[특허]Bleed air systems for use with aircrafts and related methods

The Boeing Company

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Bleed air systems for use with aircrafts and related methods

국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	B64B-001/24
미국특허분류(USC)	244/053R; 244/134R
출원번호	US-0357293 (2012-01-24)
등록번호	US-8967528 (2015-03-03)
발명자 / 주소	Mackin, Steve G. Ludlow, George A.
출원인 / 주소	The Boeing Company
대리인 / 주소	Hanley, Flight & Zimmerman, LLC
인용정보	피인용 횟수 : 12 인용 특허 : 29

초록

Bleed air systems for use with aircrafts and related methods are disclosed. An example apparatus includes a turbo-compressor including a compressor having a compressor inlet fluidly coupled to a low-pressure compressor of the aircraft engine and a compressor outlet fluidly coupled to a first system of an aircraft. The turbo-compressor also includes a turbine inlet fluidly coupled to a high-pressure compressor of the aircraft engine and a turbine outlet fluidly coupled to a second system of the aircraft.

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1. An apparatus comprising: a turbo-compressor including: a compressor having a compressor inlet fluidly coupled to a low-pressure compressor of an aircraft engine and a compressor outlet fluidly coupled to a first system of an aircraft; anda turbine having a turbine inlet fluidly coupled to a high-pressure compressor of the aircraft engine and a turbine outlet fluidly coupled to a second system of the aircraft, wherein the turbo-compressor is to increase a pressure of bleed air received from the low-pressure compressor, and wherein the high-pressure compressor has a higher pressure than the low-pressure compressor. 2. An apparatus as defined in claim 1, wherein the first system is an environmental control system of the aircraft or a wing anti-icing system. 3. An apparatus as defined in claim 1, wherein the second system is an engine anti-icing system. 4. An apparatus as defined in claim 1, further comprising a mix-flow bypass system to fluidly couple the turbine outlet and the compressor outlet, the mix-flow bypass system to mix bleed air from the turbine outlet and bleed air from the compressor outlet. 5. An apparatus as defined in claim 1, further comprising an anti-icing boost bypass system to fluidly couple the high-pressure compressor and the second system, the anti-icing boost bypass system to enable bleed air to bypass the turbine and flow directly to the second system. 6. An apparatus as defined in claim 5, wherein the anti-icing boost bypass system comprises an ejector conduit to fluidly couple bleed air from the high-pressure compressor of the aircraft engine and the second system. 7. An apparatus as defined in claim 1, wherein the turbo-compressor is disposed within a nacelle of the aircraft. 8. An apparatus as defined in claim 1, further comprising an intercooler disposed between the compressor inlet and the low-pressure compressor of an aircraft engine. 9. An apparatus comprising: a turbo-compressor comprising a compressor and a turbine;a first passageway to fluidly couple a low-pressure bleed air port from a low-pressure compressor of an aircraft engine to a compressor inlet of the turbo-compressor, the low-pressure bleed air port to provide bleed air at a first pressure from the low-pressure compressor to the compressor inlet of the turbo-compressor;a second passageway to fluidly couple a high-pressure bleed air port from a high-pressure compressor of the aircraft engine to a turbine inlet of the turbo-compressor, wherein the high-pressure compressor has a higher pressure than the low-pressure compressor; anda third passageway to fluidly couple a compressor outlet of the turbo-compressor to a system of the aircraft, the compressor to increase the pressure of the bleed air received at the compressor inlet to a second pressure at the compressor outlet, the second pressure being higher than the first pressure. 10. An apparatus as defined in claim 9, further comprising a fourth passageway to fluidly couple a turbine outlet of the turbine to an anti-icing system. 11. An apparatus as defined in claim 10, wherein the anti-icing system comprises an engine anti-icing system. 12. An apparatus as defined in claim 10, further comprising a fifth passageway to fluidly couple the turbine outlet and the compressor outlet. 13. An apparatus as defined in claim 12, further comprising a valve disposed between the turbine outlet, the fourth passageway, and the fifth passageway, the valve to enable bleed air to flow from the turbine outlet to the compressor outlet via the fifth passageway based on a comparison of a pressure of the bleed air at the turbine outlet to a pressure of the bleed air at the compressor outlet. 14. An apparatus as defined in claim 10, further comprising a fifth passageway to bypass the turbine to fluidly couple the high-pressure bleed air port and the anti-icing system. 15. An apparatus as defined in claim 14, further comprising a valve disposed between the high-pressure bleed air port and the fifth passageway, the valve to allow fluid flow from the high-pressure bleed air port to the fifth passageway based on a temperature of the bleed air at the turbine outlet. 16. An apparatus as defined in claim 9, further comprising a fourth passageway to fluidly couple the compressor outlet and an anti-icing system. 17. An apparatus as defined in claim 16, wherein the anti-icing system comprises a wing anti-icing system. 18. An apparatus as defined in claim 9, wherein the turbo-compressor, via the third passageway, is to directly provide bleed air from the low-pressure bleed air port to the system of the aircraft. 19. A method comprising: fluidly coupling a compressor inlet of a turbo-compressor to a low-pressure bleed air source provided by a low-pressure compressor of an aircraft engine, the turbo-compressor comprising a compressor and a turbine different from a compressor and a turbine, respectively, of the aircraft engine;fluidly coupling a compressor outlet of the turbo-compressor to a first system of the aircraft that receives a bleed air supply; andfluidly coupling a turbine inlet of the turbo-compressor to a high-pressure bleed air source provided by a high-pressure compressor of the aircraft engine, the high-pressure bleed air to drive a turbine operatively coupled to the compressor, wherein the turbo-compressor is to increase a pressure of bleed air received from the low-pressure compressor, wherein the high-pressure compressor has a higher pressure than the low-pressure compressor. 20. A method of claim 19, further comprising fluidly coupling a turbine outlet of the turbo-compressor to a second control system of the aircraft. 21. A method of claim 20, further comprising bypassing the turbine and directly coupling the high-pressure bleed air source and the second control system. 22. A method of claim 20, further comprising fluidly coupling the turbine outlet and the compressor outlet to mix bleed air from the turbine outlet and the compressor outlet based on a comparison of a pressure of bleed air at the turbine outlet and bleed air at the compressor outlet. 23. An apparatus as defined in claim 1, further comprising a supplemental bleed air bypass system to fluidly couple bleed air from the high-pressure compressor to the compressor outlet.

이 특허에 인용된 특허 (29)

이 특허를 인용한 특허 피인용횟수: 12

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Bleed air systems for use with aircrafts and related methods

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Bleed air systems for use with aircrafts and related methods

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이 특허에 인용된 특허 (29)

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IPC 상세정보(version : 201201)

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