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 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.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

An igniting device for igniting a mixture, in particular for an engine, comprises an energy converting device and a fluid flow injecting device. The energy converting device is configured for converting fluid flow energy of at least one fluid flow into heat, thereby igniting the mixture. The energy converting device comprises an ignition chamber for the at least one fluid flow. The fluid injecting device is configured for injecting a plurality of fluid flows into the ignition chamber. The injection takes place such that a first fluid flow is injected into the ignition chamber with a higher fluid flow velocity than a second fluid flow.

An igniting device for igniting a mixture, in particular for an engine, comprises an energy converting device and a fluid flow injecting device. The energy converting device is configured for converting fluid flow energy of at least one fluid flow into heat, thereby igniting the mixture. The energy converting device comprises an ignition chamber for the at least one fluid flow. The fluid injecting device is configured for injecting a plurality of fluid flows into the ignition chamber. The injection takes place such that a first fluid flow is injected into the ignition chamber with a higher fluid flow velocity than a second fluid flow.

A rocket engine pintle injector with optimized spray pattern and with integrated igniter design for providing construction simplicity, throttleable thrust, stop/start/restart capability, optimized operational combustion, and improved ignition combustion stability. A user can start, throttle, and stop the engine by moving the internal concentric injector sleeve forward and backward to cause the fuel/oxidizer to spray out of the pintle head at different flow rates. The concentric igniter can be deployed so that the hot gasses or spark produced are radially projected into the spray of fuel/oxidizer surrounding the igniter. Once the fuel/oxidizer spray has been ignited, the igniter is stopped and retracted to protect the device from the heat of the combustion chamber and is ready for redeployment and restart of the engine as needed. Thus, a versatile, fully integrated, and scalable device can be used to start, throttle, stop, and restart any size rocket engine during any mission phase from launch to return from space.

A rocket engine pintle injector with optimized spray pattern and with integrated igniter design for providing construction simplicity, throttleable thrust, stop/start/restart capability, optimized operational combustion, and improved ignition combustion stability. A user can start, throttle, and stop the engine by moving the internal concentric injector sleeve forward and backward to cause the fuel/oxidizer to spray out of the pintle head at different flow rates. The concentric igniter can be deployed so that the hot gasses or spark produced are radially projected into the spray of fuel/oxidizer surrounding the igniter. Once the fuel/oxidizer spray has been ignited, the igniter is stopped and retracted to protect the device from the heat of the combustion chamber and is ready for redeployment and restart of the engine as needed. Thus, a versatile, fully integrated, and scalable device can be used to start, throttle, stop, and restart any size rocket engine during any mission phase from launch to return from space.

A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.