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公开(公告)号：US20240052782A1
公开(公告)日：2024-02-15
申请号：US18144576
申请日：2023-05-08
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Carroll V. Sidwell
CPC classification number: F02C7/224 , F02C3/22 , F02C7/232 , F02C7/26 , F02C7/262 , F02C7/32 , F02C9/40 , B64F1/28
Abstract: Aircraft hydrogen fuel systems and methods and systems of starting such systems are described. The aircraft hydrogen fuel systems include a hydrogen burning main engine, a main tank configured to contain liquid hydrogen to be supplied to the main engine during a normal operation, and a starter tank configured to contain gaseous hydrogen to be used during a startup operation of the main engine. Methods and processes for starting and/or restarting such systems are described.
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公开(公告)号：US11781478B2
公开(公告)日：2023-10-10
申请号：US17571720
申请日：2022-01-10
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Joseph B. Staubach , Jesse M. Chandler , Neil Terwilliger , Gabriel L. Suciu
Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.
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公开(公告)号：US11754000B2
公开(公告)日：2023-09-12
申请号：US17379308
申请日：2021-07-19
Applicant: Raytheon Technologies Corporation
Inventor： Stephen G. Pixton , Matthew R. Feulner , Marc J. Muldoon , Xinwen Xiao
IPC: F02C7/36
CPC classification number: F02C7/36 , F05D2220/36 , F05D2260/40311
Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.
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公开(公告)号：US11746664B2
公开(公告)日：2023-09-05
申请号：US17483210
申请日：2021-09-23
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Michael E. McCune , Keith B. Allyn
CPC classification number: F01D9/041 , F01D25/162 , F01D25/28 , F05D2220/32
Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.
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公开(公告)号：US11578657B2
公开(公告)日：2023-02-14
申请号：US17081627
申请日：2020-10-27
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Marc J. Muldoon , Jonathan M. Rheaume
Abstract: The cooling system may comprise: an electric machine; a first conduit including a cable housing and an inlet; a plurality of conductive cables extending from the electric machine, the plurality of conductive cables disposed at least partially in the cable housing; and an electric fan disposed in the first conduit, the cooling system configured to passively flow air through the first conduit to cool the plurality of conductive cables during operation of the gas turbine engine, and the electric fan configured to actively cool the plurality of conductive cables after an engine shutdown of the gas turbine engine.
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Patent Agency Ranking
公开(公告)号：US20230010158A1
公开(公告)日：2023-01-12
申请号：US17860742
申请日：2022-07-08
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Joseph B. Staubach , Charles E. Lents
Abstract: A turbine engine system includes aircraft systems including at least one hydrogen fuel tank, engine systems comprising a compressor section, a combustor section having a burner, and a turbine section, and a hydrogen fuel flow supply line configured to supply hydrogen fuel from the at least one hydrogen fuel tank into the burner for combustion. The turbine engine system has a bypass ratio between 5 to 20.
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公开(公告)号：US20220412272A1
公开(公告)日：2022-12-29
申请号：US17360480
申请日：2021-06-28
Applicant: Raytheon Technologies Corporation
Inventor： John R. Farris , Marc J. Muldoon
Abstract: A hybrid electric engine including a gas turbine engine including a low speed spool, a high speed spool a fan section, a compressor section, a combustor section, and a turbine section. The hybrid electric engine further includes an electric generator configured to convert rotational power of the high or low speed spool to electricity and a variable area turbine control system electrically connected to the electric generator. The variable area turbine control system being configured to adjust a cross-sectional area of a core flow path of the hybrid electric engine. The variable area turbine control system including a plurality of variable turbine vanes located in the turbine section and a variable area turbine actuator configured to rotate each of the plurality of variable turbine vanes to adjust the cross-sectional area of the core flow path of the hybrid electric engine. The variable area turbine actuator is an electromechanical actuator.
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公开(公告)号：US20220364513A1
公开(公告)日：2022-11-17
申请号：US17321052
申请日：2021-05-14
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Joseph B. Staubach , Charles E. Lents
IPC: F02C7/224
Abstract: An aircraft propulsion systems and aircraft having the same are described. The aircraft propulsion systems have one or more aircraft systems including at least one hydrogen tank and a first heat exchanger and one or more engine systems including at least a main engine core, a second heat exchanger, and a third heat exchanger. The main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is configured to be supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the first heat exchanger of the aircraft systems, the second heat exchanger of the engine systems, and the third heat exchanger of the engine systems, and then supplied into the burner for combustion.
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公开(公告)号：US12180911B2
公开(公告)日：2024-12-31
申请号：US18137887
申请日：2023-04-21
Applicant: Raytheon Technologies Corporation
Inventor： Murat Yazici , Marc J. Muldoon
Abstract: A gas turbine engine is provided that includes fan, compressor, combustor, and turbine sections, an outer casing, an outside core annular region, and a fan air circulation system. The outer casing is disposed radially outside of the compressor, combustor, and turbine sections. The engine has a core gas path and a fan bypass air duct. The core gas path extends through the compressor, combustor, and turbine sections, and is disposed radially inside of the outer casing. The fan bypass air duct is defined by inner and outer radial flow path boundaries. The outside core annular region is disposed radially between the outer casing and the inner radial boundary flow path boundary. The fan air circulation system is configured to receive fan bypass air from the fan bypass air duct and selectively pass the received fan bypass air into the outside core annular region.
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公开(公告)号：US12152535B2
公开(公告)日：2024-11-26
申请号：US17967345
申请日：2022-10-17
Applicant: Raytheon Technologies Corporation
Inventor： Marc J. Muldoon , Russell B. Witlicki
Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating structure, a stationary structure and an electric machine. The rotating structure is configured to rotate about a rotational axis. The stationary structure circumscribes the rotating structure. The electric machine includes a rotor and a stator. The rotor circumscribes the rotating structure and is coupled to the rotating structure through a spline connection. The stator is connected to the stationary structure.
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