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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：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：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：EP4450788A1
公开(公告)日：2024-10-23
申请号：EP24171443.5
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 (68), and a fan air circulation system (52). The outer casing (43) is disposed radially outside of the compressor, combustor, and turbine sections (28, 30, 32, 34). A core gas path extends through the compressor, combustor, and turbine sections (28, 30, 32, 34), and is disposed radially inside of the outer casing (43). A fan bypass air duct (62) is defined by inner and outer radial flow path boundaries (66, 60). The outside core annular region (68) is disposed radially between the outer casing (43) and the inner radial flow path boundary (66). The fan air circulation system (52) is configured to receive fan bypass air from the fan bypass air duct (62) and selectively pass the received fan bypass air into the outside core annular region (68).
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