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.

A combustion chamber including a diverging portion. The combustion chamber extends along a longitudinal axis and includes a fluid injection system from which there extends in a downstream direction a wall presenting a throat and a diverging portion situated downstream from the throat. The chamber further includes a tubular element surrounding the wall at least in part and configured to take up most of forces generated during operation of the chamber on the downstream end of the wall to transfer the forces to a structure situated upstream from the chamber.

A combustion chamber including a diverging portion. The combustion chamber extends along a longitudinal axis and includes a fluid injection system from which there extends in a downstream direction a wall presenting a throat and a diverging portion situated downstream from the throat. The chamber further includes a tubular element surrounding the wall at least in part and configured to take up most of forces generated during operation of the chamber on the downstream end of the wall to transfer the forces to a structure situated upstream from the chamber.

A rocket engine having a co-axial, bidirectional flow arrangement is described herein. The rocket engine receives fuel and an oxidizer into the rocket engine in a first direction, whereby a portion of the fuel is combusted in a pre-burner. The flow direction of the partially combusted fuel/oxidizer mixture is reversed, whereby the mixture is introduced into a combustion chamber. The fuel and oxidizer are combusted in the combustion chamber. The combustion products exit a throat and an expansion plenum in a direction similar to the first direction, whereby the combustion products exit a nozzle of the rocket engine, providing thrust.

A rocket engine having a co-axial, bidirectional flow arrangement is described herein. The rocket engine receives fuel and an oxidizer into the rocket engine in a first direction, whereby a portion of the fuel is combusted in a pre-burner. The flow direction of the partially combusted fuel/oxidizer mixture is reversed, whereby the mixture is introduced into a combustion chamber. The fuel and oxidizer are combusted in the combustion chamber. The combustion products exit a throat and an expansion plenum in a direction similar to the first direction, whereby the combustion products exit a nozzle of the rocket engine, providing thrust.

A rotating detonation engine includes an outer body with an opening therethrough having an interior wall and an inner body received in the outer body opening and with an outer wall tapering in the flow direction of the engine and spaced from the outer body opening interior wall defining a non-cylindrical improved efficiency detonation channel between the inner body outer wall and outer body opening interior wall.

A thrust chamber liner includes a metallic combustion chamber having an annular protrusion extending radially away from an exterior surface of the combustion chamber adjacent to its injector opening. A metallic nozzle is coupled to the combustion chamber at its throat opening. A composite material encases the exterior surface of the combustion chamber, but is only bonded to the annular protrusion.

A thrust chamber liner includes a metallic combustion chamber having an annular protrusion extending radially away from an exterior surface of the combustion chamber adjacent to its injector opening. A metallic nozzle is coupled to the combustion chamber at its throat opening. A composite material encases the exterior surface of the combustion chamber, but is only bonded to the annular protrusion.

A liquid rocket engine cools a thruster body by pumping propellant through cooling channels integrated in the thruster body between internal and external surfaces. One or more of the cooling channel surfaces has a variable depth along a thrust axis to mix propellant flow and destroy thermal stratification, such as a depth that varies with a repeated contiguous sinusoidal form along the thrust axis. Fuel passed through the cooling channels injects from the combustion chamber wall towards a central portion of the combustion chamber to cross impinge with oxygen injected at the combustion chamber head so that a toroidal vortex forms to enhance propellant mixing.