Interdigitated turbine engine air bearing and method of operation

Patent metadata

Field	Value
Publication number	US-10718265-B2
Application number	US-201715605164-A
Country	US
Kind code	B2
Filing date	May 25, 2017
Priority date	May 25, 2017
Publication date	Jul 21, 2020
Grant date	Jul 21, 2020

Abstract

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The first turbine bearing defines an outer air bearing along an outer diameter of the first turbine bearing and adjacent to the hub of the low speed turbine rotor. The first turbine bearing defines an inner air bearing along an inner diameter of the first turbine bearing and adjacent to the HP shaft.

First claim

What is claimed is:

1. A gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction, the gas turbine engine comprising:
a combustion section;
a compressor section, wherein one or both of the combustion section or the compressor section forms a pressure plenum;
a first turbine rotor comprising a hub extended along the longitudinal direction;
a second turbine rotor comprising a shaft coupling the second turbine rotor to a compressor rotor of the compressor section, wherein the shaft is extended along the longitudinal direction and is radially inward of the hub of the first turbine rotor;
a turbine bearing assembly positioned between the hub of the first turbine rotor and the shaft, wherein the turbine bearing assembly comprises an outer air bearing at an outer diameter of the turbine bearing assembly, and wherein the outer air bearing is adjacent to the hub of the first turbine rotor, and wherein the turbine bearing assembly comprises an inner air bearing at an inner diameter of the turbine bearing assembly, and wherein the inner air bearing is adjacent to the shaft; and
a first manifold extended from the pressure plenum to the turbine bearing assembly, wherein the first manifold provides fluid communication between the pressure plenum and the turbine bearing assembly.

2. The gas turbine engine of claim 1 , wherein the turbine bearing assembly comprises an outer annular wall at the outer diameter of the turbine bearing assembly, and wherein the turbine bearing assembly comprises an inner annular wall at the inner diameter of the turbine bearing assembly, and wherein each of the outer annular wall and the inner annular wall are extended at least partially along the longitudinal direction.

3. The gas turbine engine of claim 1 , wherein the turbine bearing assembly comprises an outer plurality of orifices arranged at least along the longitudinal direction and circumferentially along the outer air bearing, and wherein the outer plurality of orifices is adjacent to the hub of the first turbine rotor.

4. The gas turbine engine of claim 1 , wherein the turbine bearing assembly comprises an inner plurality of orifices arranged at least along the longitudinal direction and circumferentially along the inner air bearing, and wherein the inner plurality of orifices is adjacent to the second turbine rotor.

5. The gas turbine engine of claim 1 , wherein the combustion section comprises an inner diffuser frame, and wherein the turbine bearing assembly is coupled to the inner diffuser frame, and wherein the turbine bearing assembly is cantilevered from the inner diffuser frame.

6. The gas turbine engine of claim 1 , wherein the turbine bearing assembly defines an at least partially annular groove at the outer diameter of the turbine bearing assembly, and wherein the groove is radially adjacent to the hub of the first turbine rotor.

7. The gas turbine engine of claim 1 , wherein the gas turbine engine comprises a compressor seal interface between the compressor rotor and the turbine bearing assembly, and wherein the gas turbine engine forms a first cavity between the compressor seal interface and the turbine bearing assembly, wherein the first cavity is extended at least partially circumferentially around the axial centerline of the gas turbine engine.

8. The gas turbine engine of claim 1 , wherein the gas turbine engine comprises a turbine seal interface between the second turbine rotor and the turbine bearing assembly, and wherein the gas turbine engine forms a cavity between the second turbine rotor and the turbine bearing assembly, wherein the cavity is extended at least partially circumferentially around the axial centerline of the gas turbine engine.

9. The gas turbine engine of claim 1 , the gas turbine engine comprising:
a pressure regulating valve; and
a second manifold extended from the turbine bearing assembly, wherein the second manifold provides fluid communication with the pressure regulating valve and a first cavity formed between a compressor seal interface and the turbine bearing assembly.

10. The gas turbine engine of claim 9 , wherein the outer air bearing of the turbine bearing assembly comprises a groove, and wherein the groove forms an annular cavity.

11. The gas turbine engine of claim 10 , wherein the second manifold is in fluid communication with the annular cavity at the outer air bearing, the first cavity formed at the compressor seal interface, and a second cavity formed at a turbine seal interface, and wherein the second manifold provides a flow and pressure of air therethrough from the annular cavity, the first cavity, and the second cavity.

12. The gas turbine engine of claim 9 , the gas turbine engine comprising: a third manifold extended from the pressure plenum to the pressure regulating valve, wherein the third manifold provides fluid communication of the pressure plenum and the pressure regulating valve.

13. The gas turbine engine of claim 1 , wherein the first turbine rotor and the second turbine rotor are together in alternating interdigitated arrangement along the longitudinal direction.

14. The gas turbine engine of claim 1 , wherein the first turbine rotor is interdigitated with the second turbine rotor, and wherein the first turbine rotor is operable at a first maximum rotational speed less than the second turbine rotor operable at a second maximum rotational speed, and wherein the first turbine rotor comprises at least one stage upstream of the second turbine rotor.

15. The gas turbine engine of claim 14 , wherein the first turbine rotor is immediately downstream of the combustion section.

16. The gas turbine engine of claim 1 , wherein the first manifold is connected to the turbine bearing assembly, and wherein the first manifold provides a flow of air from the pressure plenum to the turbine bearing assembly.

17. The gas turbine engine of claim 16 , the gas turbine engine comprising: a second manifold extended from the turbine bearing assembly, wherein the second manifold provides fluid communication with a pressure sink and a first cavity formed between a compressor seal interface and the turbine bearing assembly, wherein the pressure sink comprises a pressure less than the first cavity.

18. The gas turbine engine of claim 16 , wherein the turbine bearing assembly comprises an outer plurality of orifices at the outer air bearing of the turbine bearing assembly, wherein the outer plurality of orifices provides fluid communication of the flow of air from the first manifold to a first volume between the hub of the first turbine rotor and the turbine bearing assembly.

19. The gas turbine engine of claim 18 , wherein the turbine bearing assembly comprises an inner plurality of orifices at the inner air bearing of the turbine bearing assembly, wherein the inner plurality of orifices provides fluid communication of the flow of air from the first manifold to a second volume between the shaft of the second turbine rotor and the turbine bearing assembly.

Assignees

• Gen Electric

Inventors

• Stuart Alan Roy
• Clements Jeffrey Donald
• Schmidt Richard
• Moniz Thomas Ory

Classifications

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F01D25/22 Primary

  using working-fluid or other gaseous fluid as lubricant · CPC title

• F01D9/023

  Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings · CPC title

• F05D2240/53

  Hydrodynamic or hydrostatic bearings · CPC title

• F01D5/18

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title