Rotor thrust balanced turbine engine

Patent metadata

Field	Value
Publication number	US-11053797-B2
Application number	US-201715412216-A
Country	US
Kind code	B2
Filing date	Jan 23, 2017
Priority date	Jan 23, 2017
Publication date	Jul 6, 2021
Grant date	Jul 6, 2021

Abstract

The present disclosure is directed to a rotor thrust balanced turbine engine that may increase engine performance and efficiency while managing thrust mismatch or imbalance in a low pressure (LP) spool between a fan assembly and a turbine rotor assembly. The gas turbine engine defines a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a turbine rotor assembly and a turbine frame. The turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline. The first flowpath radius is disposed at the upstream end of the turbine rotor assembly, and wherein the second flowpath radius is disposed at the downstream end of the turbine rotor assembly. The turbine frame and the turbine rotor assembly together define a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein the turbine rotor assembly defines a ratio of the first flowpath radius to the seal interface radius less than or equal to approximately 1.79.

First claim

What is claimed is:

1. A gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, and an axial centerline extended along the longitudinal direction, the gas turbine engine comprising:
a low pressure spool comprising a fan assembly and a turbine rotor assembly, wherein the low pressure spool is connected to a bearing assembly, and wherein the turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline, wherein the first flowpath radius is disposed at an upstream end of the turbine rotor assembly, wherein the upstream end is at a forward edge of a blade of a first rotating stage of the turbine rotor assembly, and wherein the second flowpath radius is disposed at a downstream end of the turbine rotor assembly; and
a turbine frame, wherein the turbine frame and the turbine rotor assembly together define a seal interface, the seal interface positioned forward of or aligned along the radial direction with the upstream end of the turbine rotor assembly, and the seal interface having a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein the turbine rotor assembly defines a ratio of the first flowpath radius to the seal interface radius, the ratio of the first flowpath radius to the seal interface radius defining at least in part a thrust balance at the bearing assembly at the low pressure spool, wherein the ratio of the first flowpath radius to the seal interface radius is less than or equal to 1.79.

2. The gas turbine engine of claim 1 , wherein the turbine rotor assembly defines a ratio of the second flowpath radius to seal interface radius less than or equal to 1.74.

3. The gas turbine engine of claim 1 , wherein the turbine rotor assembly defines a range of ratios of the first flowpath radius to the seal interface radius of 1.79 to 1.50.

4. The gas turbine engine of claim 1 , wherein the seal interface radius is 130 mm to 430 mm.

5. The gas turbine engine of claim 1 , wherein the turbine rotor assembly defines a range of ratios of the second flowpath radius to the seal interface radius of 1.74 to 1.50.

6. The gas turbine engine of claim 1 , wherein the gas turbine engine defines a sea level thrust imbalance range between 44 kN and 515 kN.

7. The gas turbine engine of claim 1 , wherein the seal interface radius is 430 mm or less, and wherein the gas turbine engine defines a sea level rotor thrust imbalance of at least 44 kN.

8. The gas turbine engine of claim 1 , wherein the seal interface radius is 220 mm to 430 mm, and wherein the gas turbine engine defines a sea level thrust imbalance range of at least 340 kN.

9. The gas turbine engine of claim 1 , wherein the seal interface radius is between 130 mm and 220 mm, and wherein the gas turbine engine defines a sea level rotor thrust imbalance of at least 44 kN.

10. The gas turbine engine of claim 1 , wherein the turbine frame and turbine rotor assembly together define the seal interface at the seal interface radius.

11. The gas turbine engine of claim 10 , wherein the seal interface defines a first cavity generally inward of a second cavity along the radial direction, and wherein the second cavity defines a higher pressure than the first cavity.

12. The gas turbine engine of claim 10 , wherein the seal interface defines a shroud at the turbine frame and a knife edge seal at the turbine rotor assembly.

13. The gas turbine engine of claim 10 , wherein the seal interface defines a first cavity and a second cavity, and wherein the second cavity defines a higher pressure than the first cavity.

14. The gas turbine engine of claim 1 , wherein the second flowpath radius corresponds to a last rotating stage of the turbine rotor assembly.

15. The gas turbine engine of claim 1 , wherein the first and second flowpath radii each correspond to an inner radius of a core flowpath of the engine.

16. The gas turbine engine of claim 1 , wherein the gas turbine engine defines a direct drive engine, and wherein a fan rotor and the turbine rotor assembly rotate at the same rotational speed.

17. A gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, and an axial centerline extended along the longitudinal direction, the gas turbine engine comprising:
a low pressure spool comprising a fan assembly and a turbine rotor assembly, the fan assembly and the turbine rotor assembly rotatably connected to one another by a shaft, wherein the low pressure spool is connected to a bearing assembly at the shaft, and wherein the turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline, wherein the first flowpath radius corresponds to a forward edge of a blade of a first rotating stage of the turbine rotor assembly and the second flowpath radius corresponds to a last rotating stage of the turbine rotor assembly, and wherein the first flowpath radius is disposed at an upstream end of the turbine rotor assembly and the second flowpath radius is disposed at a downstream end of the turbine rotor assembly; and
a turbine frame, wherein the turbine frame and the turbine rotor assembly together define a seal interface, the seal interface positioned forward of or aligned along the radial direction with the upstream end of the turbine rotor assembly, and the seal interface having a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein the turbine rotor assembly defines a ratio of the first flowpath radius to the seal interface radius, the ratio of the first flowpath radius to the seal interface radius and a ratio of the second flowpath radius to the seal interface radius together defining at least in part a thrust balance at the bearing assembly at the low pressure spool, wherein the ratio of the first flowpath radius to the seal interface radius is less than or equal 1.79, and wherein the ratio of the second flowpath radius to the seal interface radius is less than or equal 1.74.

18. A gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, and an axial centerline extended along the longitudinal direction, the gas turbine engine comprising:
a low pressure spool comprising a fan assembly and a turbine rotor assembly, the fan assembly and the turbine rotor assembly rotatably connected to one another by a shaft, wherein the low pressure spool is connected to a bearing assembly at the shaft, and wherein the turbine rotor assembly defines a first flowpath radius corresponding to a forward edge of a blade of a first rotating stage and the turbine rotor assembly defines a second flowpath radius corresponding to a last rotating stage, wherein the first flowpath radius is disposed at an upstream end of the turbine rotor assembly and the second flowpath radius is disposed at a downstream end of the turbine rotor assembly, and wherein the first flowpath radius and the second flowpath radius are each extended from the axial centerline; and
a turbine frame, wherein the turbine frame and the turbine rotor assembly together define a seal interface, the seal interface positioned forward of or aligned along the radial direction with the upstream end of the turbine rotor assembly, and the seal interface having a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein th

Assignees

• Gen Electric

Inventors

• Moniz Thomas Ory
• Clements Jeffrey Donald

Classifications

• F01D3/00 Primary

  Machines or engines with axial-thrust balancing effected by working-fluid · CPC title

• F01D25/16

  Arrangement of bearings; Supporting or mounting bearings in casings (bearings per se F16C) · CPC title

• F01D11/005

  Sealing means between non relatively rotating elements · CPC title

• F02C7/04 Primary

  Air intakes for gas-turbine plants or jet-propulsion plants · CPC title

• F01D11/00

  Preventing or minimising internal leakage of working-fluid, e.g. between stages (sealings in general F16J {; sealing arrangements for transition ducts of combustor cans F01D9/023}) · CPC title