Gas turbine engine with third stream

Patent metadata

Field	Value
Publication number	US-12560132-B2
Application number	US-202418982445-A
Country	US
Kind code	B2
Filing date	Dec 16, 2024
Priority date	Aug 2, 2022
Publication date	Feb 24, 2026
Grant date	Feb 24, 2026

Abstract

A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The primary fan is driven by the turbomachine, and a secondary fan is located downstream of the primary fan within the inlet duct. One or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to increase or decrease an exit area of the fan duct.

First claim

We claim:

1 . A gas turbine engine comprising:
a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct; and one or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to increase or decrease an exit area of the fan duct.

2 . The gas turbine engine of claim 1 , wherein the one or more actuation devices comprise at least one actuation device actuable to open or close a bleed passageway extending radially inward or radially outward from the fan duct.

3 . The gas turbine engine of claim 2 , wherein the bleed passageway extends longitudinally either axially or circumferentially.

4 . The gas turbine engine of claim 1 , wherein the one or more actuation devices comprise:
a first actuation device; and a second actuation device; and wherein the first and second actuation devices are ganged together or independently actuable.

5 . The gas turbine engine of claim 1 , further comprising a frame assembly defining one or more interface locations configured to support attachment of an outer nacelle to the gas turbine engine, and wherein the one or more actuation devices are disposed:
upstream of at least one interface location of the one or more interface locations with respect to the airflow through the fan duct; downstream of the at least one interface location with respect to the airflow through the fan duct; or both upstream and downstream of the at least one interface location with respect to the airflow through the fan duct.

6 . The gas turbine engine of claim 1 , further comprising a core cowl surrounding at least a portion of the turbomachine, and wherein at least one actuation device of the one or more actuation devices is actuable to open or close a bleed passageway extending radially inward through the core cowl.

7 . The gas turbine engine of claim 6 , wherein the core cowl defines a core cowl vent, and when the bleed passageway is open, a portion of the airflow through the fan duct flows through the bleed passageway and is exhausted through the core cowl vent.

8 . The gas turbine engine of claim 7 , wherein at least a second actuation device of the one or more actuation devices is actuable to vary an area of the core cowl vent.

9 . The gas turbine engine of claim 6 , wherein the core cowl comprises one or more doors, and when the bleed passageway is open, at least a portion of the airflow through the fan duct flows through the bleed passageway and is exhausted through the one or more doors.

10 . The gas turbine engine of claim 1 , wherein the airflow through the fan duct exits the fan duct through a fan duct exhaust nozzle, and wherein the one or more actuation devices comprises at least one flap defining at least a portion of the fan duct, wherein the at least one flap is movable to vary an area of the fan duct exhaust nozzle.

11 . The gas turbine engine of claim 10 , further comprising an outer nacelle defining at least a portion of the fan duct, and wherein the at least one flap is movably coupled to the outer nacelle.

12 . The gas turbine engine of claim 1 , further comprising at least one structural element disposed at least partially within the fan duct, and wherein at least one actuation device of the one or more actuation devices comprises at least one fairing axially movable to increase or decrease an area of the fan duct between the at least one structural element and the at least one fairing.

13 . The gas turbine engine of claim 1 , further comprising a core cowl surrounding at least a portion of the turbomachine, and wherein at least one actuation device of the one or more actuation devices is actuable to move at least a portion of the core cowl radially inward.

14 . The gas turbine engine of claim 1 , further comprising a controller communicatively coupled to the one or more actuation devices and configured to control actuation of the one or more actuation devices.

15 . A method of operating a gas turbine engine, comprising:
operating the gas turbine engine at a rated speed, wherein operating the gas turbine engine at the rated speed comprises operating the gas turbine engine to define a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 5, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over a turbomachine of the gas turbine engine plus an airflow through a fan duct to an airflow through a core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct; and actuating one or more actuation devices operably associated with the fan duct to increase or decrease an exit area of the fan duct.

16 . The method of claim 15 , wherein actuating the one or more actuation devices comprises opening one or more bleed passageways fluidly connected to the fan duct.

17 . The method of claim 15 , wherein actuating the one or more actuation devices comprises moving one or more flaps defining at least a portion of the fan duct.

18 . The method of claim 15 , wherein a core cowl surrounds at least a portion of the turbomachine, and wherein actuating the one or more actuation devices comprises:
opening one or more bleed passageways fluidly connected to the fan duct into the core cowl; and opening one or more core cowl vents positioned at an aft portion of the core cowl.

19 . A gas turbine engine comprising:
a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct; and one or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to open or close one or more bleed passageways fluidly connected to the fan duct.

20 . The gas turbine engine of claim 19 , further comprising a controller communicatively coupled to the one or more actuation devices and configured to control actuation of the one or more actuation devices.

Assignees

• Gen Electric

Inventors

• Bowden William Joseph
• Blodgett Keith Edward James
• Dindar Mustafa
• Cerra David
• Grooms James Hamilton
• Miller Brandon Wayne
• Vondrell Randy M
• Ostdiek David Marion
• Higgins Craig William
• Simpson Alexander Kimberley

Classifications

• F02C3/06 Primary

  the compressor comprising only axial stages (F02C3/10 takes precedence) · CPC title

• F02K1/30

  for varying effective area of jet pipe or nozzle · CPC title

• F02K1/085

  by transversely deforming an internal member · CPC title

• F02C9/18

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

• F02K3/077

  the plant being of the multiple flow type, i.e. having three or more flows · CPC title