Flexible support structure for a geared architecture gas turbine engine

• US 10,227,893 B2
• Filed: 05/26/2017
• Issued: 03/12/2019
• Est. Priority Date: 06/08/2011
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: McCune, Michael E., Husband, Jason
• Primary Examiner(s): Sosnowski, David E
• Assistant Examiner(s): Flores, Juan G
• Application Number:
  US15/606,494
• Publication Number:
  US 20170335718A1
• Time in Patent Office: 655 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 driving a fan having fan blades. A fan shaft support supports the fan shaft and defines a support transverse stiffness. A gear system is connected to the fan shaft and includes a gear mesh defining a gear mesh transverse stiffness and a reduction ratio greater than 2.3. A gear system input is connected to the gear system and defines a gear system input lateral stiffness. A flexible support supports the gear system and defines a flexible support transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness and the flexible support transverse stiffness is less than 20% of the fan shaft support transverse stiffness.

Claims

1. A gas turbine engine, comprising:

a fan shaft driving a fan having fan blades;

a fan shaft support that supports said fan shaft defining a fan shaft support transverse stiffness;

a gear system connected to said fan shaft, said gear system includes a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness;

a gear system input connected to said gear system, said gear system input defines a gear system input lateral stiffness;

a flexible support supporting said gear system and defining a flexible support transverse stiffness; and

wherein said gear system input lateral stiffness is less than 5% of said gear mesh lateral stiffness.

2. The gas turbine engine of claim 1, wherein said gear system input defines a gear system input transverse stiffness, and said gear system input transverse stiffness is less than 5% of said gear mesh transverse stiffness.

3. The gas turbine engine of claim 2, wherein said gear system input transverse stiffness is less than 20% of said fan shaft support transverse stiffness.

4. The gas turbine engine of claim 3, wherein said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness.

5. The gas turbine engine of claim 1, further comprising a low fan pressure ratio of less than about 1.45 and said low fan pressure ratio measured across the fan blades alone, and wherein said gear system includes a reduction ratio greater than 2.3.

6. The gas turbine engine of claim 5, further comprising a two stage high pressure turbine and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than about 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.

7. The gas turbine engine of claim 6, wherein said gear system input defines a gear system input transverse stiffness and said gear system input transverse stiffness is less than 5% of said gear mesh transverse stiffness, and said flexible support transverse stiffness is less than 20% of said fan shaft support transverse stiffness.

8. The gas turbine engine of claim 7, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T_ram°

R)/(518.7°

R)]^0.5, where T represents said ambient temperature in degrees Rankine.

9. A gas turbine engine, comprising:

a fan shaft driving a fan having fan blades;

a fan shaft support that supports said fan shaft;

a gear system connected to said fan shaft, said gear system includes a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness;

a flexible support supporting said gear system and defining a flexible support lateral stiffness and a flexible support transverse stiffness; and

wherein 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.

10. The gas turbine engine of claim 9, further comprising a low fan pressure ratio of less than about 1.45, said low fan pressure ratio measured across the fan blades alone, and wherein said gear system includes a reduction ratio greater than 2.3.

11. The gas turbine engine of claim 10, further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than about 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.

12. The gas turbine engine of claim 11, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T_ram°

R)/(518.7°

R)]^0.5, where T represents said ambient temperature in degrees Rankine.

13. The gas turbine engine of claim 11, further comprising a bypass ratio greater than ten (10).

14. The gas turbine engine of claim 9, further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than about 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, and wherein said gear system includes a reduction ratio greater than 2.3.

15. The gas turbine engine of claim 9, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T_ram°

R)/(518.7°

R)]^0.5, where T represents said ambient temperature in degrees Rankine.

16. The gas turbine engine of claim 9, further comprising a ring gear defining a ring gear transverse stiffness and said ring gear transverse stiffness is less than 20% of said gear mesh transverse stiffness.

17. The gas turbine engine of claim 16, wherein both 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.

18. The gas turbine engine of claim 17, further comprising a low fan pressure ratio of less than about 1.45, said low fan pressure ratio measured across the fan blades alone, and wherein said gear system includes a reduction ratio greater than 2.3.

19. The gas turbine engine of claim 18, wherein said fan shaft support defines a fan shaft support lateral stiffness and said flexible support lateral stiffness is less than 11% of said fan shaft support lateral stiffness.

20. The gas turbine engine of claim 19, wherein said flexible support defines a flexible support transverse stiffness and said fan shaft support defines a fan shaft support transverse stiffness and said flexible support transverse stiffness is less than 20% of said fan shaft support transverse stiffness.

21. The gas turbine engine of claim 20, wherein said flexible support lateral stiffness is less than 11% of said fan shaft support lateral stiffness and said flexible support transverse stiffness is less than 11% of said fan shaft support transverse stiffness.

