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公开(公告)号：EP4450797A1
公开(公告)日：2024-10-23
申请号：EP24171748.7
申请日：2024-04-22
申请人： RTX Corporation
发明人： YAZICI, Murat , MULDOON, Marc J.
摘要： A gas turbine engine (20) includes a propulsor delivering air into a core engine inward of an outer core panel (104). The core engine has a compressor section (24), a combustor (56) and a turbine section (28). A compressor bleed system includes a bleed valve (115) on a conduit communicating with compressed air from the compressor section (24). The bleed valve (115) is configured to deliver air into a chamber (116) where it can flow outwardly of the core engine at an exit location radially inward of the outer core panel (104).
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公开(公告)号：EP4421299A3
公开(公告)日：2024-10-23
申请号：EP24155383.3
申请日：2024-02-01
申请人： RTX Corporation , Rohr, Inc.
发明人： MULDOON, Marc J. , YAZICI, Murat , SHERMAN, Brian
摘要： A propulsion system (100) of an aircraft includes a hybrid electric gas turbine engine, and a nacelle (200) at least partially enclosing the hybrid electric gas turbine engine. The nacelle (200) includes a first nacelle half (200a) and a second nacelle half (200b). Each of the first nacelle half (200a) and the second nacelle half (200b) include an outer nacelle sleeve (214), an inner nacelle sleeve (216) radially offset from the outer nacelle sleeve (214) such that a flowpath is defined between the outer nacelle sleeve (214) and the inner nacelle sleeve (216), and an upper bifurcation (202) connecting the outer nacelle sleeve (214) to the inner nacelle sleeve (216) at an upper end of the nacelle (200). The flowpath is circumferentially continuous between the upper bifurcation (202) of the first nacelle half (200a) and the upper bifurcation (202) of the second nacelle half (200b).
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公开(公告)号：EP4417795A1
公开(公告)日：2024-08-21
申请号：EP24158183.4
申请日：2024-02-16
申请人： RTX Corporation
IPC分类号： F02C6/08
CPC分类号： F02C6/08
摘要： A gas turbine engine (20) includes a compressor section (106) having a downstream most location (108) and more upstream locations. A compressor housing (149) surrounds the compressor rotors. A combustor (110) is downstream of the compressor and a turbine section (112) is downstream of the combustor (110). At least one bleed inlet passage bleeds air compressed from the compressor section (106), and delivers the air for use (122) on an associated aircraft (99). A baffle plate (132) is intermediate the at least one bleed inlet passage and a bleed exit port (131) from the gas turbine engine (20). The baffle plate (132) has a plurality of holes (140) for allowing the air to flow from the bleed inlet passage to the bleed exit port (131), with an area of the holes (140) increasing from a location adjacent the at least one bleed inlet passage and in a circumferential direction away from the at least one bleed inlet passage. An aircraft air supply system and an aircraft are also disclosed.
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公开(公告)号：EP4368817A3
公开(公告)日：2024-06-05
申请号：EP24167608.9
申请日：2021-02-02
申请人： RTX Corporation
发明人： MULDOON, Marc J.
CPC分类号： F02C7/00 , F02C7/16 , F02C7/18 , F05D2260/60820130101 , F01D9/065 , F01D25/162 , F05D2220/7620130101 , F01D21/20 , Y02T50/60
摘要： A cooling system (300; 400) for a plurality of conductive cables (310) in a gas turbine engine (110) includes a cooling source and an electric motor (210) disposed in a tail cone (122). The cooling source may comprise an electric fan (360) or an oil pump (460). The cooling source may be configured for active cooling of the plurality of conductive cables (310). The electric fan (360) may be in fluid communication with ambient air during operation of the gas turbine engine (110).
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公开(公告)号：EP4446568A1
公开(公告)日：2024-10-16
申请号：EP24170101.0
申请日：2024-04-12
申请人： RTX Corporation
发明人： SCHAFER, Bradley C. , WINEBRENNER, Brian V. , SOLOMONIDES, Jason B. , MULDOON, Marc J. , VIRTUE, Jr., John P.
摘要： A bowed rotor prevention system (100) includes a gas turbine engine (20), an electric motor (102), and a core turning controller (150). The gas turbine engine (20) includes a rotor that is rotatably coupled to a drive shaft (50). The electric motor (102) is rotatably coupled to a motor shaft (108), which is mechanically coupled to the drive shaft (50) so as to rotate therewith. The core turning controller (150) is in signal communication with the electric motor (102). The core turning controller (150) determines whether a rotor bow is present in the rotor based on an operating time of the gas turbine engine (20), determines a motoring time based on the operating time, and invokes an anti-rotor bowing mode configured to control the electric motor (102) to rotate the rotor according to the motoring time.
