A hybrid metal composite (HMC) structure comprises tiers comprising fiber composite material structures, and additional tiers longitudinally adjacent one or more of the tiers and comprising perforated metallic structures and additional fiber composite material structures laterally adjacent the perforated metallic structures. Methods of forming an HMC structure, and related rocket motors and multi-stage rocket motor assemblies are also disclosed.

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.

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.

A readily combustible portion (110) includes a readily combustible exposed surface (111) that is exposed to a flow channel (CA). A combustion-resistant portion (140), which comprises a material that is more resistant to combustion than the readily combustible portion (110), covers an outer surface of the readily combustible portion (110) on the opposite side from the readily combustible exposed surface (111) in a direction orthogonal to a length direction parallel to a direction in which a hybrid rocket is propelled. The combustion-resistant portion (140) includes a thick portion (120) that serves as a stopper that prevents peeling of the readily combustible portion (110) from the combustion-resistant portion (140) in a direction from a starting end surface (100a) toward a terminating end surface (100b).

Systems and methods for determining bi-propellant volume and adjusting a mixture ratio are discussed herein. A controller can calculate an adjusted mixture ratio and command the rocket engine to implement the adjusted mixture ratio by opening or closing valves of the propellant tanks, which changes the volumetric flow rates of each of the propellants. The adjusted mixture ratio can be calculated by an algorithm based on sensed or calculated data associated with each propellant. The adjusted mixture ratio can be used to evenly deplete the propellants to reduce the amount of each propellant remaining after a mission and to improve propellant use, which allows for an increase in a non-propellant payload.

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.

Rocket propulsion systems and associated methods are disclosed. A representative system includes a combustion chamber having an inwardly-facing chamber wall enclosing a combustion zone. The chamber has a generally spherical shape and is exposed to the combustion zone. A propellant injector is coupled to the combustion chamber and has at least one fuel injector nozzle positioned to direct a flow of cooling fuel radially outwardly along the inwardly-facing chamber wall. In addition to or in lieu of the foregoing features, the injector can include an oxidizer piston and a fuel piston that deliver oxidizer and fuel, respectively, to the combustion chamber, in a sequenced manner so that the oxidizer is introduced prior to the fuel.

A novel approach provides a small satellite propulsion system that uses vapor to generate thrust for the small satellite. The vapor naturally sits on top of liquid propellant(s), which are stored within a propellant tank. The vapor may flow from the propellant tank and through a membrane to interact with a reacting surface to generate thrust.

Embodiments of the present invention generally relate to a vapor retention device and methods of using a vapor retention device to manage propellant for upper stage space vehicles. The use of a vapor retention device, in combination with controlled acceleration, drives liquid propellant from a propellant supply line communicating with an upper stage main engine back into a propellant tank and establishes an insulating liquid/gas propellant interface that prevents the exchange of gaseous propellant across the interface.