Bleed air systems for use with aircraft and related methods
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-006/08
F02C-007/047
F02K-003/02
F02C-007/32
F02C-007/36
F02C-009/18
F02C-007/141
B64D-013/06
출원번호
US-0242493
(2014-04-01)
등록번호
US-10054051
(2018-08-21)
발명자
/ 주소
Foutch, David W.
Mackin, Steve G.
Bowman, Michael D.
출원인 / 주소
The Boeing Company
대리인 / 주소
Hanley, Flight & Zimmerman, LLC
인용정보
피인용 횟수 :
0인용 특허 :
47
초록▼
Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor is to be driven by a drive shaft extending from an engine of an aircraft. The example apparatus also includes a firs
Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor is to be driven by a drive shaft extending from an engine of an aircraft. The example apparatus also includes a first passageway to fluidly couple a first low-pressure bleed air port from the engine to the compressor inlet and a second passageway to fluidly couple the compressor outlet to a system of the aircraft.
대표청구항▼
1. An apparatus comprising: a shaft-driven compressor system comprising: a compressor having a compressor inlet and a compressor outlet, the compressor to be driven by a drive shaft extending from an engine of an aircraft;a first passageway to fluidly couple a first bleed air port of a low-pressure
1. An apparatus comprising: a shaft-driven compressor system comprising: a compressor having a compressor inlet and a compressor outlet, the compressor to be driven by a drive shaft extending from an engine of an aircraft;a first passageway to fluidly couple a first bleed air port of a low-pressure compressor of the engine to the compressor inlet;a second passageway to fluidly couple the compressor outlet to a system of the aircraft; anda first flow control member operable between a first position to enable the shaft-driven compressor system to supply bleed air to the system of the aircraft and a second position to prevent the shaft-driven compressor system from supplying bleed air to the system of the aircraft;a two-port bleed air system to fluidly couple a second bleed air port and a third bleed air port of a high-pressure compressor of the engine to the system of the aircraft, the two-port bleed air system including a second flow control member operable between a third position to enable the two-port bleed air system to supply bleed air to the system of the aircraft and a fourth position to prevent the two-port bleed air system from supplying bleed air to the system of the aircraft; anda computer control system configured to:when the aircraft is operating in a first condition, operate the first flow control member to the first position and operate the second flow control member to the fourth position; andwhen the aircraft is operating in a second condition, operate the first flow control member to the second position and operate the second flow control member to the third position. 2. The apparatus of claim 1, wherein the system of the aircraft is at least one of an environmental control system, a wing anti-icing system or an engine anti-icing system. 3. The apparatus of claim 1 further comprising a third passageway to fluidly couple the second bleed air port to the compressor inlet, the second bleed air port to provide higher pressure bleed air than the first bleed air port. 4. The apparatus of claim 1, wherein the compressor is operatively coupled to a gearbox coupled to the drive shaft. 5. The apparatus of claim 1 further comprising a third passageway to fluidly couple a fan of the engine to the compressor inlet. 6. The apparatus of claim 5, wherein the third passageway couples the fan to the first passageway such that the first passageway supplies air from the fan to the compressor inlet via the first passageway. 7. The apparatus of claim 6, further including a check valve coupled to the third passageway between the fan and a junction of the third passageway and the first passageway, the check valve to prevent higher pressure air in the first passageway from flowing through the third passageway to the fan. 8. The apparatus of claim 7, further including a third flow control member coupled to the first passageway between the first bleed air port and the junction of the third passageway and the first passageway, the third flow control member to at least one of regulate bleed air from the first bleed air port to a pre-set pressure value or provide fluid flow shut-off. 9. The apparatus of claim 1, wherein the compressor of the shaft-driven compressor system comprises at least one of a centrifugal compressor, an axial compressor or a mixed-flow compressor. 10. The apparatus of claim 1, wherein the first flow control member is coupled to the second passageway between the compressor outlet and the system of the aircraft, and wherein, in the first position, the first flow control member enables compressed bleed air to flow from the compressor outlet to the system of the aircraft via the second passageway and, in the second position, the first flow control member directs the compressed bleed air in the second passageway from the compressor outlet to at least one of a turbine of the engine, a casing of the engine, or produce thrust for the aircraft. 11. The apparatus of claim 1, wherein the two-port bleed air system is coupled at a junction to the second passageway upstream of the system of the aircraft, further including a check valve coupled to the second passageway between the compressor outlet and the junction to prevent high pressure air from flowing through the second passageway to the compressor outlet when the aircraft is operating in the second condition. 12. The apparatus of claim 1, wherein the first condition occurs when the aircraft is operating at cruise and the second condition occurs when the aircraft is operating at idle or during a descent. 13. An apparatus comprising: a compressor operatively coupled to a drive shaft of an engine of an aircraft, the compressor having an inlet fluidly coupled to a first bleed air port of a low-pressure compressor of the engine and an outlet fluidly coupled via a passageway to a system of the aircraft, a first flow control member coupled to the passageway;a bleed air system to fluidly couple a second bleed air port of a high-pressure compressor of the engine to the system of the aircraft, a second flow control member coupled to the bleed air system; anda computer control system configured to:when the aircraft is operating in a first condition, operate the first flow control member to enable the compressor to provide bleed air to the system of the aircraft and operate the second flow control member to prevent the bleed air system from providing bleed air to the system of the aircraft; andwhen the aircraft is operating in a second condition, operate the first flow control member to prevent the compressor from providing bleed air to the system of the aircraft and operate the second flow control member to enable the bleed air system to provide bleed air to the system of the aircraft. 14. The apparatus of claim 13, wherein the first condition occurs when the aircraft is operating at cruise. 15. The apparatus of claim 14, wherein the second condition occurs when the aircraft is operating at idle or during a descent. 16. The apparatus of claim 13, wherein the bleed air system is a two-port bleed air system. 17. The apparatus of claim 13, wherein the compressor operatively coupled to the drive shaft is further operatively coupled to the engine via bevel gears having a fixed gear ratio. 18. The apparatus of claim 13, wherein the drive shaft is a first drive shaft, and wherein the first drive shaft is operatively coupled to a second drive shaft of the high-pressure compressor of the engine. 19. A method comprising: coupling a compressor to a drive shaft of an engine of an aircraft;fluidly coupling a compressor inlet of the compressor to a first bleed air port of a low-pressure compressor of the engine;fluidly coupling via a passageway a compressor outlet of the compressor to a system of the aircraft that receives bleed air supply, a first flow control member coupled to the passageway;fluidly coupling via a bleed air system a second bleed air port and a third bleed air port of a high-pressure compressor of the engine to the system of the aircraft, a second flow control member coupled to the bleed air system; andoperating the first flow control member and the second flow control member with a computer control system configured to:when the aircraft is operating in a first condition, operate the first flow control member to enable the compressor to provide bleed air to the system of the aircraft and operate the second flow control member to prevent the bleed air system from providing bleed air to the system of the aircraft; andwhen the aircraft is operating in a second condition, operate the first flow control member to prevent the compressor from providing bleed air to the system of the aircraft and operate the second flow control member to enable the bleed air system to provide bleed air to the system of the aircraft. 20. The method of claim 19 further comprising fluidly coupling the compressor inlet to the second bleed air port of the high-pressure compressor of the engine. 21. The method of claim 19 further comprising fluidly coupling the compressor inlet to a fan of the engine.
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