Gear system architecture for gas turbine engine

• US 9,840,969 B2
• Filed: 12/15/2016
• Issued: 12/12/2017
• Est. Priority Date: 05/31/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Sheridan, William G.
• Primary Examiner(s): LEE JR, WOODY A
• Application Number:
  US15/379,619
• Publication Number:
  US 20170096943A1
• Time in Patent Office: 362 Days
• Field of Search: 60 3908, 60805, 415112, 4151241, 4151242, 415175, 184 611, 184 612, 1841041
• CPC Class Codes: F01D 25/162 Bearing supports
• CPC Class Codes: F01D 25/18 Lubricating arrangements of machines or engines in general F01M
• CPC Class Codes: F01D 25/20 using lubrication pumps
• 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 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F02C 3/04 having a turbine driving a compressor power transmission arrangements F02C7/36; control of working fluid flow F02C9/16
• CPC Class Codes: F02C 3/06 the compressor comprising only axial stages F02C3/10 takes precedence
• CPC Class Codes: F02C 3/062 the turbine being of the radial-flow type
• CPC Class Codes: F02C 3/107 with two or more rotors connected by power transmission
• CPC Class Codes: F02C 7/06 Arrangements of bearings bearings F16C; Lubricating of turbo machines F01D25/18; of machines or engines in general F01M
• CPC Class Codes: F02C 7/16 characterised by cooling medium
• 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 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 2220/36 specially adapted for the fan of turbofan engines
• CPC Class Codes: F05D 2240/35 Combustors or associated equipment
• CPC Class Codes: F05D 2260/4031 as in toothed gearing
• CPC Class Codes: F05D 2260/40311 : of the epicyclical, planetary or differential type
• CPC Class Codes: F05D 2260/98 : Lubrication
• CPC Class Codes: F16H 1/28 : with gears having orbital motion
• CPC Class Codes: F16H 1/48 : Special means compensating for misalignment of axes , e.g. for equalising distribution of load on the face width of the teeth in combination with distribution of load on the planet-wheels F16H1/2809
• CPC Class Codes: F16H 57/0412 : Cooling or heating; Control of temperature
• CPC Class Codes: F16H 57/0486 : with fixed gear ratio differentials F16H57/0483
• CPC Class Codes: Y02T 50/60 : Efficient propulsion technologies, e.g. for aircraft

Abstract

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system.

Claims

1. A gas turbine engine comprising:

a fan including a plurality of fan blades rotatable about an axis;

a fan pressure ratio across a fan blade alone of less than 1.45;

a bypass duct;

a compressor section;

a bypass ratio greater than ten (10), the bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section;

a combustor in fluid communication with the compressor section;

a fan drive turbine in communication with the combustor, the fan drive turbine comprising a plurality of fan drive turbine rotors and a pressure ratio greater than about five (5), wherein the fan drive turbine further includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the fan drive turbine is a ratio of the inlet pressure to the outlet pressure;

a gear system including a plurality of gears providing a speed reduction between the fan drive turbine and the fan and transferring power input from the fan drive turbine to the fan at an efficiency greater than 98%, wherein the gear system further comprises a gear reduction ratio of greater than 2.3;

a gear support system including a spring rate that accommodates deflections of the gear system;

a lubrication system providing lubricant to the gear system and removing thermal energy from the gear system, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 2% of power input into the gear system by the fan drive turbine during operation of the engine; and

wherein the plurality of fan blades is less than 26 and the plurality of fan drive turbine rotors is less than 6 fan drive turbine rotors, and a ratio between the number of fan blades and the number of fan drive turbine rotors is between 3.3 and 8.6.

2. The gas turbine engine recited in claim 1, wherein the plurality of fan blades is less than twenty (20).

3. The gas turbine engine recited in claim 2, further comprising a second turbine in communication with the combustor, wherein the second turbine is a two stage turbine, and wherein the gear system comprises a sun gear, a rotatable carrier, a plurality of planet gears supported on the carrier driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.

4. The gas turbine engine recited in claim 3, further comprising a mid-turbine frame between the fan drive turbine and the second turbine, the mid-turbine frame including at least one vane in a core airflow path functioning as an inlet guide vane for the fan drive turbine.

