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.

A combustion chamber suitable in particular for use in an engine comprises a combustion space, a combustion space wall delimiting the combustion space, and a plurality of cooling channel webs extending from a surface of the combustion space wall which faces away from the combustion space and separating mutually adjacent cooling channels from one another. The cooling channel webs are each provided with a projection extending from an end face of the cooling channel webs which faces away from the combustion space. Furthermore, the combustion chamber comprises a plurality of cover elements, wherein each cover element extends along a longitudinal axis of a cooling channel delimited by two mutually adjacent cooling channel webs between the projections of the mutually adjacent cooling channel webs and is form-fittingly connected to the projections of the two mutually adjacent cooling channel webs in order to cover the cooling channel.

A combustion chamber suitable in particular for use in an engine comprises a combustion space, a combustion space wall delimiting the combustion space, and a plurality of cooling channel webs extending from a surface of the combustion space wall which faces away from the combustion space and separating mutually adjacent cooling channels from one another. The cooling channel webs are each provided with a projection extending from an end face of the cooling channel webs which faces away from the combustion space. Furthermore, the combustion chamber comprises a plurality of cover elements, wherein each cover element extends along a longitudinal axis of a cooling channel delimited by two mutually adjacent cooling channel webs between the projections of the mutually adjacent cooling channel webs and is form-fittingly connected to the projections of the two mutually adjacent cooling channel webs in order to cover the cooling channel.

An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.

A coupling system is utilized to form a multi-part rocket engine thrust compartment that maintains inner channels within walls of the thrust compartment for regenerative cooling. The coupling system includes an insert joint arranged between joint faces of a first segment and a second segment. The first segment and the second segment include inner edges that, when jointed together, form an inner wall. The joint insert is installed between the first segment and the second segment after the inner wall is formed and coupled to the first segment and the second segment. The joint faces of the first segment and the second segment include extending feature to form a flow passage along with cavities at least partially defined by the joint insert.

A coupling system is utilized to form a multi-part rocket engine thrust compartment that maintains inner channels within walls of the thrust compartment for regenerative cooling. The coupling system includes an insert joint arranged between joint faces of a first segment and a second segment. The first segment and the second segment include inner edges that, when jointed together, form an inner wall. The joint insert is installed between the first segment and the second segment after the inner wall is formed and coupled to the first segment and the second segment. The joint faces of the first segment and the second segment include extending feature to form a flow passage along with cavities at least partially defined by the joint insert.

A method is provided for fabricating a coolant channel closeout jacket on a structure having coolant channels formed in an outer surface thereof. A line of tangency relative to the outer surface is defined for each point on the outer surface. Linear rows of a metal feedstock are directed towards and deposited on the outer surface of the structure as a beam of weld energy is directed to the metal feedstock so-deposited. A first angle between the metal feedstock so-directed and the line of tangency is maintained in a range of 20-90°. The beam is directed towards a portion of the linear rows such that less than 30% of the cross-sectional area of the beam impinges on a currently-deposited one of the linear rows. A second angle between the beam and the line of tangency is maintained in a range of 5-65°.

A method is provided for fabricating a coolant channel closeout jacket on a structure having coolant channels formed in an outer surface thereof. A line of tangency relative to the outer surface is defined for each point on the outer surface. Linear rows of a metal feedstock are directed towards and deposited on the outer surface of the structure as a beam of weld energy is directed to the metal feedstock so-deposited. A first angle between the metal feedstock so-directed and the line of tangency is maintained in a range of 20-90°. The beam is directed towards a portion of the linear rows such that less than 30% of the cross-sectional area of the beam impinges on a currently-deposited one of the linear rows. A second angle between the beam and the line of tangency is maintained in a range of 5-65°.

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.