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公开(公告)号：US20220355916A1
公开(公告)日：2022-11-10
申请号：US17504618
申请日：2021-10-19
Applicant: General Electric Company
Inventor： Dhananjaya Rao Gottapu , Kurt David Murrow , Narendra Digamber Joshi
IPC: B64C23/00
Abstract: An aircraft having distributed fans for boundary layer ingestion is provided. In one aspect, an aircraft includes a fuselage extending between a forward end and an aft end. The aircraft includes a plurality of boundary layer ingestion fans arranged in an array. Each fan of the array is mounted to and arranged circumferentially around the aft end of the fuselage. The fans are positioned so as to ingest boundary layer airflow flowing along the fuselage. At least two fans of the array are different sizes. Each fan of the fan array is operatively coupled with an electric machine. The electric machines are operable to drive their respective fans to produce thrust. The fans of the array are independently controlled in accordance with the boundary layer suction requirements of the aircraft.
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公开(公告)号：US11230387B2
公开(公告)日：2022-01-25
申请号：US17021864
申请日：2020-09-15
Applicant: General Electric Company
Inventor： Narendra Digamber Joshi , Konrad Weeber , Michael Solomon Idelchik
Abstract: A motor driven propulsor of an aircraft includes magnets disposed in fan shrouds of fan blades connected with a fan hub, a stator having individual conductive coils in a nacelle located radially outside of the fan hub, and a distributed inverter assembly having several inverter power stages and gate drivers, each of the inverter power stages coupled with a separate gate driver of the gate drivers and a separate coil of the coils in the stator. Each of the gate drivers is configured to individually control supply of direct current to the corresponding inverter power stage. Each of the inverter power stages is configured to convert the direct current supplied to the inverter power stage to an alternating current that is supplied to the corresponding coil in the stator to rotate the magnets and the fan blades around a center line of the fan hub for propelling the aircraft.
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公开(公告)号：US11053844B2
公开(公告)日：2021-07-06
申请号：US16133362
申请日：2018-09-17
Applicant: General Electric Company
IPC: F23R3/14 , F02C3/14 , F23R3/00 , F23R3/34 , F23R3/44 , F04D29/54 , F01D9/02 , F04D27/02 , F04D29/56 , F23R3/04 , F23R3/12 , F23R3/26 , F23R3/28 , F23R3/50 , B23P15/04 , F02C3/04
Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.
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公开(公告)号：US20210179283A1
公开(公告)日：2021-06-17
申请号：US17021864
申请日：2020-09-15
Applicant: General Electric Company
Inventor： Narendra Digamber Joshi , Konrad Weeber , Michael Solomon Idelchik
Abstract: A motor driven propulsor of an aircraft includes magnets disposed in fan shrouds of fan blades connected with a fan hub, a stator having individual conductive coils in a nacelle located radially outside of the fan hub, and a distributed inverter assembly having several inverter power stages and gate drivers, each of the inverter power stages coupled with a separate gate driver of the gate drivers and a separate coil of the coils in the stator. Each of the gate drivers is configured to individually control supply of direct current to the corresponding inverter power stage. Each of the inverter power stages is configured to convert the direct current supplied to the inverter power stage to an alternating current that is supplied to the corresponding coil in the stator to rotate the magnets and the fan blades around a center line of the fan hub for propelling the aircraft.
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公开(公告)号：US20210147060A1
公开(公告)日：2021-05-20
申请号：US16685504
申请日：2019-11-15
Applicant: General Electric Company
Inventor： Timothy John Sommerer , Nicholas William Rathay , Narendra Digamber Joshi , Douglas Carl Hofer , Hongbo Cao
Abstract: A hypersonic aircraft includes one or more leading edge assemblies that are designed to cool the leading edge of certain portions of the hypersonic aircraft that are exposed to high thermal loads, such as extremely high temperatures and/or thermal gradients. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. A coolant supply may be in fluid communication with at least one fluid passageway that passes through the outer wall to deliver a flow of cooling fluid, such as liquid metal, to the stagnation point. The liquid metal vaporizes when the leading edge experiences a high heat load, thereby transpiration cooling the leading edge and/or facilitating a magnetohydrodynamic process for generating thrust or electricity.
