotating blades includes a plurality of blades, each said blade having a blade tip, said gas turbine engine comprising:

an outer shroud enclosing said compressor section, said outer shroud defining a tip clearance between said blade tips of said rotating blades and said outer shroud, said outer shroud defining a plurality of grooves, each of said plurality of grooves extending substantially circumferentially through said outer shroud, each of said plurality of grooves being substantially adjacent to one respective row of said plurality of rows of rotating blades, each of said plurality of grooves being in fluid communication with said turbine section for extracting air from said tip clearance of said respective row of rotating blades of said compressor section to be routed to said turbine section to increase stability of said gas turbine engine.

12. The gas turbine engine according to claim 11, further comprising:

a plenum formed in said outer shroud and being in communication with each of said plurality of grooves and said turbine section.

13. The gas turbine engine according to claim 12 further comprising:

a channeling pipe formed in said outer shroud and being in communication with said plenum and routing said cooling air from said plenum to said turbine section of said gas turbine engine.

14. The gas turbine engine according to claim 11 further comprising:

a plurality of channeling pipes formed in said outer shroud and being in communication with each of said plurality of grooves and routing said cooling air from each of said plurality of grooves to said turbine section of said gas turbine engine.

15. The gas turbine engine according to claim 11 further comprising

a plurality of plenums formed in said outer shroud and being in with said plurality of grooves and said turbine section.

16. A gas turbine engine having a compressor section, a combustor section, and a turbine section with an annular flow path extending axially therethrough, said compressor section having a plurality of rows of rotating blades alternating with a plurality of rows of stationary vanes, said gas turbine engine comprising:

an engine case enclosing said compressor section and defining a groove, said groove extending substantially circumferentially through said engine case, said groove being disposed between a leading edge and a mid-cord section of one row of said plurality of rows of rotating blades, said groove being in fluid communication with said turbine section for channeling cooling air from said compressor section to said turbine section to reduce potential for an engine surge, thereby increasing stability of said gas turbine engine.