5. The gas turbine engine recited in claim 3, further comprising an input shaft arranged in a driving relationship to the fan drive turbine, wherein the gear support system includes a flexible support structure connecting the gear system and the input shaft.

6. The gas turbine engine recited in claim 5, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/518.7)^0.5], where T represents the ambient temperature in degrees Rankine.

7. The gas turbine engine recited in claim 6, further comprising a load limiter operative to constrain movement of the gear system during an unbalanced load condition.

8. The gas turbine engine recited in claim 6 further comprising a mid-turbine frame between the fan drive turbine and the second turbine, the mid-turbine frame including at least one vane in a core airflow path.

9. The gas turbine engine recited in claim 3, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 1% of power input into the gear system by the fan drive turbine during operation of the engine.

10. The gas turbine engine recited in claim 3, wherein the lubrication system comprises a first lubrication system providing lubricant flow to the gear system and at least one other rotating component, and a second lubrication system supplementing operation of the first lubrication system.

11. The gas turbine engine recited in claim 10, wherein the second lubrication system supplies lubricant flow to the gear system in response to an interruption in lubricant flow from the first lubrication system.

12. The gas turbine engine recited in claim 1, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 1% of power input into the gear system by the fan drive turbine during operation of the engine.

13. The gas turbine engine recited in claim 12, wherein the gear system transfers power input from the fan drive turbine to the fan at an efficiency greater than 99% , and the lubrication system provides lubricant to and removes thermal energy from at least one bearing assembly.

14. The gas turbine engine recited in claim 11, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 1% of power input into the gear system by the fan drive turbine during operation of the engine.

15. The gas turbine engine recited in claim 14, further comprising a load limiter including a stop that reacts to unbalanced loads on the gear system, and wherein the at least one other rotating component is a bearing assembly during operation of the engine.

16. The gas turbine engine recited in claim 1, further comprising an input shaft arranged in a driving relationship to the fan drive turbine, a flexible support structure supporting at least one gear of the gear system relative to the input shaft, wherein the flexible support structure includes a spring rate that allows a defined amount of deflection and misalignment of the at least one gear.

17. The gas turbine engine recited in claim 16, further comprising a mid-turbine frame between the fan drive turbine and a second turbine, the mid-turbine frame including at least one vane in a core airflow path functioning as an inlet guide vane for the fan drive turbine.

18. The gas turbine engine recited in claim 16, further comprising a load limiter operative to constrain movement of the gear system during an unbalanced load condition.

19. The gas turbine engine recited in claim 18, wherein the load limiter is operative to constrain radial loads and torsional loads within defined limits.

20. The gas turbine engine recited in claim 19, wherein the load limiter includes a stop.

21. The gas turbine engine recited in claim 16, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 1% of power input into the gear system by the fan drive turbine during operation of the engine.

22. The gas turbine engine recited in claim 21, further comprising a second turbine in communication with the combustor, wherein the second turbine is a two stage turbine and the plurality of fan blades is less than twenty (20).

23. The gas turbine engine recited in claim 22, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/518.7)^0.5], where T represents the ambient temperature in degrees Rankine.

24. The gas turbine engine recited in claim 16, wherein the plurality of fan blades is less than twenty (20).

25. The gas turbine engine recited in claim 1, wherein the lubrication system comprises a first lubrication system providing lubricant flow to the gear system, and a second lubrication system supplementing operation of the first lubrication system.

26. The gas turbine engine recited in claim 25, wherein the second lubrication system supplies lubricant flow to the gear system in response to an interruption in lubricant flow from the first lubrication system.

27. The gas turbine engine recited in claim 25, wherein the lubrication system has a maximum capacity for removing thermal energy equal to no more than about 1% of power input into the gear system by the fan drive turbine during operation of the engine.

28. The gas turbine engine recited in claim 27, further comprising a low corrected fan tip speed less than about 1150 ft/second, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram °

R)/518.7)^0.5], where T represents the ambient temperature in degrees Rankine.

29. The gas turbine engine recited in claim 25, wherein the plurality of fan blades is less than twenty (20).

30. The gas turbine engine recited in claim 1, wherein the lubrication system provides lubricant flow to the gear system and at least one other rotating component.