The subject of the invention is a detonation rocket engine comprising an annular detonation chamber (5) connected to the Aerospike nozzle (4) and lines (2, 3) for supplying propellant components connected to the detonation chamber (5). The detonation chamber (5) has a bottom (9) connecting the inner wall (10) and the outer wall (11) between which the outlet (6) is formed. At the outlet (6) of the detonation chamber (5) there are at least three evenly distributed centring elements (1) connecting the inner wall (10) and the outer wall (11) of the detonation chamber (5), with cooling channels (7) connected to one of the lines (2, 3) supplying the propellant components to the detonation chamber (5).

A rocket propulsion system comprises a combustion chamber, an oxygen supply system, comprising an oxygen supply duct and being configured to supply oxygen to the combustion chamber, and a hydrogen supply system, comprising a hydrogen supply duct and being configured to supply hydrogen to the combustion chamber. An ignition unit of the propulsion system, to which at least portions of the oxygen and the hydrogen supplied to the combustion chamber can be supplied, is configured to initiate combustion of the oxygen-hydrogen mixture in the combustion chamber. The propulsion system further comprises a cooling duct extending along an inner surface of a combustion chamber wall and through which at least a portion of the oxygen supplied to the combustion chamber, at least a portion of the hydrogen supplied to the combustion chamber or a combustion gas mixture emerging from the ignition unit flows.

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.

Power device based on alkali-water reaction, having a reaction chamber to carry out a chemical reaction between water and alkali element, having an exhaust nozzle to exhaust the reaction products generated in the reaction chamber, with a siphon tube connected to the exhaust nozzle, a nozzle shutter plate sealing the exhaust nozzle, and an external pressure inlet. The device has a water reservoir to store transfer water to the reaction chamber and an alkali reservoir to store and transfer an alkali element to the reaction chamber, and transfer device connecting the reaction chamber to both reservoirs to transmit the pressure generated in reaction chamber by part of the reaction products to the reservoirs, providing the transfer of water and alkali element from both reservoirs to the reaction chamber.

A fuel-free spacecraft propelling system having an open-ended outer cylinder of a propelling device and an atomic oxygen collecting device is disclosed. The latter is arranged at the forwardly-propelled front end of the outer cylinder and is hermetically connected with an RF generating device and an ion cyclotron wave heating device through a magnetic confinement device. A spiral wave discharge oxygen plasma inlet and a spiral wave discharge oxygen plasma outlet in the ion cyclotron wave heating device are respectively provided with another magnetic confinement device. The propulsion of the invention does not need to carry the propellant, which greatly reduces the launch costs, and enables a spacecraft to advantageously have an increased orbit life over existing spacecraft systems.

A rocket propulsion system comprises a combustion chamber, a hydrogen-oxygen supply system connected to the combustion chamber, which hydrogen-oxygen supply system is configured to conduct hydrogen and oxygen into the combustion chamber, and a coolant supply system connected to the combustion chamber, which coolant supply system is configured to conduct a combustible coolant into the combustion chamber. An ignition system of the rocket propulsion system is configured to initiate combustion of the hydrogen-oxygen-coolant mixture in the combustion chamber.

A method for operating a rocket propulsion system comprises the steps of supplying oxygen to a combustion chamber, supplying hydrogen to the combustion chamber and combusting the oxygen-hydrogen mixture in the combustion chamber. The rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.

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 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.