Exhaust nozzle arrangement for geared turbofan

Patent metadata

Field	Value
Publication number	US-10294871-B2
Application number	US-201514875750-A
Country	US
Kind code	B2
Filing date	Oct 6, 2015
Priority date	Feb 6, 2013
Publication date	May 21, 2019
Grant date	May 21, 2019

Abstract

A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine section configured to drive the geared architecture. The turbine section has an exit point, and a diameter (Dt) defined as the outer diameter of a last blade airfoil stage in the turbine section at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance (Lc or Ln) from the exit point.

First claim

What is claimed is:

1. A propulsion system, comprising:
a geared architecture that drives a fan section including a fan, said fan including a plurality of fan blades;
a turbine section that drives said geared architecture, said turbine section having an exit point, and a diameter (D t ) defined as the outer diameter of a last blade airfoil stage in said turbine section at said exit point;
a nacelle surrounding a core engine housing, said fan that delivers air into a bypass duct defined between said nacelle and said core engine housing;
a core engine exhaust nozzle downstream of said exit point, with a downstream most point of said core engine exhaust nozzle being defined at a distance (L c or L n ) from said exit point, wherein a ratio of said distance (L c or L n ) to said diameter (D t ) is greater than or equal to about 1.30;
wherein said ratio is less than or equal to 1.64; and
wherein said core engine exhaust nozzle includes a plug, said downstream most point of said core engine exhaust nozzle is defined by a downstream end of said plug.

2. The propulsion system as set forth in claim 1 , wherein said core engine exhaust nozzle includes an inner periphery adjacent to a bypass flow path and disposed about said plug, said inner periphery tapering inwardly at an angle to define a trailing edge, and a maximum value of said angle is less than 17 degrees.

3. The propulsion system as set forth in claim 1 , wherein said turbine section drives a compressor section, said turbine section including a fan drive turbine that drives said fan section, said fan drive turbine defining said exit point, and said fan drive turbine includes an inlet, an outlet and a fan drive turbine pressure ratio greater than 5, wherein said fan drive turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

4. The propulsion system as set forth in claim 1 , wherein an exhaust case is positioned between said turbine section and said core engine exhaust nozzle.

5. The propulsion system as set forth in claim 1 , wherein said nacelle includes an outer surface downstream of a point defining a maximum diameter of said nacelle, said outer surface tapering inwardly at an angle to define a trailing edge, and a maximum value of said angle is less than 14 degrees.

6. The propulsion system as set forth in claim 5 , wherein:
said fan delivers a portion of air into a compressor section and a portion of air into said bypass duct, said fan defining a pressure ratio less than 1.45 across the plurality of fan blades alone; and
a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is greater than 6.

7. The propulsion system as set forth in claim 6 , wherein said bypass ratio is greater than 10, and said geared architecture defines a gear reduction ratio greater than 2.3.

8. The propulsion system as set forth in claim 1 , wherein said ratio is greater than 1.39.

9. A propulsion system, comprising:
a geared architecture that drives a fan section including a fan, said fan including a plurality of fan blades;
a turbine that drives said geared architecture, said turbine having an exit point, and a diameter (D t ) defined as the outer diameter of a last blade airfoil stage in said turbine at said exit point;
a nacelle surrounding a core engine housing, said fan delivers air into a bypass duct defined between said nacelle and said core engine housing;
a core engine exhaust nozzle downstream of said exit point, said core engine exhaust nozzle including a plug; and
wherein a downstream most point of said core engine nozzle is defined by a downstream end of said plug at a distance (L c ) from said exit point, and a ratio of said distance (L c ) to said diameter (D t ) is between 1.30 and 1.64.

10. The propulsion system as set forth in claim 9 , wherein said core engine exhaust nozzle includes an inner periphery disposed about said plug and bounding a bypass flow path, said inner periphery tapering inwardly at an angle to define a trailing edge region, and a maximum value of said angle is less than 17 degrees.

11. The propulsion system as set forth in claim 9 , wherein said nacelle includes an outer surface downstream of a maximum diameter of said nacelle, said outer surface tapering inwardly at an angle to define a trailing edge region, and a maximum value of said angle is less than 14 degrees.

12. The propulsion system as set forth in claim 11 , wherein said turbine drives said fan section, said turbine includes an inlet, an outlet and a turbine pressure ratio greater than 5, and said turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

13. The propulsion system as set forth in claim 12 , wherein:
said fan delivers a portion of air into a compressor section and a portion of air into said bypass duct, said fan defining a pressure ratio less than 1.45 across the plurality of fan blades alone and
a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is greater than 10.

14. The propulsion system as set forth in claim 9 , wherein an exhaust case is positioned between said exit point of said turbine and an inlet of said engine exhaust nozzle such that said exhaust case slopes radially outward from said exit point to said inlet.

15. The propulsion system as set forth in claim 9 , wherein said ratio is greater than 1.39.

16. A propulsion system, comprising:
a geared architecture that drives a fan section including a fan, said fan including a plurality of fan blades;
a turbine section including a first turbine that drives a first compressor and a fan drive turbine that drives said geared architecture, said turbine section having an exit point, and a diameter (D t ) defined as the outer diameter of a last blade airfoil stage in said fan drive turbine at said exit point;
a nacelle surrounding a core engine housing, said core engine housing surrounding at least said turbine section and said first compressor, said fan delivers air into a bypass duct defined between said nacelle and said core engine housing;
a core engine exhaust nozzle downstream of said exit point, said core engine exhaust nozzle including an inner periphery disposed about a plug; and
wherein a downstream most point of said core engine nozzle is defined by a downstream end of said plug at a distance (L c ) from said exit point, and a ratio of said distance (L c ) to said diameter (D t ) is between 1.30 and 1.64.

17. The propulsion system as set forth in claim 16 , wherein said inner periphery tapers inwardly at an angle to define a trailing edge region, and a maximum value of said angle is less than or equal to 15 degrees.

18. The propulsion system as set forth in claim 16 , wherein said fan drive turbine drives a second compressor, said fan drive turbine includes an inlet, an outlet and a fan drive turbine pressure ratio greater than 5, and said fan drive turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

19. The propulsion system as set forth in claim 16 , wherein said ratio is greater than 1.39.

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M
• Malecki Robert E

Classifications

• F02C7/36 Primary

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title

• F02K3/06

  with front fan · CPC title

• B64D33/02

  of combustion air intakes · CPC title

• F02K3/068

  being characterised by a short axial length relative to the diameter · CPC title

• F02C3/107

  with two or more rotors connected by power transmission · CPC title