[특허]Gear system architecture for gas turbine engine

United Technologies Corporation

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특허 상세정보

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Gear system architecture for gas turbine engine

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F01D-025/18 F02C-007/36 F02C-003/04 F02C-007/06 F02C-007/16 F01D-005/06 F01D-009/04 F04D-029/32 F04D-019/02 F02K-003/06 F16H-001/48 F16H-057/04
출원번호	US-0379619 (2016-12-15)
등록번호	US-9840969 (2017-12-12)
발명자 / 주소	Sheridan, William G.
출원인 / 주소	United Technologies Corporation
대리인 / 주소	Carlson, Gaskey & Olds, P.C.
인용정보	피인용 횟수 : 1 인용 특허 : 174

초록

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.

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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; andwherein 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.

이 특허에 인용된 특허 (174)

이 특허를 인용한 특허 피인용횟수: 1

Sheridan, William G.. Fundamental gear system architecture. USP20190310221770.

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Gear system architecture for gas turbine engine

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Gear system architecture for gas turbine engine

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