A turbojet engine capable of operation in an Air Turbo Rocket (ATR) mode includes a compressor, a rotatable turbine wheel comprising turbine blades, a non-rotating guide vane ring comprising guide vanes, a turbine shaft configured to power said compressor, a combustor, a gas generator, and a main combustor. The main combustor is configured to combust hot, fuel rich gas from the gas generator in air compressed by the compressor. Hot, fuel rich gas from the gas generator is directed towards the turbine blades by a directing means.

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).

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 engine system having a heating system configured to heat an oxidizer; a combustion section having a flow path from an upstream inlet section through a restricted throat section to a downstream outlet section, the combustion section configured to accelerate the heated oxidizer as an oxidizer stream within the flow path in response to flow dynamics to supersonic speed; and a fuel system configured to introduce a fuel into the flow path to mix supersonically with the heated oxidizer to define a combined fuel and oxidizer stream at a first supersonic speed. The combined fuel and oxidizer stream undergoes a deflagration to denotation transition in the combustion section defined by an oblique shock wave and a normal shock wave that interact to achieve a standing detonation wave generally at an upstream portion of the restricted throat section configured such that combustion exits the downstream outlet section to provide thrust.

A rocket engine system having a heating system configured to heat an oxidizer; a combustion section having a flow path from an upstream inlet section through a restricted throat section to a downstream outlet section, the combustion section configured to accelerate the heated oxidizer as an oxidizer stream within the flow path in response to flow dynamics to supersonic speed; and a fuel system configured to introduce a fuel into the flow path to mix supersonically with the heated oxidizer to define a combined fuel and oxidizer stream at a first supersonic speed. The combined fuel and oxidizer stream undergoes a deflagration to denotation transition in the combustion section defined by an oblique shock wave and a normal shock wave that interact to achieve a standing detonation wave generally at an upstream portion of the restricted throat section configured such that combustion exits the downstream outlet section to provide thrust.

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 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.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

Closed-loop refrigeration cycles for liquid oxygen densification are disclosed. The disclosed refrigeration cycles may be turbine-based refrigeration cycles or a Joule-Thompson (JT) expansion valve based refrigeration cycles and include a refrigerant or working fluid comprising a mixture of neon or helium together with nitrogen and/or oxygen.