Geared architecture for high speed and small volume fan drive turbine

Patent metadata

Field	Value
Publication number	US-9752511-B2
Application number	US-201715484441-A
Country	US
Kind code	B2
Filing date	Apr 11, 2017
Priority date	Jun 8, 2011
Publication date	Sep 5, 2017
Grant date	Sep 5, 2017

Abstract

A gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan rotor. Aspects of the gear system are provided with some flexibility. The fan drive turbine has a first exit area and rotates at a first speed. A second turbine section has a second exit area and rotates at a second speed, which is faster than said first speed. A performance quantity can be defined for both turbine sections as the products of the respective areas and respective speeds squared. A performance quantity ratio of the performance quantity for the fan drive turbine to the performance quantity for the second turbine section is relatively high.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a fan shaft driving a fan having fan blades;
a frame which supports said fan shaft, said frame having a frame lateral stiffness and a frame transverse stiffness;
an epicyclic gear system having a gear reduction ratio of greater than 2.3;
a flexible support which supports said epicyclic gear system and has a flexible support lateral stiffness and a flexible support transverse stiffness, and at least one of said flexible support lateral stiffness and said flexible support transverse stiffness is less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness;
a first turbine section providing a drive input into a sun gear in an epicyclic gear system;
a second turbine section, and
wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed,
wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than said first speed, said first and second speeds being redline speeds,
wherein a first performance quantity is defined as the product of said first speed squared and said first area,
wherein a second performance quantity is defined as the product of said second speed squared and said second area,
wherein a performance quantity ratio of said first performance quantity to said second performance quantity is between 0.5 and 1.5; and
a fan section including said fan is configured to include a low fan pressure ratio of less than 1.45 measured across said fan blades alone.

2. The gas turbine engine as set forth in claim 1 , wherein a mid-turbine frame is positioned between said first and second turbine sections, and includes a plurality of airfoils positioned in a flow path.

3. The gas turbine engine as set forth in claim 2 , wherein said first turbine section includes a first turbine having an inlet, an outlet, and a first turbine pressure ratio greater than 5, wherein said first turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet and prior to any exhaust nozzle.

4. The gas turbine engine as set forth in claim 3 , including a power ratio of a flat-rated Sea Level Take-Off thrust provided by said engine in lbf, to a volume of a turbine section including both said first and second turbine sections in inch 3 being greater than or equal to 4.0 lbf/inch 3 .

5. The gas turbine engine as set forth in claim 4 , wherein said flexible support lateral stiffness and said flexible support transverse stiffness both are less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness.

6. The gas turbine engine as set forth in claim 5 , wherein said fan section has a low corrected fan tip speed less than 1,150 ft/sec, wherein said low corrected fan tip speed is an actual fan tip speed divided by [(Tram ° R)/(518.7° R)] 0.5 .

7. The gas turbine engine as set forth in claim 6 , wherein said performance quantity ratio is less than or equal to 1.075.

8. The gas turbine engine as set forth in claim 3 , including said power ratio being greater than or equal to 1.5 lbf/inch 3 .

9. The gas turbine engine as set forth in claim 8 , wherein said fan section has a low corrected fan tip speed less than 1,150 ft/sec, wherein said low corrected fan tip speed is an actual fan tip speed divided by [(Tram ° R)/(518.7° R)] 0.5 .

10. The gas turbine engine as set forth in claim 3 , wherein said power ratio being greater than or equal to 4.0 and less than or equal to 5.5 lbf/inch 3 .

11. The gas turbine engine as set forth in claim 1 , wherein said fan section has a low corrected fan tip speed less than 1,150 ft/sec, wherein said low corrected fan tip speed is an actual fan tip speed divided by [(Tram ° R)/(518.7° R)] 0.5 .

12. The gas turbine engine as set forth in claim 11 , wherein said performance quantity ratio is less than or equal to 1.075.

13. The gas turbine engine as set forth in claim 1 , wherein said performance quantity ratio is greater than or equal to 0.8.

14. The gas turbine engine as set forth in claim 13 , wherein said flexible support lateral stiffness and said flexible support transverse stiffness both are less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness.

15. The gas turbine engine as set forth in claim 14 , wherein said plurality of gears have a gear mesh transverse stiffness, and said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness.

16. A gas turbine engine comprising:
a fan shaft driving a fan having fan blades and a frame supporting said fan shaft and having a frame lateral stiffness and a frame transverse stiffness;
a plurality of gears which drive said fan shaft;
a flexible support which supports said plurality of gears, and has a flexible support lateral stiffness and a flexible support transverse stiffness and at least one of said flexible support lateral stiffness and said flexible support transverse stiffness is less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness;
a first turbine section providing a drive input into said plurality of gears;
a second turbine section,
wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed,
wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than said first speed, said first and second speeds being redline speeds,
wherein a first performance quantity is defined as the product of said first speed squared and said first area,
wherein a second performance quantity is defined as the product of said second speed squared and said second area,
wherein a performance quantity ratio of said first performance quantity to said second performance quantity is between 0.5 and 1.5; and

a bypass ratio greater than 10.

17. The gas turbine engine as set forth in claim 16 , including a power ratio of a flat-rated Sea Level Take-Off thrust provided by said engine in lbf, to a volume of a turbine section including both said first and second turbine sections in inch 3 being greater than or equal to 1.5 and less than or equal to 5.5 lbf/inch 3 .

18. The gas turbine engine as set forth in claim 16 , wherein both said flexible support lateral stiffness and said flexible support transverse stiffness are less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness.

19. The gas turbine engine set forth in claim 18 , wherein a fan section including said fan is configured to include a low fan pressure ratio of less than 1.45, said low fan pressure ratio measured across said fan blades alone.

20. The gas turbine engine as set forth in claim 19 , wherein said plurality of gears is an epicyclic gear system having a gear reduction ratio of greater than 2.3.

21. The gas turbine engine as set forth in claim 20 , wherein said performance quantity being greater than or equal to 0.8 and said second speed is more than twice said first speed.

22. The gas turbine engine as set forth in claim 20 , wherein said plurality of gears have a gear mesh transverse stiffness, and said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness.

23. The gas turbine engine as set forth in claim 17 , wherein said p

Assignees

• United Technologies Corp

Inventors

• Mccune Michael E
• Husband Jason
• Schwarz Frederick M
• Kupratis Daniel Bernard
• Suciu Gabriel L
• Ackermann William K

Classifications

• F02K3/06 Primary

  with front fan · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F04D29/325

  for axial flow fans (blade mountings F04D29/34, blades F04D29/38) · CPC title

• F05D2260/40311

  of the epicyclical, planetary or differential type · CPC title

• F01D5/06

  Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections {(F01D5/022, F01D5/023 take precedence)} · CPC title