A rocket nozzle is made from an optimized metal lattice structure, with a hardened material applied onto the metal lattice structure so as to coat the structure and fill voids in the lattice by chemical vapor deposition. A rocket nozzle is further provided having one or more bypass lines for taking expanding gas from a combustion chamber of the rocket nozzle and redirecting the expanding gas to a skirt of the rocket nozzle to thereby manipulate the shape of a plume of expanding gas exiting the rocket nozzle. A rocket nozzle is also provided having one or more main injectors extending into the combustion chamber for injecting fuel for combustion into the combustion chamber and one or more opposing injectors oriented to direct an opposing flow of energy and gas expansion towards the main injectors, or having one or more secondary injectors arranged around the combustion chamber proximal the throat.

A rocket catapult assembly for an ejection seat may comprise a drive motor, a metering tube, a first cartridge, and a second cartridge. The metering tube may include an outer wall having a gas pervious section and a gas impervious section. The drive motor may be configured to translate the metering tube and align the gas pervious section or gas impervious section with a first cartridge and a second cartridge to produce a desired thrust of the rocket catapult assembly.

A rocket catapult assembly for an ejection seat may comprise a drive motor, a metering tube, a first cartridge, and a second cartridge. The metering tube may include an outer wall having a gas pervious section and a gas impervious section. The drive motor may be configured to translate the metering tube and align the gas pervious section or gas impervious section with a first cartridge and a second cartridge to produce a desired thrust of the rocket catapult assembly.

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 attitude control and thrust boosting system (100) for a space launcher is disclosed, wherein the space launcher is equipped with a rocket engine (303) provided with an exhaust nozzle. The exhaust nozzle comprises a divergent portion (302) so designed as to make a supersonic gas flow exit through an exit section defined by a given angle of divergence with respect to a longitudinal axis of the rocket engine. The attitude control and thrust boosting system (100) comprises flaps (110, 111, 112, 113) that are arranged around the exit section, are shaped so as to extend the divergent portion of the exhaust nozzle, are mechanically decoupled from said exhaust nozzle and can be actuated to take different angular positions with respect to the longitudinal axis of the rocket engine. Control means (130) are also provided to receive quantities indicative of an actual attitude of the space launcher and an ambient static pressure, and to make the flaps (110,111,112,113) take a neutral angular position where the flaps (110,111,112,113) are inclined, with respect to the longitudinal axis of the rocket engine, according to an inclination angle greater than, or equal to, the given angle of divergence, in order to control the neutral angular position taken by the flaps (110,111,112,113) according to the ambient static pressure and to make one or more flaps (110,111,112,113) take an angular position different than the neutral angular position according to the actual attitude of the space launcher and to a required attitude for said space launcher.

A method of controlling a multi-engine bay in which the following steps are performed: a) controlling the bay so that it delivers desired thrust and each engine is operated in compliance with a set of operating limits for the engine; b) periodically evaluating a level of damage for each of the engines, the level of damage of an engine being information representative of a probability of the engine failing; c) for each engine, periodically evaluating whether its level of damage exceeds a predetermined value; and d) if the level of damage of an engine, referred to as a “damaged” engine, exceeds a predetermined value, modifying at least one operating limit of the damaged engine so that the rate of damage of the damaged engine is less than a predetermined maximum rate of engine damage.

A method of controlling a multi-engine bay in which the following steps are performed: a) controlling the bay so that it delivers desired thrust and each engine is operated in compliance with a set of operating limits for the engine; b) periodically evaluating a level of damage for each of the engines, the level of damage of an engine being information representative of a probability of the engine failing; c) for each engine, periodically evaluating whether its level of damage exceeds a predetermined value; and d) if the level of damage of an engine, referred to as a “damaged” engine, exceeds a predetermined value, modifying at least one operating limit of the damaged engine so that the rate of damage of the damaged engine is less than a predetermined maximum rate of engine damage.

An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.

An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.

Mining apparatuses, systems and methods related to the use of a rocket engine's plume and a collection manifold to efficiently displace, collect, process and store frozen volatiles embedded within or below a surface is disclosed. The plume contacts and churns up the surface. The frozen volatiles are displaced and/or evaporated within a closed environment under a collection manifold. The collection manifold has related components for addressing these frozen or gaseous volatiles downstream. Various apparatuses and subsystems are also disclosed including a rover, processing plants, collection manifold, and vapor manifold.