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.

Abstract: A satellite has thrusters that are integral parts of its frame. The frame defines cavities therein where thrusters are located. The thrusters may include an electrically-operated propellant and electrodes to activate combustion in the electrically-operated propellant. The frame may be additively manufactured, and the propellant and/or the electrodes may also be additively manufactured, with the frame and the propellant and/or the electrodes also being manufactured in a single process. In addition the thrusters may have nozzle portions through which combustion gases exit the thrusters. The thrusters may be located at corners and/or along edges of the frame, and may be used to accomplish any of a variety of maneuvers for the satellite. The satellite may be 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.

Abstract: 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).

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.

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.

Abstract: 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.

Abstract: A missile includes a rocket motor that has a flow deflector in an expansion region of the rocket motor's nozzle. The flow deflector diverts flow of combustion products away from a safe region that is aft of the missile. The safe region protects an operator of a launcher used to fire the missile, such as a shoulder-fired launcher, from harm caused by the combustion products. The flow deflector may be small enough such that it does not significantly adversely affect the performance of the rocket motor. The presence of the flow diverter may allow for the rocket motor to be started sooner in the flight of the missile, or at a distance closer to the operator, while still keeping the operator safe. The flow deflector and supporting structure, such as struts, may be additively manufactured with at least an aft part of the nozzle, as a continuous single-piece part.

Abstract: A missile includes a rocket motor that has a flow deflector in an expansion region of the rocket motor's nozzle. The flow deflector diverts flow of combustion products away from a safe region that is aft of the missile. The safe region protects an operator of a launcher used to fire the missile, such as a shoulder-fired launcher, from harm caused by the combustion products. The flow deflector may be small enough such that it does not significantly adversely affect the performance of the rocket motor. The presence of the flow diverter may allow for the rocket motor to be started sooner in the flight of the missile, or at a distance closer to the operator, while still keeping the operator safe. The flow deflector and supporting structure, such as struts, may be additively manufactured with at least an aft part of the nozzle, as a continuous single-piece part.

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.

Abstract: 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).

Abstract: 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.

Abstract: A satellite has thrusters that are integral parts of its frame. The frame defines cavities therein where thrusters are located. The thrusters may include an electrically-operated propellant and electrodes to activate combustion in the electrically-operated propellant. The frame may be additively manufactured, and the propellant and/or the electrodes may also be additively manufactured, with the frame and the propellant and/or the electrodes also being manufactured in a single process. In addition the thrusters may have nozzle portions through which combustion gases exit the thrusters. The thrusters may be located at corners and/or along edges of the frame, and may be used to accomplish any of a variety of maneuvers for the satellite. The satellite may be 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.

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.

Abstract: A device may include an electrically-operated propellant or energetic gas-generating material, additively manufactured together with electrodes for producing a reaction in the material. The device may also include a casing that is additively manufactured with the other components. The additive manufacturing may be accomplished by extruding or otherwise depositing raw materials for the different components where desired. The electrodes may be made of a conductive polymer material, for example using an electrically-conductive fill in a polymer.

Abstract: A satellite has thrusters that are integral parts of its frame. The frame defines cavities therein where thrusters are located. The thrusters may include an electrically-operated propellant and electrodes to activate combustion in the electrically-operated propellant. The frame may be additively manufactured, and the propellant and/or the electrodes may also be additively manufactured, with the frame and the propellant and/or the electrodes also being manufactured in a single process. In addition the thrusters may have nozzle portions through which combustion gases exit the thrusters. The thrusters may be located at corners and/or along edges of the frame, and may be used to accomplish any of a variety of maneuvers for the satellite. The satellite may be 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.

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.

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.

Abstract: 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.

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.

Abstract: 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.