Gas turbine engine arrangement with ultra high pressure compressor

Patent metadata

Field	Value
Publication number	US-10823191-B2
Application number	US-201815922153-A
Country	US
Kind code	B2
Filing date	Mar 15, 2018
Priority date	Mar 15, 2018
Publication date	Nov 3, 2020
Grant date	Nov 3, 2020

Abstract

The present disclosure is directed to a gas turbine engine including a first frame comprising a first bearing assembly, a second frame comprising a second bearing assembly, and a compressor rotor. A first stage compressor airfoil is defined at an upstream-most stage of the compressor rotor. The compressor rotor is rotatable via the first bearing assembly and the second bearing assembly. The first stage compressor airfoil is disposed between the first bearing assembly and the second bearing assembly.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
a first frame comprising a first bearing assembly;
a compressor rotor comprising a first stage compressor airfoil defined at an upstream-most stage of the compressor rotor; and
a second frame comprising a second bearing assembly,
wherein the compressor rotor is rotatable via the first bearing assembly and the second bearing assembly,
wherein the first stage compressor airfoil is disposed between the first bearing assembly and the second bearing assembly,
wherein an entirety of the first stage compressor airfoil is disposed within a core flowpath of the engine, and
wherein the first stage compressor airfoil overlaps at least partially with a bearing of the first bearing assembly in an axial direction of the gas turbine engine.

2. The engine of claim 1 , wherein a radial plane is defined extended from an axial centerline of the compressor rotor, and wherein the second bearing assembly is disposed co-planar to the compressor rotor along the radial plane.

3. The engine of claim 2 , wherein the second bearing assembly is disposed aft of the first stage compressor airfoil of the compressor rotor.

4. The engine of claim 1 , wherein the first frame defines a first airfoil upstream in fluid communication with the compressor rotor.

5. The engine of claim 1 , wherein the second frame comprises a structural member extended radially across a core flowpath of the engine.

6. The engine of claim 5 , wherein the second frame further comprises a second airfoil extended radially across the core flowpath.

7. The engine of claim 6 , wherein the second airfoil defines a variable vane at least partially rotatable around a radial axis of the second airfoil.

8. The engine of claim 1 ,
wherein the second frame comprises a plurality of structural members extended radially across a core flowpath of the engine, and
wherein the second frame further defines a second airfoil disposed between the plurality of structural members.

9. The engine of claim 1 , further comprising:
a combustor assembly;
a first turbine rotor; and
a third bearing assembly, wherein the third bearing assembly provides rotatable support to the compressor rotor and the first turbine rotor, and further wherein the third bearing assembly is downstream of the second bearing assembly.

10. The engine of claim 9 , wherein the third bearing assembly is disposed radially inward of the combustor assembly or the first turbine rotor.

11. The engine of claim 9 , further comprising:
a fan assembly in serial flow arrangement upstream of the compressor rotor, wherein the compressor rotor is in direct fluid communication with the fan assembly; and
a second turbine rotor coupled to the fan assembly via a second shaft, wherein the second turbine rotor and the fan assembly are together rotatable via the second shaft, and further wherein the gas turbine engine defines the fan assembly, the compressor rotor, the combustor assembly, the first turbine rotor, and the second turbine rotor in direct serial flow arrangement.

12. The engine of claim 11 , further comprising:
an outer casing generally surrounding the first turbine rotor and the compressor rotor, wherein the outer casing defines a core flow inlet into a core flowpath, and further wherein the first stage compressor airfoil of the compressor rotor is in direct fluid communication with the core flow inlet.

13. The engine of claim 1 , wherein the first stage compressor airfoil defines a first stage pressure ratio of at least approximately 1.7 during operation of the gas turbine engine at a tip speed of at least approximately 472 meters per second.

14. The engine of claim 13 , wherein the first stage compressor airfoil defines a maximum first stage pressure ratio of approximately 1.9.

15. The engine of claim 13 , wherein the first stage compressor airfoil defines a radius ratio of an inner radius of the first stage compressor airfoil within a core flowpath versus an outer radius of the first stage compressor airfoil within the core flowpath, and wherein the radius ratio is less than approximately 0.4.

16. The engine of claim 15 , wherein the first stage compressor airfoil defines the radius ratio between approximately 0.2 and approximately 0.4.

17. The engine of claim 1 , wherein the compressor rotor defines a maximum tip speed of approximately 564 meters per second or less.

18. The engine of claim 1 , wherein the first stage compressor airfoil comprises a first material defining a tensile strength to density ratio of approximately 0.18 or greater.

19. The engine of claim 18 , wherein the first material of the first stage compressor airfoil further defines a tensile strength equal to or greater than approximately 1000 Mpa.

20. The engine of claim 1 , wherein the compressor rotor defines a compressor pressure ratio between approximately 20:1 and approximately 39:1.

Assignees

• Gen Electric

Inventors

• Vanapalli Veeraraju
• Mondal Bhaskar Nanda
• Wankhade Rajendra Mahadeorao
• Kollam Ramana Reddy

Classifications

• F02C7/06 Primary

  Arrangements of bearings (bearings F16C); Lubricating ({of turbo machines F01D25/18; of machines or} engines in general F01M) · CPC title

• F05D2240/54

  Radial bearings · CPC title

• F04D29/26

  Rotors specially for elastic fluids · CPC title

• F05D2300/17

  Alloys · CPC title

• F02K3/06

  with front fan · CPC title