172 results about "Spacecraft propulsion" patented technology

A thruster (1) has a main chamber (6) defined within a tube (2). The tube has a longitudinal axis which defines an axis (4) of thrust; an injector (8) injects ionizable gas within the tube, at one end of the main chamber. An ionizer (124) is adapted to ionize the injected gas within the main chamber (6). A first magnetic field generator (12, 14) and an electromagnetic field generator (18) are adapted to generate a magnetized ponderomotive accelerating field downstream of said ionizer (124) along the direction of thrust on said axis (4),

The thruster (1) ionizes the gas, and subsequently accelerates both electrons and ions by the magnetized ponderomotive force.

A high-efficiency spacecraft propulsion system, including includes electric pumps inserted in the oxidizer and fuel lines that increase liquid apogee engine (“LAE”) operating pressure and reduce tank-operating pressure. An on-board computer generates pump drive signals, in response to measured oxidizer and fuel line pressures, that are input to the pump controller electronics. The controller electronics provides current drives to the pump motors. The system uses an LAE that can operate at higher thrust chamber pressures (e.g., 500 psia) than standard LAEs, where pump-fed rocket motors have integrated turbopumps that are fuel operated. The turbopump increases LAE complexity and cost, reduces fuel efficiency, and is not compatible with active thrust and mixture ratio control.

An aircraft propulsion system in which a combustion engine is arranged to drive an electrical generator. An electrical energy store is provided within the system. A propulsive rotor is arranged to be driven by an electric motor and a controller selectively varies the supply of power to the electric motor from the generator and/or energy store in dependence on one or more property of a vapour trail resulting from the engine exhaust flow. The controller may also control the supply of power to the energy store for charging.

A watercraft propulsion system has a control lever for giving instructions regarding operating mode and output power from a source of driving force. A controller sets a propeller driving mode according to instructions given from the control lever. In an embodiment, an electric motor and an engine are included in the propulsion system, and the controller simultaneously controls both the engine and motor based on inputs from the control lever and sensed conditions.

A method is provided for wirelessly monitoring an aircraft propulsion system having a plurality of rotor blades, at least some of which rotor blades have an internal compartment. The method includes: providing a primary transducer having a primary coil, and one or more secondary transducers, each secondary transducer having a secondary coil connected to a sensor, wherein the secondary transducers are respectively disposed within the internal compartments; wirelessly powering the secondary transducers using a magnetic field generated by the primary transducer; monitoring at least one operational parameter from within each internal compartment using a respective sensor; and transmitting output signals from the secondary transducers to the primary transducer through the magnetic field, wherein the output signals are indicative of the monitored parameters.

Spacecraft payload orientation steering is provided for an orbiting spacecraft in motion along an orbit track around a celestial body, the orbit track having a nominal inclination with respect to an equatorial orbit, a substantial eccentricity, and a drift angle with respect to the nominal inclination. Coordinates of an optimal payload target location as a function of a spacecraft position along the orbit track are determined, the target location being on the surface of the celestial body and having a substantial motion with respect to the surface and with respect to a spacecraft nadir. A payload of the spacecraft is substantially aligned with the determined coordinates by steering the satellite body to correct for at least one of the inclination drift angle, and the eccentricity, thereby adjusting the spacecraft orientation as a function of the spacecraft position along the orbit track.

A thruster (1) has a main chamber (6) defined within a tube (2). The tube has a longitudinal axis which defines an axis (4) of thrust; an injector (8) injects ionizable gas within the tube, at one end of the main chamber. An ionizer (124) is adapted to ionize the injected gas within the main chamber (6). A first magnetic field generator (12, 14) and an electromagnetic field generator (18) are adapted to generate a magnetized ponderomotive accelerating field downstream of said ionizer (124) along the direction of thrust on said axis (4). The thruster (1) ionizes the gas, and subsequently accelerates both electrons and ions by the magnetized ponderomotive force.

The invention discloses a watercraft propulsion system. Two sets of propulsion motors symmetrically arranged at the left and the right, a full swivelling propeller type swing mechanism and a hydraulic system constitute an electric propulsion power device for the watercraft. A propulsion controller, a swing controller and a steering control board constitute an electric propulsion control device. The control of the whole system during sailing includes propulsion control and turning control. The propulsion control process is that the propulsion controller receives sailing speed instruction of the watercraft, propulsion instruction of left and right screw propellers and control instructions to open or close the loop and then outputs propulsion set values of left and right screw propellers and deviation value of left and right screw propellers; the turning control process is that the swing controller receives sailing direction instruction and steering instruction of the watercraft from the control board and then outputs set turning angles of left and right screw propellers. The watercraft propulsion system can realize propulsion operation intelligentization and control simplification and is more precise in control.

The invention provides a new concept high-speed aerocraft propulsion system layout method. Two cycles are built in a propulsion system, namely a Brayton cycle with the air as the working medium and a closed cycle with supercritical state fluid as the working medium. The two cycles are coupled through a supercritical microscopic scale heat exchange technology, a supercritical state fluid turbine and compressor power balance. By adjusting related valves, the propulsion system can be in a turbofan engine model when taking off or flying at a low speed and in a turbine rocket engine model when flying at a high Mach number, so that it is guaranteed that the aerocraft can effectively cruise for a long time in both a subsonic state and a supersonic state. Through the supercritical microscopic scale heat exchange technology, the gas flow temperature at an inlet of a compressor can be effectively reduced when the propulsion system is flying at a high speed, and when the supercritical microscopic scale heat exchange technology is applied in combination with a closed cycle technology, optical distribution of energy of the propulsion system can be achieved. By means of the method, the defects of the propulsion system of an existing high-speed aerocraft are overcome, and working performance of the high-speed aerocraft propulsion system is remarkably improved when Ma ranges from 0 to 5.

