988results about "Cosmonautic propulsion system apparatus" patented technology

A modular satellite system is provided comprising a main body member, one or more controllers, a communication system, a datastore, a power unit, and one or more orbital cameras. The controller(s) may be operable to perform a mission instruction. The communication system may be in communication with the controller(s) and with a client terminal. The datastore may store data accessible by the controller(s). The orbital camera(s) may be operable to capture one or more image(s) associated with the mission instruction, which may be defined as captured image(s). The controller(s) may be operable to detect and identify one or more predetermined object(s) in the one or more captured image(s). The controller(s) may generate a mission analytics packet, and the controller(s) may be operable to transmit the mission analytics packet to the client terminal.

This application presents ideas for devices that solve energy generation problems and space exploration. This application show idea that is using device such using conductor / element or device. And this application show idea that is using device such using conductive line or carrier device. This application show idea that is using device such using photon or particle related device.

To study a system that transmutes atomic nuclei using muons, devise a light source device and a propulsion device including a nuclear transmutation system, and examine a centrifugal force application unit which may be accelerated by the propulsion device, and devise an electric field generator or a magnetic field generator.SOLUTION: A system is devised that decelerates cosmic muons produced from cosmic rays or high-speed muons derived from mesons produced from atomic collisions with electric field and deceleration means and uses the muons for nuclear transmutation. Also, a decelerator using a laser wake field and a decelerator using an element forming an electric field are proposed. A radiation light source having the nuclear transmutation unit is disclosed. An ionization device, a centrifugal gun device, a centrifugal force and pressure applying device, a pressure generation unit, a pressure applying material, a high-pressure induction superconductivity experimenting system, a magnetic device, and a magnetic sail are disclosed.SELECTED DRAWING: Figure 18

Implementations of the disclosed subject matter may provide an adaptive thrust structure for a space launch vehicle that may include an interface block that is fluidically coupled to a fuel supply, a fuel manifold, an oxidizer supply, and an oxidizer manifold. A first adapter plate may be coupled to a first engine assembly. First adapter lines may be fluidically coupled to the first engine assembly and the interface block, via the first adapter plate. A second engine assembly, second adapter lines, and a second adapter plate are capable of replacing the first engine assembly, first adapter lines, and first adapter plate. Implementations of the disclosed subject matter may provide a reusable and reconfigurable launch vehicle that may include a reusable second stage, and a reusable and reconfigurable transporter having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.

The invention provides an air pressure regulation and control device, a cold air micro-propulsion system and a satellite, and relates to the technical field of cold air micro-propulsion. The air pressure regulation and control device comprises a high-pressure valve, a two-stage pressure reducing valve, a precise pressure regulating valve, a buffer gas cylinder and a pressure sensor; one end of the high-pressure valve is connected with the high-pressure gas cylinder, and the other end of the high-pressure valve is connected with the two-stage pressure reducing valve through a first pipeline; the precise pressure regulating valve is provided with a first connector, a second connector, a third connector and a fourth connector, the first connector is connected with the two-stage pressure reducing valve through a second pipeline, the second connector is connected with the pressure sensor through a third pipeline, the third connector is connected with the buffer gas cylinder through a fourth pipeline, and the fourth connector is connected with the thruster. Compared with the prior art, the air pressure regulation and control device has the advantages that the mode of combining two-stage pressure reduction, air capacity buffering and air pressure closed-loop precise regulation and control is adopted, stable and low-noise air can be supplied to the thruster, and therefore high-stability and low-noise thrust output is achieved.

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

The invention provides a swing gas transmission mechanism based on a metal tube spring. The swing gas transmission mechanism comprises the metal tube spring, a tooth-shaped supporting structure and a tooth comb support. A plurality of supporting columns of the tooth comb bracket are arranged along the circumferential direction; tooth-shaped supporting structures protruding outwards in the radial direction are distributed on a supporting column of the tooth comb support in the axial direction. Gaps exist between the tooth-shaped supporting structures in the axial direction, and the metal tube spring penetrates through the gaps and is supported. The metal tube spring is of a hollow structure and is provided with an inlet and an outlet. The metal pipe spring is provided with a spiral pipeline part, and the spiral pipeline part is arranged on the outer side of the tooth comb support in a surrounding mode through the gap. The invention provides a structural combination of a single metal spiral pipeline and a comb-shaped support, and the process difficulty of multi-pipe binding and winding is reduced.

