A rocket motor has an electrically operated propellant initiator for a propellant grain that includes an electrode arrangement configured to concentrate an electric field at an ignition electrode for igniting an electrically operated propellant. The rocket motor includes a combustion chamber containing at least one propellant grain and an electrically operated propellant initiator operatively coupled to the propellant grain to initiate combustion of the propellant grain. The electrically operated propellant initiator includes the electrically operated propellant and at least one pair of electrodes configured to ignite the electrically operated propellant. The pair of electrodes includes a ground plane electrode and an ignition electrode. When an electrical input is applied to the electrically operated propellant initiator, the electric field is concentrated at the ignition electrode to ignite the electrically operated propellant at the location where the ignition electrode is arranged.

A rocket motor has an electrically operated propellant initiator for a propellant grain that includes an electrode arrangement configured to concentrate an electric field at an ignition electrode for igniting an electrically operated propellant. The rocket motor includes a combustion chamber containing at least one propellant grain and an electrically operated propellant initiator operatively coupled to the propellant grain to initiate combustion of the propellant grain. The electrically operated propellant initiator includes the electrically operated propellant and at least one pair of electrodes configured to ignite the electrically operated propellant. The pair of electrodes includes a ground plane electrode and an ignition electrode. When an electrical input is applied to the electrically operated propellant initiator, the electric field is concentrated at the ignition electrode to ignite the electrically operated propellant at the location where the ignition electrode is arranged.

An ignition safety device (ISD) used in an ignition system of a missile is configured to selectively control the ignition of two or more pulses or stages of a rocket motor propulsion system, based on a flight profile mode selection of a flight velocity mode, in which the missile is configured to travel at an optimized flight velocity, or a flight distance mode, in which the missile is configured to travel an optimized flight distance. The ISD is configured to selectively ignite the pulses or stages substantially simultaneously upon selection of the flight velocity mode, or in a delayed sequential manner upon selection of the flight distance mode. The ISD is also configured to selectively inhibit the delayed sequential ignition of the pulses or stages in the event of incidental ground or water impact of the missile after ignition of the primary pulse or stage.

An ignition safety device (ISD) used in an ignition system of a missile is configured to selectively control the ignition of two or more pulses or stages of a rocket motor propulsion system, based on a flight profile mode selection of a flight velocity mode, in which the missile is configured to travel at an optimized flight velocity, or a flight distance mode, in which the missile is configured to travel an optimized flight distance. The ISD is configured to selectively ignite the pulses or stages substantially simultaneously upon selection of the flight velocity mode, or in a delayed sequential manner upon selection of the flight distance mode. The ISD is also configured to selectively inhibit the delayed sequential ignition of the pulses or stages in the event of incidental ground or water impact of the missile after ignition of the primary pulse or stage.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

An injection device for injecting an oxidizer for a liquid rocket includes a housing, a plate disposed inside the housing and having an injection hole to eject an oxidizer, a duct disposed above the plate to guide the oxidizer, and a manifold with one end connected to the injection hole of the plate and the other end connected to the duct, wherein the oxidizer may be distributed to the injection hole at an equal flow rate.

An injection device for injecting an oxidizer for a liquid rocket includes a housing, a plate disposed inside the housing and having an injection hole to eject an oxidizer, a duct disposed above the plate to guide the oxidizer, and a manifold with one end connected to the injection hole of the plate and the other end connected to the duct, wherein the oxidizer may be distributed to the injection hole at an equal flow rate.

A multi-pulse propulsion system includes at least one pulse chamber containing at least one propellant for igniting during at least one pulse of the multi-pulse propulsion system, at least one additional pulse chamber containing at least one additional propellant for igniting during at least one additional pulse of the multi-pulse propulsion system, and at least one passive fuzing system configured to initiate the at least one additional pulse. The at least one passive fuzing system includes a sensor and an igniter. The sensor is configured to sense an environmental condition and/or a ballistic condition. The igniter is configured to provide a stimulus that causes ignition of the at least one additional propellant in response to the sensor sensing that the environmental condition and/or the ballistic condition has reached or exceeded one or more threshold values.

A multi-pulse propulsion system includes at least one pulse chamber containing at least one propellant for igniting during at least one pulse of the multi-pulse propulsion system, at least one additional pulse chamber containing at least one additional propellant for igniting during at least one additional pulse of the multi-pulse propulsion system, and at least one passive fuzing system configured to initiate the at least one additional pulse. The at least one passive fuzing system includes a sensor and an igniter. The sensor is configured to sense an environmental condition and/or a ballistic condition. The igniter is configured to provide a stimulus that causes ignition of the at least one additional propellant in response to the sensor sensing that the environmental condition and/or the ballistic condition has reached or exceeded one or more threshold values.