1423results about "Cosmonautic propulsion system apparatus" patented technology

The disclosed methods, systems, and kits provide the ability to deliver entire clean sheet designs from concept to first hot fire in under six weeks with instant specific impulses above 330 seconds in some of our engines. In examples, thrusters can be delivered that are at less than half of the mass budget allowable for them and they can be delivered in weeks.

Material transfer interfaces for space vehicles, and associated systems and methods are disclosed. A representative system includes a first coupler configured to be carried by a first space vehicle, and a first valve device carried by the first coupler. The system further includes a second coupler configured to be carried by a second space vehicle and a second valve device carried by the second coupler. The first coupler includes rotatable and translatable latch arms positioned to engage with and connect to the second coupler. The first valve device incudes a moveable probe that is insertable into the second valve device when the latch arms of the first coupler are connected to the second coupler to transfer fluid between the first and second valve devices.

The invention relates to a propulsion system, in particular to an air-breathing radio frequency ion electric propulsion system with mixed working medium power and a multi-mode working method thereof, and aims to solve the problems that an existing air-breathing electric propulsion system only takes atmosphere as a single working medium, the atmosphere trapping efficiency is low, and the performance of a spacecraft is affected. The device comprises a propelling and gas trapping unit, a radio frequency neutralizer, a power and control unit and a xenon storage and supply unit, the propelling and gas trapping unit comprises a radio frequency ion thruster and an atmosphere trapping gas inlet channel; the power and control unit comprises a direct current module, a control module and a radio frequency module; the xenon storage and supply unit comprises a xenon bottle and a valve and sensor assembly; the control module and the valve are electrically connected with the sensor assembly. The radio frequency ion thruster is electrically connected with the radio frequency neutralizer through the radio frequency module. The first output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency neutralizer, and the second output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency ion thruster.

The invention provides a maintenance-free dual-mode attitude control power system which is used for being installed on an aerospace vehicle and comprises a positive feedback circulation system, a thruster set, a fuel gas generator, a fuel storage box and an oxidizing agent storage box. The positive feedback circulation system is used for self-pressurization of the attitude control power system; the positive feedback circulation system comprises a gunpowder starter, a differential storage tank, a liquid path pressure stabilizing valve, a fuel gas generator and a pressurizing one-way valve; the thruster group comprises a gas thruster, a single-component thruster and a double-component thruster; an inlet of the fuel gas generator is communicated with the positive feedback circulation system; the outlet is communicated with the gas thruster, the fuel storage tank and the oxidant storage tank; the fuel tank and the oxidant tank are in communication with the thruster set. The advantages of a single-component engine and a double-component engine are organically combined, the overall performance of the power system is improved, the use and maintenance requirements of the attitude control power system before rocket launching are lowered, and the launching preparation time is shortened.

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.

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The propulsion assembly is coupled to the boom by a rotating joint. A second wing is optionally coupled to the rotating joint. The first wing is coupled to the boom by a rotating joint. The first wing is coupled to the rotating joint by a hinge. A vehicle with roll, pitch, and yaw maneuverability able to mirror the aircraft movements may be coupled to the second end of the boom. The vehicle body may be picked up with a vehicle chassis disconnected from the vehicle body. The boom houses an energy source to power the propulsion assembly. A rudder is coupled to the second end of the boom. A paddle is disposed between the propulsion assembly and the boom.

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

A Hall effect thruster system includes a thruster body and a diffuser configured to apply a torque to the thruster body during operation. The diffuser applies the torque by ejecting the propellant in a non-axial direction, such as a direction tangent to helical or curvilinear channels formed within a body of the diffuser. The applied torque can be used to counteract a swirl torque that is induced on the thruster body by the ionizing Hall current flowing in an annular channel of the thruster body. This effective counter-torque is useful in deep space applications outside the Earth's magnetic field.

