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1. 发明申请

WO2018175037A1 ELECTRICALLY OPERATED PROPELLANT THRUST ASSIST FOR SUPPLEMENTING AIRPLANE TAKEOFF, LANDING OR IN-FLIGHT MANEUVERABILITY 审中-公开
公开(公告)号：WO2018175037A1
公开(公告)日：2018-09-27
申请号：PCT/US2018/019423
申请日：2018-02-23
申请人： RAYTHEON COMPANY
发明人： SUMMERS, Matt H. , KOEHLER, Frederick B. , VILLARREAL, James K. , LANGHENRY, Mark T.
IPC分类号： B64D27/02 , C06B33/06 , C06D5/06 , F02K9/94 , F02K9/95 , B64G1/40
CPC分类号： B64D27/023 , B64C2201/046 , B64C2201/167 , B64D31/02 , B64G1/403 , B64G1/404 , C06B33/06 , C06D5/06 , F02K9/08 , F02K9/94 , F02K9/95 , F02K9/97 , F05D2220/80 , F05D2240/35
摘要： Electrically operated propellant thrust assist supplements an airplane's takeoff, landing or inflight maneuvers. Unlike conventional SRM propellants, the burn rate of the electrically operated propellant can be varied via an electrical input and even extinguished by interrupting the electrical to control a secondary thrust profile (e.g., amplitude, transition rates) to fulfill the needs of a given takeoff, inflight or landing maneuver and provide a smooth transition in and out of the maneuver. Multiple pairs of fixed thrusters (opposite sides of the fuselage), a single pair of gimbaled thrusters or a hybrid of fixed and gimbaled thrusters may be configured to provide all such maneuvers. Flight control inputs are passed back and forth through an interface to enable the thrust assist.

2. 发明申请

WO2018034795A1 MICROWAVE IGNITION OF ELECTRICALLY OPERATED PROPELLANTS 审中-公开
标题翻译：电动推进器的微波点火
公开(公告)号：WO2018034795A1
公开(公告)日：2018-02-22
申请号：PCT/US2017/043418
申请日：2017-07-22
申请人： RAYTHEON COMPANY
发明人： VILLARREAL, James K. , DANFORTH, Jeremy C. , SUMMERS, Matt H. , JOHNSON, Daniel K. , LANGHENRY, Mark T.
IPC分类号： C06B23/00 , C06B45/10 , F02K9/94 , F02K9/95
CPC分类号： F02K9/95 , C06B23/00 , C06B27/00 , C06B29/00 , C06B45/10 , F02K9/08 , F02K9/94
摘要： Microwave energy is used to ignite and control the ignition of electrically operated propellant to produce high-pressure gas. The propellant includes conductive particles that act as a free source of electrons. Incoming microwave energy accumulates electric charge in an attenuation zone, which is discharged in the form of dielectric breakdowns to create local randomly oriented currents. The propellant also includes polar molecules. The polar molecules in the attenuation zone absorb microwave energy causing the molecules to rapidly vibrate thereby increasing the temperature of the propellant. The increase in temperature and the local current densities together establish an ignition condition to ignite and sustain ignition of an ignition surface of the attenuation zone as the zone regresses without igniting the remaining bulk of the propellant.
摘要翻译：微波能量用于点燃和控制电动推进剂的点火以产生高压气体。推进剂包括充当电子自由源的导电粒子。输入的微波能量在衰减区累积电荷，衰减区以介质击穿形式放电，产生局部随机取向电流。推进剂还包括极性分子。衰减区中的极性分子吸收微波能量，使分子快速振动，从而增加推进剂的温度。温度升高和局部电流密度一起形成点火条件，以点燃和维持衰减区的点火表面点火，因为区域退化而不点燃剩余的推进剂体积。

