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.

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.

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.

A fuel grain for a hybrid rocket engine includes multiple, concentric cylindrical layers of fuel grain material defining a combustion port extending axially through the fuel grain, in which each layer includes multiple beads of fuel grain material, in which the multiple beads in a given layer are disposed adjacent to one another and bonded together, and in which adjacent concentric layers are bonded together. Each bead of fuel grain material includes a polymer based rocket fuel material and a nanoscale metallic material, and a composition of the beads of the fuel grain material varies along a radius of the cylindrical fuel grain.

A fuel grain for a hybrid rocket engine includes multiple, concentric cylindrical layers of fuel grain material defining a combustion port extending axially through the fuel grain, in which each layer includes multiple beads of fuel grain material, in which the multiple beads in a given layer are disposed adjacent to one another and bonded together, and in which adjacent concentric layers are bonded together. Each bead of fuel grain material includes a polymer based rocket fuel material and a nanoscale metallic material, and a composition of the beads of the fuel grain material varies along a radius of the cylindrical fuel grain.

A method of making a fuel grain for a hybrid rocket engine includes deposing beads of fuel grain material onto mandrel using additive manufacturing to form a cylindrical fuel grain, each bead including a polymer based rocket fuel material and a nanoscale metallic material. The deposing includes deposing multiple, adjacent beads to form concentric layers of beads, wherein a composition of the beads of the fuel grain material differs between the beads of a first layer and the beads of a second layer of the fuel grain.

A method of making a fuel grain for a hybrid rocket engine includes deposing beads of fuel grain material onto mandrel using additive manufacturing to form a cylindrical fuel grain, each bead including a polymer based rocket fuel material and a nanoscale metallic material. The deposing includes deposing multiple, adjacent beads to form concentric layers of beads, wherein a composition of the beads of the fuel grain material differs between the beads of a first layer and the beads of a second layer of the fuel grain.

A rocket motor is disclosed that can include a combustion chamber containing a solid fuel that is operable to burn during operation of the rocket motor to generate combustion gas and unburned gaseous fuel. The rocket motor can also include a propellant supply containing an energy-rich oxidizer with a decomposition energy greater than or equal to 1.0 MJ/kg. In addition, the rocket motor can include a thrust augmented nozzle (TAN) operably coupled to the combustion chamber to receive the combustion gas from the combustion chamber and direct a flow of the combustion gas through the TAN. The TAN can have a divergent portion downstream of a throat, and a propellant injection port associated with the divergent portion and in communication with the propellant supply to inject the energy-rich oxidizer into the divergent portion. Only the energy-rich oxidizer, independent of another propellant, may be introduced into the flow of the combustion gas and the unburned gaseous fuel for secondary combustion of the unburned gaseous fuel and thermal decomposition of the energy-rich oxidizer within the divergent portion.

A rocket motor is disclosed that can include a combustion chamber containing a solid fuel that is operable to burn during operation of the rocket motor to generate combustion gas and unburned gaseous fuel. The rocket motor can also include a propellant supply containing an energy-rich oxidizer with a decomposition energy greater than or equal to 1.0 MJ/kg. In addition, the rocket motor can include a thrust augmented nozzle (TAN) operably coupled to the combustion chamber to receive the combustion gas from the combustion chamber and direct a flow of the combustion gas through the TAN. The TAN can have a divergent portion downstream of a throat, and a propellant injection port associated with the divergent portion and in communication with the propellant supply to inject the energy-rich oxidizer into the divergent portion. Only the energy-rich oxidizer, independent of another propellant, may be introduced into the flow of the combustion gas and the unburned gaseous fuel for secondary combustion of the unburned gaseous fuel and thermal decomposition of the energy-rich oxidizer within the divergent portion.

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.