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1. 发明申请

US20170284339A1 THRUSTER WITH SEGMENTED PROPELLANT 审中-公开
公开(公告)号：US20170284339A1
公开(公告)日：2017-10-05
申请号：US15091245
申请日：2016-04-05
申请人： Raytheon Company
发明人： Frederick B. Koehler , Jeremy C. Danforth , Ward D. Lyman , Mark T. Langhenry , Matt H. Summers , Paul E. Pontius , Brian M. Pape , Jared D. Stallings , James K. Villarreal , Thomas Villarreal
IPC分类号： F02K9/28
CPC分类号： F02K9/28 , F02K9/26 , F02K9/32 , F02K9/94 , F02K9/95
摘要： A thruster includes multiple segments of electrically-operated propellant, electrodes for igniting one or a few of the electrically-operated propellant segments at a time, and a propellant feeder for moving further propellant segments into engagement with the electrodes. The segments may be configured to provide equal increments of thrust, or different amounts of thrust. The segments may each include an electrically-operated propellant material surrounded by a sealing material, so as to keep the propellant material away from moisture and other contaminants (and/or the vacuum of space) before each individual segment is to be used. The thruster may be included in any of a variety of flight vehicles, for example in a small satellite such as a CubeSat satellite, for instance having a volume of about 1 liter, and a mass of no more than about 1.33 kg.

2. 发明申请

US20170097213A1 ELECTRICALLY OPERATED PULSE INITIATORS AND IGNITION 审中-公开
公开(公告)号：US20170097213A1
公开(公告)日：2017-04-06
申请号：US14875741
申请日：2015-10-06
申请人： Raytheon Company
发明人： James Kendall Villarreal , Mark T. Langhenry , Matt H. Summers , Daniel V. Macinnis
IPC分类号： F42B3/04
CPC分类号： F42B3/04 , F02K9/28 , F02K9/95
摘要： A gas generation system for generating gases, such as for use as or as part of a rocket motor in propelling a projectile, includes two or more propellant charges and electrically operated propellant initiators operatively coupled to respective of the propellant charges, to initiate combustion in the propellant charges, wherein the propellant charges are operatively isolated from one another such that the propellant charges can be individually initiated and are not ignited due to gases generated from other of the propellant charges being combusted.

3. 发明申请

US20160355447A1 ADDITIVE MANUFACTURED COMBUSTIBLE ELEMENT WITH FUEL AND OXIDIZER 审中-公开
标题翻译：燃料和氧化剂的添加剂制造可燃组分
公开(公告)号：US20160355447A1
公开(公告)日：2016-12-08
申请号：US14729429
申请日：2015-06-03
申请人： Raytheon Company
发明人： Jeremy C. Danforth , Mark T. Langhenry , Matt H. Summers , Teresa Perdue
IPC分类号： C06B45/10 , C06B21/00
CPC分类号： C06B45/10 , B33Y10/00 , B33Y80/00 , C06B21/0033 , F02K9/12 , F05D2230/31
摘要： 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.
摘要翻译：可燃元件包括散布有氧化剂材料区域的燃料材料区域。该元件可以通过添加剂制造工艺（例如三维印刷）制成，其中燃料材料区域和氧化剂材料区域放置在可燃元件的层中的适当位置。例如，可以使用不同的挤出机来挤出和沉积燃料丝和氧化剂丝的部分在可燃元件的每层中的不同位置。可燃元件可以在元件内限定燃烧室，其中当可燃元件被点燃时发生燃烧。可以选择燃料材料和氧化剂材料，并且可以控制它们的相对量，使得燃烧所需的相对量的燃料和氧化剂存在于具有所需特性如燃烧速率的燃烧中。

4. 发明授权

US11326556B2 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
申请人： Raytheon Company
发明人： Matt H. Summers , Jeremy C. Danforth
IPC分类号： F02K9/72 , F02K9/32 , F02K9/10 , B33Y10/00 , B33Y80/00
摘要： 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.

5. 发明申请

US20200080518A1 RESIN TRANSFER MOLDED ROCKET MOTOR NOZZLE WITH ADAPTIVE GEOMETRY 审中-公开
公开(公告)号：US20200080518A1
公开(公告)日：2020-03-12
申请号：US16126225
申请日：2018-09-10
申请人： Raytheon Company
发明人： Matt H. Summers , David M. Case , John J. Raymond, IV , Gray Fowler , Mark T. Langhenry
IPC分类号： F02K9/97 , F02K9/86
摘要： A rocket motor has a nozzle that is reconfigurable by erosion or ablation of the material around the throat of the nozzle. The nozzle throat has layers of materials with different erosion characteristics, with the erosion occurring so as to achieve the desired nozzle characteristics (configurations) during different parts of the fuel burn. The nozzle throat includes relatively-high-erosion material layers and relatively-low-erosion material layers, with some layers of the throat resisting erosion, while other of the layers erode or ablate relatively quickly. The relatively-low-erosion material layers may act as thermal barriers to fix the throat at relatively stable geometry for long periods of time, such as during most of the burn of different fuel segments, with the relatively-high-erosion material layers allowing rapid transition of the throat from one geometry to the next. The layers may be made by resin transfer molding (RTM).

6. 发明授权

US10563617B2 Electrically operated propellant for solid rocket motor thrust management 有权
公开(公告)号：US10563617B2
公开(公告)日：2020-02-18
申请号：US15274669
申请日：2016-09-23
申请人： Raytheon Company
发明人： James K. Villarreal , Mark T. Langhenry , Matt H. Summers
IPC分类号： F02K9/24 , F02K9/12 , F02K9/95 , F02K9/86 , F02K9/82 , B64G1/40 , F02K9/08 , F02K9/94 , F02K9/26 , F02K9/28 , F02K9/34 , F02K9/97
摘要： Electrically operated propellant is used to supplement the thrust provided by solid rocket motor (SRM) propellant to manage thrust produced by a SRM. The gas produced by burning the electrically operated propellant may be injected upstream of the nozzle to add mass and increase chamber pressure Pc, injected at the throat of the nozzle to reduce the effect throat area At to increase chamber pressure Pc or injected downstream of the throat to provide thrust vector control or a combination thereof. Certain types of electrically operated propellants can be turned on and off provided the chamber pressure Pc does not exceed a self-sustaining threshold pressure eliminating the requirement for physical control valves.

7. 发明申请

US20190002362A1 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
申请人： Raytheon Company
发明人： Jeremy C. Danforth , Mark T. Langhenry , Matt H. Summers , Teresa Perdue
IPC分类号： C06B45/10 , C06B21/00 , F02K9/12 , B33Y10/00 , B33Y80/00
摘要： 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.

8. 发明授权

US10145337B2 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
申请人： Raytheon Company
发明人： Matt H. Summers , James K. Villarreal , Mark T. Langhenry , Jeremy C. Danforth , John W. Walter
IPC分类号： F02K9/26 , B60R21/264 , F02K9/95 , F02K9/94 , F42C19/08 , B60R21/26
摘要： 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.

9. 发明授权

US10023505B2 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
申请人： Raytheon Company
发明人： Jeremy C. Danforth , Matt H. Summers , David G. Garrett
IPC分类号： C06B21/00 , C06B45/00 , F02K9/10 , B29C64/106 , B33Y10/00
摘要： 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.

10. 发明申请

US20170113817A1 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
申请人： Raytheon Company
发明人： Frederick B. Koehler , Ward D. Lyman , Matt H. Summers , Jeremy C. Danforth
IPC分类号： B64G1/22 , B64G1/10
CPC分类号： B64G1/222 , B64G1/10 , F03G7/065
摘要： 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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