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公开(公告)号：EP4421299A2
公开(公告)日：2024-08-28
申请号：EP24155383.3
申请日：2024-02-01
申请人： RTX Corporation , Rohr, Inc.
发明人： MULDOON, Marc J. , YAZICI, Murat , SHERMAN, Brian
摘要： A propulsion system (100) of an aircraft includes a hybrid electric gas turbine engine, and a nacelle (200) at least partially enclosing the hybrid electric gas turbine engine. The nacelle (200) includes a first nacelle half (200a) and a second nacelle half (200b). Each of the first nacelle half (200a) and the second nacelle half (200b) include an outer nacelle sleeve (214), an inner nacelle sleeve (216) radially offset from the outer nacelle sleeve (214) such that a flowpath is defined between the outer nacelle sleeve (214) and the inner nacelle sleeve (216), and an upper bifurcation (202) connecting the outer nacelle sleeve (214) to the inner nacelle sleeve (216) at an upper end of the nacelle (200). The flowpath is circumferentially continuous between the upper bifurcation (202) of the first nacelle half (200a) and the upper bifurcation (202) of the second nacelle half (200b).
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公开(公告)号：EP4345252A1
公开(公告)日：2024-04-03
申请号：EP23201006.6
申请日：2023-09-29
申请人： RTX Corporation
发明人： CLARK, Thomas E. , MULDOON, Marc J.
摘要： A gas turbine engine includes a primary flow path (230) fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath (240) is disposed radially outward of the primary flowpath (230). A first seal (210) spans from an inner diameter of the cooling air flowpath (240) to an outer diameter of the cooling air flowpath (240). The first seal (210) includes at least one axial convolution (212) and a plurality of pass through features (214) defining a purge airflow (244). A heat shield (220) is positioned immediately downstream of the first seal (210) and is configured in relation to the first seal (210) such that the purge airflow (244) enters a mixing plenum (256) defined between the heat shield (220) and the first seal (210).
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公开(公告)号：EP4303418A1
公开(公告)日：2024-01-10
申请号：EP23184291.5
申请日：2023-07-07
申请人： RTX Corporation
摘要： An aircraft propulsion systems (400) has an aircraft systems (404) including at least one hydrogen tank (432) and an aircraft-systems heat exchanger (458) and engine systems (402) includes at least a main engine core, a high pressure pump (416), a hydrogen-air heat exchanger (424), and a turbo expander (420). The main engine core includes a compressor section, a combustor section having a burner (410), and a turbine section arranged along an engine shaft. Hydrogen is supplied from the at least one hydrogen tank (432) through a hydrogen flow path (444), passing through the aircraft-systems heat exchanger (458), the high pressure pump (416), the hydrogen-air heat exchanger (424), and the turbo expander (420), prior to being injected into the burner (410) for combustion. The turbo expander (420) is configured to impart power to the engine shaft.
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公开(公告)号：EP4450790A1
公开(公告)日：2024-10-23
申请号：EP24171461.7
申请日：2024-04-19
申请人： RTX Corporation
发明人： YAZICI, Murat , MULDOON, Marc J.
摘要： A gas turbine engine (20) includes fan, compressor, combustor, and turbine sections (28, 30, 32, 34), an outer casing (43), an outside core annular region (64) and a fan air circulation system (52). A fan bypass air duct (60) is defined by inner and outer radial flow path boundaries (62, 58). The outside core annular region (64) is disposed radially between the outer casing (43) and the inner radial flow path boundary (62). The fan air circulation system (52) has a plurality of inlet ports (72), valves (74), and exit ports (76) and is configured such that a respective inlet port (72) is in fluid communication with a respective valve (74), and the respective valve (74) is in fluid communication with a respective exit port (76). Each valve (74) is selectively controllable to control fan bypass air flow therethrough and into the outside core annular region (64).
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公开(公告)号：EP4446567A1
公开(公告)日：2024-10-16
申请号：EP24170097.0
申请日：2024-04-12
申请人： RTX Corporation
发明人： MULDOON, Marc J. , SCHAFER, Bradley C. , VIRTUE, Jr., John P. , WINEBRENNER, Brian V. , SOLOMONIDES, Jason B.
摘要： A bowed rotor prevention system (100) includes a gas turbine engine (20), an electric motor (102), and a core turning controller (150) in signal communication with the electric motor (102). The gas turbine engine (20) includes a rotor that is rotatably coupled to a drive shaft (50). The electric motor (102) is rotatably coupled to a motor shaft (108), which is mechanically coupled to the drive shaft (50) so as to rotate therewith. The core turning controller (150) is configured to invoke an anti-rotor bowing mode, and to control the electric motor (102) to periodically rotate the rotor into a plurality of rotor positions during a given time period in response to invoking the anti-rotor bowing mode.
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