Solid propellant systems include a main propellant and a secondary propellant in contact with the first propellant that exhibits autoignition temperatures of at least about 100° F. lower than the autoignition temperature of the main propellant. The secondary propellant of the present invention is most advantageously employed with conventional AP-containing solid propellant formulations as the main propellant, especially formulations containing both AP, an energetic solid, and a binder. In especially preferred forms, the secondary propellant will include a nitramine which is at least one selected from nitroguanidine (NQ), cyclotrimethylene trinitramine (RDX) and cyclotetramethylenetetranitramine (HMX), and a binder which is at least one selected from HTPB, HTPE or glycidyl azide polymer (GAP). Most preferably, the secondary propellant will include a combination of nitramines which includes NQ and one of RDX or HMX.

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.

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.

A spark plug having an insulating body defining a longitudinal axis and including a base portion and an obstruction portion, a first electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the first electrode extending through the base portion of the insulating body, and a second electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the second electrode extending through the base portion and the obstruction portion of the insulating body, wherein the obstruction portion axially extends beyond the distal portion of the first electrode.

A spark plug having an insulating body defining a longitudinal axis and including a base portion and an obstruction portion, a first electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the first electrode extending through the base portion of the insulating body, and a second electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the second electrode extending through the base portion and the obstruction portion of the insulating body, wherein the obstruction portion axially extends beyond the distal portion of the first electrode.

A catalyzing reactor comprising a reactor entrance and a reactor exit and an internal structure arranged for flowing a reacting medium through the reactor from the reactor entrance to the reactor exit. The reactor structure comprising at least one thin walled reactor channel arranged between the entrance and the exit of the reactor. The channel having a channel wall that includes a catalyst and that defines a flow path, in which channel in use, a catalyzed exothermic reaction takes place in the medium as it flows along the flow path. The at least one channel is looped to have a portion of its flow path that is downstream with respect to the reactor entrance in heat exchanging contact with a portion of a flow path that is that is more upstream with respect to the reactor entrance, so as to transfer heat between a downstream portion of the reacting medium to an upstream portion thereof.

A catalyzing reactor comprising a reactor entrance and a reactor exit and an internal structure arranged for flowing a reacting medium through the reactor from the reactor entrance to the reactor exit. The reactor structure comprising at least one thin walled reactor channel arranged between the entrance and the exit of the reactor. The channel having a channel wall that includes a catalyst and that defines a flow path, in which channel in use, a catalyzed exothermic reaction takes place in the medium as it flows along the flow path. The at least one channel is looped to have a portion of its flow path that is downstream with respect to the reactor entrance in heat exchanging contact with a portion of a flow path that is that is more upstream with respect to the reactor entrance, so as to transfer heat between a downstream portion of the reacting medium to an upstream portion thereof.

The rapid assistance device applies to a free turbine engine of an aircraft having at least a first free turbine engine provided with a gas generator and associated with an electrical machine capable of operating both as a starter and as a generator, the first engine being capable of being put into a standby mode or into an unwanted shut-down mode, the electrical machine being powered from on on-board electrical energy power supply network. The device further includes at least one electrical energy storage member adapted to be electrically connected to the electrical machine associated with the first engine in order to provide a burst of assistance to the gas generator of that engine. The electrical energy storage member constitutes a non-rechargeable “primary” energy storage member that can be used once only. The device includes a system for activating the electrical energy storage member and a device for coupling the electrical energy storage member with an electrical power supply system of the electrical machine.

A method for generating electric power for a rocket system includes burning a primary solid propellant grain to create a primary high pressure gas for providing thrust to the rocket, opening a first valve to divert a portion of the high pressure gas to an auxiliary solid propellant grain for igniting the auxiliary solid propellant grain, wherein the auxiliary solid propellant grain is disposed in a housing separate from the primary solid propellant grain, and burning the auxiliary solid propellant grain to create an auxiliary high pressure gas for turning a turbine. The method further includes driving a generator with the turbine and generating an electric power with the generator.

A method for generating electric power for a rocket system includes burning a primary solid propellant grain to create a primary high pressure gas for providing thrust to the rocket, opening a first valve to divert a portion of the high pressure gas to an auxiliary solid propellant grain for igniting the auxiliary solid propellant grain, wherein the auxiliary solid propellant grain is disposed in a housing separate from the primary solid propellant grain, and burning the auxiliary solid propellant grain to create an auxiliary high pressure gas for turning a turbine. The method further includes driving a generator with the turbine and generating an electric power with the generator.