Abstract: The method serves to place a space vehicle, such as a satellite, on a target orbit such as the orbit adapted to normal operation of the space vehicle and starting from an elliptical initial orbit that is significantly different from, and in particular more eccentric than the target orbit.

Abstract: The method serves to place a space vehicle, such as a satellite, on a target orbit such as the orbit adapted to normal operation of the space vehicle and starting from an elliptical initial orbit that is significantly different from, and in particular more eccentric than the target orbit.

Abstract: A first satellite is placed practically directly by a launcher on a first final orbit. A second satellite is placed on the same launcher is initially transferred to a highly elliptical waiting orbit whose semi-major axis points to intercept the torus formed by the sphere of influence of the moon on its orbit, and then during a maneuver at the perigee of the highly elliptical orbit, the second satellite is transferred to lunar transfer orbit. Changes in the perigee altitude and the inclination of an intermediate orbit on which the second satellite is to be found are obtained mainly by gravitational reaction in the sphere of influence of the moon, and during a last maneuver, the second satellite is placed on a final orbit having orbital parameter values that are quite different from those of the orbital parameters of the first final orbit.

Abstract: A launcher places a first satellite practically directly on a final orbit. A second satellite carried by the same launcher is initially transferred onto a waiting orbit that is highly elliptical, having a semi-major axis situated in the initial orbital plane. The inclination and the perigee of the waiting orbit are then changed in the vicinity of the apogee of the waiting orbit, thereby placing the second satellite on an intermediate orbit. A maneuver is performed that includes at least one step making use of atmospheric braking in the vicinity of perigee of the intermediate orbit so as to lower the altitude of apogee of the intermediate orbit, and an impulse is supplied to the second satellite at apogee of the intermediate orbit so as to raise its perigee and transform the intermediate orbit into a final orbit having orbital parameters whose values are substantially different from those of the orbital parameters of the final orbit of the first satellite.

Abstract: A liquid propellant rocket engine comprising a chamber body connected to a converging/diverging thrust nozzle and to at least one lightly pressurized propellant feed duct, said body including a body bore in which an injector-forming piston can move longitudinally between a rest first position and an operating second position, the piston comprising a piston head and a piston rod, and subdividing the chamber body bore into a combustion chamber in front of the piston, and at least one injection chamber surrounding the piston rod, the chamber body further including a flow section that is closed in said rest position to prevent any propellant being fed from the injection chamber(s) and that is opened when the piston moves from said rest position to said operating position in order to allow the injection chamber(s) to be filled via said flow section directly from the feed duct(s), the propellant(s) being injected into the combustion chamber via injection channels formed between the combustion chamber and the inject

Abstract: A liquid propellant rocket engine comprising a chamber body connected to a converging/diverging thrust nozzle and to at least one lightly pressurized propellant feed duct, said body including a body bore in which an injector-forming piston can move longitudinally between a rest first position and an operating second position, the piston comprising a piston head and a piston rod, and subdividing the chamber body bore into a combustion chamber in front of the piston, and at least one injection chamber surrounding the piston rod, the chamber body further including a flow section that is closed in said rest position to prevent any propellant being fed from the injection chamber(s) and that is opened when the piston moves from said rest position to said operating position in order to allow the injection chamber(s) to be filled via said flow section directly from the feed duct(s), the propellant(s) being injected into the combustion chamber via injection channels formed between the combustion chamber and the inject

Abstract: A device for pressurizing at least one propellant tank feeding the combustion members of a space vehicle, the tank(s) being connected to one or more pipes for feeding said combustion members, and at least one of the propellants being catalytically or thermally decomposable, the device comprising: tapping means placed at the outlet from the decomposable propellant tank to tap a portion of said propellant, drive gas generation means for delivering drive gases, suction means formed by a converging-diverging nozzle connected firstly to the drive gas generation means and secondly to the tapping means to suck decomposable propellant from the tapping means by lowering the pressure of the drive gases to a value that is less than that required for pressurizing the tank(s), and decomposition means placed at the outlet from the suction means for decomposing the sucked propellant into pressurization gas for pressurizing the tank(s).

Abstract: The present invention relates to a system for feeding the engines of a space vehicle with at least one liquid propellant, the system comprising for the purpose of storing said liquid propellant under a pressure greater than the feed pressure, firstly a main tank having no liquid expulsion device, and secondly an auxiliary tank that is smaller than the main tank and that is provided with a device for separating the liquid phase from the gas phase, switching means enabling the auxiliary tank to be selectively refilled with liquid propellant from the main tank; such that liquid propellant that is free from pressurization gas bubbles can always be delivered regardless of the flight conditions of the vehicle, and in particular when it is not possible to take said liquid propellant from the main tank.