IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0016436
(2013-09-03)
|
등록번호 |
US-8714913
(2014-05-06)
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발명자
/ 주소 |
- Topol, David A.
- Morin, Bruce L.
|
출원인 / 주소 |
- United Technologies Corporation
|
대리인 / 주소 |
Carlson, Gaskey & Olds, PC
|
인용정보 |
피인용 횟수 :
1 인용 특허 :
18 |
초록
▼
A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compres
A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: (number of blades×rotational speed)/60 s≧5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.
대표청구항
▼
1. A gas turbine engine comprising: a fan and a turbine section having a fan drive turbine rotor, said fan drive turbine rotor driving a compressor rotor;a gear reduction effecting a reduction in the speed of said fan relative to an input speed from said fan drive turbine rotor that drives said comp
1. A gas turbine engine comprising: a fan and a turbine section having a fan drive turbine rotor, said fan drive turbine rotor driving a compressor rotor;a gear reduction effecting a reduction in the speed of said fan relative to an input speed from said fan drive turbine rotor that drives said compressor rotor;said compressor rotor having a number of compressor blades in at least one of a plurality of rows of said compressor rotor, and said blades operating at least some of the time at a rotational speed, and said number of compressor blades in said at least one row and said rotational speed being such that the following formula holds true for said at least one row of the compressor rotor (said number of blades×said rotational speed)/60 s≧5500 Hz; andsaid rotational speed being in revolutions per minute. 2. The gas turbine engine as set forth in claim 1, wherein the formula results in a number greater than or equal to about 6000 Hz. 3. The gas turbine engine as set forth in claim 2, wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 4. The gas turbine engine as set forth in claim 1, wherein the formula holds true for the majority of the blade rows of the compressor rotor. 5. The gas turbine engine as set forth in claim 1, wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 6. The gas turbine engine as set forth in claim 1, wherein said gear reduction has a gear ratio of greater than about 2.3. 7. The gas turbine engine as set forth in claim 6, wherein said gear reduction has a gear ratio of greater than about 2.5. 8. The gas turbine engine as set forth in claim 1, wherein said fan delivers air into a bypass duct, and a portion of air into said compressor rotor, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor rotor, and said bypass ratio being greater than about 6. 9. The gas turbine engine as set forth in claim 8, wherein said bypass ratio is greater than about 10. 10. The gas turbine engine as set forth in claim 9, wherein the formula results in a number greater than or equal to about 6000 Hz. 11. The gas turbine engine as set forth in claim 1, wherein said rotational speed being an approach speed. 12. The gas turbine engine as set forth in claim 1, wherein said turbine section including a higher pressure turbine rotor and a lower pressure turbine rotor, and said fan drive turbine rotor being said lower pressure turbine rotor. 13. The gas turbine engine as set forth in claim 12, wherein said compressor rotor is a lower pressure compressor rotor, and said higher pressure turbine rotor driving a higher pressure compressor rotor. 14. A method of designing a gas turbine engine comprising the steps of: including a fan drive turbine rotor to drive a compressor rotor and for driving a fan through a gear reduction, and selecting a number of blades in at least one row of the compressor rotor, in combination with a rotational speed of the compressor rotor, such that the following formula holds true for said at least one row of the compressor rotor: (said number of blades×said rotational speed)/60 s≧5500 Hz; andsaid rotational speed being in revolutions per minute. 15. The method as set forth in claim 14, wherein the formula results in a number greater than or equal to about 6000 Hz. 16. The method as set forth in claim 14, wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 17. The method as set forth in claim 14, wherein said rotational speed is an approach speed. 18. The method as set forth in claim 14, wherein the turbine section including a higher pressure turbine rotor and a lower pressure turbine rotor, and said fan drive turbine rotor being said lower pressure turbine rotor. 19. The method as set forth in claim 18, wherein said compressor rotor is a lower pressure compressor rotor, and said higher pressure turbine rotor driving a higher pressure compressor rotor. 20. A compressor module comprising: a compressor rotor having a first blade row that includes a number of blades, the first blade row being capable of rotating at a rotational speed, so that when measuring said rotational speed in revolutions per minute: (said number of blades×said rotational speed)/60 s≧5500 Hz. 21. The compressor module as set forth in claim 20, wherein said rotational speed is an approach speed. 22. The compressor module as set forth in claim 20, wherein the formula results in a number greater than or equal to about 6000 Hz.
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