Gas turbine engine with third stream

Patent metadata

Field	Value
Publication number	US-12565866-B2
Application number	US-202418991142-A
Country	US
Kind code	B2
Filing date	Dec 20, 2024
Priority date	Aug 2, 2022
Publication date	Mar 3, 2026
Grant date	Mar 3, 2026

Abstract

A gas turbine engine includes a 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; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10. The thrust to power airflow ratio is a ratio of airflow through a bypass passage over the turbomachine plus airflow through the fan duct to airflow through the core duct. The core bypass ratio is a ratio of airflow through the fan duct to airflow through the core duct. The fan duct includes an exhaust nozzle having a plurality of chevrons disposed at its aft end to define an exhaust outlet.

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; and 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, wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct, and wherein the fan duct includes a fan exhaust nozzle having a plurality of chevrons disposed at an aft end of the fan exhaust nozzle to define an exhaust outlet.

2 . The gas turbine engine of claim 1 , wherein the thrust to power airflow ratio and the core bypass ratio are defined when the gas turbine engine is operated at a rated speed during standard day operating conditions.

3 . The gas turbine engine of claim 1 , wherein the thrust to power airflow ratio is between 14 and 75.

4 . The gas turbine engine of claim 1 , wherein the thrust to power airflow ratio is between 35 and 50.

5 . The gas turbine engine of claim 1 , wherein the core bypass ratio is between 0.3 and 5.

6 . The gas turbine engine of claim 1 , wherein the core bypass ratio is between 0.5 and 3.

7 . The gas turbine engine of claim 1 , wherein the secondary fan is a single stage secondary fan.

8 . The gas turbine engine of claim 1 , wherein the secondary fan is a multi-stage secondary fan.

9 . The gas turbine engine of claim 1 , wherein the primary fan is an unducted primary fan.

10 . The gas turbine engine of claim 1 , wherein core duct includes a core exhaust nozzle having at least one mixing device.

11 . The gas turbine engine of claim 1 , wherein the plurality of chevrons are each triangular in configuration.

12 . The gas turbine engine of claim 1 , wherein the plurality of chevrons include first chevrons having a first geometric shape and second chevrons having a second geometric shape different than the first geometric shape.

13 . The gas turbine engine of claim 1 , wherein the plurality of chevrons include first chevrons that are concave in configuration and second chevrons that are convex in configuration.

14 . The gas turbine engine of claim 1 , wherein the plurality of chevrons include first chevrons having a first length and second chevrons having a second length different than the first length.

15 . The gas turbine engine of claim 1 , wherein the plurality of chevrons include first chevrons having a first width and second chevrons having a second width different than the first width.

16 . 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, wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct, and wherein the fan duct includes an exhaust nozzle having a plurality of chevrons disposed at an aft end of the exhaust nozzle to define an exhaust outlet.

17 . The method of claim 16 , wherein the thrust to power airflow ratio is between 14 and 75.

18 . The method of claim 16 , wherein the primary fan is an unducted primary fan, and wherein the thrust to power airflow ratio is between 30 and 60.

19 . The method of claim 16 , wherein the core bypass ratio is between 0.3 and 5.

20 . The method of claim 16 , wherein the plurality of chevrons include first chevrons having a first geometric shape and second chevrons having a second geometric shape different than the first geometric shape.

Assignees

• Gen Electric

Inventors

• Miller Brandon W
• Vondrell Randy M
• Ostdiek David M
• Higgins Craig W
• Simpson Alexander
• Bowden William

Classifications

• F02C3/06 Primary

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

• F05D2250/611

  undulated · CPC title

• F05D2250/61

  corrugated · CPC title

• F05D2250/11

  triangular · CPC title

• F02K1/34

  for attenuating noise · CPC title