A combustion chamber for a rocket engine, the combustion chamber including a combustion chamber body enclosing a combustion chamber volume and a nozzle portion tapering in a longitudinal direction of the combustion chamber and adjoining the combustion chamber body. The combustion chamber body has at least one first portion and a second portion, wherein an inner surface of the at least one first portion facing the combustion chamber volume is closer to a cross-sectional center of the combustion chamber body than an inner surface of the second portion of the combustion chamber body. Furthermore, a additive layer manufacturing method for manufacturing such a combustion chamber is described.

A combustion chamber for a rocket engine, the combustion chamber including a combustion chamber body enclosing a combustion chamber volume and a nozzle portion tapering in a longitudinal direction of the combustion chamber and adjoining the combustion chamber body. The combustion chamber body has at least one first portion and a second portion, wherein an inner surface of the at least one first portion facing the combustion chamber volume is closer to a cross-sectional center of the combustion chamber body than an inner surface of the second portion of the combustion chamber body. Furthermore, a additive layer manufacturing method for manufacturing such a combustion chamber is described.

Provided are an igniter-integrated thrust chamber for a rocket engine using a cryogenic fuel and liquid oxygen and a rocket including the thrust chamber. The thrust chamber includes a combustion chamber and a mixing head assembly, which is disposed at one side of the combustion chamber and is integrated with the combustion chamber.

Provided are an igniter-integrated thrust chamber for a rocket engine using a cryogenic fuel and liquid oxygen and a rocket including the thrust chamber. The thrust chamber includes a combustion chamber and a mixing head assembly, which is disposed at one side of the combustion chamber and is integrated with the combustion chamber.

Rocket propulsion systems and associated methods are disclosed. A representative system includes a combustion chamber having an inwardly-facing chamber wall enclosing a combustion zone. The chamber has a generally spherical shape and is exposed to the combustion zone. A propellant injector is coupled to the combustion chamber and has at least one fuel injector nozzle positioned to direct a flow of cooling fuel radially outwardly along the inwardly-facing chamber wall. In addition to or in lieu of the foregoing features, the injector can include an oxidizer piston and a fuel piston that deliver oxidizer and fuel, respectively, to the combustion chamber, in a sequenced manner so that the oxidizer is introduced prior to the fuel.

According to one contemplated embodiment of the rocket engine invention, water is first pumped from a water tank through a rocket nozzle cooling heat exchanger wherein it is evaporated into said superheated steam. A generator supplies electricity to an electrolyzer that electrolyzes superheated steam into gaseous hydrogen and gaseous oxygen. The gaseous hydrogen and gaseous oxygen is employed for forming an annular curtain of secondary combustion in a divergent rocket engine. The secondary combustion gas surrounds a central thrust of combustion gas produced in an upstream combustion chamber by a primary injection of hydrogen/oxygen supplied from a liquid hydrogen tank and liquid oxygen tank. The rocket liquid hydrogen tank and liquid oxygen tank are pressurized by gaseous hydrogen and gaseous oxygen generated by the electrolyzer.

According to one contemplated embodiment of the rocket engine invention, water is first pumped from a water tank through a rocket nozzle cooling heat exchanger wherein it is evaporated into said superheated steam. A generator supplies electricity to an electrolyzer that electrolyzes superheated steam into gaseous hydrogen and gaseous oxygen. The gaseous hydrogen and gaseous oxygen is employed for forming an annular curtain of secondary combustion in a divergent rocket engine. The secondary combustion gas surrounds a central thrust of combustion gas produced in an upstream combustion chamber by a primary injection of hydrogen/oxygen supplied from a liquid hydrogen tank and liquid oxygen tank. The rocket liquid hydrogen tank and liquid oxygen tank are pressurized by gaseous hydrogen and gaseous oxygen generated by the electrolyzer.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

A rotary detonation rocket engine generator system can include an axial drive shaft operably coupleable to an electrical generator. At least one support arm is radially coupled to the axial drive shaft and has corresponding rotary detonation rocket engines. An air-fuel mixing chamber receives ambient air and fuel to form an air-fuel mixture and deliver the air-fuel mixture to an annular combustion chamber. At least one pulse detonation combustion chamber is in fluid communication with the annular combustion chamber to receive an oxidizer and fuel to form an oxidizer-fuel mixture. The at least one pulse detonation combustion chamber creates a detonation wave that travels along the at least one pulse detonation chamber to the annular combustion chamber and ignites the air-fuel mixture as the detonation wave travels around the annular combustion chamber to generate thrust force that causes rotation of the axial drive shaft to drive the electrical generator.