Carbide-based fuel assembly includes outer structural member of ceramic matrix composite material (e.g., SiC—SiC composite), insulation layer of porous refractory ceramic material (e.g., zirconium carbide with open-cell foam structure or fibrous zirconium carbide), and interior structural member of refractory ceramic-graphite composite material (e.g., zirconium carbide-graphite or niobium carbide-graphite). Spacer structures between various layers provide a defined and controlled spacing relationship. A fuel element bundle positioned between support meshes includes a plurality of distributively arranged fuel elements or a solid, unitary fuel element with coolant channels, each having a fuel composition including high assay, low enriched uranium (HALEU). Fuel assemblies are distributively arranged in a moderator block and the upper end of the outer structural member is attached to a metallic inlet tube for hydrogen propellant and the lower end of the outer structural member is interfaced with a support plate, forming a nuclear thermal propulsion reactor.

The invention belongs to the technical field of composite materials, and particularly relates to a carbon fiber composite material for preventing low-temperature medium leakage and a preparation method and application thereof. The carbon fiber composite material comprises carbon fibers, a resin matrix, modified nano-hydroxyapatite and an anti-seepage film, the mass percentage content of the carbon fiber is 45-69%, the mass percentage content of the resin matrix is 26-37%, the mass percentage content of the impermeable film is 2-10%, and the balance is modified nano-hydroxyapatite; the modified nano-hydroxyapatite, the impermeable film and other raw materials are added into the formula of the carbon fiber composite material, and the impermeable performance of the carbon fiber composite material in a low-temperature medium is comprehensively improved through the modified nano-hydroxyapatite and the impermeable film. Experimental results show that the thickness of the carbon fiber composite material is smaller than 3 mm, and the leakage rate under 3 MPa is smaller than 10 <-5 > Pa.m < 3 > / s.

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 discloses a multifunctional electric propulsion system of a spacecraft, and relates to the technical field of space propulsion systems. The multifunctional electric propulsion system comprises a gas cylinder module, a fluid management module and a thruster module. The downstream of the gas cylinder module is connected with the fluid management module; the downstream of the fluid management module is connected with the thruster module; the thruster module comprises a Hall thruster, an electric heating type thruster and a cold air thruster. The low-cost working medium gas is utilized, the attitude control system and the orbit control system of the satellite are integrated through modular design and heat flow design, the design and manufacturing difficulty of the satellite is lowered, the launching cost of the satellite is reduced, and meanwhile the integrated control system is combined with an orbit control algorithm, so that intelligent control over the satellite can be achieved.

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 present invention relates to a new form of air, land, underwater, or space propulsion, achieved by the use of suitable electromagnetic interactions. When using a longitudinal electromagnetic field emitter (1), which emits longitudinal electric or magnetic fields, with asymmetric electric or magnetic field derivatives, through space in the direction of element (2), optionally focused or amplified by element (3), we develop directional forces in elements (1) and (2). This is possible due to a new electromagnetic propulsion mechanism that uses the conservation of total momentum where the sum of mechanical momentum and field momentum must always be conserved resulting in a constant and zero total sum of the two components, where the variation in electric or magnetic field momentum will generate a corresponding change in the mechanical momentum of the assembly, thus generating propulsion forces.

Spacecraft servicing systems include a spacecraft servicing device and at least one mission extension pod comprising at least one spacecraft servicing component. The spacecraft servicing device is configured to transfer the at least pod to a target spacecraft in order to service the target spacecraft with the at least one spacecraft servicing component of the at least one pod. Spacecraft servicing pods configured to be supplied to a spacecraft with a spacecraft servicing device include at least one spacecraft servicing component.

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.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

The invention relates to a method for stabilization of the orbital motion set-up of a group of satellites (2, 3, 4, 5, 6,.... n) bound to a tether (10). To such end, the tension (T) of the tether (10) that connects two or more satellites is controlled in order to keep them in the desired flight formation. This effect is obtained by perturbing a satellite (2, 3, 4) to generate a. tension component (T) in the tether, which results in a recoil force that brings the system back into a condition of equilibrium. According to a preferred embodiment, the tension (T) of the tether (10) is obtained by adjusting the disposition of one or more aerodynamic surfaces (20, 21, 30, 31) provided on the satellites (2, 3) at the ends of the tether.

