503 results about "Space debris" patented technology

An apparatus for orbital change of a satellite incorporates a capture mechanism for space debris and a tether connecting the capture mechanism to a satellite. The tether is extendable to position the capture mechanism relative to the space debris. A controller is employed for timed release of the space debris by the capture mechanism for orbital change by the satellite.

The invention discloses a rope net type space debris capturing and cleaning system which comprises a transmission subsystem and a fly net subsystem. The transmission subsystem comprises a transmitter, a net cabin, a net cabin cover ejecting device, a buffer device, a tying rope tension control device and a tying rope cutter, and the fly net subsystem comprises a rope net and four mass blocks; a net cabin cover can be ejected out of an outlet of the net cabin, and the rope net can be driven by the net cabin cover to be drawn out of the net cabin; the transmitter works at the moment when the fly net and a tying rope are straightened, the sides of the rope net are driven by four mass blocks according to preset transmitting opening angles, so that the rope net can fly out and is opened simultaneously until the rope net is fully opened, an object is netted after the rope net reaches the objective position, a port of the rope net is tightened under the control of the mass bocks, and the object can be completely captured; after the tying rope draws the captured object aircraft away from an orbit, the tying rope cutter cuts off the tying rope, and the object aircraft can be separated from a satellite platform. The rope net type space debris capturing and cleaning system has the advantages that a secondary transmission mode is adopted, so that twining and knotting can be prevented in rope net spreading procedures; the mass blocks are pushed to be synchronously transmitted, so that a millisecond-level synchronous transmission effect can be realized, and the shape of the rope net can be controlled advantageously.

A debris removal system that includes at least a control module that includes a debris contact module storage area, a maneuverable arm and camera storage area, a maneuverable arm, a camera, and a universal coupler socket; a guidance, control, and communications unit; a maneuvering module, wherein the maneuvering module includes fuel / propellant storage containers that provide fuel / propellant, via a control and fuel / propellant delivery conduit, to one or more control thrusters; a propulsion module, wherein the propulsion module includes fuel / propellant storage containers that provide fuel / propellant to at least one primary thrust device; and a debris contact module, wherein the debris contact module includes a debris contact member attached or coupled to a debris contact element, and wherein a universal coupler pin extends from the debris contact member so as to allow the debris contact module to be attached or coupled to the universal coupler socket of the control module.

A debris removal management system and a method of removing space debris. In one embodiment, the system includes: (1) a frame, (2) a plurality of net sections coupled to the frame and (3) a plurality of microvehicles coupled to the plurality of net sections and configured to be ejected relative to the frame to deploy the plurality of net sections, the plurality of net sections cooperating to form a net configured to capture space debris.

Disclosed herein, in certain embodiments, is a method of altering the stability of unstable space debris. In some embodiments, the method further comprises changing the orbit of the unstable space debris.

The invention discloses a method and device for cleaning space debris. According to the method, electromagnetic force is applied to the space debris running on an original orbit; the speed and / or direction of the space debris are / is changed under the action of the electromagnetic force; the space debris of which the speed is changed is changed to run on a new orbit under the action of gravity; at least the height of the perigee of the space debris on the new orbit is smaller than the height of the perigee of the space debris on the original orbit; the space debris running on the new orbit finally falls into the atmosphere of the earth under the action of atmosphere resistance. The device comprises a satellite bearing platform which is provided with an electric field device cleaning the space debris or a magnetic field device cleaning the space debris or a combination of the electric field device and the magnetic field device. The method and device for cleaning the space debris do not have high requirements for the accuracy of position detection of the space debris, micro space debris can be effectively cleaned, solar energy can be fully utilized for power generation, and self-carried energy is saved. The device for cleaning the space debris has the advantages of being simple in structure, low in manufacturing cost, and better in pulse action effect.

The invention discloses a child-mother satellite space debris clearing platform. The child-mother satellite space debris clearing platform includes a plurality of child satellites for clearing space debris; each child satellite is detachably connected the same mother satellite; the mother satellite is used for driving the child satellites to move among a multiple pieces of space debris; a space debris clearing device is arranged on each child satellite; and each space debris clearing device includes a child measurement device for measuring a pose of the space debris and a child satellite capturing device. A microsatellite is used as the mother satellite, and a plurality of micro-nano satellites are used as the child satellites; operations such as orbit transferring, closing and detailed survey, and capturing and de-orbiting can be completed coordinately; the child-mother satellite space debris clearing platform is low in cost, is fast in response speed, and is high in reliability; after the child satellites are released, the combined space debris observation network can acquire accurate debris pose information through coordinate observation; and the child satellites can independently complete de-orbiting operations, the mother satellite can directly go to next space debris to be cleared, time and energy can be saved, the benefits of space debris clearing can be increased, and the cost is reduced.

