TWO SPOOL GAS TURBINE ENGINE WITH INTERDIGITATED TURBINE SECTION

• US 20180274365A1
• Filed: 01/23/2017
• Published: 09/27/2018
• Est. Priority Date: 01/23/2017
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Stuart, Alan Roy, Clements, Jeffrey Donald, Schmidt, Richard, Moniz, Thomas Ory
• Granted Patent: US 10,539,020 B2
• Time in Patent Office: Days
• CPC Class Codes: F01D 1/24 characterised by counter-rotating rotors subjected to same working fluid stream without intermediate stator blades or the like
• CPC Class Codes: F01D 1/26 traversed by the working-fluid substantially axially
• CPC Class Codes: F01D 5/03 Annular blade-carrying members having blades on the inner periphery of the annulus and extending inwardly radially, i.e. inverted rotors
• CPC Class Codes: F01D 5/06 Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections F01D5/022, F01D5/023 take precedence
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F01D 5/146 of blades with tandem configuration, split blades or slotted blades
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F02C 3/067 having counter-rotating rotors F02C3/073 takes precedence
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F02K 3/072 with counter-rotating , e.g. fan rotors
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2240/24 for turbines
• CPC Class Codes: F05D 2240/50 Bearings
• CPC Class Codes: F05D 2250/38 angled, e.g. sweep angle

Abstract

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section that includes a first rotating component and a second rotating component. The first rotating component includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction. The outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction. The first rotating component further includes at least one connecting airfoil coupling the inner shroud and the outer shroud. The second rotating component is upstream of the one or more connecting airfoils of the first rotating component along the longitudinal direction. The second rotating component includes a plurality of second airfoils extended outward in the radial direction. The first rotating component defines at least one stage of the plurality of outer shroud airfoils upstream of the second rotating component.

Claims

1. A gas turbine engine, wherein the gas turbine engine defines a longitudinal direction, a radial direction, and a circumferential direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction, the gas turbine engine comprising:

a turbine section comprising a first rotating component and a second rotating component, wherein the first rotating component includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction, wherein the outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction, and wherein the first rotating component further includes at least one connecting airfoil coupling the inner shroud and the outer shroud, and wherein the second rotating component is upstream of the one or more connecting airfoils of the first rotating component along the longitudinal direction, the second rotating component includes a plurality of second airfoils extended outward in the radial direction, and wherein the first rotating component defines at least one stage of the plurality of outer shroud airfoils upstream of the second rotating component.

2. The gas turbine engine of claim 1, wherein the first rotating component defines a lean angle relative to the axial centerline, and wherein the plurality of outer shroud airfoils and/or the one or more connecting airfoils each define an obtuse lean angle approximately perpendicular to the axial centerline or extending toward the upstream end along the longitudinal direction from inward to outward along the radial direction.

3. The gas turbine engine of claim 1, wherein the first rotating component defines a lean angle relative to the axial centerline, and wherein the plurality of outer shroud airfoils and/or the one or more connecting airfoils each define an acute lean angle extending toward the downstream end along the longitudinal direction from inward to outward along the radial direction.

4. The gas turbine engine of claim 1, wherein the first rotating component and the second rotating component are in interdigitation along the longitudinal direction.

5. The gas turbine engine of claim 1, wherein the second rotating component defines a high speed turbine and the first rotating component defines a low speed turbine.

6. The gas turbine engine of claim 1, wherein the gas turbine engine defines, in serial flow arrangement along the longitudinal direction from the upstream end to the downstream end, the plurality of outer shroud airfoils of the first rotating component, the plurality of second airfoils of the second rotating component, and the one or more connecting airfoils of the first rotating component.

7. The gas turbine engine of claim 1, wherein the inner shroud of the first rotating component defines a plurality of inner shroud airfoils extended outward along the radial direction.

8. The gas turbine engine of claim 7, wherein the inner shroud extends from the connecting airfoil toward the downstream end.

9. The gas turbine engine of claim 1, further comprising a combustion section arranged in serial flow arrangement with the turbine section, and wherein the combustion section, the first stage of the first rotating component, and the second rotating component are in serial flow arrangement along the longitudinal direction from the upstream end to the downstream end.

10. The gas turbine engine of claim 9, wherein the gas turbine engine defines, in serial flow arrangement along the longitudinal direction from the upstream end to the downstream end, the combustion section, the first stage of the first rotating component, the second rotating component, and then the first rotating component.

11. The gas turbine engine of claim 1, wherein the plurality of outer shroud airfoils at the first stage are coupled to an axially extended hub disposed inward along the radial direction of the plurality of outer shroud airfoils.

12. The gas turbine engine of claim 11, wherein the plurality of outer shroud airfoils at the first stage is further coupled to an arm extended generally inward along the radial direction, and wherein the arm is coupled to the axially extended hub, and wherein the axially extended hub extends generally in the longitudinal direction toward the upstream end of the engine.

13. The gas turbine engine of claim 11, further comprising a first turbine bearing, and wherein the second rotating component is further coupled to a second shaft extended toward the upstream end, and wherein the first turbine bearing is disposed along the radial direction between the second shaft and the axially extended hub of the first rotating component.

14. The gas turbine engine of claim 13, wherein the first turbine bearing defines an air bearing, a foil bearing, a roller bearing, or a ball bearing.

15. The gas turbine engine of claim 1, wherein the inner shroud of the first rotating component defines an inner shroud diameter and the outer shroud of the first rotating component defines an outer shroud diameter, and wherein the inner shroud diameter is approximately 115% or less of the outer shroud diameter.

16. The gas turbine engine of claim 15, wherein the inner shroud diameter is approximately equal to the outer shroud diameter.

17. The gas turbine engine of claim 1, wherein the first rotating component defines between about 3 and 10 stages inclusively.

18. The gas turbine engine of claim 17, wherein the first rotating component defines at least two stages of the plurality of outer shroud airfoils upstream of the connecting airfoil.

19. The gas turbine engine of claim 1, wherein the first rotating component rotates in a first direction and the second rotating component rotates in a second direction opposite of the first direction.

20. The gas turbine engine of claim 9, further comprising:

a compressor section comprising a high pressure compressor; and

a fan assembly defining one or more stages of a plurality of blades, wherein the fan assembly, the compressor section, the combustion section, and the turbine section are in serial flow arrangement along the longitudinal direction from the upstream end to the downstream end, and wherein the first rotating component is connected and rotatable with the fan assembly by a first shaft, and wherein the second rotating component is connected and rotatable with the high pressure compressor by a second shaft.