Flexible support structure for a geared architecture gas turbine engine

• US 10,301,968 B2
• Filed: 09/21/2015
• Issued: 05/28/2019
• Est. Priority Date: 06/08/2011
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: McCune, Michael E., Husband, Jason
• Primary Examiner(s): Sosnowski, David E
• Assistant Examiner(s): Flores, Juan G
• Application Number:
  US14/859,381
• Publication Number:
  US 20160053635A1
• Time in Patent Office: 1,345 Days
• Field of Search: None
• CPC Class Codes: F01D 15/12 Combinations with mechanical gearing driven by multiple engines F01D13/00
• CPC Class Codes: F01D 25/16 Arrangement of bearings; Supporting or mounting bearings in casings bearings per se F16C
• CPC Class Codes: F01D 25/164 Flexible supports; Vibration damping means associated with the bearing
• CPC Class Codes: F01D 25/28 Supporting or mounting arrangements, e.g. for turbine casing
• CPC Class Codes: F01D 5/06 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 Class Codes: F01D 9/02 Nozzles; Nozzle boxes; Stator blades; Guide conduits , e.g. individual nozzles nozzle boxes F01D9/047
• CPC Class Codes: F02C 7/32 Arrangement, mounting, or driving, of auxiliaries
• CPC Class Codes: F02C 7/36 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 Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F04D 19/02 Multi-stage pumps
• CPC Class Codes: F04D 25/045 the pump wheel carrying the fluid driving means, e.g. turbine blades
• CPC Class Codes: F04D 29/053 Shafts
• CPC Class Codes: F04D 29/056 Bearings
• CPC Class Codes: F04D 29/321 for axial flow compressors
• CPC Class Codes: F04D 29/325 for axial flow fans blade mountings F04D29/34, blades F04D29/38
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2240/60 Shafts
• CPC Class Codes: F05D 2260/40311 of the epicyclical, planetary or differential type
• CPC Class Codes: F05D 2260/96 Preventing, counteracting or reducing vibration or noise
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft
• CPC Class Codes: Y10T 29/49321 : Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.

Claims

1. A gas turbine engine comprising:

a fan shaft driving a fan having fan blades;

a frame which supports said fan shaft having a frame lateral stiffness and a frame transverse stiffness;

a plurality of gears in driving engagement with said fan shaft; and

a flexible support which supports said plurality of gears, said flexible support having at least one of a flexible support lateral stiffness and a flexible support transverse stiffness that is less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness.

2. The gas turbine engine as set forth in claim 1, wherein said plurality of gears includes a gear mesh, said gear mesh having a gear mesh lateral stiffness and a gear mesh transverse stiffness, and at least one of said flexible support lateral stiffness and said flexible support transverse stiffness is less than 8% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.

3. The gas turbine engine as set forth in claim 2, including a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:

1, wherein said low pressure 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 gas turbine engine as set forth in claim 3, including a ring gear having a ring gear lateral stiffness and a ring gear transverse stiffness and at least one of said ring gear lateral stiffness and said ring gear transverse stiffness is less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.

5. The gas turbine engine as set forth in claim 4, wherein both said flexible support transverse stiffness and said flexible support lateral stiffness are less than 11% of a respective one of said frame transverse stiffness and said frame lateral 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 both said ring gear lateral stiffness and said ring gear transverse stiffness are less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.

8. The gas turbine engine as set forth in claim 1, including a mid-turbine frame positioned between a first turbine section and a second turbine section having a plurality of airfoils positioned in a flow path.

9. The gas turbine engine as set forth in claim 1, wherein said plurality of gears includes a gear mesh, said gear mesh having a gear mesh lateral stiffness and a gear mesh transverse stiffness, and said flexible support lateral stiffness and said flexible support transverse stiffness are less than 8% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.

10. The gas turbine engine as set forth in claim 9, including a ring gear having a ring gear lateral stiffness and a ring gear transverse stiffness, at least one of said ring gear lateral stiffness and said ring gear transverse stiffness being less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.

11. The gas turbine engine as set forth in claim 10, 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 10, further comprising a bypass ratio greater than 10.

13. The gas turbine engine as set forth in claim 12, including a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:

1, wherein said low pressure 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.

14. The gas turbine engine as set forth in claim 13, wherein a flexible coupling connects a sun gear of the plurality of gears in driving engagement with a first turbine section, and said flexible coupling having at least one of a flexible coupling transverse stiffness and a flexible coupling lateral stiffness that is less than 5% of a respective one of said gear mesh transverse stiffness and said gear mesh lateral stiffness.

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

16. The gas turbine engine of claim 15, further comprising a two stage high pressure turbine.

17. A gas turbine engine, comprising:

a fan having fan blades;

a fan shaft drivingly connected to said fan;

a gear system connected to said fan shaft in driving engagement with a first turbine section;

a gear system flex mount arrangement, wherein said gear system flex mount arrangement accommodates misalignment of said fan shaft during operation and includes a frame for supporting said fan shaft defining a frame lateral stiffness and a frame transverse stiffness; and

a flexible support which supports said gear system, said flexible support having both a flexible support lateral stiffness and a flexible support transverse stiffness that is less than 11% of a respective one of said frame lateral stiffness and said frame transverse stiffness.

18. The gas turbine engine as set forth in claim 17, further comprising a fan pressure ratio of less than 1.45 measured across said fan blades alone and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:

1, wherein said low pressure 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 gas turbine engine as set forth in claim 18, further comprising a bypass ratio greater than 10 and wherein said gear system includes a gear reduction ratio of greater than 2.3.

20. The gas turbine engine as set forth in claim 19, 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.

21. The gas turbine engine of claim 19, wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear transverse stiffness and a gear mesh defining a gear mesh transverse stiffness and said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness.

22. The gas turbine engine of claim 21, wherein said gear mesh defines a gear mesh lateral stiffness and said ring gear defines a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness.

23. The gas turbine engine of claim 17, wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness and a gear mesh defining a gear mesh lateral stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness.

24. The gas turbine engine of claim 23, wherein said ring gear define a ring gear transverse stiffness and said gear mesh defines a gear mesh transverse stiffness and said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness.

25. The gas turbine engine as set forth in claim 24, further comprising a bypass ratio greater than 10.

26. The gas turbine engine as set forth in claim 25, wherein said gear system includes a gear reduction ratio of greater than 2.3.

27. The gas turbine engine as set forth in claim 26, further including a fan pressure ratio of less than 1.45 measured across said fan blades alone.

28. The gas turbine engine as set forth in claim 27, further including a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:

1, wherein said low pressure 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.

29. The gas turbine engine as set forth in claim 28, 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.