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    ROCKET MOTOR PRODUCED BY ADDITIVE MANUFACTURING
    22.
    发明申请
    ROCKET MOTOR PRODUCED BY ADDITIVE MANUFACTURING 审中-公开
    由添加剂制造生产的摇臂电机

    公开(公告)号：US20160356245A1

    公开(公告)日：2016-12-08

    申请号：US14729390

    申请日：2015-06-03

    Applicant: Raytheon Company

    Inventor： Jeremy C. Danforth ,  Teresa Perdue ,  Mark T. Langhenry ,  Matt H. Summers

    IPC: F02K9/95

    CPC classification number: F02K9/95 ,  F02K9/72 ,  F02K9/76

    Abstract: A nozzleless hybrid rocket motor includes a fuel element that defines a combustion chamber therewithin, in which combustion of the fuel and an oxidizer occurs. The combustion gases produced by the combustion between the fuel and the oxidizer transition to supersonic flow before leaving the fuel element, eliminating the need for a separate nozzle. The fuel element may be a part of a structural element of a vehicle, for example being a part of a fuselage, wing, fairing, or other part of a space vehicle or an air vehicle, with the fuel element an integral and continuous part of the structural element. Combustion of part of the fuel element may allow vehicle structure to be used to provide thrust, such as for maneuver, consuming part of the structure. The fuel element may be made by an additive manufacturing process.

    Abstract translation: 无喷嘴混合火箭发动机包括在其中限定燃烧室的燃料元件，其中发生燃料和氧化剂的燃烧。 由燃料和氧化剂之间的燃烧产生的燃烧气体在离开燃料元件之前转变为超音速流动，消除了对单独喷嘴的需要。 燃料元件可以是车辆的结构元件的一部分，例如是机身，机翼，整流罩或空间飞行器或空中交通工具的其他部分的一部分，其中燃料元件是整体和连续的部分 结构元素。 燃料元件的一部分的燃烧可以允许车辆结构用于提供推力，例如用于操纵，消耗部分结构。 燃料元件可以通过添加剂制造工艺制成。

    Hybrid rocket motor with integral oxidizer tank
    23.
    发明授权
    Hybrid rocket motor with integral oxidizer tank 有权

    公开(公告)号：US11326556B2

    公开(公告)日：2022-05-10

    申请号：US17035880

    申请日：2020-09-29

    Applicant: Raytheon Company

    Inventor： Matt H. Summers ,  Jeremy C. Danforth

    IPC: F02K9/72 ,  F02K9/32 ,  F02K9/10 ,  B33Y10/00 ,  B33Y80/00

    Abstract: A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.

    Integrated thruster
    24.
    发明授权
    Integrated thruster 有权

    公开(公告)号：US10662898B2

    公开(公告)日：2020-05-26

    申请号：US15259945

    申请日：2016-09-08

    Applicant: Raytheon Company

    Inventor： Matthew H. Summers ,  Jeremy C. Danforth ,  David G. Garrett ,  Dmitry V. Knyazev ,  Stephen M. Baggs ,  Gaines S. Gibson

    IPC: F02K9/34 ,  F42B10/66 ,  F02K9/30 ,  F02K9/88 ,  F02K9/97 ,  F02K9/38 ,  F02K9/08 ,  F02K9/36 ,  F02K9/95

    Abstract: A thruster has an additively-manufactured housing that includes an integrally-formed nozzle with a burst disk in it. The housing is part of a casing that surrounds and encloses a propellant that is burned to produce pressurized gases that burst the burst disk and produce thrust. The thruster may be placed in a receptacle that defines a recess for receiving the thruster. The receptacle also may be additively manufactured. The thruster and the recess both may be cylindrical, with the housing being closely fit with the cylindrical walls of the receptacle. This may allow some of the structural loads on the housing, such as loads produced by the combustion of the propellant, to be transferred to the adjoining walls of the receptacle. This enables the housing to have less structural strength than if it were to have to contain the pressure from the propellant all on its own.

