A high density, generally recognized as safe hybrid rocket motor is described which has a density-specific impulse similar to a solid rocket motor, with good performance approaching or equal to a liquid rocket motor. These high density hybrid motors resolve the packaging efficiency/effectiveness problems limiting the application of safe, low cost hybrid motor technology.

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).

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).

Apparatus for trajectory control and/or position control of a missile (99), comprising a controllable gas generator (109, 200) with a fuel flow control valve (124, 213), an injector head (112, 202), a combustion chamber (111) and at least one outflow nozzle (103, 204) or at least one throttle.

Apparatus for trajectory control and/or position control of a missile (99), comprising a controllable gas generator (109, 200) with a fuel flow control valve (124, 213), an injector head (112, 202), a combustion chamber (111) and at least one outflow nozzle (103, 204) or at least one throttle.

A rocket engine has: a combustion chamber having a chamber inlet for receiving an oxidizer and a chamber outlet for expelling combustion gases in an environment outside the combustion chamber; a manifold having a manifold inlet fluidly connectable to a source of the oxidizer and a manifold outlet; a catalyst having a catalyst inlet fluidly connected to the manifold outlet and a catalyst outlet; and an injector plate having a injector inlet fluidly connected to the catalyst outlet and an injector outlet fluidly connected to the chamber inlet.

A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.

A hybrid rocket motor includes a solid fuel element, and an oxidizer tank containing an oxidizer. The solid fuel element adjoins and at least partially defines a combustion chamber in which the solid fuel and the oxidizer are burned, to produce thrust from the hybrid rocket motor. The oxidizer tank is at least partially within the combustion chamber, and the entire oxidizer tank may be within the combustion chamber. The oxidizer tank may be protected by an insulating material, which may also serve as a structural material that contains the pressure of the oxidizer. The insulating material and the fuel material may both be polymer-based materials, although they may be different materials having different characteristics, for example including different additives to the same polymer material. The fuel element and the oxidizer tank may be made by additive manufacturing processes, for example by adding different materials in different locations.

A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer. Combustion of the fuel material in the combustion chamber may generate thrust of no more than 5 N at an oxidizer flow rate of no more than 5 g/s.

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.