Fundamental gear system architecture

• US 10,221,770 B2
• Filed: 02/26/2014
• Issued: 03/05/2019
• Est. Priority Date: 05/31/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Sheridan, William G.
• Primary Examiner(s): Lee, Jr., Woody A
• Application Number:
  US14/190,159
• Publication Number:
  US 20140178180A1
• Time in Patent Office: 1,833 Days
• Field of Search: 415112, 4151241, 4151242, 415175, 60 3908, 60805
• CPC Class Codes: F01D 25/18 Lubricating arrangements of machines or engines in general F01M
• 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 7/36 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: F05D 2260/4031 as in toothed gearing
• 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 pressure ratio greater than about 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 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;

an input shaft arranged in a driving relationship to the fan drive turbine;

a gear support system including a flexible support structure supporting at least one gear of the gear system relative to the input shaft;

a lubrication system providing lubricant to rotating components of the engine including the gear system and removing thermal energy from the lubricant, 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 in response to an interruption in lubricant flow from the first lubrication system; and

a low corrected fan tip speed is 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.

2. 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 (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 pressure ratio greater than about 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 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 lubrication system providing lubricant to rotating components of the engine including the gear system and removing thermal energy from the lubricant, 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 in response to an interruption in lubricant flow from the first lubrication system; and

a second turbine is in communication with the combustor, wherein the second turbine is a two stage turbine, and the fan drive turbine includes less than six turbine rotors.

3. The gas turbine engine recited in claim 2, wherein a ratio between the number of fan blades and the number of fan drive turbine rotors is between 3.3 and 8.6.

4. The gas turbine engine recited in claim 3, wherein the plurality of fan blades is less than twenty (20).

5. The gas turbine engine recited in claim 4, 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.

6. The gas turbine engine recited in claim 5, wherein the gear system comprises a sun gear, a rotatable carrier, a plurality of planet gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.

7. 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 functioning as an inlet guide vane for the fan drive turbine.

8. The gas turbine engine recited in claim 7, further comprising a gear support system including a spring rate that allows a defined amount of deflection and misalignment among at least some of the plurality of gears of the gear system.

9. The gas turbine engine recited in claim 8, wherein the gear support system includes a mount flexibly supporting the gear system relative to a static structure of the engine.

10. The gas turbine engine recited in claim 9, 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 supporting at least one gear of the gear system relative to the input shaft.

11. The gas turbine engine recited in claim 10, wherein the mount extends radially from a static structure of the engine with respect to a central axis to accommodate radial movement between the gear system and the static structure.

12. 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 (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 pressure ratio greater than about 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 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 lubrication system providing lubricant to rotating components of the engine including the gear system and removing thermal energy from the lubricant, 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 in response to an interruption in lubricant flow from the first lubrication system; and

a second turbine is in communication with the combustor, wherein the second turbine is a two stage turbine and the fan drive turbine includes less than six 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.

13. The gas turbine engine recited in claim 12, wherein the plurality of fan blades is less than twenty (20).

14. The gas turbine engine recited in claim 13, 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.

15. The gas turbine engine recited in claim 14, wherein the gear system comprises a sun gear, a rotatable carrier, a plurality of planet gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.

16. 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 (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 pressure ratio greater than about 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 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 lubrication system providing lubricant to rotating components of the engine including the gear system and removing thermal energy from the lubricant, 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 in response to an interruption in lubricant flow from the first lubrication system; and

a second turbine is in communication with the combustor, wherein the second turbine is a two stage turbine, and the fan drive turbine includes less than six (6) fan drive turbine rotors.

17. The gas turbine engine recited in claim 16, wherein a ratio between the number of fan blades and the number of fan drive turbine rotors is between 3.3 and 8.6.

18. The gas turbine engine recited in claim 17, wherein the gear system comprises a sun gear, a rotatable carrier, a plurality of planet gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of planet gears, and wherein the plurality of fan blades is less than twenty (20).

19. The gas turbine engine recited in claim 18, 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.

20. The gas turbine engine recited in claim 19, 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.

21. The gas turbine engine recited in claim 19, wherein the gear system comprises a sun gear in driving connection to the fan drive turbine, a non-rotatable carrier, a plurality of intermediate gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of intermediate gears, and wherein the plurality of fan blades is less than twenty (20).