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.

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 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 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 rocket motor is disclosed that can include a combustion chamber containing a solid fuel that is operable to burn during operation of the rocket motor to generate combustion gas and unburned gaseous fuel. The rocket motor can also include a propellant supply containing an energy-rich oxidizer with a decomposition energy greater than or equal to 1.0 MJ/kg. In addition, the rocket motor can include a thrust augmented nozzle (TAN) operably coupled to the combustion chamber to receive the combustion gas from the combustion chamber and direct a flow of the combustion gas through the TAN. The TAN can have a divergent portion downstream of a throat, and a propellant injection port associated with the divergent portion and in communication with the propellant supply to inject the energy-rich oxidizer into the divergent portion. Only the energy-rich oxidizer, independent of another propellant, may be introduced into the flow of the combustion gas and the unburned gaseous fuel for secondary combustion of the unburned gaseous fuel and thermal decomposition of the energy-rich oxidizer within the divergent portion.

A rocket motor is disclosed that can include a combustion chamber containing a solid fuel that is operable to burn during operation of the rocket motor to generate combustion gas and unburned gaseous fuel. The rocket motor can also include a propellant supply containing an energy-rich oxidizer with a decomposition energy greater than or equal to 1.0 MJ/kg. In addition, the rocket motor can include a thrust augmented nozzle (TAN) operably coupled to the combustion chamber to receive the combustion gas from the combustion chamber and direct a flow of the combustion gas through the TAN. The TAN can have a divergent portion downstream of a throat, and a propellant injection port associated with the divergent portion and in communication with the propellant supply to inject the energy-rich oxidizer into the divergent portion. Only the energy-rich oxidizer, independent of another propellant, may be introduced into the flow of the combustion gas and the unburned gaseous fuel for secondary combustion of the unburned gaseous fuel and thermal decomposition of the energy-rich oxidizer within the divergent portion.

A selectable ramjet propulsion system for propelling a rocket or missile includes a gas generator adjacent a booster. A frangible diaphragm is disposed between the gas generator and the booster. The booster and fuel gas generator can be operated in normal sequence, or operated at the same time in order to increase the thrust produced for short-range missions. A logic circuit contained on the rocket or missile determines a time to rupture the frangible diaphragm based on whether or not the distance to the target exceeds a threshold distance.

A selectable ramjet propulsion system for propelling a rocket or missile includes a gas generator adjacent a booster. A frangible diaphragm is disposed between the gas generator and the booster. The booster and fuel gas generator can be operated in normal sequence, or operated at the same time in order to increase the thrust produced for short-range missions. A logic circuit contained on the rocket or missile determines a time to rupture the frangible diaphragm based on whether or not the distance to the target exceeds a threshold distance.

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 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.