COOLING AIR SYSTEM FOR MID TURBINE FRAME

• US 20110079019A1
• Filed: 10/01/2009
• Published: 04/07/2011
• Est. Priority Date: 10/01/2009
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: LEGARE, Pierre-Yves, DUROCHER, Eric
• Granted Patent: US 8,371,127 B2
• Time in Patent Office: Days
• US Class Current: 60/796
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 11/02 by non-contact sealings, e.g. of labyrinth type for sealing space between rotor blade tips and stator F01D11/08
• CPC Class Codes: F01D 25/125 of bearings
• CPC Class Codes: F01D 25/162 Bearing supports
• CPC Class Codes: F01D 5/082 on the side of the rotor disc
• CPC Class Codes: F01D 9/065 Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids see also F01D25/16, F01D25/24 and F01D25/26
• 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: F02C 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence

Abstract

A mid turbine frame is disposed between high and low pressure turbine assemblies. A cooling air system defined in the mid turbine frame of a gas turbine engine includes internal cavities for containing pressurized cooling air to cool the inter-turbine duct and the hollow struts, and discharges the used cooling air to further cool respective high and low pressure turbine assemblies.

Claims

1. A gas turbine engine comprising:

a first turbine rotor assembly and a second turbine rotor assembly axially spaced apart from each other;

a mid turbine frame (MTF) disposed axially between the first and second turbine rotor assemblies, including an annular outer case, annular inner case and annular bearing housing with bearing seals, the bearing housing being connected to the inner case, a plurality of load spokes radially extending between and interconnecting the outer and inner cases to transfer loads from the bearing housing to the outer case;

an annular inter-turbine duct (ITD) disposed radially between the outer and inner cases of the MTF, the ITD including an annular outer duct wall and annular inner duct wall, thereby defining an annular hot gas path between the outer and inner duct walls for directing hot gases from the first turbine rotor assembly to the second turbine rotor assembly, a plurality of hollow struts radially extending between and interconnecting the outer and inner duct walls, the load transfer spokes radially extending through at least a number of the hollow struts; and

wherein the MTF defines a cooling air system, the system being formed with a first cavity between the outer case and the outer duct walls of the ITD with a first inlet defined in the outer case, a second cavity between the inner duct wall and the inner case, the first cavity, second cavity and the respective hollow struts being in fluid communication with the first inlet for receiving pressurized cooling air, the cooling air system including a cooling air discharge device at respective upstream and downstream sides of the MTF for discharging cooling air from the system to further cool the respective first and second rotor assemblies, and a flow restrictor configured for metering a cooling air flow escaping from the second cavity in order to provide a pressure margin the cavities in the MTF and the hot gas path to impede hot gas ingestion into the cavities of the MTF.

2. The gas turbine engine as defined in claim 1 wherein the cooling air system further comprises a third cavity between the inner case and the bearing housing, and a cooling air passage in the bearing housing, the third cavity and the cooling air passage in the bearing housing being in fluid communication with a second inlet defined in a rotating shaft of the engine for receiving pressurized cooling air.

3. The gas turbine engine as defined in claim 2 wherein the flow restrictor defines an annular gap of a predetermined size, in fluid communication with the second and third cavities to allow cooling air to escape from the second cavity while maintaining the first cavity, second cavity and the respective hollow struts pressurized with the cooling air.

4. The gas turbine engine as defined in claim 3 wherein the cooling air discharge device at the upstream side of the MTF defines a flow restricting/sealing arrangement for splitting the cooling air escaping from the flow restrictor, into a first air flow to be discharged from the cooling air system to cool the first turbine rotor assembly and a second air flow to be directed into the third cavity.

5. The gas turbine engine as defined in claim 4 wherein the flow restricting/sealing arrangement is formed in cooperation with the first turbine rotor assembly and the MTF, to allow the first air flow discharged from the cooling air system to pressurize a blade rim seal and to cool a blade back cavity of the first turbine rotor assembly.

6. The gas turbine engine as defined in claim 4 wherein the flow restricting/sealing arrangement comprises at least one annular, axial surface defined in each of the inner duct wall and the inner case, in cooperation with a seal component defined on an annular rear plate mounted on a rotor disc of the first turbine rotor assembly.

7. The gas turbine engine as defined in claim 4 wherein the second air flow directed from the flow restricting/sealing arrangement is mixed with a second inlet cooling air flow introduced from the second inlet, to form a mixed cooling air flow.

8. The gas turbine engine as defined in claim 7 wherein the cooling air discharge device at the downstream side of the MTF is configured to discharge the mixed cooling air flow from the third cavity to supply cooling air for the second turbine rotor assembly.