    ADDITIVE MANUFACTURED COMBUSTIBLE ELEMENT WITH FUEL AND OXIDIZER
    26.
    发明申请
    ADDITIVE MANUFACTURED COMBUSTIBLE ELEMENT WITH FUEL AND OXIDIZER 审中-公开

    公开(公告)号：US20190002362A1

    公开(公告)日：2019-01-03

    申请号：US16125871

    申请日：2018-09-10

    Applicant: Raytheon Company

    Inventor： Jeremy C. Danforth ,  Mark T. Langhenry ,  Matt H. Summers ,  Teresa Perdue

    IPC: C06B45/10 ,  C06B21/00 ,  F02K9/12 ,  B33Y10/00 ,  B33Y80/00

    Abstract: A combustible element includes regions of fuel material interspersed with regions of oxidizer material. The element may be made by additive manufacturing processes, such as three-dimensional printing, with the fuel material regions and the oxidizer material regions placed in appropriate locations in layer of the combustible element. For example, different extruders may be used to extrude and deposit portions of a fuel filament and an oxidizer filament at different locations in each layer of the combustible element. The combustible element may define a combustion chamber within the element, where combustion occurs when the combustible element is ignited. The fuel material and the oxidizer material may be selected, and their relative amounts may be controlled, such that desired relative amounts of fuel and oxidizer are present for combustion with desired characteristics, such as combustion rate.

    Electrode ignition and control of electrically operated propellants
    27.
    发明授权
    Electrode ignition and control of electrically operated propellants 有权

    公开(公告)号：US10145337B2

    公开(公告)日：2018-12-04

    申请号：US15197421

    申请日：2016-06-29

    Applicant: Raytheon Company

    Inventor： Matt H. Summers ,  James K. Villarreal ,  Mark T. Langhenry ,  Jeremy C. Danforth ,  John W. Walter

    IPC: F02K9/26 ,  B60R21/264 ,  F02K9/95 ,  F02K9/94 ,  F42C19/08 ,  B60R21/26

    Abstract: Electrical ignition of electrically operated propellant in a gas generation system provides an ignition condition at an ignition surface between a pair of electrodes that satisfies three criteria of a current density J that exhibits a decreasing gradient along an axis normal to an ignition surface, is substantially constant across the ignition surface and exceeds an ignition threshold at the ignition surface. These criteria may be satisfied by one or more of an angled electrode configuration, a segmented electrode configuration or an additive to the electrically operated propellant that modifies its conductivity. These configurations improve burn rate control and consumption of the available propellant and are scalable to greater propellant mass to support larger gas generation systems.

    Method of producing solid propellant element
    28.
    发明授权
    Method of producing solid propellant element 有权

    公开(公告)号：US10023505B2

    公开(公告)日：2018-07-17

    申请号：US15057587

    申请日：2016-03-01

    Applicant: Raytheon Company

    Inventor： Jeremy C. Danforth ,  Matt H. Summers ,  David G. Garrett

    IPC: C06B21/00 ,  C06B45/00 ,  F02K9/10 ,  B29C64/106 ,  B33Y10/00

    Abstract: A method of producing a propellant material element, such as an electrically-operated propellant material, includes extruding a propellant material through a heated nozzle. The nozzle may be heated to a temperature that is above the boiling point of a solvent that is part of the propellant material, yet is below a decomposition temperature of the propellant material. This allows some of the solvent to be driven off during the extruding process, while still preventing initiation of an energy-creating reaction within the material. The heating of the material in the extruding process, and especially the heating of the nozzle that the material is extruded through, may be controlled to remove an amount of solvent that results in the extruded material having desirable properties.

    SPACECRAFT WITH SHAPE MEMORY POLYMER DEPLOYMENT MECHANISM
    30.
    发明申请
    SPACECRAFT WITH SHAPE MEMORY POLYMER DEPLOYMENT MECHANISM 审中-公开

    公开(公告)号：US20170113817A1

    公开(公告)日：2017-04-27

    申请号：US14921545

    申请日：2015-10-23

    Applicant: Raytheon Company

    Inventor： Frederick B. Koehler ,  Ward D. Lyman ,  Matt H. Summers ,  Jeremy C. Danforth

    IPC: B64G1/22 ,  B64G1/10

    CPC classification number: B64G1/222 ,  B64G1/10 ,  F03G7/065

    Abstract: A spacecraft, such as a satellite, uses a shape memory polymer actuator to deploy one or more deployable parts. The shape memory polymer actuator may be formed integrally with a deployable part and/or with a fuselage or other structure of the spacecraft, with the shape memory polymer actuator being for example a relatively thin portion of the shape memory polymer material of the integral structure. The shape memory actuator allows deployment of the deployable part(s) upon heating of the shape memory polymer material of the actuator, such as after the satellite has been launched into space. The heating may be caused by a heat source that is part of the spacecraft itself, or may be merely the result of exposing the spacecraft to solar heating after launch. The deployable part of the spacecraft may include any of a wide variety of parts that are used after launch.

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