Abstract: A part of a spacecraft, for instance an upper stage, is configured to re-enter into the atmosphere and to be reused for several missions. The part is equipped with a reusable kit comprising non-ablative heat shields, for example constructed with an outer surface formed of Ceramic Matrix Composites, such as Ultra High Temperature Ceramic Matrix Composites, and optionally at least one decelerator and/or at least one parachute and/or parafoil system.

Abstract: A connection system for a cryogenic tank with a tank wall formed with a first material. A connector is provided for a component to be connected to the tank. The connector is formed with a second material, and the connector is positioned on an exterior side of the tank and essentially congruent with a passage opening of the tank wall. At least one sealing element is provided, and the first and the second material have different thermal expansion coefficients. A counterpart formed with the second material is positioned on an interior side of the tank, and is connected with the connector via at least two fastening elements, such that the tank wall is clamped between the counterpart, the connector and the at least one sealing element, and a slight displaceability of the connector and the counterpart parallel to the tank wall remains to compensate for thermally induced mechanical stresses.

Abstract: A rocket engine section includes a combustion chamber body having an inner wall and a channel carrying a cooling medium extending outside and along the inner wall. The rocket engine section further comprises a porous portion integrally formed with the inner wall and integral with the inner wall and adapted to allow the cooling medium carried in the channel to pass from the channel to the interior of the combustion chamber body. A porosity of the porous portion determines a volume flow rate and/or mass flow rate of the cooling medium let through into the interior of the combustion chamber body.

Abstract: A rocket propulsion system comprising a first cryogenic tank and a second cryogenic tank, wherein the first cryogenic tank is filled with a first propellant, and the second cryogenic tank is filled with a second propellant, for purposes of feeding at least one repeatedly ignitable main propulsion unit in a propulsion phase of the rocket propulsion system. For purposes of tank pressurization via at least a low level of acceleration in a ballistic phase, a first auxiliary propulsion unit can be operated by means of a first gas pressure accumulator, and at least one further auxiliary propulsion unit can be operated by means of a further gas pressure accumulator, and the rocket propulsion system is assigned an energy conversion unit, which is designed at least to charge the first and the second gas pressure accumulator, preferably in the ballistic phase.

Abstract: A seat system for use in a crew escape system of a space transport vehicle comprises a seat having a backrest with a supporting surface. The seat is configured to be installed in a cabin area of the crew escape system so it can rotate about a first axis between a first position where the seat faces a first direction, and a second position where faces a second direction that is opposed to the first direction. The seat system further comprises a control unit configured to control a rotation of the seat at least between the first position and the second position in dependence on at least one signal which is indicative of a trajectory of the crew escape system.

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 tank for the storage of liquid or gaseous substances with a fiber-reinforced plastic sheathing, which limits at least one interior space for holding the materials to be stored. At least one circumferential section of the fiber-reinforced plastic sheathing comprises an evacuable sandwich structure, which completely encloses the circumferential section. As a result, the tank has a low weight with a high thermal insulation capacity. Furthermore, a reliable explosion protection of the tank is provided. Furthermore, a method for manufacturing a tank for the storage of liquid or gaseous substances is provided.

Abstract: A heater apparatus configured to provide heat to at least one component of a spacecraft. The heater apparatus comprises a combustion chamber for a hypergolic propellant, and a heat radiator configured to radiate heat from the combustion chamber towards the at least one component to be heated. A spacecraft comprises at least one component to be heated and a heater apparatus configured to heat the at least one component to be heated. A method for heating at least one component of a spacecraft. The method comprises generating heat in a combustion chamber for a hypergolic propellant, and radiating at least a portion of the heat towards the at least one component.

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 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.

Abstract: A launch vehicle to transport at least one payload into an earth orbit, wherein the launch vehicle comprises a plurality of solar cells on its outer surface. Furthermore, a manufacturing method for a launch vehicle with a plurality of solar cells on its outer surface and a transport method for at least one payload using a launch vehicle with a plurality of solar cells on its outer surface are provided.

Abstract: A rocket engine section includes a combustion chamber body having an inner wall and a channel carrying a cooling medium extending outside and along the inner wall. The rocket engine section further comprises a porous portion integrally formed with the inner wall and integral with the inner wall and adapted to allow the cooling medium carried in the channel to pass from the channel to the interior of the combustion chamber body. A porosity of the porous portion determines a volume flow rate and/or mass flow rate of the cooling medium let through into the interior of the combustion chamber body.

Abstract: A guide vane arrangement configured for use in a turbo pump, wherein the guide vane arrangement includes a first guide vane and a second guide vane, wherein the second guide vane is arranged adjacent to the first guide vane such that a flow channel is defined between a leading surface of the first guide vane and a trailing surface of the second guide vane.

Abstract: 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).

Abstract: The invention relates to a mounting assembly with a spherical bearing for mounting a rocket engine and a rocket having such mounting assembly. The spherical bearing includes a spherical bearing base, a spherical retaining ring and a suspension link with a spherical end arranged in a space between the spherical bearing base and the spherical retaining ring. The spherical bearing base is part of an injector head of the rocket engine or is part of a fuel tank, the parts having a spherical shape.

Abstract: 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.

Abstract: A catalyst chamber with a catalyst bed embedded therein for a monopropellant thruster of a rocket engine. The catalyst chamber comprises an inlet having a first cross-sectional area through which a propellant can be introduced into the catalyst chamber and an outlet having a second cross-sectional area through which the propellant and/or resulting reaction products can be introduced into a combustion chamber of the thruster. The outlet is connected to the inlet via a catalyst volume of the catalyst chamber. At least one helical wall member is arranged within the catalyst chamber and is dividing the catalyst volume into two or more segments such that an effective length of the catalyst bed of each segment passed through by the propellant and/or its reaction products is larger than a geometrical length of the catalyst chamber defined between the inlet and the outlet along a direction of extension of the catalyst chamber.

Abstract: A rocket propulsion system comprises a combustion chamber, a hydrogen-oxygen supply system connected to the combustion chamber, which hydrogen-oxygen supply system is configured to conduct hydrogen and oxygen into the combustion chamber, and a coolant supply system connected to the combustion chamber, which coolant supply system is configured to conduct a combustible coolant into the combustion chamber. An ignition system of the rocket propulsion system is configured to initiate combustion of the hydrogen-oxygen-coolant mixture in 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: A method for operating a rocket propulsion system comprises the steps of supplying oxygen to a combustion chamber, supplying hydrogen to the combustion chamber and combusting the oxygen-hydrogen mixture in the combustion chamber. The rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.