The invention discloses an electric energy autonomous distribution method for a direct-drive electric propulsion power supply system of a spacecraft, and belongs to the technical field of spacecraft power supplies. In the spacecraft electric energy distribution process, the power of a solar cell array, the power of a low-voltage load and the current and voltage of a storage battery pack are collected, the SOC of the storage battery pack is predicted, reasoning is carried out by utilizing a fuzzy rule, and the working gear and the working power of the electric propulsion load are obtained through adjustment on the basis of the reference power of the electric propulsion load. And meanwhile, a model prediction control method is utilized, the state space of the storage battery pack model is discretized, an objective function and constraint conditions capable of tracking the battery current of the storage battery pack with high precision are designed, the SOC is controlled within a safe range, and the service life of the SOC is prolonged.

A system for mitigating backsputter to a thruster in a propulsion test facility includes a plate disposed in an output plume of the thruster and a magnetic field generator disposed about or at the plate and configured to generate a magnetic field between the thruster and the plate. The magnetic field generator is oriented such that a magnetic field is generated transverse to the output plume.

The invention provides an air suction type electric propulsion air suction chamber. The air suction type electric propulsion air suction chamber comprises grids, an air inlet channel and a rectifying device. The bottom of the grid is connected with the front end of the air inlet channel, the inner wall of the air inlet channel is of a parabola-shaped structure, the rear end of the air inlet channel is connected with the long pipe, the rectifying device is installed in the air inlet channel through the supporting frame, the rectifying device is of a convex-mirror-like structure, the inner wall of the rectifying device is of a parabola-shaped structure, and the focus of the rectifying device is the same as the focus of the parabola-shaped structure of the inner wall of the air inlet channel in position. The opening direction of the rectifying device is opposite to the opening direction of the air inlet channel. According to the invention, the incoming flow gas is refracted twice and is finally injected into the long pipe at the rear end of the air suction chamber in parallel, so that collision and corrosion of air particles to the wall surface of the long pipe are reduced, the service life and the performance of the air suction type electric propulsion are greatly improved, the gas collection rate is improved, and a new technical approach is provided for the development of the air suction type electric propulsion.

An electron propulsion engine utilizes the acceleration of electrons to propel a spacecraft through space. The acceleration of the electrons in space emits electromagnetic radiation which can be used to propel the spacecraft. The radiation emission also decelerates the electrons, which allows the decelerated electrons to be recycled for reuse by the electron propulsion engine. The electron propulsion engine includes a first engine module, a second engine module, and an engine control system. The first engine module and the second engine module correspond to two mirror structures that form the electron propulsion engine. The engine control system facilitates the automatic control of the operation of the electron propulsion engine. The first engine module and the second engine module each includes a vacuum housing. The vacuum housing of each engine module is a D-shaped structure that facilitates the acceleration of the electrons and the resulting radiation emission to propel the spacecraft.

The present invention relates to a new form of air, land, underwater, or space propulsion, achieved by the use of suitable electromagnetic interactions. When using coils (1), with internal core (2) and support piece (3), subjected to current pulses with asymmetric current derivative and magnetic field, we obtain directional propulsion forces. This is possible due to a new electromagnetic propulsion mechanism that uses the conservation of total momentum where the sum of the mechanical moment with the moment of the magnetic field must always be conserved, resulting in a constant and null total sum of the two components, where the variation of the magnetic field moment will generate a corresponding change in the mechanical moment of the coil, thus generating propulsion forces. When magnetic fields with asymmetric derivative are produced in an external volume, they may also generate force fields.

The invention provides a rocket engine gimbal seat which comprises a pair of upper supports used for being connected with a rocket engine rack, a gimbal ring and a pair of lower supports, the bottoms of the pair of upper supports are hinged to the first two opposite side walls of the gimbal ring in a one-to-one correspondence mode, and the bottoms of the pair of upper supports are hinged to the second two opposite side walls of the gimbal ring in a one-to-one correspondence mode. The tops of the pair of lower supports are hinged to the second two opposite side walls of the gimbal ring in a one-to-one correspondence mode, and the middles of the pair of upper supports and the middles of the pair of lower supports make contact with the gimbal ring through cylindrical surfaces.