22. The gas turbine engine of claim 20, further comprising a gear system input to said gear system, said gear system input defines a gear system input lateral stiffness, wherein said gear system input lateral stiffness is less than 20% of said fan shaft support lateral stiffness.

23. The gas turbine engine of claim 22, wherein said gear system input defines a gear system input transverse stiffness and said gear system input transverse stiffness is less than 20% of said fan shaft support transverse stiffness.

24. The gas turbine engine of claim 23, wherein said gear system input lateral stiffness is less than 11% of said fan shaft support lateral stiffness.

25. The gas turbine engine of claim 24, wherein said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness.

26. The gas turbine engine of claim 23, wherein said gear system input lateral stiffness is less than 5% of said gear mesh lateral stiffness.

27. The gas turbine engine of claim 20, further comprising a gear system input to said gear system defining a gear system input lateral stiffness, wherein said gear system input lateral stiffness is less than 5% of said gear mesh lateral stiffness.

28. The gas turbine engine of claim 27, wherein said gear system input defines a gear system input transverse stiffness and said gear system input transverse stiffness is less than 5% of said gear mesh transverse stiffness and said gear system input both transfers torque and facilitates segregation of vibrations.

29. The gas turbine engine of claim 20, further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than about 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 and a low fan pressure ratio of less than about 1.45 with said low fan pressure ratio measured across the fan blades alone.

30. The gas turbine engine of claim 29, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(T_ram°

R)/(518.7°

R)]^0.5, where T represents said ambient temperature in degrees Rankine.

31. The gas turbine engine of claim 9 wherein said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness.

32. The gas turbine engine of claim 31, further comprising a bypass ratio greater than ten (10) and said gear system includes a reduction ratio greater than 2.3.

33. The gas turbine engine of claim 32, further comprising 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.

34. The gas turbine engine of claim 33, further comprising a two stage high pressure turbine.

35. The gas turbine engine of claim 34, wherein said flexible support is mounted to an engine static structure.

36. The gas turbine engine of claim 32, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/(518.7 °

R)]0.5, where T represents said ambient temperature in degrees Rankine.

37. The gas turbine engine of claim 32, further comprising a gear system input to said gear system defining a gear system input transverse stiffness and said fan shaft support defining a fan shaft support transverse stiffness, wherein said gear system input transverse stiffness is less than 20% of said fan shaft support transverse stiffness.

38. The gas turbine engine of claim 37, wherein said flexible support is mounted to an engine static structure.

39. The gas turbine engine of claim 38, further comprising a mid-turbine frame including at least one airfoil extending into a flow path, and wherein said fan shaft support is a K-frame bearing support.

40. The gas turbine engine of claim 37, wherein said gear mesh defines a gear mesh lateral stiffness and said gear system input defines a gear system input lateral stiffness and said gear system input lateral stiffness is less than 5% of said gear mesh lateral stiffness.

41. The gas turbine engine of claim 40, wherein said fan shaft support defines a fan shaft support lateral stiffness and said gear system input lateral stiffness is less than 11% of said fan shaft support lateral stiffness.

42. The gas turbine engine of claim 37, further comprising a two stage high pressure turbine and a low fan pressure ratio of less than 1.45 and said low fan pressure ratio is measured across the fan blades alone.

43. The gas turbine engine of claim 42, further comprising 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.

44. The gas turbine engine of claim 43, wherein said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness.

45. The gas turbine engine of claim 43, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/(518.7 °

R)]0.5, where T represents said ambient temperature in degrees Rankine.

46. The gas turbine engine of claim 45, wherein said gear system input transverse stiffness is less than 11% of said fan shaft support transverse stiffness, and said gear system input both transfers torque and facilitates segregation of vibrations.

47. The gas turbine engine of claim 31, further comprising a two stage high pressure turbine, a low fan pressure ratio of less than 1.45 and said low fan pressure ratio is measured across the 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 and said gear system includes a reduction ratio greater than 2.3.

48. The gas turbine engine of claim 47, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein said low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/(518.7 °

R)]0.5, where T represents said ambient temperature in degrees Rankine, and wherein said flexible support is mounted to an engine static structure.

49. The gas turbine engine of claim 48, further comprising an input coupling to said gear system defining an input coupling lateral stiffness, wherein said fan shaft support defines a fan shaft support lateral stiffness and said input coupling lateral stiffness is less than 11% of said fan shaft support lateral stiffness.

50. The gas turbine engine of claim 9, further comprising a ring gear defines a ring gear lateral stiffness, wherein said gear mesh defines a gear mesh lateral stiffness, and said flexible support defines a flexible support lateral stiffness and said flexible support lateral stiffness is less than 8% of said gear mesh lateral stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness.