Gas turbine engine with ultra high pressure compressor

Patent metadata

Field	Value
Publication number	US-11231043-B2
Application number	US-201815900891-A
Country	US
Kind code	B2
Filing date	Feb 21, 2018
Priority date	Feb 21, 2018
Publication date	Jan 25, 2022
Grant date	Jan 25, 2022

Abstract

The present disclosure is directed to a gas turbine engine including a compressor rotor. The compressor rotor includes a first stage compressor airfoil defined at an upstream-most stage of the compressor rotor. 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.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
a compressor rotor comprising a first stage compressor airfoil defined at an upstream-most stage of the compressor rotor,
wherein the first stage compressor airfoil comprises a nickel-based material,
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,
wherein the radius ratio is greater than or equal to 0.2 and less than 0.4, and
wherein the first stage compressor airfoil is structured to operate at a first stage pressure ratio of between 1.7 and 1.9 at a tip speed of between 472 and 564 meters per second.

2. The gas turbine engine of claim 1 , wherein the first stage compressor airfoil defines a substantially hollow airfoil.

3. The gas turbine engine of claim 1 , wherein the first stage compressor airfoil defines the radius ratio greater than or equal to 0.33 and less than 0.4.

4. The gas turbine engine of claim 3 , wherein the first stage compressor airfoil defines a substantially solid airfoil.

5. The gas turbine engine of claim 1 , wherein the compressor rotor comprises twelve or fewer stages.

6. The gas turbine engine of claim 5 , wherein the compressor rotor defines a compressor pressure ratio between 20:1 and 39:1.

7. The gas turbine engine of claim 1 , further comprising:
a first turbine rotor coupled to the compressor rotor via a first shaft, wherein the first turbine rotor and the compressor rotor are together rotatable via the first shaft; and
an outer casing generally surrounding the first turbine rotor and the compressor rotor, wherein the outer casing defines a core flow inlet into the core flowpath, and further wherein the first stage compressor airfoil of the compressor rotor is in direct fluid communication with the core flow inlet.

8. The gas turbine engine of claim 7 , 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.

9. The gas turbine engine of claim 1 , wherein the nickel-based material defines a tensile strength to density ratio of 0.18 or greater.

10. The gas turbine engine of claim 9 , wherein the nickel-based material of the first stage compressor airfoil further defines a tensile strength equal to or greater than 1000 Mpa.

11. The gas turbine engine of claim 1 , wherein the compressor rotor is not part of a centrifugal compressor.

12. The gas turbine engine of claim 1 , wherein the gas turbine engine does not include a centrifugal compressor.

13. The gas turbine engine of claim 1 , wherein the first stage compressor airfoil has a scallop structure at a hub portion thereof defined by the hub portion curving downward from a trailing edge to a leading edge such that the leading edge is closer to an axial centerline than the trailing edge.

14. The gas turbine engine of claim 1 , wherein the radius ratio is at a leading edge of the first stage compressor airfoil.

15. A gas turbine engine, comprising:
a core engine comprising:
a compressor rotor comprising a first stage compressor airfoil defined at an upstream-most stage of the compressor rotor, wherein the first stage compressor airfoil is structured to operate at a first stage pressure ratio between 1.7 and 1.9 at a tip speed between 472 and 564 meters per second;
a first turbine rotor coupled to the compressor rotor via a first shaft, wherein the first turbine rotor and the compressor rotor are together rotatable via the first shaft; and
a combustor assembly disposed between the compressor rotor and the first turbine rotor in direct serial flow arrangement,

wherein the first stage compressor airfoil comprises a nickel-based material,
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,
wherein the radius ratio is greater than or equal to 0.2 and less than 0.4.

16. The gas turbine engine of claim 15 , 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 core engine, and the second turbine rotor in serial flow arrangement.

17. The gas turbine engine of claim 15 , wherein the compressor rotor defines a compressor pressure ratio between 20:1 and 39:1.

18. The gas turbine engine of claim 15 , wherein the compressor rotor is not part of a centrifugal compressor.

19. The gas turbine engine of claim 15 , wherein the gas turbine engine does not include a centrifugal compressor.

20. The gas turbine engine of claim 15 , wherein the first stage compressor airfoil has a scallop structure at a hub portion thereof defined by the hub portion curving downward from a trailing edge to a leading edge such that the leading edge is closer to an axial centerline than the trailing edge.

Assignees

• Gen Electric

Inventors

• Vanapalli Veeraraju
• Mondal Bhaskar Nanda
• Narasimharao Jagata Laxmi
• Nakano Tsuguji
• Narayanan Subramanian

Classifications

• F01D5/141 Primary

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F05D2240/30

  Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor · CPC title

• F01D5/147

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title

• F01D25/24

  Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) · CPC title