The invention relates to a welding seam reinforcing structure for a stainless steel storage tank, and belongs to the technical field of aerospace carrier propellant storage tank structures. A first box bottom, a first short shell, a cylinder section, a second short shell and a second box bottom; the first box bottom and the second box bottom are each of a hemispherical shell structure. The first short shell, the second short shell and the barrel section are all of barrel section structures; the first short shell, the barrel section and the second short shell are horizontally and coaxially butted in sequence; the first box bottom is coaxially mounted at the axial outer end of the first short shell; the second box bottom is coaxially mounted at the axial outer end of the second short shell; the opening end of the first box bottom extends into the first short shell, and the outer wall of the first box bottom is in butt joint with the inner wall of the first short shell. The opening end of the second box bottom extends into the second short shell, and the outer wall of the second box bottom is in butt joint with the inner wall of the second short shell. The problem that the welding performance of reinforcing materials cannot be weakened through partition thickness design is solved, and the purposes of high performance, low cost manufacturing and the like of the stainless steel storage box are achieved.

A mobile fusion device (MFD) comprises three main sections: the MFD reactor (467), MFD support sub-systems (468), and the MFD support and fuel injection sub-systems (469), which are indicated by broken lines in Figure 1 The MFD reactor (467) is where the Deuterium and Tritium fuel is heated to high enough temperatures and densities and confinement times that fusion occurs, releasing neutrons that are absorbed in and heats up the 17Li-83Pb breeder blanket where more Tritium fuel is bred and where heat is transferred to Helium coolant. The MFD support sub-systems (468) houses auxiliary power units that generate electricity to power the electrical sub-systems of the MFD, and houses the MFD control computer, and where the Helium coolant is cryogenically cooled and purified, and where the 17Li-83Pb breeder blanket fluid is purified by separating and storing Deuterium fuel and Tritium fuel and Helium coolant, and where a Helium heat exchanger is located. The MFD support and fuel injection sub-systems (469) houses Deuterium and Tritium neutral beam injection plasma heating systems and a Deuterium and Tritium fuel gas puffer system and a Deuterium and Tritium fuel storage and pellet manufacturing system, and where the 17Li-83Pb breeder blanket fluid is purified by separating and storing Deuterium fuel and Tritium fuel and Helium coolant, and where a Helium heat exchanger is located, and where the Helium coolant is cryogenically cooled and purified, and houses auxiliary power units that generate electricity to power the electrical sub-systems of the MFD. The MFD is housed within a multi-layered pressurised outer containment pod MLP, which has a multi-layered pressurised rear access hatch MLH. The MFD is mounted on a sliding raft SR for ease of access.

A magnetic field propulsion unit generates a drive force when being arranged in an external magnetic field. The magnetic field propulsion unit includes a support structure, a magnetic field generating device attached to the support structure and configured to generate a first magnetic field, an energy supply unit, and a control unit. The energy supply unit provides the magnetic field generating device with electrical energy. The control unit controls the energy supply unit so that the energy supply unit supplies energy to the magnetic field generating device. The control unit supplies varying current to the magnetic field generating device. The magnetic field generating device has a central axis and is arranged in the external magnetic field such that the central axis is inclined with respect to magnetic field lines of the external magnetic field at an angle different from 0°, 90°, 180°, and 270°.

The invention discloses a pulse detonation attitude control engine system which comprises an ignition detonation straight pipe and at least one S-shaped attitude control pipe. The ignition detonating straight pipe comprises an ignition section and an obstacle detonating section which are sequentially arranged in the axial direction; wherein an explosion promotion barrier is arranged in the barrier explosion initiation section; each S-shaped attitude control tube comprises a branch tube and two bent push tubes; the center of the branch pipe is vertically arranged at the tail end of the obstacle detonating section and is communicated with the obstacle detonating section through a valve; the two bent push pipes are symmetrically arranged relative to the center of the branch pipe to form an S shape; each bent push pipe comprises a bent pipe section and an exhaust nozzle section, and the two ends of each bent pipe section are connected with the corresponding branch pipe and the corresponding exhaust nozzle section respectively. According to the invention, the advantages of high-efficiency combustion and high-precision adjustable pulse thrust of the pulse detonation engine can be played as far as possible while the microminiaturization of the attitude control engine is met, the thermodynamic cycle theory can be improved by more than 20% compared with common chemical propulsion, and the on-orbit life of a spacecraft is effectively prolonged.