A gutter cover is adapted for covering an opening of a generally U-shaped eaves gutter used for catching and distributing rainwater runoff from a pitched roof. The gutter cover includes an elongated cover panel having a longitudinal upstream section adapted for residing proximate a top rear edge of the gutter adjacent an eaves of the roof, and a longitudinal downstream section located away from the roof eaves. The downstream section is adapted for engaging an outwardly-spaced leading edge of the gutter. The cover panel defines a first row of longitudinally-spaced debris filter openings adapted for receiving rainwater runoff from the roof into the gutter. A longitudinal weir ridge is formed downstream of the first row of debris filter openings, and is adapted for slowing and flattening a downstream flow of rainwater runoff bypassing the first row of debris filter openings. The cover panel further defines a second row of longitudinally-spaced debris filter openings downstream of the weir ridge. The second row of openings is adapted to further receive rainwater runoff from the roof into the gutter.

An apparatus for orbital change of a satellite incorporates a capture mechanism for space debris and a tether connecting the capture mechanism to a satellite. The tether is extendable to position the capture mechanism relative to the space debris. A controller is employed for timed release of the space debris by the capture mechanism for orbital change by the satellite.

The invention provides a dynamic process based debris flow dynamic risk analysis system and method. The method comprises the steps of: establishing a drainage basis model; obtaining spatial and temporal distribution of rainfalls; calculating debris flow characteristic parameters of the drainage basis space; and establishing a risk assessment model for analyzing dynamic risk of the debris flow and outputting a debris flow dynamic risk chart. The dynamic process based debris flow dynamic risk analysis system and method have the beneficial effects that: for considering the current situations and for solving the problems in the current debris flow risk researches, a new method for debris flow dynamic risk analysis based on debris flow dynamic process by considering dynamic boundary conditions is provided; a general framework model is established; a 3D dynamic risk chart can be output by inputting any space debris flow dynamic process and the characteristic parameters (flowing speed, flowing depth, flow process and the like) in various kinds of rainfall processes (at different design frequencies, and under monitoring); and therefore, the dynamic process based debris flow dynamic risk analysis system and method can be widely applied to debris flow monitoring and early warning, design of key parameters of prevention and treatment projects, disaster management and the like.

The invention relates to the field of the protective structures of spacecrafts, and discloses a lightweight filled composite protective structure for space debris, comprising a first layer of protective screen which is a corrugated board, a second layer of protective screen which is a corrugated board / density gradient metal foam; and spacing is left among the first layer of protective screen, the second layer of protective screen and a stiffener structure. The first layer of protective screen is in a corrugated shape, and the corrugated protective screen can effectively decompose debris cloud produced when the space debris with different incident angles collide with the protective structure at high speed, so that the damage energy is scattered. The corrugated surface plate of the second layer of protective screen decomposes the debris cloud again produced when the space debris collide with the first layer of protective structure, and a density gradient metal foam filling layer can further crush and melt the hit space debris, reduce the speed of the cloud debris, and minimize the damage to the external structure of a spacecraft and a sub-system.

Fiber-based debris interceptors are used to intercept and / or contain space debris. The debris interceptors may be made up of fibers that are formed in space from a material supply on a space vehicle. The fibers for the debris interceptors may be formed by extrusion, with a heat source, such as a mirror to focus sunlight, used to heat the material of the material supply that is carried on the space vehicle. The debris interceptor may be separated from the space vehicle, and used to remove debris from an orbit, or otherwise prevent debris entering an orbit from damaging a satellite or other spacecraft that travels in that orbit. The debris interceptor may be deployed prior to later launch of a valuable spacecraft, in order to “cleanse” the intended orbit of debris. Debris objects may pass through the debris interceptor, but in so doing may lose energy so as to de-orbit.

Provided are a space debris removing device and a method which enable easy installation of a deceleration device to space debris undergoing a tumbling motion. The space debris removing device includes: a propulsion device (3) for performing approach and attitude control on target debris (1); a capture device (4) having a harpoon (41) which can be ejected toward the target debris (1); an observation device (5) for calculating a capture position (E) and a capture attitude at which the harpoon (41) can be driven into a tank (11) (hollow portion) of the target debris (1) by observing a motion of the target debris (1); a deceleration device (6) directly or indirectly connected to the harpoon (41), for decelerating the target debris (1); and a body part (21) on which the propulsion device (3), the capture device (4), the observation device (5), and the deceleration device (6) are mounted.

The invention provides an antenna array forming technology-based ground-based radar system for space debris, which comprises a signal emitting part, a signal receiving part and a frequency and time unified signal board, wherein the signal emitting part comprises a signal generating source, a digital beam former, a phase synchronous circuit, a plurality of up-converters, a plurality of power amplifiers and an emitting antenna array; the emitting antenna array is a one-dimensional antenna array consisting of a plurality of emitting antennas along the east and west direction; each emitting antenna is used for emitting a signal amplified by a corresponding power amplifier to the space debris; the signal receiving part comprises a receiving antenna array, a plurality of low noise amplifiers, a plurality of down-converters, a plurality of analog-to-digital converters, a space debris accurate measurement module, a space debris common measurement module and a data postprocessing interaction module; and the receiving antenna array is a two-dimensional antenna array consisting of a plurality of receiving antennas arranged according to a plane, and is used for receiving signals reflected by the space debris.

