Active turbine or compressor tip clearance control

Patent metadata

Field	Value
Publication number	US-9598974-B2
Application number	US-201313775806-A
Country	US
Kind code	B2
Filing date	Feb 25, 2013
Priority date	Feb 25, 2013
Publication date	Mar 21, 2017
Grant date	Mar 21, 2017

Abstract

A gas turbine engine includes an annular plenum defined with an outer skin and a perforated inner skin for receiving selective air flow to impinge a support case which supports shrouds of the rotor assemblies of the engine therein for active tip clearance control of the rotor assemblies. In one embodiment a bobbin-type transfer tube for supplying cooling air into the plenum, is provided between an outer case of the engine an the plenum such that the thermally induced relative movement of the outer case and the plenum is permitted.

First claim

The invention claimed is:

1. A gas turbine engine comprising a rotor encircled by an annular shroud supported within an annular support case, an annular plenum provided within an annular outer case of the engine and defined radially between an outer skin and a perforated inner skin of the annular plenum, an annular cavity surrounding the outer skin of the annular plenum and defined within the outer case, the annular plenum being placed around and supported on the annular support case, a first transfer tube extending radially through and fluidly isolated from the annular cavity having a radial inner end sealingly received in a first port in the outer skin and a radial outer end sealingly received in a second port in the outer case, the second port in the outer case being in fluid communication with a source of pressurized cooling air to allow a first flow of pressurized cooling air to pass through the first transfer tube and then enter the plenum in order to pass through the perforated inner skin for impingement cooling of the support case, wherein the respective radial inner and outer ends of the first transfer tube each comprise a portion of a spherical surface in contact with a cylindrical inner surface of the respective first and second ports to allow thermally induced axial, radial and circumferential relative movement of the outer case and the plenum, a second transfer tube extending radially through the annular plenum to form at least a part of an air passage isolated from the plenum, the air passage communicating with said annular cavity surrounding the outer skin of the annular plenum, and with a space defined within the annular support case for continuously introducing a second flow of pressurized cooling air from said cavity into the annular support case to cool the annular shroud.

2. The gas turbine engine as defined in claim 1 wherein the second port in the outer case is connected through a control valve to the source of the pressurized cooling air for selectively cooling the support case.

3. The gas turbine engine as defined in claim 1 wherein the first flow of pressurized cooling air from the first source is cooler than the second flow of pressurized cooling air contained in the annular cavity.

4. The gas turbine engine as defined in claim 1 wherein the second transfer tube extends through the inner skin and comprises a radial inner end sealingly received in a third port in the annular support case and a radial outer end sealingly received in a fourth port in the outer skin, wherein the respective radial inner and outer ends of the second transfer tube each comprise a portion of a spherical surface in contact with a cylindrical inner surface of the respective third and fourth ports to allow thermally induced axial, radial and circumferential relative movement of the plenum and the support case.

5. The gas turbine engine as defined in claim 1 wherein the cylindrical inner surface of the first port extends radially inwardly and terminates at an annular shoulder of the first port, the annular shoulder extending inwardly from the cylindrical inner surface to support the radially inner end of the first transfer tube and to define an opening of the first port in fluid communication with the first transfer tube and the annular plenum.

6. The gas turbine engine as defined in claim 5 wherein the cylindrical inner surface of the second port extends radially through a body of the second port to allow the first transfer tube to be inserted radially and inwardly through the second port until the radial inner end of the first transfer tube rests on the annular shoulder of the first port.

7. The gas turbine engine as defined in claim 4 wherein the third port is formed as a counterbore defined in the annular support case, an enlarged portion of the counterbore defining the cylindrical inner surface of the third port in contact with the portion of the spherical surface of the radial inner end of the second transfer tube.

8. The gas turbine engine as defined in claim 7 wherein the cylindrical inner surface of the fourth port extends radially through a body of the fourth port to allow the second transfer tube to be inserted radially and inwardly through the fourth port until the radial inner end of the second transfer tube rests in the counterbore of the third port.

9. A gas turbine engine comprising a first turbine stage assembly, a second turbine stage assembly and a stator vane ring assembly disposed axially between the first and second turbine stage assemblies, the first turbine stage assembly including a first turbine rotor encircled by an annular first turbine shroud, the second turbine stage assembly including a second turbine rotor encircled by an annular second turbine shroud, the stator vane ring assembly defining a hot gas path between the first and second turbine stage assemblies, an annular support case positioned around and supporting the first turbine stage assembly, the stator vane ring assembly and the second turbine stage assembly, a perforated impingement skin being positioned within the annular support case, adjacent and radially spaced apart from a radial outer surface of the respective first, second pressure turbine shrouds and stator vane ring assembly, an annular plenum defined radially between an outer skin and a perforated inner skin of the annular plenum, the annular plenum being provided within an annular outer case of the engine and surrounding and being supported on the annular support case, a first transfer tube radially extending through and being fluidly isolated from an annular cavity surrounding the annular plenum and defined in the outer case, the first transfer tube being connected at a radial inner end to the outer skin and being in communication with the annular plenum, the first transfer tube being connected at a radial outer end to the annular outer case and being connected to a source of pressurized cooling air through a control valve to thereby form an active tip clearance control system to selectively conduct a first flow of pressurized cooling air from the first source to enter the annular plenum and to pass through the perforated inner skin in order to impinge on the support case, a plurality of second transfer tubes extending radially through the annular plenum to form at least part of air passages isolated from the annular plenum, said air passages being in fluid communication with the annular cavity and a space in the annular support case in order to introduce a second flow of pressurized cooling air from the annular cavity into the annular support case and to then pass through the respective perforated impingement skins to continuously cool the first and second turbine shrouds and the stator vane ring assembly during engine operation, wherein the respective radial inner and outer ends of the first transfer tube each comprise a portion of a spherical surface in contact with a cylindrical inner surface of the respective first and second ports to allow thermally induced axial, radial and circumferential relative movement of the outer case and the plenum.

10. The gas turbine engine as defined in claim 9 wherein each of the second transfer tubes extends through the inner skin and comprises a radial inner end sealingly received in one of a plurality of third ports in the annular support case and a radial outer end sealingly received in one of a plurality of fourth ports in the outer skin.

11. The gas turbine engine as defined in claim 10 wherein the respective radial inner and outer ends of the second transfer tubes each comprise a portion of a spherical surface in contact with a cylindrical inner surface of the respective third and fourth ports to allow thermally induced axial, rad

Assignees

• Pratt & Whitney Canada

Inventors

• Gekht Eugene
• Lucas Terrence
• Haslam-Jones Thomas Francis
• Mills Danny
• Bouchard Guy

Classifications

• F01D25/12

  Cooling · CPC title

• F01D11/24 Primary

  by selectively cooling-heating stator or rotor components · CPC title

• F05D2250/711

  convex · CPC title

• F05D2220/321

  for a special turbine stage · CPC title

• F05D2250/241

  spherical · CPC title