The energy recycling heat control device provided by the embodiment of the present application comprises a liquid collecting chamber, an engine assembly and a spring, the engine assembly and the spring are arranged between the liquid collecting chamber and an injector, the pushing action of a piston by the target gas filling parameters in a sealed space of the engine assembly is combined with the spring mechanical parameters of the spring to control the distance between a heat plate and the outer wall of the liquid collecting chamber, so that the heat energy is recycled and reused, which is beneficial to improve the rapid response speed of the engine, reduce the power consumption, and thus reduce the risk of explosion after shutdown.

The invention provides a channel type flow guide plate based on an inverted-cone-shaped pore plate structure. The channel type flow guide plate comprises a capillary porous plate and a groove structure. And the channel type guide plate is connected with a liquid accumulator at the bottom of the storage box through a bracket. The channel type flow guide plate is parallel to the inner wall face of the storage box and extends to the top of the storage box along the inner wall face. The cross section of the channel type guide plate is of a trapezoidal structure, the channel type guide plate is hollow, and a closed channel is formed for propellant flowing. A capillary perforated plate is arranged on the side, close to the wall face of the storage box, of the channel type flow guide plate, a plurality of micron-sized inverted-cone-shaped capillary pores are formed in the capillary perforated plate in a single-side laser drilling mode, and the functions of storing liquid and serving as a gas barrier can be achieved. The channel type guide plate is suitable for a satellite propulsion system with strong maneuverability and high acceleration, has the advantages of simple structure, portability, high reliability, strong flow conductivity and the like, and can provide guidance for the development of a third-generation surface tension storage tank PMD.

One or more embodiments relates to an air-breathing plasma thruster including a thruster wall, an anode, a cathode, and at least one ring electrode. The thruster wall defines a cylindrical channel, the cylindrical channel having a first end and an opposing second end in fluid communication with the first end, where the cylindrical channel is adapted to receive incoming airflow. The anode is at the first end of the channel and the cathode is at the second end of the channel opposite the first end. The at least one ring electrode is positioned on the thruster wall.

A rocket motor comprising a non-permeable outer shell, an intra-wall channel between the non-permeable outer shell and an annular section, and a center chamber formed at least in part by a first propellant is disclosed. The first propellant is made up of a multi-grained fuel grain, and the center chamber is further formed by an annular section. The intra-wall channel is configured to permit combustion gases to flux between the non-permeable outer shell and the annular section. The annular section configured to permit combustion gases to flux through pores embedded within the annular section.

To solve the problem that there is a case in which it is difficult to recover a flying object.SOLUTION: A recovery device has an engine part which burns supplied propellant to obtain a thrust in a predetermined direction, and is configured to be mounted on an airframe of a recovery object such that the recovery device can obtain, with the engine part, the thrust in a direction opposite to a direction in which the airframe of the recovery object obtains a thrust.SELECTED DRAWING: Figure 1

Provided are a method and device for driving movement of a target object by a gravitational field generated by a varying electromagnetic field which relate to the technical field of physics. The method includes: generating, by a variable electromagnetic field generator, a time-varying electromagnetic field in space where a target is located; adjusting a frequency, amplitude and phase of the time-varying electromagnetic field so as to generate an effect of a varying magnetic field in the space; and based on the time-varying electromagnetic field, generating an acting force on the target object to drive the target object to move along a predetermined trajectory or direction. The present invention breaks through limitation that the traditional electromagnetic mechanics can only drive metal materials, utilizes the varying characteristics of the electromagnetic field to drive and realize accurate control of movement of all objects comprising a non-magnetic substance and a non-conductive substance.

Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

The invention discloses a calculation method for calibrating the on-orbit thrust performance of a thruster, which comprehensively considers each design parameter, working parameter, orbit control parameter and orbit control result of a propulsion system, considers the working conditions of all thrusters with components in the on-orbit control direction, and combines the prediction condition and the actual condition of orbit change to calibrate the on-orbit thrust performance of the thruster. The calculation of the on-orbit thrust performance of the thruster is achieved, the method is very suitable for a propelling system with the requirement for accurate calibration of the on-orbit thrust performance of the thruster, and the calculation result can be used for correcting the ground calibration result of the thruster.

