Vortex spoiler for delivery of cooling airflow in a turbine engine

• US 7,708,519 B2
• Filed: 03/26/2007
• Issued: 05/04/2010
• Est. Priority Date: 03/26/2007
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Mignano, Frank A.
• Primary Examiner(s): Kershteyn; Igor
• Application Number:
  US11/691,307
• Publication Number:
  US 20090067986A1
• Time in Patent Office: 1,135 Days
• Field of Search: 415/115, 415/157, 415/171, 415/202, 415/208.2
• US Class Current: 415/157
• 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 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence
• CPC Class Codes: F04D 29/321 for axial flow compressors
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A vortex spoiler (180) for delivery of a cooling airflow (192) in a turbine (108) engine (100) including a plurality of inlet ports (182) formed circumferentially about a radial exterior sidewall (188), and a plurality of outlet ports (184) formed circumferentially about a radial interior sidewall (190). The plurality of inlet ports (182) are coupled to the plurality of outlet ports (184) via a plurality of ducts (186). Each of the ducts is formed having an interior diameter at the inlet port and the outlet port formed at a preselected angle normal to the surface of the each of the radial sidewalls to form a radially curved profile such that a cooling airflow (192) may pass radially inwardly through each of the plurality of ducts (186) with minimal tangential stress and minimal static pressure loss.

Claims

1. A vortex spoiler for delivery of a cooling airflow in a turbine engine comprising:

a radial exterior sidewall positioned to rotate about an axis of rotation;

a radial interior sidewall disposed radially inward of, and surrounded by, the radial exterior sidewall;

a core area coupled between the radial exterior sidewall and the radial interior sidewall, the core area extending inwardly perpendicularly to the axis of rotation, the radial exterior sidewall, the radial interior sidewall and the core area defining a discrete vortex spoiler;

a plurality of inlet ports formed in and spaced circumferentially about the radial exterior sidewall;

a plurality of outlet ports formed in and spaced circumferentially about the radial interior sidewall, the plurality of outlet ports configured to provide an outlet for the cooling airflow from the vortex spoiler; and

a plurality of bores extending through the core area coupling each of the plurality of outlet ports to one of the plurality of inlet ports and defining a plurality of ducts that fluidly communicate the inlet ports and the outlet ports, each duct configured such that a cooling airflow may pass radially inwardly with minimal tangential stress and minimal static pressure loss.

2. A vortex spoiler as claimed in claim 1, wherein each of the plurality of bores extends radially inward in a curved rotated profile from one of the plurality of inlet ports to one of the plurality of outlet ports.

3. A vortex spoiler as claimed in claim 1, wherein the plurality of ducts defined by the plurality of bores extends radially from the radial exterior sidewall, through the core area, to the radial interior sidewall.

4. A vortex spoiler as claimed in claim 1, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of inlet ports formed at a preselected angle in a range of 65-75 degrees normal to a surface of the radial exterior sidewall.

5. A vortex spoiler as claimed in claim 4, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of inlet ports formed at a preselected angle of 70 degrees normal to the surface of the radial exterior sidewall.

6. A vortex spoiler as claimed in claim 1, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of outlet ports formed at a preselected angle in a range of 15-25 degrees normal to surface of the radial interior sidewall.

7. A vortex spoiler as claimed in claim 6, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of outlet ports formed at a preselected angle of 20 degrees normal to a surface of the radial interior sidewall.

8. A gas turbine engine having a compressor section that includes a plurality of rotary compressor stages interconnected by a rotary shaft and a secondary cooling airflow system comprising:

a collecting chamber arranged to collect pressurized, substantially nonswirling, cooling airflow from the compressor section; and

a vortex spoiler in fluidic communication with the collecting chamber and interconnected to rotate with the rotary shaft, the vortex spoiler comprising;

a radial exterior sidewall;

a radial interior sidewall disposed radially inward of, and surrounded by, the radial exterior sidewall;

a core area coupled between the radial exterior sidewall and the radial interior sidewall and extending inwardly perpendicularly to the axis of rotation, the radial exterior sidewall, the radial interior sidewall and the core area defining a discrete vortex spoiler component;

a plurality of inlet ports formed in and spaced circumferentially about the radial exterior sidewall

a plurality of outlet ports formed in and spaced circumferentially about the radial interior sidewall, the plurality of outlet ports configured to provide an outlet for the cooling airflow from the vortex spoiler; and

a plurality of bores formed through the core area coupling each of the plurality of outlets ports to one of the plurality of inlet ports, the plurality of bores defining a plurality of ducts that fluidly communicate the inlet ports and the outlet ports, each duct configured such that a cooling airflow may pass radially inwardly with minimal tangential stress and minimal static pressure loss.