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.

The invention provides a light agile remote sensing satellite platform configuration. The light agile remote sensing satellite platform configuration comprises a satellite body, a radar antenna, a solar cell array, a control moment gyroscope group, a hoisting structure, an electric propulsion system and a heat dissipation device, the star body comprises an integrated carbon fiber truss and a cellular board, the integrated carbon fiber truss serves as an installation foundation and is of a cuboid structure, the cellular board is arranged in the integrated carbon fiber truss, and the control moment gyroscope group is of a pentagonal pyramid structure and is fixedly installed on the cellular board; the solar cell array and the radar antenna are oppositely installed on the two sides of the integrated carbon fiber truss, the radar antenna comprises a plurality of movably connected sub-arrays, and the solar cell array comprises a plurality of movably connected sub-plates; a hoisting structure is arranged on the top of the integrated carbon fiber truss, the electric propulsion system is integrally installed on a star body in a modularized mode, and a heat dissipation device is arranged on the outer side face of the integrated carbon fiber truss. According to the invention, good effects of light weight, small envelope and rapid attitude maneuver are realized.

A spacecraft-borne propulsion device includes a propellant storage mechanism including a propellant storage container that stores a propellant in a vapor-liquid equilibrium state or a liquid phase, the propellant being ethanol or an aqueous ethanol solution; a propellant transport mechanism configured to supply, with an electric pump, the propellant under pressurization to a pressure exceeding 1 atm at room temperature; a gas heating mechanism including a heater including a separate heater for heating up connected via a check valve; a thruster head mechanism having a nozzle that generates a thrust with a heated gas; and a power supply mechanism including a storage battery for driving the electric pump and the heater for heating up. The propellant storage mechanism, the propellant transport mechanism, the gas heating mechanism, and the thruster head mechanism are connected in series.

A conductive liquid-fed pulsed plasma thruster includes a first electrode having a conductive solid portion and a conductive liquid portion, a second electrode separated from the first electrode to define an ignition space therebetween, at least one electric insulator separating the first and second electrodes, and a conductive-liquid passage extending within the conductive solid portion through which the conductive liquid portion flows from an inlet to an outlet located at the ignition space. The first and second electrodes are configured so that a drop of the conductive liquid portion forms and grows at the outlet when the conductive liquid portion flows through the conductive liquid passage until the drop of the conductive liquid causes an arc discharge between the drop and the second electrode that ignites the drop to produce a plasma cloud that generates thrust when exhausted.

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 is applied to the technical field of aerospace, and particularly discloses a satellite attitude and orbit integrated control device for step-length control thrusters, the satellite attitude and orbit integrated control device comprises a body, the mass center coordinate system of the body is 0-XYZ, a third thruster and a fourth thruster are rotatably mounted on the XOZ face of the body, a rotating meshing mechanism is arranged on the outer surface of the body, and the third thruster and the fourth thruster are meshed with each other through the rotating meshing mechanism. The posture of the body can be accurately adjusted by rotating the meshing mechanism. According to the satellite attitude and orbit integrated control device for controlling the thruster through the step length, when the body needs to be subjected to orbit adjustment, a nozzle in a certain axial direction is adjusted, so that the thrust direction of the nozzle is perpendicular to the axial direction of the advancing direction, the speed is provided, and orbit transfer is conducted according to the Kepler law; when the satellite needs to be subjected to attitude adjustment around a certain axis, the nozzles are adjusted to be parallel to the bottom surface, the two engines on the same plane are started to generate torque, the satellite is made to roll to adjust the attitude, and when the body needs attitude and orbit control, the engine step length can be adjusted to control thrust working time adjustment.

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.