Bleed air systems for use with aircrafts and related methods
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-009/18
F02C-006/08
F02C-007/32
F02C-003/13
F02C-006/04
B64D-013/06
출원번호
US-0632322
(2015-02-26)
등록번호
US-9765700
(2017-09-19)
발명자
/ 주소
Mackin, Steve G.
Ludlow, George A.
출원인 / 주소
The Boeing Company
대리인 / 주소
Hanley Flight & Zimmerman, LLC
인용정보
피인용 횟수 :
2인용 특허 :
37
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
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.
대표청구항▼
1. An apparatus comprising: a turbo-compressor including: a compressor having a compressor inlet fluidly coupled, via a first bleed air port, 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 inl
1. An apparatus comprising: a turbo-compressor including: a compressor having a compressor inlet fluidly coupled, via a first bleed air port, 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, via a second bleed air port, to a high-pressure turbine of the aircraft engine and a turbine outlet fluidly coupled to a second system of the aircraft. 2. The apparatus of claim 1, wherein the second system is a low-pressure turbine of the aircraft engine. 3. The apparatus of claim 2, wherein the turbine outlet is fluidly coupled to a casing of the low-pressure turbine of the aircraft engine to provide cooled bleed air to the casing and blades within the casing. 4. The apparatus of claim 1, wherein the first system includes at least one of a thermal anti-icing system or an environmental control system. 5. The apparatus of claim 1, wherein the compressor of the turbo-compressor is to increase a pressure of bleed air received at the compressor inlet to a higher pressure at the compressor outlet. 6. 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 first stage of a high-pressure turbine of the aircraft engine and to a second stage of the high-pressure turbine of the aircraft engine, the first stage to provide higher pressure bleed air than the second stage, and the turbine having a turbine outlet fluidly coupled to a second system of the aircraft. 7. The apparatus of claim 1, wherein the turbine inlet is fluidly coupled to a high-pressure compressor of the aircraft engine, the high-pressure turbine of the aircraft engine is to provide higher pressure bleed air than the high-pressure compressor of the aircraft engine. 8. An aircraft comprising: a turbo-compressor comprising: a compressor and a turbine, the compressor having a compressor inlet and a compressor outlet, and the turbine having a turbine inlet and a turbine outlet;a first passageway to fluidly couple a first bleed air port from a compressor of an engine of the aircraft to the compressor inlet, the first passageway to provide bleed air at a first pressure from the compressor of the engine to the compressor inlet;a second passageway to fluidly couple a second bleed air port from a turbine of the engine to the turbine inlet, the second passageway to provide bleed air at a second pressure from the turbine of the engine to the turbine inlet, the second pressure higher than the first pressure; anda third passageway to fluidly couple the compressor outlet to a system of the aircraft that is to receive compressed air. 9. The aircraft of claim 8, wherein the turbine of the engine is a high-pressure turbine. 10. The aircraft of claim 9, wherein the system is a first system, the turbo-compressor further comprising a fourth passageway to fluidly couple the turbine outlet to a second system of the aircraft for cooling the second system. 11. The aircraft of claim 10, wherein the fourth passageway is to provide cooled bleed air to a casing of a low-pressure turbine of the engine. 12. The aircraft of claim 10 further comprising at least one of a thermal anti-icing system or an environmental control system, wherein the third passageway is to fluidly couple the compressor outlet to at least one of the thermal anti-icing system or the environmental control system. 13. The aircraft of claim 8, wherein the compressor is to increase a pressure of the bleed air received at the compressor inlet to a third pressure at the compressor outlet, the third pressure higher than the first pressure. 14. A method comprising: receiving, at a compressor inlet of a turbo-compressor, a first bleed air from a first bleed air port of a compressor of an aircraft engine, the turbo-compressor comprising a turbine operatively coupled to a compressor;receiving, at a turbine inlet of the turbo-compressor, a second bleed air from a second bleed air port of a high-pressure turbine of the aircraft engine to drive the turbine of the turbo-compressor;increasing a pressure of the first bleed air in the compressor of the turbo-compressor; andafter increasing the pressure of the first bleed air, discharging the first bleed air through a compressor outlet of the turbo-compressor to a system of the aircraft. 15. The method of claim 14 further comprising discharging cooled bleed air through a turbine outlet of the turbo-compressor to a low-pressure turbine of the aircraft engine. 16. The method of claim 14 further comprising receiving, at the turbine inlet, a third bleed air having lower pressure than a pressure of the second bleed air. 17. The method of claim 14 further comprising receiving the second bleed air from the high-pressure turbine during a first operating state of the aircraft engine and receiving, at the turbine inlet, a third bleed air from a high-pressure compressor of the aircraft engine during a second operating state of the aircraft engine. 18. The method of claim 14 further comprising receiving the second bleed air from a first stage of the high-pressure turbine during a first operating state of the aircraft engine and receiving, at the turbine inlet, a third bleed air from a second stage of the high-pressure turbine during a second operating state of the aircraft engine, the second stage located downstream of the first stage. 19. The method of claim 14, wherein receiving the first bleed air comprises receiving bleed air from a low pressure compressor of the aircraft engine during a first period of operation, the method further comprising receiving, at the compressor inlet, bleed air from a high pressure compressor of the aircraft engine during a second period of operation.
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