Compressor bleed for gas turbine engine

Patent metadata

Field	Value
Publication number	US-2023073647-A1
Application number	US-202117470261-A
Country	US
Kind code	A1
Filing date	Sep 9, 2021
Priority date	Sep 9, 2021
Publication date	Mar 9, 2023
Grant date	—

Abstract

A gas turbine engine includes a turbomachine defining a core flowpath extending through a first compressor and a second compressor, wherein a first compressor frame defines a compressor bypass passage extending from the core flowpath at the first compressor frame; and a forward compressor frame positioned between the first compressor and the second compressor, wherein the forward compressor frame defines at least in part a second portion of the core flowpath at a location downstream of the first portion of the core flowpath, wherein the forward compressor frame defines a compressor bleed passage extending from the core flowpath at a location downstream of the first compressor and upstream of the second compressor to egress a flow of air away from the core flowpath.

First claim

What is claimed is:

1 . A gas turbine engine, the engine comprising:
a turbomachine comprising
a first compressor and a second compressor in serial flow arrangement, the turbomachine defining a core flowpath extending through the first compressor and the second compressor, wherein the first compressor comprises a first compressor rotor assembly and a first compressor frame configured to support the first compressor rotor assembly, wherein the first compressor frame defines at least in part a first portion of the core flowpath at a location upstream of the second compressor and in fluid communication with the second compressor, and wherein the first compressor frame defines a compressor bypass passage extending from the core flowpath at the first compressor frame; and
a forward compressor frame positioned between the first compressor and the second compressor, wherein the forward compressor frame defines at least in part a second portion of the core flowpath at a location downstream of the first portion of the core flowpath, wherein the forward compressor frame defines a compressor bleed passage extending from the core flowpath at a location downstream of the first compressor and upstream of the second compressor to egress a flow of air away from the core flowpath.

2 . The engine of claim 1 , wherein the forward compressor frame defines a bleed outlet opening, and wherein the compressor bleed passage extends through the bleed outlet opening to egress the flow of air away from the core flowpath.

3 . The engine of claim 1 , wherein the engine defines a radial direction, wherein the first compressor frame forms a first portion of the compressor bypass passage, wherein the forward compressor frame comprises:
a first circumferential wall forming an outer radius of the second portion of the core flowpath; a second circumferential wall forming an inner radius of a second portion of the compressor bypass passage, the second portion of the compressor bypass passage located downstream of the first portion of the compressor bypass passage; and a second radial wall extended along the radial direction from the first circumferential wall to the second circumferential wall.

4 . The engine of claim 3 , wherein a compressor bleed outlet opening is defined through the second radial wall, and wherein a bleed duct extends from the compressor bleed outlet opening to the compressor bypass passage.

5 . The engine of claim 4 , wherein a bleed passage outlet is defined at a junction of the compressor bleed passage and the compressor bypass passage to egress the flow of air from the compressor bleed passage to the compressor bypass passage.

6 . The engine of claim 5 , wherein the bleed passage outlet is defined downstream of the forward compressor frame.

7 . The engine of claim 3 , wherein a compressor bleed outlet opening is defined through the second radial wall, and wherein a bleed duct extends from the compressor bleed outlet opening to allow fluid communication of the compressor bypass passage to a fan bypass passage.

8 . The engine of claim 7 , wherein a bleed passage outlet is defined at a junction of the compressor bleed passage and the fan bypass passage to egress the flow of air from the compressor bleed passage to the fan bypass passage.

9 . The engine of claim 8 , wherein the engine comprises:
a strut extended through the compressor bypass passage to allow for fluid communication of the compressor bleed passage to the fan bypass passage.

10 . The engine of claim 9 , wherein the strut is extended through the compressor bypass passage aft of the forward compressor frame.

11 . The engine of claim 3 , wherein a compressor bleed outlet opening is defined through the second radial wall, and wherein a bleed duct extends along an axial direction from the compressor bleed outlet opening, and wherein the bleed duct surrounds one or more of the second compressor, a combustion section, or a turbine section of the engine.

12 . The engine of claim 11 , wherein a bleed passage outlet is formed at an outlet end of an exhaust nozzle of the compressor bypass passage.

13 . The engine of claim 3 , wherein a compressor bleed outlet opening is defined through the second circumferential wall to allow for fluid communication from the compressor bleed passage to the compressor bypass passage.

14 . The engine of claim 13 , wherein the forward compressor frame forms the compressor bleed passage as a serpentine passage.

15 . The engine of claim 13 , wherein the forward compressor frame is an integral, unitary structure.

16 . The engine of claim 1 , further comprising:
a flow control device positioned at the compressor bleed passage, wherein the flow control device is configured to modulate an amount of air allowed into the compressor bleed passage from the core flowpath.

17 . The engine of claim 1 , wherein the engine is an open rotor engine configuration.

18 . The engine of claim 1 , the engine comprising:
a fan section comprising a plurality of fan blades, wherein a fan bypass passage is defined downstream of the plurality of fan blades, and wherein the fan section comprises a plurality of outlet guide vanes extended from the forward compressor frame; and wherein the compressor bypass passage comprises an exhaust nozzle at an outlet end, and wherein the compressor bypass passage is configured to generate a pressure ratio at the compressor bypass passage higher than a pressure ratio at the fan bypass passage during operation of the engine.

19 . The engine of claim 1 , further comprising an outer casing surrounding the first and second compressors, wherein the outer casing comprises:
a first outer casing forming, at least in part, a fan bypass passage; and a second outer casing forming, at least in part, the compressor bypass passage between the first outer casing and the second outer casing.

20 . The engine of claim 1 , wherein the first compressor frame comprises a first vane extended through the compressor bypass passage, and wherein the forward compressor frame comprises a second vane extended through the compressor bypass passage aft of the first vane.

Assignees

• Gen Electric

Inventors

• Glessner John Carl
• Oehrle Daniel John
• Linz Mark Edward

Classifications

• F02C6/08 Primary

  the gas being bled from the gas-turbine compressor · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F02C9/18 Primary

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F05D2220/32

  in gas turbines · CPC title