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    Combustor swirler
    12.
    发明授权
    Combustor swirler 有权

    公开(公告)号：US11226101B2

    公开(公告)日：2022-01-18

    申请号：US16264775

    申请日：2019-02-01

    Applicant: General Electric Company

    Inventor： Karthikeyan Sampath ,  Saket Singh ,  Gurunath Gandikota ,  Perumallu Vukanti ,  Pradeep Naik ,  Steven Clayton Vise

    IPC: F23R3/14 ,  F23R3/28

    Abstract: A gas turbine engine swirler that includes a tubular body having a forward face, an aft end, and a throat. A plurality of primary swirl vanes that is positioned between the aft end and the forward face. A plurality of secondary swirl vanes that is positioned between the primary swirl vanes and the aft end. The plurality of primary swirl vanes and the plurality of secondary swirl vanes are configured such that the throat is fluidly connected to a plenum that is positioned outside of the tubular body. A tubular ferrule is positioned such that it joins the body at the forward face thereof. Each of the primary swirl vanes extend radially inwardly to a vane lip. The secondary swirl vanes extend radially inwardly for swirling air therefrom. The body also includes a tubular Venturi that extends aft from between the primary swirler vanes and the secondary swirler vanes for radially separating air swirled therefrom. Wherein the primary swirl vanes are configured to swirl air along a passageway and through an outlet that is oriented axially aft.

    COMBUSTOR LINER FOR A GAS TURBINE ENGINE AND AN ASSOCIATED METHOD THEREOF
    15.
    发明申请
    COMBUSTOR LINER FOR A GAS TURBINE ENGINE AND AN ASSOCIATED METHOD THEREOF 有权

    公开(公告)号：US20210003281A1

    公开(公告)日：2021-01-07

    申请号：US17029587

    申请日：2020-09-23

    Applicant: GENERAL ELECTRIC COMPANY

    Inventor： Mayank Krisna Amble ,  Perumallu Vukanti ,  Steven Clayton Vise ,  Michael Anthony Benjamin

    IPC: F23R3/06 ,  F23R3/00 ,  F23R3/26 ,  F02K1/82 ,  F02C7/18

    Abstract: A method for regulating jet wakes in a combustor including: directing compressed fluid into a passageway between a casing and a combustor liner; directing a combustion gas along a combustion zone; discharging a first portion of compressed fluid from the passageway into the combustion zone via a first through-hole which is disposed on a section of a combustor panel in a first circumferential row; and discharging a second portion of the compressed fluid from the passageway into the combustion zone via second through-holes, wherein the second through-holes are disposed on a section of the panel, spaced apart axially and circumferentially, and adjacent to the first through-hole, wherein the second through-holes include a first set of through-holes and a second set of through-holes, the first set of through-holes in a second row and the second set of through-holes in a third row, wherein the second and third rows extend circumferentially.

    Combustion Section Heat Transfer System for a Propulsion System
    16.
    发明申请
    Combustion Section Heat Transfer System for a Propulsion System 审中-公开

    公开(公告)号：US20200088101A1

    公开(公告)日：2020-03-19

    申请号：US16675787

    申请日：2019-11-06

    Applicant: General Electric Company

    Inventor： Joseph Zelina ,  Sibtosh Pal ,  Arthur Wesley Johnson ,  Clayton Stuart Cooper ,  Steven Clayton Vise

    IPC: F02C7/18 ,  F02C3/14 ,  F23R3/00 ,  F02C7/10 ,  F01D1/24 ,  F01D9/02

    Abstract: The present disclosure is directed to a propulsion system including a wall defining a combustion chamber inlet, a combustion chamber outlet, and a combustion chamber therebetween, a nozzle assembly disposed at the combustion chamber inlet, the nozzle assembly configured to provide a fuel/oxidizer mixture to the combustion chamber, a turbine nozzle coupled to the wall and positioned at the combustion chamber outlet, wherein the turbine nozzle defines a cooling circuit within the turbine nozzle, and a casing positioned radially adjacent to the wall, wherein a channel structure is positioned between the casing and the wall, the channel structure in fluid communication with the cooling circuit within the turbine nozzle, and wherein a flowpath is formed between the wall and the casing, the flowpath in fluid communication from the cooling circuit at the turbine nozzle to the nozzle assembly to provide a flow of oxidizer to the thereto.

    MULTIPLE CHAMBER ROTATING DETONATION COMBUSTOR
    18.
    发明申请
    MULTIPLE CHAMBER ROTATING DETONATION COMBUSTOR 审中-公开

    公开(公告)号：US20180355822A1

    公开(公告)日：2018-12-13

    申请号：US15618289

    申请日：2017-06-09

    Applicant: General Electric Company

    Inventor： Steven Clayton Vise ,  Joseph Zelina ,  Arthur Wesley Johnson ,  Clayton Stuart Cooper ,  Sibtosh Pal

    IPC: F02K7/00

    CPC classification number: F02K7/005

    Abstract: The present disclosure is directed to a rotating detonation combustion system for a propulsion system including a plurality of combustors in adjacent arrangement along the circumferential direction. Each combustor defines a combustor centerline extended through each combustor, and each combustor comprises an outer wall defining a combustion chamber and a combustion inlet. Each combustion chamber is defined by an annular gap and a combustion chamber length together defining a volume of each combustion chamber. Each combustor defines a plurality of nozzle assemblies each disposed at the combustion inlet in adjacent arrangement around each combustor centerline. Each nozzle assembly defines a nozzle wall extended along a lengthwise direction, a nozzle inlet, a nozzle outlet, and a throat therebetween, and each nozzle assembly defines a converging-diverging nozzle. A first array of combustors defines a first volume and a second array of combustors defines a second volume different from the first volume.

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