Device and method for regulating a flow rate of gas intended to supply a propulsion apparatus for a spacecraft comprising xenon tank, a circuit comprising a withdrawing pipe having an upstream end connected to the tank and a downstream end connected to a propulsion member, the withdrawing pipe comprising an isolation first valve, a regulating second valve and a member for measuring the pressure downstream of the regulating second valve. The regulating second valve regulates the flow rate and/or the determined pressure according to the pressure measured. The regulating second valve is a proportional valve of electrically operated variable throughout PCV type.

A turbine pump assembly has a turbine, a centrifugal pump, a passive electrical speed control system, and a pneumatic circuit breaker. The pneumatic circuit breaker has a plurality of elements that are configured to move to a position blocking an outlet duct of the turbine when a flow velocity exceeds a predetermined threshold. A rocket thrust vector control system is also disclosed.

A turbine pump assembly has a turbine, a centrifugal pump, a passive electrical speed control system, and a pneumatic circuit breaker. The pneumatic circuit breaker has a plurality of elements that are configured to move to a position blocking an outlet duct of the turbine when a flow velocity exceeds a predetermined threshold. A rocket thrust vector control system is also disclosed.

A rocket motor and rocket motor feed system are disclosed. The rocket motor feed system includes a sonic choke which passively regulates the mass flow rate of gaseous propellant passing through the sonic choke. An injector is provided and isolates the upstream feed line of the rocket motor feed system from a combustor. Regenerative cooling circuits are disclosed. Self-pressurised gaseous propellants may be used with the rocket motor and rocket motor feed system. Suitable propellants are disclosed. Bi-propellants may be used. The sonic choke may provide a ratio of oxidiser:fuel to a combustor. Rocket motor feed systems with separate fuel and oxidiser branches are also disclosed. A rocket motor utilising such a feed system is disclosed.

A distributed fuel module includes a fuel pressure vessel with a gas port and a fuel port, a hydraulic circuit breaker connected to the fuel port, and a gaseous circuit breaker. The gaseous circuit breaker is connected to the gas port, is fluidly coupled to the hydraulic circuit breaker through the fuel pressure vessel, and is cooperatively associated with the gaseous circuit breaker to isolate the fuel pressure vessel from a compressed gas header and a fuel header according to pressure differential within the hydraulic circuit breaker and pressure differential within the gaseous circuit breaker. Power modules and methods of controlling fuel flow in fuel modules are also described.

A system includes a flow inlet conduit. A primary conduit branches from the flow inlet conduit for delivering flow to a set of primary nozzles. An equalization bypass valve (EBV) connects between the flow inlet conduit and a secondary conduit for delivering flow to a set of secondary nozzles. The EBV is connected to be controlled to apportion flow from the flow inlet conduit to the secondary conduit. A regulator valve is connected to be controlled by a pressure differential between pressure in the flow inlet conduit (PIN) and pressure in the secondary conduit (PSEC) to apportion flow from an inlet of the regulator valve to the flow inlet conduit.

According to one embodiment, there is provided a propulsion apparatus of liquid propellant rocket engine. The propulsion apparatus of liquid propellant rocket engine, the propulsion apparatus including: a body in which liquid propellant flows; an injector core located inside the body; at least one outlet connected to the injector core to discharge combustion gas; and an injector for discharging the liquid propellant flowing into the body, wherein the injector is located in an area adjacent to the outlet, wherein the liquid propellant moves between a frame of the body and a frame of the injector core.

According to one embodiment, there is provided a propulsion apparatus of liquid propellant rocket engine. The propulsion apparatus of liquid propellant rocket engine, the propulsion apparatus including: a body in which liquid propellant flows; an injector core located inside the body; at least one outlet connected to the injector core to discharge combustion gas; and an injector for discharging the liquid propellant flowing into the body, wherein the injector is located in an area adjacent to the outlet, wherein the liquid propellant moves between a frame of the body and a frame of the injector core.

A system and methods are disclosed for an upper stage space launch vehicle that uses gases from the propellant tanks to power an internal combustion engine that produces mechanical power for driving other components including a generator for generation of electrical current for operating compressors and fluid pumps and for charging batteries. These components and others comprise a thermodynamic system from which system enthalpy may be leveraged by extracting and moving heat to increase the efficient use of propellant and the longevity and performance of the launch vehicle.

A system and methods are disclosed for an upper stage space launch vehicle that uses gases from the propellant tanks to power an internal combustion engine that produces mechanical power for driving other components including a generator for generation of electrical current for operating compressors and fluid pumps and for charging batteries. These components and others comprise a thermodynamic system from which system enthalpy may be leveraged by extracting and moving heat to increase the efficient use of propellant and the longevity and performance of the launch vehicle.