Elongated geared turbofan with high bypass ratio

Patent metadata

Field	Value
Publication number	US-10436121-B2
Application number	US-201314038886-A
Country	US
Kind code	B2
Filing date	Sep 27, 2013
Priority date	Feb 6, 2013
Publication date	Oct 8, 2019
Grant date	Oct 8, 2019

Abstract

A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined 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 provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

First claim

The invention claimed is:

1. A propulsion system comprising:
a fan section including a fan;
a gear system, wherein said gear system is an epicyclic gear train;
a turbine section incorporating a high pressure turbine and a fan drive turbine that drives said gear system that drives said fan, said turbine section having an exit point at a last blade airfoil stage, and a diameter (D t ) defined as the diameter of said last blade airfoil stage in the turbine section at said exit point;
a compressor section including at least two compressor rotors with said high pressure turbine driving a high pressure compressor rotor and there being a lower pressure compressor rotor also being driven;
a nacelle surrounding a core engine housing, said fan delivers air into a bypass duct defined between said nacelle and said core engine housing to said low pressure compressor, and a bypass ratio defined as a volume of air delivered into said bypass duct compared to the volume of air delivered into 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 downstream of an internal plug received within said core engine exhaust nozzle and said downstream most point of said core engine exhaust nozzle being defined at a distance (L n ) from the exit point,
wherein a distance/diameter ratio of said distance (L n ) to said diameter (D t ) is greater than or equal to 0.90; and
the bypass ratio is greater than 6.

2. The propulsion system as set forth in claim 1 , wherein said distance/diameter ratio is greater than or equal to 1.02.

3. The propulsion system as set forth in claim 2 , wherein said distance/diameter ratio is greater than or equal to 1.17.

4. The propulsion system as set forth in claim 2 , wherein an exhaust case is positioned between said fan drive turbine and said core engine exhaust nozzle.

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

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

7. The propulsion system as set forth in claim 6 , wherein said core engine exhaust nozzle includes an inner periphery adjacent to a bypass flow path and disposed about said internal plug, said inner periphery tapering inwardly at a second angle to define a trailing edge, and a maximum value of said second angle is less than 17 degrees.

8. The propulsion system as set forth in claim 7 , wherein the bypass ratio is greater than 10, and a gear ratio of said gear system is greater than or equal to 2.3.

9. The propulsion system as set forth in claim 8 , wherein said high pressure turbine is a two stage high pressure turbine.

10. The propulsion system as set forth in claim 9 , wherein said fan section has only a single fan stage comprising said fan.

11. The propulsion system as set forth in claim 10 , wherein said ratio of said distance (L n ) to said diameter (D t ) is less than or equal to 1.29.

12. The propulsion system as set forth in claim 8 , wherein said ratio of said distance (L n ) to said diameter (D t ) is greater than or equal to 1.20.

13. The propulsion system as set forth in claim 12 , wherein said second angle is greater than 12 degrees.

14. The propulsion system as set forth in claim 13 , wherein said high pressure turbine is a two stage high pressure turbine.

15. The propulsion system as set forth in claim 14 , wherein said fan section has only a single fan stage comprising said fan.

16. The propulsion system as set forth in claim 15 , wherein said ratio of said distance (L n ) to said diameter (D t ) is less than or equal to 1.29.

17. A propulsion system comprising:
a fan section including a fan;
a gear system;
a turbine section incorporating a high pressure turbine and a fan drive turbine that drives said gear system that drives said fan, said turbine section having an exit point at a last blade airfoil stage, and a diameter (D t ) defined as the diameter of the last blade airfoil stage in the turbine section at said exit point;
a compressor section including at least two compressor rotors with said high pressure turbine driving a high pressure compressor rotor and there being a lower pressure compressor rotor also being driven;
a nacelle surrounding a core engine housing, said fan delivers air into a bypass duct defined between said nacelle and said core engine housing to said low pressure compressor, and a bypass ratio defined as a volume of air delivered into said bypass duct compared to the volume of air delivered into 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 downstream of an internal plug received within said core engine exhaust nozzle and said downstream most point being defined at a distance (L n ) from the exit point,
wherein a distance/diameter ratio of said distance (L n ) to said diameter (D t ) is greater than or equal to 0.90; and
a bypass ratio is greater than 10.

18. The propulsion system as set forth in claim 17 , wherein said distance/diameter ratio is greater than or equal to 1.02.

19. The propulsion system as set forth in claim 18 , wherein said distance/diameter ratio is greater than or equal to 1.17.

20. The propulsion system as set forth in claim 18 , wherein an exhaust case is positioned between said fan drive turbine and said core engine exhaust nozzle.

21. The propulsion system as set forth in claim 17 , wherein an exhaust case is positioned between said exit of said fan drive turbine and an entrance to said engine exhaust nozzle.

22. The propulsion system as set forth in claim 17 , wherein a gear ratio of said gear system is greater than or equal to 2.3.

23. The propulsion system as set forth in claim 22 , wherein an exhaust case is positioned between said fan drive turbine and said core engine exhaust nozzle.

24. The propulsion system as set forth in claim 23 , wherein said nacelle includes an outer surface downstream of a maximum diameter of said nacelle, said outer surface tapering inwardly at a first angle to define a trailing edge, and a maximum value of said first angle is less than or equal to 14 degrees.

25. The propulsion system as set forth in claim 24 , wherein said core engine exhaust nozzle includes an inner periphery adjacent to a bypass flow path and disposed about said internal plug, said inner periphery tapering inwardly at a second angle to define a trailing edge, and a maximum value of said second angle is less than 17 degrees.

26. The propulsion system as set forth in claim 25 , wherein said high pressure turbine is a two stage high pressure turbine.

27. The propulsion system as set forth in claim 26 , wherein said fan section has only a single fan stage comprising said fan.

28. The propulsion system as set forth in claim 27 , wherein said ratio of said distance (L n ) to said diameter (D t ) is greater than or equal to 1.20.

29. The propulsion system as set forth in claim 28 , further comprising a mid-turbine frame arranged between

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M
• Malecki Robert E

Classifications

• F02K3/06

  with front fan · CPC title

• F02K3/068

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

• 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

• B64D33/02

  of combustion air intakes · CPC title

• F02C3/10

  with another turbine driving an output shaft but not driving the compressor · CPC title