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Patent Agency Ranking
公开(公告)号：US10814991B2
公开(公告)日：2020-10-27
申请号：US14935814
申请日：2015-11-09
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Manoj Ramprasad Shah , Narendra Digamber Joshi
Abstract: Embodiments of a propulsion system are provided herein. In some embodiments, a propulsion system for an aircraft may include an electrical power supply; a motor coupled to the electrical power supply, wherein the electrical power supply provides power to the motor; and a fan disposed proximate a rear portion of an aircraft and rotatably coupled to the motor, wherein the fan is driven by the motor.
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公开(公告)号：US20200248905A1
公开(公告)日：2020-08-06
申请号：US16267473
申请日：2019-02-05
Applicant: General Electric Company
Inventor： Kapil Kumar Singh , Narendra Digamber Joshi
Abstract: The present disclosure is directed to a rotating detonation combustor that includes a forward wall, a radially inner wall, and a radially outer wall. The forward wall is disposed at an inlet end of the rotating detonation combustor. The radially inner wall surrounds a longitudinal axis and extends downstream from the forward wall to an outlet end of the rotating detonation combustor. The radially outer wall extends downstream from the forward wall to the outlet end and surrounds the radially inner wall to define at least one annular plenum between the radially inner wall and the radially outer wall. At least one partition is proximate to the inlet end and defines at least two mixing zones. A plurality of oxidizer inlets and a plurality of fuel inlets are disposed at the inlet end in fluid communication with the at least two mixing zones.
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公开(公告)号：US10724432B2
公开(公告)日：2020-07-28
申请号：US15805878
申请日：2017-11-07
Applicant: General Electric Company
Inventor： Andrew Philip Shapiro , Narendra Digamber Joshi
IPC: F02C3/20 , H01M8/124 , F02C6/10 , H01M8/04111 , H01M8/1246 , H01M8/04082 , F02C6/06 , H01M4/86
Abstract: An integrated fuel cell and engine combustor assembly includes an engine combustor having a combustion chamber fluidly coupled with a compressor and a turbine. The assembly also includes a fuel cell stack circumferentially extending around the combustion chamber of the combustor. The fuel cell stack includes fuel cells configured to generate electric current. The fuel cell stack is positioned to receive discharged air from the compressor and fuel from a fuel manifold. The fuel cells in the fuel cell stack generate electric current using the discharged air and at least some of the fuel. The fuel cell stack is positioned to radially direct partially oxidized fuel from the fuel cells into the combustion chamber of the combustor. The combustor combusts the partially oxidized fuel into one or more gaseous combustion products that are directed into and drive the downstream turbine.
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公开(公告)号：US10677158B2
公开(公告)日：2020-06-09
申请号：US14982793
申请日：2015-12-29
Applicant: General Electric Company
IPC: F02C3/107 , F02K3/065 , F02K3/077 , F01D5/02 , F01D25/30 , F02C7/04 , F02C7/36 , F04D29/32 , F02K3/062 , F02K3/06
Abstract: A gas turbine engine system and method of operating gas turbine engines are provided. The gas turbine engine assembly includes a gas turbine engine includes a power shaft configured to rotate about an axis of rotation. The gas turbine engine assembly also includes a first fan and a second fan coupled to the power shaft coaxially with the gas turbine engine. The gas turbine engine assembly also includes a first fan duct configured to direct a first stream of air to the first fan. The gas turbine engine assembly also includes a second fan duct configured to direct a second stream of air to the second fan. The gas turbine engine assembly also includes an exhaust duct configured to direct a stream of exhaust gases of the gas turbine engine in a direction of the axis of rotation.
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公开(公告)号：US20200149496A1
公开(公告)日：2020-05-14
申请号：US16185235
申请日：2018-11-09
Applicant: General Electric Company
Inventor： Kapil Kumar Singh , Narendra Digamber Joshi
Abstract: A combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one fluid inlet disposed in the annular tube proximate the inlet end, the fluid inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; and at least one inlet fluid plenum disposed upstream of the fluid inlet. The inlet fluid plenum includes at least one reflective surface.
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