3. 发明申请

WO2017176311A1 THRUSTER WITH SEGMENTED PROPELLANT 审中-公开
标题翻译：推进器与分段推进剂
公开(公告)号：WO2017176311A1
公开(公告)日：2017-10-12
申请号：PCT/US2016/056387
申请日：2016-10-11
申请人： RAYTHEON COMPANY
发明人： KOEHLER, Frederick B. , DANFORTH, Jeremy C. , LYMAN, Ward D. , LANGHENRY, Mark T. , SUMMERS, Matt H. , PONTIUS, Paul E. , PAPE, Brian M. , STALLINGS, Jared D. , VILLARREAL, James K. , VILLARREAL, Thomas
IPC分类号： 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.
摘要翻译：推进器包括多个电动推进剂段，用于一次点燃一个或几个电动推进剂段的电极，以及推进剂供给器，用于将另外的推进剂段移动到与电极。这些段可以被配置成提供相等的推力增量或不同的推力量。这些区段可以各自包括由密封材料围绕的电操作推进剂材料，以便在每个单独的区段将被使用之前使推进剂材料远离湿气和其他污染物（和/或空间的真空）。推进器可以包括在各种飞行器中的任何一种中，例如在诸如CubeSat卫星的小卫星中，例如具有大约1升的体积和不大于大约1.33kg的质量。

4. 发明申请

WO2020055625A1 RESIN TRANSFER MOLDED ROCKET MOTOR NOZZLE WITH ADAPTIVE GEOMETRY 审中-公开
公开(公告)号：WO2020055625A1
公开(公告)日：2020-03-19
申请号：PCT/US2019/049311
申请日：2019-09-03
申请人： RAYTHEON COMPANY
发明人： SUMMERS, Matt H. , CASE, David M. , RAYMOND, John J. , FOWLER, Gray , LANGHENRY, Mark T.
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).

5. 发明申请

WO2018022203A1 ELECTRODE IGNITION AND CONTROL OF ELECTRICALLY OPERATED PROPELLANTS 审中-公开
标题翻译：电动推进器的电极点火和控制
公开(公告)号：WO2018022203A1
公开(公告)日：2018-02-01
申请号：PCT/US2017/036803
申请日：2017-06-09
申请人： RAYTHEON COMPANY
发明人： SUMMERS, Matt H. , VILLARREAL, James K. , LANGHENRY, Mark T. , DANFORTH, Jeremy C. , WALTER, John W.
IPC分类号： F42C19/08 , F02K9/94 , F02K9/95
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.
摘要翻译：气体发生系统中的电操作推进剂的电点火在一对电极之间的点火表面处提供满足电流密度J的三个标准的点火条件，所述电流密度J表现出沿着法向轴的递减梯度到点火表面，在整个点火表面上基本恒定并超过点火表面处的点火阈值。这些标准可以通过改变其电导率的成角度的电极构造，分段电极构造或电动推进剂的添加剂中的一种或多种来满足。这些配置提高了燃烧速率控制和可用推进剂的消耗，并且可扩展到更大的推进剂质量以支持更大的气体发生系统。

6. 发明申请

WO2017062077A1 ELECTRICALLY OPERATED PULSE INITIATORS AND IGNITION 审中-公开
标题翻译：电动操作脉冲发动机和点火
公开(公告)号：WO2017062077A1
公开(公告)日：2017-04-13
申请号：PCT/US2016/038620
申请日：2016-06-22
申请人： RAYTHEON COMPANY
发明人： VILLARREAL, James Kendall , LANGHENRY, Mark T. , SUMMERS, Matt H. , MACINNIS, Daniel V.
IPC分类号： F02K9/95 , F02K9/28
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.
摘要翻译：用于产生气体的气体发生系统，例如用作或作为推动射弹的火箭发动机的一部分，包括两个或更多个推进剂装料和可操作地联接到相应的推进剂装料的电动推进剂起动器，以在推进剂装料，其中推进剂装料彼此有效地隔离，使得推进剂装料可以单独启动，并且由于其他燃烧的推进剂装料产生的气体而不被点燃。

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