Abstract: Cryogenic propellant tank (1) for a spacecraft engine, comprising an external enclosure (10) defining an internal volume, characterized in that the internal volume of the tank comprises a primary volume (V1) and a secondary volume (V2) connected to the primary volume (V1) via a valve (20) configured to selectively allow a passage of fluid from the primary volume (V1) to the secondary volume (V2), or to isolate the secondary volume (V2) from the primary volume (V1), the primary volume (V1) having a primary orifice (11) adapted to be connected to a first pressurization source (41), the secondary volume (V2) having a supply orifice (4) adapted to be connected to a supply line of a spacecraft engine (30), and a secondary orifice (12) adapted to be connected to a second pressurization source (42).

Abstract: Cryogenic propellant tank (1) for a spacecraft engine, comprising an external enclosure (10) defining an internal volume, characterized in that the internal volume of the tank comprises a primary volume (V1) and a secondary volume (V2) connected to the primary volume (V1) via a valve (20) configured to selectively allow a passage of fluid from the primary volume (V1) to the secondary volume (V2), or to isolate the secondary volume (V2) from the primary volume (V1), the primary volume (V1) having a primary orifice (11) adapted to be connected to a first pressurization source (41), the secondary volume (V2) having a supply orifice (4) adapted to be connected to a supply line of a spacecraft engine (30), and a secondary orifice (12) adapted to be connected to a second pressurization source (42).

Abstract: A rocket engine benefiting from better behavior during its starting stage, the rocket engine (1) including a diverging section (13) and a lashing system (40) configured to hold the diverging section (13) while starting the rocket engine (1), the lashing system (40) comprising: a plurality of radial cables (42) connected at respective first ends to a plurality of points of the diverging section (13), and a peripheral cable (41) connected to the second ends of the radial cables (42) and configured to co-operate with attachment points (43) of a launch platform (3).

Abstract: The device comprises two heat exchangers (74, 90) suitable respectively for vaporizing first and second propellants before they are reintroduced in gaseous form into their tanks (16, 18). The heat exchangers co-operate respectively with first and second gas generators (60, 84) suitable for being fed with a mixture of propellants in order to produce combustion, the second gas generator (84) being suitable for being fed at least in part by the exhaust from the first gas generator (60).

Abstract: The device comprises a primary heater (58) suitable for heating the propellant coming from the tank (16) prior to it being reintroduced into its tank. The primary heater uses the heat of combustion from the engine (10) and the device further comprises a secondary heater (66) having its source of heat independent from the operation of the engine, the secondary heater being arranged downstream from the primary heater (58) in order to heat the propellant between its outlet from the primary heater and being reintroduced into the tank. The device also has means (62) between the feed to the primary heater (58) and the return of the propellant to the tank for putting the propellant under pressure.

Abstract: A rocket engine benefiting from better behavior during its starting stage, the rocket engine (1) including a diverging section (13) and a lashing system (40) configured to hold the diverging section (13) while starting the rocket engine (1), the lashing system (40) comprising: a plurality of radial cables (42) connected at respective first ends to a plurality of points of the diverging section (13), and a peripheral cable (41) connected to the second ends of the radial cables (42) and configured to co-operate with attachment points (43) of a launch platform (3).

Abstract: The invention relates to the aerospace field, and in particular to the field of vehicles propelled by rocket engines. In particular, the invention relates to a feed circuit (6) for feeding a rocket engine (2) at least with a first liquid propellant, the feed circuit including at least one first heat exchanger (18) suitable for being connected to a cooling circuit (17) for cooling at least one heat source in order to cool said heat source by transferring heat to the first propellant, and, in addition, downstream from said first heat exchanger, a branch passing through a second heat exchanger.

Abstract: A rocket engine igniter 10, characterized in that it comprises a spark plug 12 that is movable between a retracted position and a deployed position, and a pressure compensation chamber 14 presenting a first internal cavity 14A and a second internal cavity 14B distinct from the first cavity, the first cavity and the second cavity being separated by at least one movable base 16, the spark plug being secured to the movable base, the first cavity communicating with a first vent 22 while the second cavity communicates with a second vent 24 that is distinct from the first vent.

