Turbomachine with alternatingly spaced rotor blades

Patent metadata

Field	Value
Publication number	US-11753939-B2
Application number	US-201916280531-A
Country	US
Kind code	B2
Filing date	Feb 20, 2019
Priority date	Feb 20, 2019
Publication date	Sep 12, 2023
Grant date	Sep 12, 2023

Abstract

A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.

First claim

What is claimed is:

1. A gas turbine engine defining an axial direction and comprising:
a turbine section comprising a turbine having a plurality of first speed turbine rotor blades;
a compressor section comprising a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades, the plurality of first speed compressor rotor blades alternatingly spaced with the plurality of second speed compressor rotor blades along the axial direction;
a fan section comprising a fan; and
a first spool rotatable by the plurality of first speed turbine rotor blades and coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades, coupled to the plurality of second speed compressor rotor blades for driving the plurality of second speed compressor rotor blades, and coupled to the fan for driving the fan,
wherein the turbine further comprises a plurality of second speed turbine rotor blades,
wherein the plurality of first speed turbine rotor blades are alternatingly spaced with the plurality of second speed turbine rotor blades along the axial direction,
wherein the gas turbine engine further comprises a second spool coupled to and rotatable with the plurality of second speed turbine rotor blades and the fan for driving the fan,
wherein the plurality of second speed compressor rotor blades are further coupled to the fan for driving the fan.

2. The gas turbine engine of claim 1 ,
wherein a forward-most stage of the plurality of second speed compressor rotor blades is a forward-most stage of compressor rotor blades of the compressor section and an aft-most stage of the plurality of second speed compressor rotor blades is an aft-most stage of the compressor rotor blades of the compressor section,
wherein the forward-most stage of compressor rotor blades is coupled to the fan, and
wherein the aft-most stage of compressor rotor blades is directly coupled to a gearbox so as to be coupled to the first spool through the gearbox.

3. The gas turbine engine of claim 1 ,
wherein the plurality of first speed turbine rotor blades is a plurality of intermediate speed turbine rotor blades,
wherein the plurality of first speed compressor rotor blades is a plurality of intermediate speed compressor rotor blades, and
wherein the plurality of second speed compressor rotor blades is a plurality of low-speed compressor rotor blades.

4. The gas turbine engine of claim 1 , further comprising:
a compressor frame member positioned downstream of the compressor;
a support member coupled to the compressor frame and coupled to a gearbox;
a first bearing positioned between the support member and the first spool for supporting the first spool; and
a second bearing positioned between the first spool and the second spool and substantially aligned with the first bearing along the axial direction of the gas turbine engine.

5. The gas turbine engine of claim 1 ,
wherein the compressor is a low pressure compressor,
wherein the compressor section further comprises a high pressure compressor positioned downstream of the low pressure compressor,
wherein the high pressure compressor comprises a plurality of first speed HP compressor rotor blades and a plurality of third speed HP compressor rotor blades, and
wherein the plurality of first speed HP compressor rotor blades is coupled to the first spool.

6. The gas turbine engine of claim 5 ,
wherein the turbine is a low pressure turbine,
wherein the turbine section further comprises a high pressure turbine,
wherein the high pressure turbine comprises a plurality of third speed HP turbine rotor blades, and
wherein the gas turbine engine further comprises:
a third spool rotatable by the plurality of third speed HP turbine rotor blades, the third spool coupled to the plurality of third speed HP compressor rotor blades for driving the plurality of third speed HP compressor rotor blades.

7. The gas turbine engine of claim 6 , wherein the plurality of third speed HP compressor rotor blades and the plurality of first speed HP compressor rotor blades are alternatingly spaced along the axial direction of the gas turbine engine.

8. The gas turbine engine of claim 5 ,
wherein the turbine is a low pressure turbine,
wherein the turbine section further comprises a high pressure turbine, and
wherein the gas turbine engine further comprises:
a third spool rotatable by a plurality of third speed HP turbine rotor blades; and
an accessory drive system comprising an accessory gearbox and an accessory drive member, the accessory drive member drivingly coupled to the high pressure turbine at a location aft of the high pressure turbine.

9. The gas turbine engine of claim 1 ,
wherein the plurality of first speed turbine rotor blades is configured to rotate in an opposite circumferential direction than the plurality of second speed turbine rotor blades.

10. The gas turbine engine of claim 1 , further comprising a gearbox, wherein an entirety of the gearbox is disposed radially inward of the plurality of first speed compressor rotor blades and the plurality of second speed compressor rotor blades.

11. The gas turbine engine of claim 1 , further comprising a second speed connection member extending from the plurality of second speed compressor rotor blades in a forward axial direction to a fan disk on which the fan is disposed.

Assignees

• Gen Electric

Inventors

• Stuart Alan Roy
• Nakano Tsuguji
• Moniz Thomas Ory
• Breeze-Stringfellow Andrew
• Schmidt Richard

Classifications

• F01D5/06 Primary

  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 title

• F01D1/26 Primary

  traversed by the working-fluid substantially axially · CPC title

• F02C3/113

  with variable power transmission between rotors · CPC title

• F01D5/022

  with concentric rows of axial blades · CPC title

• F01D5/03

  Annular blade-carrying members having blades on the inner periphery of the annulus and extending inwardly radially, i.e. inverted rotors · CPC title