A liquid rocket engine integrates tap-off openings at a combustion chamber wall to direct exhaust from the combustion chamber to a tap-off manifold that provides the exhaust to one or more auxiliary systems, such as a turbopump that pumps oxygen and/or fuel into the combustion chamber. The tap-off opening passes through a fuel channel formed in that combustion chamber exterior wall and receives fuel through a fuel opening that interfaces the fuel channel and tap-off opening. The tap-off manifold nests within a fuel manifold for thermal management. The fuel channel directs fuel into the combustion chamber through fuel port openings formed in the combustion chamber, the fuel port openings located closer to a headend of the combustion chamber than the tap-off openings.

A combustor of a liquid rocket engine includes a nozzle unit including a regenerative cooling channel, in which the nozzle unit includes a fuel manifold outer shell, a combustor inner shell, and a combustor outer shell having a downward channel inlet, and the combustor includes a fuel inlet connected to a nozzle neck of the nozzle unit, a fuel manifold formed between the fuel manifold outer shell and the combustor outer shell, and in which fuel introduced from the fuel inlet flows, a downward channel connected in communication with the fuel manifold through the downward channel inlet, and extending in a downward direction from an upper portion of the combustor, a diverting manifold provided at a distal end of the nozzle unit and connected in communication with the downward channel, and an upward channel connected in communication with the diverting manifold and extending in an upward direction of the combustor.

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.

A hybrid propulsion unit, for a space vehicle or rocket, includes an external body defining a combustion chamber configured to receive a block of solid propellant, a pressurized reservoir configured to receive a liquid propellant, and a nozzle to eject the combustion gases produced by the reaction of the propellants. The pressurized reservoir being placed inside the external body, surrounded by the block of solid propellant. The propulsion unit further includes a plurality of liquid-propellant injectors disposed axially between parts of the solid propellant to improve the efficacy of the combustion. A space vehicle, of the launch vehicle type, equipped with such hybrid propulsion unit.

A hybrid propulsion unit, for a space vehicle or rocket, includes an external body defining a combustion chamber configured to receive a block of solid propellant, a pressurized reservoir configured to receive a liquid propellant, and a nozzle to eject the combustion gases produced by the reaction of the propellants. The pressurized reservoir being placed inside the external body, surrounded by the block of solid propellant. The propulsion unit further includes a plurality of liquid-propellant injectors disposed axially between parts of the solid propellant to improve the efficacy of the combustion. A space vehicle, of the launch vehicle type, equipped with such hybrid propulsion unit.

A hybrid rocket engine with a vortex flow field injection system that produces a high-speed sustained vortex flow field is described. The hybrid rocket engine includes a generally cylindrical injection chamber with an inner circumference to comprise an outer edge of a solid propellant grain in the hybrid rocket engine. The engine also includes an injection system that has a throttle valve and an injector that injects injection fluid into the engine and produces a vortex flow-field for the injected fluid. The injector includes at least one primary feed line that distributes the injection fluid throughout a pre-swirl chamber and multiple orifices along an inner edge of the injection chamber. The pre-swirl chamber connects to the injection chamber and at least one of the primary feed lines and redirects a primary fluid flow of the injected fluid from a primary axial direction to a centrifugal direction.

A hybrid rocket engine with a vortex flow field injection system that produces a high-speed sustained vortex flow field is described. The hybrid rocket engine includes a generally cylindrical injection chamber with an inner circumference to comprise an outer edge of a solid propellant grain in the hybrid rocket engine. The engine also includes an injection system that has a throttle valve and an injector that injects injection fluid into the engine and produces a vortex flow-field for the injected fluid. The injector includes at least one primary feed line that distributes the injection fluid throughout a pre-swirl chamber and multiple orifices along an inner edge of the injection chamber. The pre-swirl chamber connects to the injection chamber and at least one of the primary feed lines and redirects a primary fluid flow of the injected fluid from a primary axial direction to a centrifugal direction.

Various embodiments of a vortex hybrid motor are described herein. In some embodiments, the vortex hybrid motor may include a housing with a solid propellant positioned within the housing, and an injector ring positioned at a proximal end of the housing. The injector ring can include a plurality of angled injector units each including a first injector and a second injector angled towards an impingement point. A first fluid stream of a liquid propellant dispensed from the first injector can collide with a second fluid stream of the liquid propellant dispensed from the second injector to atomize the liquid propellant and form a spray fan formation. At least one of the first injector and the second injector can be positioned at an injection angle relative to the sidewall to create a swirl flow of the atomized injector fluid

The combustor for a rocket engine includes a combustion room configured to cause a combustion reaction between a fuel and an oxidant, an injector configured to inject the fuel and the oxidant into the combustion room, and a nozzle skirt configured to inject combustion gas generated by the combustion reaction to an outside, and an inertance increasing portion configured to increase an equivalent inertance in a vibration equivalent circuit of the combustor for the rocket engine.