Abstract: After filling a tank (18) with liquid oxygen (20) that is to be used to feed a rocket engine (10) with oxidizer, but prior to causing the engine (10) to operate, the tank is pressurized by injecting gaseous nitrogen (N) into the tank. While the engine (10) is in operation, liquid oxygen (20) is taken off, the taken-off oxygen is heated so as to obtain gaseous oxygen, and the gaseous oxygen is injected into the gas blanket (42) of the tank, with the pre-pressurization nitrogen forming a nitrogen buffer (40) between the liquid oxygen present in the gaseous oxygen tank injected into the gas blanket.

Abstract: A feed device (20) for feeding propellant to at least one rocket engine propulsion chamber (19) of a propulsion assembly (21) comprises a main propellant tank (1, 2), and a main feed pipe (3, 4) extending between the main tank (1, 2) and the propulsion chamber (19) and having a main feed valve (VAH, VAO) placed thereon. The feed device further comprises an auxiliary tank (11, 12) for said propellant and an auxiliary feed pipe (13, 14) provided with an auxiliary feed valve (VH1, VO1) connecting the auxiliary tank (11, 12) to the main feed pipe (3, 4) downstream from the main feed valve (VAH, VAO).

Abstract: A propulsion assembly (1, 201) for a rocket, comprising a tank (2) configured to contain a propellant, an engine having a combustion chamber (9), a propellant feed pipe (11) extending between the tank (2) and the combustion chamber (9) and having an isolation valve (24) arranged therein, and a heater (15) having an inlet connected to the feed pipe (11) and an outlet connected to the tank (2). The inlet of the heater comprises an inlet pipe (13a) connected firstly to the feed pipe (11) downstream from the isolation valve (24) and secondly to a neutral fluid feed (31).

Abstract: The device comprises two heat exchangers (74, 90) suitable respectively for vaporizing first and second propellants before they are reintroduced in gaseous form into their tanks (16, 18). The heat exchangers co-operate respectively with first and second gas generators (60, 84) suitable for being fed with a mixture of propellants in order to produce combustion, the second gas generator (84) being suitable for being fed at least in part by the exhaust from the first gas generator (60).

Abstract: The device comprises a primary heater (58) suitable for heating the propellant coming from the tank (16) prior to it being reintroduced into its tank. The primary heater uses the heat of combustion from the engine (10) and the device further comprises a secondary heater (66) having its source of heat independent from the operation of the engine, the secondary heater being arranged downstream from the primary heater (58) in order to heat the propellant between its outlet from the primary heater and being reintroduced into the tank. The device also has means (62) between the feed to the primary heater (58) and the return of the propellant to the tank for putting the propellant under pressure.

Abstract: The invention relates to the field of rocket engines, and more particularly to a system for feeding a rocket engine (2) with propellant, the system comprising a first tank (3), a second tank (4), a first feed circuit (6) connected to the first tank (3), and a second feed circuit (7) connected to the second tank (4). In order to cool the propellant contained in the second tank (4), the first circuit (6) includes a branch (12) passing through a first heat exchanger (14) incorporated in the second tank (4). The invention also provides a method of feeding the rocket engine (2).

Abstract: The invention relates to the aerospace field, and in particular to the field of vehicles propelled by rocket engines. In particular, the invention relates to a feed circuit (6) for feeding a rocket engine (2) at least with a first liquid propellant, the feed circuit including at least one buffer-tank (20) for said first liquid propellant and a first heat exchanger (18) that is incorporated in said buffer tank (20) and suitable for being connected to a cooling circuit (17) for cooling at least one heat source in order to cool said heat source by transferring heat to the first propellant.

Abstract: The invention relates to the aerospace field, and in particular to the field of vehicles propelled by rocket engines. In particular, the invention relates to a feed circuit (6) for feeding a rocket engine (2) at least with a first liquid propellant, the feed circuit including at least one first heat exchanger (18) suitable for being connected to a cooling circuit (17) for cooling at least one heat source in order to cool said heat source by transferring heat to the first propellant, and, in addition, downstream from said first heat exchanger, a branch passing through a second heat exchanger.