[특허]Geared turbofan gas turbine engine architecture

United Technologies Corporation

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특허 상세정보

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Geared turbofan gas turbine engine architecture

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02K-003/06
미국특허분류(USC)	060/226.1; 060/268; 060/039.162; 060/792; 060/039.163; 416/124
출원번호	US-0645606 (2012-10-05)
등록번호	US-8935913 (2015-01-20)
발명자 / 주소	Kupratis, Daniel Bernard Schwarz, Frederick M.
출원인 / 주소	United Technologies Corporation
대리인 / 주소	Carlson, Gaskey & Olds, P.C.
인용정보	피인용 횟수 : 0 인용 특허 : 32

초록
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A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

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1. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, the turbine section including a fan drive turbine and a second turbine, wherein the second turbine is disposed forward of the fan drive turbine and the fan drive turbine includes a plurality of fan drive turbine stages with a ratio between the number of fan blades and the number of fan drive turbine stages is greater than about 2.5; anda speed change system driven by the fan drive turbine for rotating the fan about the axis;wherein the fan drive turbine has a first exit area and is configured to rotate at a first speed, the second turbine section has a second exit area and is configured to rotate at a second speed, which is faster than the first speed,wherein the turbine section includes a volume defined within an inner periphery and an outer periphery between a leading edge of a most upstream vane to a trailing edge of a most downstream rotating airfoil and is configured to provide a thrust density greater than 1.5 lbf/in3 and less than or equal to 5.5 lbf/in3 at Sea Level Takeoff Thrust. 2. The engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan and the fan drive turbine both rotate in a first direction about the axis and the second turbine section rotates in a second direction opposite the first direction. 3. The engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan, the fan drive turbine, and the second turbine section all rotate in a first direction about the axis. 4. The engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan and the second turbine both rotate in a first direction about the axis and the fan drive turbine rotates in a second direction opposite the first direction. 5. The engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan is rotatable in a first direction and the fan drive turbine, and the second turbine section rotate in a second direction opposite the first direction about the axis. 6. The engine as recited in claim 1, wherein the speed change system comprises a gear reduction having a gear ratio greater than 2.3. 7. The engine as set forth in claim 1, wherein said fan delivers a portion of air into a bypass duct, and a 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, with the bypass ratio being greater than 6.0. 8. The engine as set forth in claim 7, wherein the bypass ratio is greater than 10.0. 9. The engine as set forth in claim 1, wherein a fan pressure ratio across the fan is less than 1.5. 10. The engine as set forth in claim 1, wherein said fan has 26 or fewer blades. 11. The engine as set forth in claim 10, wherein said fan drive turbine section has up to 6 stages. 12. The engine as set forth in claim 1, wherein the ratio between the number of fan blades and the number of fan drive turbine stages is less than 8.5. 13. The engine as set forth in claim 1, wherein a pressure ratio across the fan drive turbine is greater than about 5:1. 14. The engine as recited in claim 1, wherein the fan drive turbine includes a first aft rotor attached to a first shaft, the second turbine includes a second aft rotor attached to a second shaft, and a first bearing assembly and a second bearing assembly are disposed aft of the combustor, wherein the first bearing assembly is disposed axially aft of a first connection between the first aft rotor and the first shaft, and the second bearing assembly is disposed axially aft of a second connection between the second aft rotor and the second shaft. 15. The engine as recited in claim 1, wherein the fan drive turbine includes a first aft rotor attached to a first shaft, the second turbine includes a second aft rotor attached to a second shaft, and a first bearing assembly and a second bearing assembly are disposed aft of the combustor, wherein the first bearing assembly is disposed axially aft of a first connection between the first aft rotor and the first shaft, and the second bearing assembly is disposed axially forward of a second connection between the second aft rotor and the second shaft. 16. The engine as recited in claim 1, wherein the fan drive turbine includes a first aft rotor attached to a first shaft, the second turbine includes a second aft rotor attached to a second shaft, and a first bearing assembly and a second bearing assembly are disposed aft of the combustor, wherein the first bearing assembly is disposed axially aft of a first connection between the first aft rotor and the first shaft, and the second bearing assembly is disposed within the annular space defined between the first shaft and the second shaft. 17. The engine as recited in claim 1, wherein the fan drive turbine includes a first aft rotor attached to a first shaft, the second turbine includes a second aft rotor attached to a second shaft, and a first bearing assembly and a second bearing assembly are disposed aft of the combustor, wherein the first bearing assembly is disposed axially forward of a first connection between the first aft rotor and the first shaft, and the second bearing assembly is disposed axially aft of a second connection between the second aft rotor and the second shaft.

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Geared turbofan gas turbine engine architecture

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Geared turbofan gas turbine engine architecture

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이 특허에 인용된 특허 (32)

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