Abstract: A combustion chamber section for a combustion chamber for a rocket engine, the combustion chamber section including a combustion chamber body enclosing a combustion chamber volume and having coolant channels disposed therein, and at least one baffle integrally formed with the combustion chamber body and projecting from the combustion chamber body into the interior of the combustion chamber. The at least one baffle includes at least one coolant channel fluidly connected to at least one of the coolant channels in the combustion chamber body. Furthermore, an additive layer manufacturing method for manufacturing such a combustion chamber section is described.

Abstract: 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.

Abstract: 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.

Abstract: A combustion chamber section (110) for a combustion chamber (100) for a rocket engine (10) is described, the combustion chambersection (110) comprising a combustion chamber body (120) enclosing a combustion chamber volume and having coolant channels (130) disposed therein, and at least one baffle (140) integrally formed with the combustion chamber body (120) and projecting from the combustion chamber body (120) into the interior of the combustion chamber. The at least one baffle (140) comprises at least one coolant channel (133-135) fluidly connected to at least one of the coolant channels (130) in the combustion chamber body (120). Furthermore, an additive layer manufacturing method for manufacturing such a combustion chamber section is described.

Abstract: An injection apparatus for a rocket engine comprises an injection plate which delimits a combustion chamber upstream. The injection apparatus further comprises a plurality of coaxial injection elements distributed in the injection plate, each of which injection elements forms an inner outlet opening, delimited by a central sleeve body, for a first propellant component and an outer outlet opening for a second propellant component. The outer outlet opening is formed between the central sleeve body and a wall section which surrounds the central sleeve body in an annular manner. At least in a partial number of the injection elements, the central sleeve body projects from the wall section in the direction towards the combustion chamber. In particular, the central sleeve body projects relative to the injection plate into the combustion chamber.

Abstract: An injection apparatus for a rocket engine comprises an injection plate which delimits a combustion chamber upstream. The injection apparatus further comprises a plurality of coaxial injection elements distributed in the injection plate, each of which injection elements forms an inner outlet opening, delimited by a central sleeve body, for a first propellant component and an outer outlet opening for a second propellant component. The outer outlet opening is formed between the central sleeve body and a wall section which surrounds the central sleeve body in an annular manner. At least in a partial number of the injection elements, the central sleeve body projects from the wall section in the direction towards the combustion chamber. In particular, the central sleeve body projects relative to the injection plate into the combustion chamber.

Abstract: The invention relates to an injection element having an inner element with a first outlet opening and an outer element. The outer element includes at least one second outlet opening structured and arranged for receiving and injecting fuel in a combustion space, and arranged coaxially to the first outlet opening. The outer element further includes third outlet openings composed of bores structured and arranged for forming a cooling liquid film layer, wherein the bores are arranged coaxially to the first outlet opening and the at least one second outlet opening.

Abstract: A method for manufacturing a regeneratively cooled nozzle extension of a rocket combustion chamber, the nozzle extension including a first wall and a second wall arranged coaxially to each other and between which a number of cooling channels is configured that are laterally delimited by cooling channel webs. The first wall and the second wall are connected to each other by a positive fit by cooling channel webs of the first wall engaging with corresponding recesses of the second wall for forming the positive fit. The positive fit is produced by a forming process in the region of the cooling channels of the second wall having the recesses.

Abstract: A liquid fuel rocket engine with a closed direct auxiliary flow cycle has a combustion chamber connected to each liquid fuel source through a respective main liquid fuel supply line. A turbo pump is provided in each main fuel supply line. The turbo pump is driven by a gas generator which is connected to a first bypass fuel supply line for supplying fuel to the gas generator. An exhaust gas duct (8, 9, 10) leads from the gas generator into at least one main fuel supply line for feeding gas generator exhaust gas into the main fuel supply line upstream of the turbo pump.

Abstract: A liquid fuel rocket engine with a closed direct auxiliary flow cycle has a combustion chamber connected to each liquid fuel source through a respective main liquid fuel supply line. A turbo pump is provided in each main fuel supply line. The turbo pump is driven by a gas generator which is connected to a first bypass fuel supply line for supplying fuel to the gas generator. An exhaust gas duct (8, 9, 10) leads from the gas generator into at least one main fuel supply line for feeding gas generator exhaust gas into the main fuel supply line upstream of the turbo pump.