A rocket engine system comprising a rocket engine, coolant and a coolant source, propellant and a propellant source, a turbopump, and a heat source. The coolant is pressurized and then heated through a heat source to a supercritical state for augmented heat transfer. The heat source may be a heat exchanger with returning coolant, or a preburner. The rocket engine system may further comprise at least one additional rocket engine with a pump to provide pressure for multiple engine. The rocket engine system may further comprise multiple turbopump shafts for independent control of propellants.

A rocket engine comprises a combustion chamber and a rocket nozzle that communicates with the combustion chamber. A metal wire supply device supplies a metal wire, which is fuel, to the combustion chamber. A water vapor generator supplies water vapor as an oxidant to the combustion chamber. An ignition device ignites the metal wire in a water vapor atmosphere.

Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

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 combustor of a liquid rocket engine includes a nozzle unit including a regenerative cooling channel, in which the nozzle unit includes a fuel manifold outer shell, a combustor inner shell, and a combustor outer shell having a downward channel inlet, and the combustor includes a fuel inlet connected to a nozzle neck of the nozzle unit, a fuel manifold formed between the fuel manifold outer shell and the combustor outer shell, and in which fuel introduced from the fuel inlet flows, a downward channel connected in communication with the fuel manifold through the downward channel inlet, and extending in a downward direction from an upper portion of the combustor, a diverting manifold provided at a distal end of the nozzle unit and connected in communication with the downward channel, and an upward channel connected in communication with the diverting manifold and extending in an upward direction of the combustor.

A combustor of a liquid rocket engine includes a nozzle unit including a regenerative cooling channel, in which the nozzle unit includes a fuel manifold outer shell, a combustor inner shell, and a combustor outer shell having a downward channel inlet, and the combustor includes a fuel inlet connected to a nozzle neck of the nozzle unit, a fuel manifold formed between the fuel manifold outer shell and the combustor outer shell, and in which fuel introduced from the fuel inlet flows, a downward channel connected in communication with the fuel manifold through the downward channel inlet, and extending in a downward direction from an upper portion of the combustor, a diverting manifold provided at a distal end of the nozzle unit and connected in communication with the downward channel, and an upward channel connected in communication with the diverting manifold and extending in an upward direction of the combustor.

A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12). The first coolant chamber (30) is fluidly connected to each of the plurality of first coolant channels (50) and the second coolant chamber (32) is fluidly connected to each of the plurality of second coolant channels (52).

A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12). The first coolant chamber (30) is fluidly connected to each of the plurality of first coolant channels (50) and the second coolant chamber (32) is fluidly connected to each of the plurality of second coolant channels (52).

The present disclosure relates to a rocket engine, and more particularly, a rocket engine with an integrated combustor head and turbopump in which a turbopump of the rocket engine is formed integrally with a combustor head.