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    AIRCRAFT COMBUSTION SYSTEMS
    41.
    发明公开
    AIRCRAFT COMBUSTION SYSTEMS 审中-公开

    公开(公告)号：US20240209792A1

    公开(公告)日：2024-06-27

    申请号：US18211809

    申请日：2023-06-20

    Applicant: ROLLS-ROYCE PLC

    Inventor： Christopher P. MADDEN ,  Craig W. BEMMENT ,  Andrew T. SMITH ,  Peter SWANN

    IPC: F02C7/264 ,  B64D27/10 ,  F23R3/28

    CPC classification number: F02C7/264 ,  B64D27/10 ,  F23R3/28

    Abstract: A method of operating a gas turbine engine having a combustor having a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The fuel spray nozzles have a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes providing a fuel to the plurality of fuel spray nozzles having a calorific value of at least 43.5 MJ/kg. A gas turbine engine can be for an aircraft.

    FUELLING SCHEDULE
    44.
    发明公开
    FUELLING SCHEDULE 审中-公开

    公开(公告)号：US20230193835A1

    公开(公告)日：2023-06-22

    申请号：US17853074

    申请日：2022-06-29

    Applicant: ROLLS-ROYCE PLC

    Inventor： Peter SWANN ,  Craig W. BEMMENT ,  Alastair G. HOBDAY ,  Benjamin J. KEELER ,  Christopher P. MADDEN

    IPC: F02C9/28 ,  G08G5/00 ,  B64D27/10 ,  B64D37/04

    CPC classification number: F02C9/28 ,  B64D27/10 ,  B64D37/04 ,  G08G5/0034

    Abstract: A method of operating an aircraft including a gas turbine engine and a plurality of fuel tanks arranged to provide fuel to the gas turbine engine, where at least two of the fuel tanks contain fuels with different fuel characteristics. The method includes obtaining a flight profile for a flight of the aircraft; and determining a fuelling schedule for the flight based on the flight profile and the fuel characteristics. The fuelling schedule governs the variation with time of how much fuel is drawn from each tank. Fuel input to the gas turbine engine may then be controlled in operation in accordance with the fuelling schedule.

    EFFICIENT JET
    45.
    发明申请
    EFFICIENT JET 有权

    公开(公告)号：US20210071586A1

    公开(公告)日：2021-03-11

    申请号：US17060797

    申请日：2020-10-01

    Applicant: ROLLS-ROYCE plc

    Inventor： Pascal DUNNING ,  Craig W. BEMMENT

    IPC: F02C7/36 ,  F02K3/06

    Abstract: A gas turbine engine includes: an engine core, compressor system, and core shaft. A compressor exit pressure is defined as an average airflow pressure at the exit of the highest pressure compressor at cruise conditions. The core has an annular splitter and bypass flow. Stagnation streamlines around the engine circumference form a streamsurface. A fan is upstream the core with blades having leading and trailing edges, and a radially inner portion within the streamtube. A fan root entry pressure is an average airflow pressure across the radially inner portion leading edge of each fan blade at cruise conditions. An overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure. A bypass jet velocity is defined as the jet velocity of air flow exiting the bypass exhaust nozzle at cruise conditions. A jet velocity is in a range between 4.7 m/s and 7.7 m/s.

    GAS TURBINE ENGINE
    46.
    发明申请
    GAS TURBINE ENGINE 有权

    公开(公告)号：US20210071572A1

    公开(公告)日：2021-03-11

    申请号：US16680668

    申请日：2019-11-12

    Applicant: ROLLS-ROYCE plc

    Inventor： Pascal DUNNING ,  Craig W. BEMMENT

    IPC: F02C3/06 ,  F01D17/10

    Abstract: A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine system comprising one or more turbines (17, 19), a compressor system comprising one or more compressors (14,15), and a core shaft (26) connecting the turbine system to the compressor system, wherein a compressor exit pressure (P30) is defined as an average pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core (11) further comprises an annular splitter (70) at which flow is divided between a core flow (A) that flows through the engine core and a bypass flow (B) that flows along a bypass duct (22), wherein stagnation streamlines (110) around the circumference of the engine (10), stagnating on a leading edge of the annular splitter (70), form a streamsurface (110) forming a radially outer boundary of a streamtube that contains all of the core flow (A); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub, each fan blade (64) having a leading edge (64a) and a trailing edge (64b), each fan blade (64) having a radially inner portion (65a) lying within the streamtube that contains the core flow (A), and wherein a fan root entry pressure (P20) is defined as an average pressure of airflow across the leading edge (64a) of the radially inner portion of each fan blade (64) at cruise conditions; and a nacelle (21) surrounding the engine core (11), the nacelle (21) defining the bypass duct (22) and a bypass exhaust nozzle (18). An overall pressure ratio is defined as the compressor exit pressure (P30) divided by the fan root entry pressure (P20). A bypass nozzle pressure ratio is defined as the nozzle pressure ratio of the bypass exhaust nozzle (18) at cruise conditions. A combined pressure ratio defined as: overall   pressure   ratio bypass   nozzle   pressure   ratio is in a range between 20 and 29. A method of operating a gas turbine engine on an aircraft is also disclosed.

    GAS TURBINE ENGINE EXHAUST
    47.
    发明申请
    GAS TURBINE ENGINE EXHAUST 审中-公开

    公开(公告)号：US20200370512A1

    公开(公告)日：2020-11-26

    申请号：US16526479

    申请日：2019-07-30

    Applicant: ROLLS-ROYCE plc

    Inventor： Craig W. BEMMENT

    IPC: F02K3/075 ,  F02C9/18

    Abstract: A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; and a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle, wherein the gas turbine engine is configured such that a first velocity ratio between an axial exhaust flow velocity from the turbine and a fully expanded axial exhaust flow velocity from the bypass exhaust nozzle is greater than around 0.655 under maximum take-off conditions.

    GAS TURBINE ENGINE CORE ARRANGEMENT
    48.
    发明申请
    GAS TURBINE ENGINE CORE ARRANGEMENT 审中-公开

    公开(公告)号：US20200370435A1

    公开(公告)日：2020-11-26

    申请号：US16513192

    申请日：2019-07-16

    Applicant: ROLLS-ROYCE plc

    Inventor： Craig W. BEMMENT

    IPC: F01D5/14 ,  F02K3/06

    Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan including a plurality of fan blades, wherein the gas turbine engine is configured such that a flow velocity ratio between a first flow velocity at an exit of the engine core and a second flow velocity at an inlet of the engine core is within a range from around 0.82 to around 1.1 at cruise conditions.

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