The present invention relates to a spacecraft electric thruster particularly suitable for both propellant-fed and air-breathing electric space propulsion comprising an ionization chamber (11) for plasma generation which develops along a central development axis (X), wherein the ionization chamber (11) has a chamber inlet (12) configured to allow gas inflow and a chamber outlet (13) configured to allow extraction of ionized gas particles; a power injection assembly (30) configured to perform electron heating for plasma generation inside the ionization chamber (11); a magnetic assembly (20) configured to generate a magnetic field for plasma confinement in the ionization chamber (11); and an ion extraction and acceleration assembly (40) configured to extract ionized gas particles from the ionization chamber (11) through the chamber outlet (13) and accelerate them along an ion acceleration direction (Z), wherein the power injection assembly (30) comprises a microwave power delivery system (31) configured to deliver an electric field oscillating at a selected microwave frequency (f) and to inject the generated oscillating electric field in the ionization chamber (11), and wherein the ionization chamber (11) comprises a plurality of walls (11a, 15,16,41) closed to the passage of microwaves apart from a portion of at least one lateral wall (11a) of the plurality of walls (11a, 15,16,41) at which the power delivery system (31) is coupled to inject the generated oscillating electric field in the ionization chamber (11), and is resonant for electromagnetic fields oscillating at the selected microwave frequency.

The embodiment of the utility model relates to the technical field of space propulsion, and discloses a combination valve for cold air propulsion and a cold air propulsion system. The combination valve for cold air propulsion comprises a combination valve module shell, a first ventilation path and a second ventilation path are arranged in the combination valve module shell, the two ends of the first ventilation path penetrate through the combination valve module shell, one end of the first ventilation path is provided with an input connector, and the other end of the first ventilation path is provided with an output connector; an electromagnetic valve assembly is arranged on the first ventilation path, the electromagnetic valve assembly controls connection and disconnection of the first ventilation path, and a pressure reducing valve is further arranged on the first ventilation path. The embodiment of the utility model further provides a cold air propelling system. The cold air propelling system comprises the combination valve for cold air propelling. According to the embodiment of the invention, the integrated structural layout is realized through the integrated design and the plug-in valve design, and the locking stability, the sealing performance and the safety of the whole system are improved.

A spacecraft with retractable solar sails is a spacecraft that is designed as a semi-autonomous spacecraft that can maneuver and propel using radiation pressure. The spacecraft includes a spherical hull and several propulsion mechanisms. The spherical hull houses the propulsion mechanisms, the payload, and other systems that enable the semi-autonomous operation of the spacecraft. The spherical hull also enables the semi-autonomous operation of the propulsion mechanisms. The propulsion mechanisms enable the propulsion and maneuvering of the spacecraft using radiation pressure in deep space. Each of the propulsion mechanisms includes a solar sail assembly, a sail receptacle, a pneumatic actuator, and a quantity of fluid. The solar sail assembly supports several retractable solar sails. The sail receptacle allows the retraction of the solar sail assembly into the spherical hull. The pneumatic actuator allows the deployment of the solar sail assembly. The quantity of fluid enables the operation of the pneumatic actuator.

The invention discloses a split rotor gyroscope stability hysteresis propeller, and relates to the technical field of airspace engines. According to the thought of the invention, the stability and precession of the split rotor gyroscope are different from those of a common gyroscope, when a rotor motor (10) rotates along with a connecting rod (5), generated precession torque acts on a pendulum bob (9), so that the rotating surface of the pendulum bob (9) is separated from the rotating surface of a rotor base (8) to generate an included angle, and the rotor base (8) rotates along with the rotor motor (10); the precession moment firstly acts on the pendulum bob (9) to consume a part and then is transmitted to the rotating shaft of the rotor motor (10), and the precession moment borne by the rotating shaft of the rotor motor (10) is delayed and reduced, so that when the connecting rod (5) rotates in the axial stage (16), a large force is needed to overcome the stability of the gyroscope, and a large upward counter-acting force is generated.

The invention relates to an air suction type cusped field electric propulsion system with multiple working medium supply modes and a supply method. The invention relates to the technical field of air suction type electric propulsion systems. The system comprises an electric thruster and collection device module, a microwave cathode module, a xenon storage and supply module and a control and energy supply module. According to the air-breathing cusped field electric propulsion system with the multiple working medium supply modes, the adaptive capacity of the electric thruster under the complex working condition is improved through the xenon working medium carried by the air-breathing cusped field electric propulsion system while the ultra-low orbit rarefied atmospheric environment is fully utilized. In addition to an ultra-low rail air-breathing type electric propulsion task, the xenon working medium carried by the ultra-low rail air-breathing type electric propulsion device can enable the ultra-low rail air-breathing type electric propulsion device to give consideration to a higher rail range, and a traditional xenon electric propulsion rail interval and an ultra-low rail space of 150 km or above are covered.