9. The gas turbine engine as defined in claim 2 wherein the cooling air passage in the bearing housing comprises a radial section to direct a second inlet cooling air flow introduced from the second inlet, to pass between a rear seal of a first bearing of the first turbine rotor assembly and a front seal of a second bearing of the second turbine rotor assembly, thereby pressurizing the respective seals.

10. The gas turbine engine as defined in claim 9 wherein the cooling air passage in the bearing housing comprises a forward section and a rearward section to split the cooling air flow having passed through the radial section, into a forward cooling air flow for pressurizing a front seal of the first bearing of the first turbine rotor assembly and a rearward cooling air flow for pressurizing a rear seal of the second bearing of the second turbine rotor assembly.

11. The gas turbine engine as defined in claim 2 wherein the inner case comprises an arch structure defining a chamber for receiving a portion of a first inlet cooling air flow which has passed through the first cavity, the hollow struts and the second cavity and a portion of a second inlet cooling air flow which has passed through the cooling air passage in the bearing housing, the portions of first and second inlet cooling air flows being mixed in the chamber and discharged through the third cavity for supplying cooling air to the second turbine rotor assembly.

12. The gas turbine engine as defined in claim 1 wherein the cooling air discharge device at the upstream side of the MTF is located at a front axial end of the outer duct wall for supplying cooling air to a turbine shroud of the first turbine rotor assembly.

13. The gas turbine engine as defined in claim 1 wherein the cooling air discharge device at the upstream side of the MTF is located at a front axial end of the inner duct wall for supplying cooling air to pressurize a blade rim seal and to cool a back cavity of the first turbine rotor assembly.

14. The gas turbine engine as defined in claim 1 wherein the cooling air discharge device at the downstream side of the MTF is located at a rear axial end of the outer duct wall for supplying cooling air to a turbine shroud of the second turbine rotor assembly.

15. The gas turbine engine as defined in claim 1 wherein the cooling air discharge device at the downstream side of the MTF is located at a rear axial end of the inner duct wall for supplying cooling air to a rotor of the second turbine rotor assembly.

16. A gas turbine engine comprising:

a first turbine rotor assembly and a second turbine rotor assembly axially spaced apart from each other;

a cooling air system including a first inlet defined in the mid turbine frame (MTF) and a second inlet defined in a rotating shaft of the engine for receiving pressurized cooling air from separate passages, the cooling air system having a first cavity between an outer case and an outer duct wall of an inter-turbine duct (ITD) disposed inside the outer case, a second cavity between an inner duct wall of the ITD and an inner case, a third cavity between the inner case and a bearing housing mounted to the inner case, a cooling air passage in the bearing housing, and a chamber defined by an arch structure integrated with the inner case and communicating with the first inlet through a first flow path, the first flow path extending from the first cavity through a hollow passage in the ITD to the second cavity, the chamber communicating with the second inlet through a second flow path, the second flow path extending through the cooling air passage in the bearing housing, the chamber also communicating with the third cavity, and at least one of the cavities including a cooling air discharge device located at respective upstream and downstream sides of the MTF for discharging cooling air from the system to the respective first and second rotor assemblies.

17. The gas turbine engine as defined in claim 16 wherein the cooling air passage in the bearing housing comprises a radial section, a forward section and a rearward section, the radial section directing the second inlet cooling air flow to pass between a rear seal of a first bearing of the first turbine rotor assembly and a front seal of a second bearing of the second turbine rotor assembly, thereby pressurizing the respective seals, the forward section and the rearward section splitting the second inlet cooling air flow which has passed through the radial section, into a forward cooling air flow for pressurizing a front seal of the first bearing of the first turbine rotor assembly and a rearward cooling air flow for pressurizing a rear seal of the second bearing of the second turbine rotor assembly.

18. The gas turbine engine as defined in claim 16 wherein the cooling air discharge device at the upstream side of the MTF is located at a front axial end of the inner duct wall for supplying cooling air to pressurize a blade rim seal and to cool a back cavity of the first turbine rotor assembly.

19. The gas turbine engine as defined in claim 16 wherein the cooling air discharge device at the downstream side of the MTF is located at a rear axial end of the outer duct wall for supplying cooling air to a turbine shroud of the second turbine rotor assembly.

20. The gas turbine engine as defined in claim 16 wherein the mixed portions of the first inlet cooling air flow and the second inlet cooling air flow are discharged from the third cavity for supplying cooling air to a rotor of the second turbine rotor assembly.