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    Multiple chamber rotating detonation combustor
    33.
    发明授权
    Multiple chamber rotating detonation combustor 有权

    公开(公告)号：US11674476B2

    公开(公告)日：2023-06-13

    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: F02K7/06 ,  F02C5/12 ,  F02K7/08 ,  F02C5/11 ,  F23R7/00 ,  F02K7/10

    CPC classification number: F02K7/06 ,  F02C5/11 ,  F02C5/12 ,  F02K7/08 ,  F23R7/00 ,  F02K7/10 ,  F05D2220/80

    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.

    ENGINE WITH ROTATING DETONATION COMBUSTION SYSTEM
    34.
    发明申请
    ENGINE WITH ROTATING DETONATION COMBUSTION SYSTEM 有权

    公开(公告)号：US20230020803A1

    公开(公告)日：2023-01-19

    申请号：US17935646

    申请日：2022-09-27

    Applicant: General Electric Company

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

    IPC: F23R7/00 ,  F02K7/08 ,  F02K7/10 ,  F02K7/16

    Abstract: A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

    Combustion section heat transfer system for a propulsion system
    36.
    发明授权
    Combustion section heat transfer system for a propulsion system 有权

    公开(公告)号：US11143106B2

    公开(公告)日：2021-10-12

    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: F23R7/00 ,  F02C7/18 ,  F01D9/02 ,  F02C3/14 ,  F23R3/00 ,  F02C7/10 ,  F01D1/24 ,  F23R3/28 ,  F01D25/12 ,  F23R3/50

    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.

    Turbine Engine With Rotating Detonation Combustion System
    37.
    发明申请
    Turbine Engine With Rotating Detonation Combustion System 审中-公开

    公开(公告)号：US20190271268A1

    公开(公告)日：2019-09-05

    申请号：US15909196

    申请日：2018-03-01

    Applicant: General Electric Company

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

    IPC: F02C7/22 ,  F23R3/28 ,  F23R3/10 ,  F01D9/02 ,  F02C3/08

    Abstract: A turbine engine including a compressor rotor and a rotating detonation combustion (RDC) system. The compressor rotor includes a compressor airfoil defining a trailing edge disposed within a core flowpath of the turbine engine. The core flowpath defines a radial distance between an outer radius and an inner radius at the compressor rotor. The RDC system includes an outer wall and an inner wall each extended along a lengthwise direction and defining a detonation chamber therebetween. The RDC system further includes a strut defining a nozzle assembly and a fuel injection opening providing a flow of fuel to the detonation chamber. The compressor rotor provides a flow of oxidizer in direct fluid communication to the nozzle assembly of the RDC system.

    Combustion Section Heat Transfer System for a Propulsion System
    38.
    发明申请
    Combustion Section Heat Transfer System for a Propulsion System 审中-公开

    公开(公告)号：US20180363555A1

    公开(公告)日：2018-12-20

    申请号：US15623773

    申请日：2017-06-15

    Applicant: General Electric Company

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

    IPC: F02C7/18 ,  F23R3/28 ,  F01D9/02

    Abstract: The present disclosure is directed to a propulsion system including an annular inner wall and an annular outer wall, a nozzle assembly, a turbine nozzle, and an inner casing and an outer casing. The inner wall and outer wall together extend at least partially along a longitudinal direction and together define a combustion chamber inlet, a combustion chamber outlet, and a combustion chamber therebetween. The nozzle assembly is disposed at the combustion inlet and provides a mixture of fuel and oxidizer to the combustion chamber. The turbine nozzle defines a plurality of airfoils in adjacent circumferential arrangement disposed at the combustion chamber outlet. The turbine nozzle is coupled to the outer wall and the inner wall. The inner casing is disposed inward of the inner wall and the outer casing is disposed outward of the outer wall. Each of the inner casing and the outer casing are coupled to the turbine nozzle. A primary flowpath is defined between the inner casing and the inner wall, through the turbine nozzle, and between the outer casing and the outer wall, and in fluid communication with the combustion chamber.

    VARIABLE GEOMETRY ROTATING DETONATION COMBUSTOR
    39.
    发明申请
    VARIABLE GEOMETRY ROTATING DETONATION COMBUSTOR 审中-公开

    公开(公告)号：US20180356094A1

    公开(公告)日：2018-12-13

    申请号：US15618431

    申请日：2017-06-09

    Applicant: General Electric Company

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

    IPC: F23R3/26 ,  F23R3/00

    CPC classification number: F23R3/26 ,  F23R3/002

    Abstract: The present disclosure is directed to a method of operating a propulsion system at an approximately constant detonation cell quantity in the combustion chamber of a detonation combustion system. The propulsion system defines an inlet section upstream of the rotating detonation combustion system and an exhaust section downstream of the rotating detonation combustion system. The method includes providing an outer wall and an inner wall together defining an annular gap and a combustion chamber length extended from a combustion chamber inlet proximate to the fuel-oxidizer mixing nozzle to a combustion chamber exit proximate to the exhaust section of the propulsion system, the annular gap and the combustion chamber length together defining a first volume at a first operating condition defining a lowest steady state pressure and temperature at the rotating detonation combustion system; providing a mixture of a fuel and an oxidizer to the combustion chamber via the fuel-oxidizer mixing nozzle; detonating the fuel and oxidizer mixture in the combustion chamber, wherein the detonation produces a detonation cell size; and adjusting the volume of the combustion chamber via articulating one or more of the outer wall, the inner wall, and the fuel-oxidizer mixing nozzle such that one or more of the annular gap and the combustion chamber length is changed based on one or more operating conditions.

    METHODS OF OPERATING A ROTATING DETONATION COMBUSTOR AT APPROXIMATELY CONSTANT DETONATION CELL SIZE
    40.
    发明申请
    METHODS OF OPERATING A ROTATING DETONATION COMBUSTOR AT APPROXIMATELY CONSTANT DETONATION CELL SIZE 审中-公开

    公开(公告)号：US20180356093A1

    公开(公告)日：2018-12-13

    申请号：US15618380

    申请日：2017-06-09

    Applicant: General Electric Company

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

    IPC: F23R3/04 ,  F02C5/12

    CPC classification number: F23R3/04 ,  F02C5/12 ,  F05D2220/32 ,  F05D2240/35

    Abstract: The present disclosure is directed to a method of operating a propulsion system including a rotating detonation combustion (RDC) system. The RDC system defines a combustion inlet at an upstream end, a combustion outlet at a downstream end, a combustion chamber therebetween, and a nozzle defined at the combustion inlet upstream of the combustion chamber, and a secondary flowpath extended from upstream of the nozzle to downstream of the nozzle. The method includes providing the combustion chamber of the rotating detonation combustion system to produce a detonation cell size configured for a first operating condition defining a lowest steady state operating condition of the propulsion system; generating a flow of oxidizer to the combustion inlet of the combustion section; providing a first portion of the flow of oxidizer to the combustion chamber and mixing the first portion of the flow of oxidizer with a fuel; providing a second portion of the flow of oxidizer to the secondary flowpath, wherein the secondary flowpath bypasses the combustion chamber; and adjusting a ratio of the first portion of the flow of oxidizer through the combustion chamber versus the second portion of the flow of oxidizer through the secondary flowpath based at least on a commanded power output of the propulsion system.

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