Bleed valve assembly for a gas turbine engine
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-009/18
F04D-027/02
F02C-007/20
F01D-009/02
F01D-025/24
F02C-003/04
F04D-027/00
F04D-029/52
F02C-006/08
F04D-015/00
F01D-025/16
출원번호
US-0287782
(2016-10-07)
등록번호
US-10233845
(2019-03-19)
발명자
/ 주소
Moniz, Thomas Ory
Rose, Joseph George
Sage, Patrick Sean
Clements, Jeffrey Donald
Glover, Jeffrey
출원인 / 주소
General Electric Company
대리인 / 주소
General Electric Company
인용정보
피인용 횟수 :
0인용 특허 :
12
초록▼
A gas turbine engine includes a first compressor, a casing surrounding the first compressor, and a liner extending forward from the first compressor. The gas turbine engine also includes a core turbine frame assembly extending between the liner and the casing and a bleed air assembly. The bleed air
A gas turbine engine includes a first compressor, a casing surrounding the first compressor, and a liner extending forward from the first compressor. The gas turbine engine also includes a core turbine frame assembly extending between the liner and the casing and a bleed air assembly. The bleed air assembly includes a bleed valve positioned in the liner, and a duct in airflow communication with the bleed valve and defining in outlet. The duct is positioned within the core turbine frame assembly and extends to the casing.
대표청구항▼
1. A gas turbine engine defining an axial direction, the gas turbine engine comprising: a first compressor;a second compressor located downstream of the first compressor:a casing surrounding the first compressor;a liner extending aft from the first compressor to the second compressor, the liner at l
1. A gas turbine engine defining an axial direction, the gas turbine engine comprising: a first compressor;a second compressor located downstream of the first compressor:a casing surrounding the first compressor;a liner extending aft from the first compressor to the second compressor, the liner at least partially defining a core air flowpath from the first compressor to the second compressor;a core turbine frame assembly extending between the liner and the casing, the core turbine frame assembly comprising a forward frame member and an aft frame member, the forward frame member and the aft frame member each being positioned aft of the first compressor and forward of the second compressor and extending from the liner to the casing; anda bleed air assembly comprising; a bleed valve positioned in the liner, the bleed valve comprising a door assembly configured to move between an open position and a closed position, wherein the door assembly comprises a transition passage which pivots with the door assembly as the door assembly moves between the open position and the closed position; anda duct in airflow communication with the bleed valve, the duct including a liner structure positioned within the core turbine frame assembly and between the forward frame member and the aft frame member, the liner structure defining an outlet of the duct, and the duct extending to the casing,wherein the aft frame member intersects the casing at a first position, the forward frame member intersects the casing at a second position, and wherein the outlet is positioned at the casing at a location forward of the first position and aft of the second position, the outlet being configured to exhaust a bleed air flow through the casing at the location of the outlet. 2. The gas turbine engine of claim 1, wherein the transition passage is aligned with the duct when the door assembly is in the open position. 3. The gas turbine engine of claim 2, wherein the transition passage is misaligned with the duct when the door assembly is in the closed position. 4. The gas turbine engine of claim 2, wherein the transition passage is in airflow communication with the core air flowpath and the duct when the door assembly is in the open position. 5. The gas turbine engine of claim 1, wherein the bleed valve is a first bleed valve, wherein the bleed air assembly further comprises a second bleed valve, and wherein the duct is in airflow communication with the first bleed valve and the second bleed valve. 6. The gas turbine engine of claim 1, wherein the liner is an outer flowpath liner, the gas turbine engine further comprises an inner flowpath liner, and the inner flowpath liner and the outer flowpath liner together define a portion of the core air flowpath extending between the first compressor and the second compressor. 7. The gas turbine engine of claim 1, wherein the first compressor, the casing, the liner, and the bleed air assembly are configured as part of an engine core of the gas turbine engine, and wherein the gas turbine engine further comprises: an outer nacelle at least partially surrounding the engine core and defining a bypass passage with the casing, wherein the outlet of the duct is in airflow communication with the bypass passage through the casing. 8. The gas turbine engine of claim 1, wherein the casing defines an opening, wherein the outlet of the duct is positioned at the opening of the casing. 9. The gas turbine engine of claim 7, wherein the bleed air assembly comprises a plate positioned in, or over, the outlet of the duct, and wherein the plate defines a plurality of airflow holes.
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