Spacecraft thrusters capable of dual-mode operation, and a method of applying s propulsion to a spacecraft using a dual-mode thruster are provided. In one embodiment, the thrusters of the current invention can operate as a chemical motor to provide high thrust and low propellant exhaust velocity to achieve fast maneuverability, or as an electric propulsion thruster to provide low thrust and high exhaust velocity to perform maneuvers with a minimal amount of propellant.

The invention discloses a spacecraft propellant in-orbit filling system and a spacecraft propellant in-orbit filling method. A technology of performing primary exhaust and then performing pressurized filling is adopted, a target conduit head and a filling conduit head do not need simultaneous connection of a gas path and a liquid path, and only a single liquid path is used for connection, so that a docking mechanism is simplified; a supplementary spacecraft is capable of performing pressurized filling without a mounting pump, not only can an attitude disturbing torque generated by rotation of a motor in the pump be avoided, but also the boost pressure during filling is reduced, and the task reliability is improved; before filling, a measure of performing pressure relief on the target spacecraft till reaching a pressure which is below a propellant saturated vapor pressure is adopted, thereby ensuring that the filling rate of the target spacecraft achieves a higher level; in addition, the system has a reusable characteristic; the supplementary spacecraft can be used for performing in-orbit supplement on multiple targets or one target for multiple times, so that a foundation for further implementing in-orbit filling service is laid.

A propulsion system comprising a nacelle substantially tubular around a longitudinal axis, having an inner wall extending from a front to a rear of the nacelle and by an outer wall, external of the inner wall, extending from the front to the rear of the nacelle, a turbojet comprising a fan and situated internally of the inner wall of the nacelle, at least one tank containing an extinguishing fluid, and a network of pipes hydraulically connected to the tank. The propulsion system comprises each tank being located in the nacelle, around the inner wall and internally of the outer wall. Such a propulsion system makes it possible to shift the one or more tanks, which no longer occupy the space at the mast, and makes it possible to carry a large volume of extinguishing fluid so as to be able to accommodate future regulations.

A spacecraft is disclosed. The spacecraft (10) includes a pressure hull (12) for containing a gas (13). The spacecraft further includes a propulsion system (18) coupled to the pressure hull so as to be disposed within the pressure hull, the propulsion system operable to generate a propulsion force for propelling the spacecraft through space.

A method for actively regulating mixture ratio of a spacecraft propellant comprises the following steps: conducting trial run before installing an engine with high consumption into a spacecraft propelling system, determining working parameters, and reaching a standard mixture ratio of an oxidant to fuel at a standard inlet pressure by regulating system hardware; arranging a pressure sensor at an oxidant and fuel inlet, and arranging a self-locking gas supply valve on pressurizing pipelines at the upstream of an oxidant storage tank and a fuel storage tank respectively; and in the working process of the engine, acquiring pressure parameters of the fuel and the oxidant at the inlet of the engine in real time through a propelling system controller, and comparing and regulating the pressure parameters to control the pressures of the two propellants at the inlet of the engine to be very close, so that the deviation level of the total mixture ratio of the propellant can be controlled. By the method, the utilization ratio of the propellant of the spacecraft propelling system is effectively improved, so that the service life of a spacecraft is prolonged, the effective load mass of the spacecraft is improved and the production cost of the spacecraft is reduced.

An aircraft propulsion system has a propulsive rotor assembly rotatable about an axis of rotation and comprising a plurality of blades and a rotationally fixed vane assembly located adjacent to the propulsive rotor assembly and arranged circumferentially around the axis of rotation. As airflow enters the propulsive rotor assembly, a portion of the airflow passes over the vane assembly which is configured to direct the airflow away from the rotor blades so as to reduce the relative velocity of the redirected airflow over the rotor blades. This results in a reduced tendency of the airflow through the propulsive rotor assembly to become choked.

An aircraft includes a fuselage, an aircraft propulsion system including an open rotor, and a pylon mounting the propulsion system to the fuselage. The fuselage includes an acoustic panel configured to attenuate noise generated by the open rotor.

A linking device including two hinge pins and having flexibility allowing vibrations between these two shafts to be filtered. The device includes a stack of plates connected to one another by layers of elastomer material. Plates of a first type are coupled to one of the shafts in relation to the movements parallel to the plates, while the plates of a second type are coupled to the other shaft in relation to the movements parallel to the plates, so as to allow relative movement of the shafts parallel to the plates. This linking device may be used in particular as a connecting rod, spreader beam or three-point shackle used in connecting a turbine engine to a pylon in a propulsion system of an aircraft such as an airplane.

The invention discloses an optimized configuration method for reentry and return spacecraft propulsion system, which solves a problem that a conventional configuration method is difficult to meet the configuration quality of the small-scale return spacecraft power system configuration. Based on a task analysis, in order to adapt to the characteristics of the small return spacecraft, a type of a propulsion system and a propellant is selected reasonably, a effective propellant weight is calculated, a selection of extrusion gas type is optimized, a rail control engine and a attitude control engine thrust and a number of that is preliminary determined by further analyses, the attitude control engine thrust and cylinder tank layout is determined. According to the calculated control capacity to optimize the attitude control engine thrust until the control requirements to meet the optimal system to achieve optimal configuration.