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    Turbo engine having blade pitch actuation system for variable pitch fan
    92.
    发明授权
    Turbo engine having blade pitch actuation system for variable pitch fan 有权

    公开(公告)号：US12060812B1

    公开(公告)日：2024-08-13

    申请号：US18184785

    申请日：2023-03-16

    Applicant: General Electric Company

    Inventor： James H. Grooms ,  Keith A. Miedema ,  Randy M. Vondrell

    IPC: F01D7/00

    CPC classification number: F01D7/00 ,  F05D2220/323 ,  F05D2240/30 ,  F05D2260/70

    Abstract: An example turbo engine includes a rotary hub and a propeller including a blade coupled to the rotary hub. The blade is rotatable about a radial axis to change a pitch of the blade. The turbo engine also includes a blade pitch actuation system including an actuator coupled to the blade. The actuator is moveable in a first direction to rotate the blade to a feather position and in a second direction to rotate the blade away from the feather position. The blade pitch actuation system also includes a hydraulic system to provide hydraulic fluid to one or more hydraulic chambers to move the actuator and a pressurized pneumatic chamber formed by a portion of the actuator. The pressurized pneumatic chamber biases the actuator in the first direction to move the blade to the feather position in an event of failure of the hydraulic system.

    GAS TURBINE ENGINE ACCESSORY PACKAGING
    93.
    发明公开
    GAS TURBINE ENGINE ACCESSORY PACKAGING 审中-公开

    公开(公告)号：US20240133345A1

    公开(公告)日：2024-04-25

    申请号：US17972720

    申请日：2022-10-24

    Applicant: General Electric Company

    Inventor： Daniel John Oehrle ,  Randy M. Vondrell

    IPC: F02C7/32 ,  B64D27/12

    CPC classification number: F02C7/32 ,  B64D27/12

    Abstract: A gas turbine engine defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.

    GAS TURBINE ENGINE WITH THIRD STREAM
    95.
    发明公开
    GAS TURBINE ENGINE WITH THIRD STREAM 审中-公开

    公开(公告)号：US20240044304A1

    公开(公告)日：2024-02-08

    申请号：US17879384

    申请日：2022-08-02

    Applicant: General Electric Company

    Inventor： Brandon Wayne Miller ,  Randy M. Vondrell ,  David Marion Ostdiek ,  Craig William Higgins ,  Alexander Kimberley Simpson

    IPC: F02K3/065 ,  F02C3/06

    CPC classification number: F02K3/065 ,  F02C3/06

    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

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