An electric power-assisted liquid-propellant rocket propulsion system includes: a pre-burner for generating combustion gas of a fuel and an oxidant; a main combustor for burning mixed gas of the fuel and the combustion gas discharged from the pre-burner; a turbopump including a turbine rotated by a flow of the combustion gas and a first pump and a second pump driven by a rotation of the turbine, the turbopump supplying the fuel from a fuel tank to the pre-burner and supplying the oxidant from an oxidant tank to the pre-burner and the main combustor; an electric motor for rotating the turbine before combustion of the pre-burner and the main combustor; and a clutch for connecting the electric motor and the turbine and releasing the connection between the electric motor and the turbine.

A turbine pump assembly has a turbine, a centrifugal pump, and a passive electrical speed control system. The turbine has a peak efficiency at a first speed that is lower than a second speed at which the centrifugal pump is operating at a peak power requirement. A rocket thrust vector control system is also disclosed.

A system for feeding a liquid fuel and a liquid oxidiser to combustion chamber of an engine, includes two containers kept under a low temperature, one for holding liquid hydrogen, and the other for holding liquid oxygen. Each of two expansion valves is connected to one separate container of said two containers, for transforming liquid hydrogen and liquid oxygen which pass through said valves, into gaseous forms. A fuel cell is connected to the valves for receiving the gaseous hydrogen and gaseous oxygen, and generating electric power. The system includes two other containers, one for holding the fuel, and the other for holding the oxidizer. Two electric pumps are connected to the two other containers, such that one of the pumps is capable of feeding the fuel, and the other is capable of feeding the oxidizer to the chamber, using the power generated by the fuel cell.

A flow control system (22) includes a fuel passage network (34) that has first (36) and second (38) network portions that are in a parallel flow arrangement with each other. A fueldraulic device (40) is located in the first network portion. Operation of the fueldraulic device varies flow through the first network portion. A flow restriction orifice (42) is located in the fuel passage network and is arranged in series with, and upstream of, the fueldraulic device. The flow restriction orifice is operable to generate a pressure differential that varies responsive to the flow through the first network portion. A flow control valve (44) is located in the second network portion. The flow control valve is operable responsive to the pressure differential across the flow restriction orifice to control flow through the second network portion.

An improved liquid-propellant engine utilizing a combination of a gas turbine, a generator, an electric pump fed engine, and a power management system. When in use, the gas turbine and generator are configured for providing power to an electric motor for operating a fuel pump and an oxidizer pump to pump fuel and oxidizer from individual tanks into a combustion chamber for mixture and combustion. The combustion of the fuel and oxidizer is configured for providing thrust from the base of a rocket to propel it. The generator of the present invention provides continuous electrical power to the electric pump(s) to continue the supply of fuel and oxidizer to the combustion chamber. The power management system of the present invention is electrically connected to the generator, the electric motor(s) configured for operating the fuel and oxidizer pumps, a flight computer, and fuel and oxidizer valves.

Disclosed herein are various technologies pertinent to rocket engines, including injector, thrust chamber, and electrical turbopump devices that may be combined to provide a more efficient rocket engine.

A method of regulating pressure within a first propellant tank of a rocket engine having a first propellant tank containing a first propellant and a second propellant tank containing a second propellant, and a regulator device for regulating pressure within the first tank, the regulator device comprising a gas generator and a heat exchanger co-operating with the gas generator so as to vaporize at least part of the first propellant prior to reintroducing it into the first tank (16), the gas generator and the heat exchanger both being fed with the first propellant by a single first motor-driven pump, while the gas generator is fed with the second propellant by a single second motor-driven pump, wherein the flow rate of the first motor-driven pump is controlled as a function of a first parameter, while the flow rate of the second motor-driven pump is controlled as a function of a second parameter.

The invention relates to the field of rocket engines, and in particular to a method and a circuit for regulating a rocket engine (1) comprising at least one combustion chamber (3) and a first liquid propellant feed circuit (10) with a first pump (12) for pumping a first liquid propellant and a first turbine (14) for actuating the first pump (12), a first feed valve (15), and a regulator device for regulating the first turbine (14). In the regulation method and circuit, an opening command (DVCH) for opening the first feed valve (15) is calculated from an external command (C.sub.ext) in application of an open loop control relationship; a command for the regulator device of the first turbine (14) is calculated on the basis of said external command (C.sub.ext) and at least one feedback value, in application of a closed loop control relationship; and the first feed valve (15) and the regulator device of the first turbine (14) are controlled in application of the respective commands.