Societe Europeenne De Propulsion

Snecma

Arianegroup Sas

Snecma Moteurs

Airbus Safran Launchers Sas

Didier Vuillamy

Andre Beaurain

Etienne Tiret

Pierre Desclos

Gérard Roz

Jean-Michel Sannino

Jean-Luc Barthoulot

Pierre-Andre Baudart

Emmanuel Edeline

Robert Gazave

David Hayoun

Claude E Bonniot

François Lassoudiere

Jean-Claude Harlay

Andre Beurain

Jean-Paul Dumont

Claude Pacou

Cryogenic propellant tank () for a spacecraft engine, comprising an external enclosure () defining an internal volume, characterized in that the internal volume of the tank comprises a primary volume (V) and a secondary volume (V) connected to the primary volume (V) via a valve () configured to selectively allow a passage of fluid from the primary volume (V) to the secondary volume (V), or to isolate the secondary volume (V) from the primary volume (V), the primary volume (V) having a primary orifice () adapted to be connected to a first pressurization source (), the secondary volume (V) having a supply orifice () adapted to be connected to a supply line of a spacecraft engine (), and a secondary orifice () adapted to be connected to a second pressurization source ().

Tank for a spacecraft engine

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

Device for pressurizing a propellant tank of a rocket engine

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 () for feeding a rocket engine () at least with a first liquid propellant, the feed circuit including at least one first heat exchanger () suitable for being connected to a cooling circuit () 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.

Propellant feed circuit and a cooling method

A device heating a fluid and usable in a rocket launcher to pressurize a liquefied propellant. The device includes a first burner performing first combustion between a limiting propellant and an excess propellant; a first heat exchanger in which first burnt gas from the first combustion transfers heat to the fluid; at least one second burner into which both the first burnt gas and some limiting propellant are injected to perform second combustion between the limiting propellant and at least a portion of unburnt excess propellant present in the first burnt gas. The second burnt gas from the second combustion flows through a second heat exchanger to transfer heat to the fluid. Burnt gas from each combustion flows in respective burnt gas tubes within a common overall heat exchanger including the heat exchange units, the gas transferring heat to the fluid, the fluid flowing between the burnt gas tubes.

Device for heating a fluid

A cryogenic thruster assembly including: a reignitable main thruster; a first cryogenic tank connected to the main thruster to feed the main thruster with a first propellant; a first gas tank; at least one settling thruster; and a first feed circuit for feeding the first gas tank. The feed circuit of the first gas tank is connected to the first cryogenic tank and includes a heat exchanger for using heat given off by the at least one settling thruster to vaporize a liquid flow of the first propellant as extracted from the first cryogenic tank to feed the first gas tank with the first propellant in the gaseous state. A method feeds the first gas tank with the first propellant in the gaseous state.

Cryogenic thruster assembly using regenerative heating from main and settling thrusters

The rocket engine nozzle system comprises a set of individual nozzles distributed in a ring around a central core of axially symmetrical shape that presents a central axis. Each individual nozzle placed at the periphery of the central core comprises a circular section throat for receiving gas from a combustion chamber, and a diverging portion that is tangential to the central core. The diverging portion has an outlet section that presents first and second lateral sides that converge radially towards the central axis of the central core, a curved outer side having its convex side facing outwards, and an inner side situated in the vicinity of the central core.

Rocket engine nozzle system

To control the separation of a hot gas ejection jet within a rocket engine nozzle, the device includes a jettisonable annular structure for placing around the outside wall of the nozzle at the gas cutlet section thereof, which structure defines a radial extension around the nozzle so that in the presence of an outlet jet coming from the nozzle, it creates a low pressure zone in the vicinity of the bottom face of the device, thereby displacing in the gas outlet direction jet separation from the interior wall of the nozzle.

Device for displacing to nozzle outlet or eliminating jet separation in rocket engine nozzles, and a nozzle including the device

The downstream portion of the diverging part of the nozzle of the rocket engine is extended by an ejectable diffuser which, downstream from the zone where it is connected to the diverging part, presents a zone of smaller cross-section that acts during a first stage of flight in the presence of significant outside pressure to recompress the flow of hot gases and to prevent a flow separation from appearing along with the wall of the diverging part. The ejectable diffuser may have a streamlined throat which constitutes the zone of smaller cross-section.

Rocket engine nozzle with an engine diffuser

In a rocket engine nozzle with a selectively smaller outlet cross-section, the nozzle comprising a converging portion which receives the gases produced in a combustion chamber, a nozzle throat of small cross-section and a diverging part connected to the nozzle throat and terminating at its downstream portion in a gas jet outlet cross-section defining a high cross-section ratio, a releasable annular obstacle is connected to the downstream end of the diverging part so as to define at the outlet of the diverging part a zone of selectively smaller cross-section which ensures stability of the separation of the flow of hot gases relative to the wall of the diverging part during a first phase of flight.

Rocket engine nozzle with selectively smaller outlet cross-section

In a rocket engine nozzle comprising: a converging portion that receives the gases produced in a combustion chamber; a throat of small section; and a diverging portion connected to the nozzle throat; the diverging portion includes a set of downstream notches running longitudinally from its downstream end and uniformly distributed around the entire periphery of the diverging portion, the total width of all of the notches within any section representing about 1% to 6% of the periphery of the diverging portion. The notches serve to avoid having a jet separation line that is irregular and unstable, and they facilitate matching the diverging portion to ambient conditions that vary in flight.

Rocket engine nozzle having a notched diverging portion

The gas generator comprises an annular main porous piece, at least one calibration lining for defining a flow rate of a first liquid propellant component that is injected through the porous piece towards a combustion chamber that is delimited by one of the faces of the porous piece, and means for radially injecting a flow of a second propellant component into the immediate proximity of the face of the porous piece that delimits the combustion chamber in such a manner as to cause the first and second propellant components to be mixed in the immediate proximity of the face of the porous piece that delimits the combustion chamber.

Close combustion gas generator

The rocket engine is fitted with an injection device comprising a single injection plate having a central portion that defines the end of the burn zone of the combustion chamber and that contains first propellant injectors having a high mixing ratio adapted to the conditions under which hot gases are produced in the burn zone, and a peripheral portion in the form of a ring that defines the end of an annular cavity of a gas generator and that contain second propellant injectors having a low mixing ratio and adapted to the conditions under which hot gases are produced in the gas generator. The gas generator is thus integral with the combustion chamber and the annular cavity of the gas generator is separated from the burn zone of the combustion chamber by a cylindrical wall that extends axially perpendicularly to the injection plate and that is fixed thereto in sealed manner.

Growing community of inventors

Quincampoix, France

Didier Vuillamy