9. A gas turbine engine as claimed in claim 8, wherein each of the plurality of bores extends radially inward in a curved rotated profile from one of the plurality of inlet ports to one of the plurality of outlet ports.

10. A gas turbine engine as claimed in claim 8, wherein each of the plurality of bores extends radially from the radial exterior sidewall, through the core area, to the radial interior sidewall.

11. A gas turbine engine as claimed in claim 8, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of inlet ports formed at a preselected angle in a range of 65-75 degrees normal to a surface of the radial exterior sidewall.

12. A gas turbine engine as claimed in claim 11, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of inlet ports formed at a preselected angle of 70 degrees normal to the surface of the radial exterior sidewall.

13. A gas turbine engine as claimed in claim 8, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of outlet ports formed at a preselected angle in a range of 15-25 degrees normal to surface of the radial interior sidewall.

14. A gas turbine engine as claimed in claim 13, wherein each of the plurality of ducts is formed having an interior diameter at each of the plurality of outlet ports formed at a preselected angle of 20 degrees normal to a surface of the radial interior sidewall.

15. A gas turbine engine, comprising:

a compressor section including an intermediate pressure compressor and a high pressure compressor;

a rotary shaft interconnecting the intermediate pressure compressor and the high pressure compressor in torque transmitting relationship about an axis of rotation;

a collecting chamber arranged to collect pressurized, substantially nonswirling, cooling airflow from the compressor section at a location radially outward of the rotary shaft; and

a discrete vortex spoiler interconnected to rotate with the rotary shaft and in fluidic communication with the collecting chamber, the discrete vortex spoiler defined by a radial exterior sidewall and a radial interior sidewall disposed radially inward of, and surrounded by the radial exterior sidewall; and

a core area coupled between the radial exterior sidewall and the radial interior sidewall, the core area extending inwardly perpendicularly to the axis of rotation, the discrete vortex spoiler including a plurality of inlet ports formed in and spaced circumferentially about the radial exterior sidewall and a plurality of outlet ports formed in and spaced circumferentially about the radial interior sidewall, each of the plurality of inlet ports coupled to one of the plurality of outlet ports via a bore, and defining a plurality of ducts that fluidly communicate the inlet ports and the outlet ports, each duct configured having an interior diameter at each of the plurality of inlet ports formed at a preselected angle in a range of approximately 65-75 degrees normal to a surface of the radial exterior sidewall and at each of the plurality of outlet ports having an interior diameter formed at a preselected angle in a range of approximately 15-25 degrees normal to a surface of the radial interior sidewall such that a cooling airflow passes radially inwardly through each of the plurality of ducts with minimal tangential stress and minimal static pressure loss.

16. A gas turbine engine as claimed in claim 15, wherein the compressor section further includes a diffuser vane positioned between the intermediate pressure compressor and the high pressure compressor, the collecting chamber positioned to receive a cooling airflow from the intermediate pressure compressor as it discharges from the diffuser vane.

17. A gas turbine engine as claimed in claim 15, wherein the interior diameter of each of the plurality of ducts at each of the plurality of inlet ports is formed at a preselected angle of 70 degrees normal to a surface of the radial exterior sidewall.

18. A gas turbine engine as claimed in claim 17, wherein the diameter of each of the plurality of ducts at each of the plurality of outlet ports is formed at a preselected angle of 20 degrees normal to a surface of the radial interior sidewall.

19. A gas turbine engine as claimed in claim 15, wherein the diameter of each of the plurality of ducts at each of the plurality of outlet ports is formed at a preselected angle of 20 degrees normal to a surface of the radial interior sidewall.