A method of controlling the pressure (PGC) and a mixture ratio of a rocket engine from a pressure setpoint (PGCc) and from a mixture ratio setpoint (RMc), the method comprising regulation delivering control signals for two control valves (VR1, VR2) of said engine, the regulation using a pressure feedback loop. The method further comprises determining an estimated value for the mixture ratio (RMe) used by said regulation, the estimated value for the mixture ratio being obtained by a model that delivers mixture ratio values as estimated from at least one of the two control valve control signals and/or from the measured pressure.

The invention also provides a control device.

An energetic ignition arrangement may comprise a pressure vessel arrangement, comprising a pressure vessel, and an ignition system arrangement. The ignition system arrangement may comprise an ignitor housing coupled to the pressure vessel and defining a sealing aperture, a seal disposed in the sealing aperture, and an electric match extending through the ignitor housing, wherein at least the ignitor housing and the pressure vessel are reusable after the energetic ignition arrangement has been employed in an ignition event.

A stage of a space launcher including an engine core, a tank secured above the engine core, and a nozzle including a first deployable nozzle section secured below the engine core, the nozzle further including a second and a third section configured to be situated extending from each other when the nozzle is in a deployed propulsion configuration. The stage further includes a structure for temporarily supporting nozzle sections mounted on the tank, configured to change the nozzle to a configuration for accessing the engine core wherein the nozzle sections are inserted into each other, with the third section bearing against a lower support of the temporary structure.

The invention relates to the field of technical testing, and more particularly to a method of testing a machine, the method comprising: at least one step (S101) of determining a plurality of operating points for said machine, each operating point being defined by a duration and a specific value of at least one operating parameter of the machine; a step of calculating a set of distances between pairs of operating points; a step (S106) of selecting an optimum sequence of operating points by applying an algorithm for solving the traveling salesman problem to said set of distances between pairs of operating points; and a step (S107) of controlling the operation of said machine according to said optimum sequence of operating points.

The invention relates to the field of technical testing, and more particularly to a method of testing a machine, the method comprising: at least one step (S101) of determining a plurality of operating points for said machine, each operating point being defined by a duration and a specific value of at least one operating parameter of the machine; a step of calculating a set of distances between pairs of operating points; a step (S106) of selecting an optimum sequence of operating points by applying an algorithm for solving the traveling salesman problem to said set of distances between pairs of operating points; and a step (S107) of controlling the operation of said machine according to said optimum sequence of operating points.

A non-destructive testing system for propellant for a rocket is provided. The system includes an insertion tube having a first end and a second end in addition to a support sleeve configured to mate with a casing of the rocket. The insertion tube includes a channel that extends from the first end to the second end of the insertion tube to receive a probe and also includes a probe tip generally disposed at the second end of the channel. The support sleeve includes an insertion tube opening configured to receive the insertion tube and allow the insertion tube to slide along the support sleeve.

A non-destructive testing system for propellant for a rocket is provided. The system includes an insertion tube having a first end and a second end in addition to a support sleeve configured to mate with a casing of the rocket. The insertion tube includes a channel that extends from the first end to the second end of the insertion tube to receive a probe and also includes a probe tip generally disposed at the second end of the channel. The support sleeve includes an insertion tube opening configured to receive the insertion tube and allow the insertion tube to slide along the support sleeve.

The present invention relates to improved rocket engine systems. In one embodiment, an improved rocket engine system includes a propellant source, at least one power source, at least one power source motor, a rocket engine, and at least one pump. The improved rocket engine system may further include at least one of the following: at least one controller, at least one propellant valve, and a propellant pressurizing source.

The present invention relates to improved rocket engine systems. In one embodiment, an improved rocket engine system includes a propellant source, at least one power source, at least one power source motor, a rocket engine, and at least one pump. The improved rocket engine system may further include at least one of the following: at least one controller, at least one propellant valve, and a propellant pressurizing source.

A solid rocket motor includes a case, a solid propellant, and a seal attached to the motor. The seal isolates the solid propellant from a fluid external of the case, and the seal is at least partially transparent.