The invention discloses a device and a method for capturing and cleaning space debris. The device comprises a satellite group. The satellite group is connected with a capturing net by conductive cables; the satellite group comprises a plurality of small satellites, cable collecting and releasing devices of the small satellites are connected with the capturing net by the conductive cables, and thesatellite group can drive the capturing net via the conductive cables to rotate with the debris; the capturing net comprises an upper-layer net and a lower-layer net, the lower-layer net is of a structure with a sealed bottom, a capturing net inlet capable of being contracted / expanded is formed in the upper-layer net, and the conductive cables are connected with the bottom of the lower-layer net.The device and the method have the advantages that the space debris can be captured and cleaned by the aid of the multiple small satellites, accordingly, the device and the method are low in cost andhigh in capturing efficiency, and fuel consumption can be effectively reduced.

The invention provides a characteristic relay method for relative position and attitude based on multi-vision. The method comprises the following steps: step 1. calculating position coordinates of the observed and extracted target characteristic points under a tracking spacecraft coordinate system by virtue of stereoscopic vision three-dimensional reconstruction to finish relative position calculation; step 2. establishing a coordinate system at each moment by virtue of any three points in the observed four characteristic points, and calculating an attitude matrix of every two coordinate systems; and step 3. calculating the attitude of the target body coordinate system at the current moment relative to the tracking spacecraft coordinate system, namely finishing the relative attitude relay calculation of a non-cooperative target. According to the method disclosed by the invention, the calculation of the relative position and attitude of non-cooperative targets such as spacecraft and space debris can be finished.

The invention discloses a rope net unfolding device based on a thin-wall type stretching arm, which is used for ontrack cleaning of space debris. The device does not need to be driven by external power, and a net body structure is driven to unfold by strain energy stored by the thin-wall type stretching arm, so as to capture the space debris. Through adoption of a thin-wall structure, the stretching arm has high flexibility and light weight, the stretching arm stretches out and contracts back to store strain energy and to enable the mechanism to achieve a high magnification ratio. Due to the design that the stretching arm drives the net body structure to unfold, the unfolding process of the net body structure has good controllability; and due to the design of a press plate and a limiting pin, the unfolding synchronism is ensured.

The invention provides a repeatable and flexible capture structure based on dielectric elastomer and shape memory polymer and a method for capturing space debris. According to the repeatable and flexible capture structure based on the dielectric elastomer and the shape memory polymer and the method for capturing the space debris, the problems that traditional aerospace machines and motor driving capture structures are complex and large in weight are solved. The repeatable and flexible capture structure based on the dielectric elastomer and the shape memory polymer is mainly composed of the unfolding structure of a shape memory polymer composite material sheet layer, namely a three-wing type unfolding beam, and the dielectric elastomer. The structure constituted by the shape memory polymer composite material sheet layer is used for driving and controlling the capture structure to the designated spot, and the structure constituted by the shape memory polymer and the dielectric elastomer is used for capturing the space debris. The repeatable and flexible capture structure based on the dielectric elastomer and the shape memory polymer is simple in structure and high in reliability, and is applied to repeatable and flexible capture structures.

The invention discloses a space debris rope system dragging despinning and removal method, relates to space debris despinning and removal methods and belongs to the technical field of spaceflight. By the adoption of the method, a trailer spacecraft is launched into an orbit, so that the rendezvous with debris is finished; a trailer is connected with the debris through multiple tying ropes; a dynamical model is established, a control law is designed to adjust the nominal rope length of each tying rope to change the tension of the tying ropes, despinning of the debris is achieved through the cooperation of the multiple tying ropes, the situation that the tying ropes are cut and broken through protruding components on the debris is avoided, and the stability and the redundancy of a rope system trailer system are improved; and a trailer engine is used for ignition, reverse dragging force is applied, the space debris is pulled into the atmosphere to be burnt, and therefore debris removal is achieved. By the adoption of the method, despinning of the space debris can be achieved, the situation that the tying ropes are cut and broken through the protruding components on the debris is avoided, and the stability of the system is improved; and in addition, the redundancy of the tying ropes is improved by increasing the number of the tying ropes. The method can be used for space debris despinning and removal and can also be used for despinning of and control over other non-cooperative targets.

The invention relates to a diffuse reflection laser ranging and high-resolution imaging synchronous measurement photoelectric telescope system capable of being used for space debris and belongs to the technical field of photoelectric telescopes. The system comprises a laser emitting part, an echo receiving part and a self-adaptive optical imaging part. According to the laser emitting part, laser light emitted by a laser enters a telescope through a laser emitting light path to be emitted to a space target. According to the echo receiving part, an echo light beam is received by the telescope and then enters an echo receiving and detecting light path, echo signals are generated and then transmitted to a laser ranging control circuit, and a laser ranging distance value is obtained after software processing. According to the self-adaptive optical imaging part, the space target light is received by the telescope, enters a self-adaptive optical system and is shrunk by a beam shrinking light path, a Hartmann sensor and a wavefront detector are used for detecting overall wavefront inclination and wavefront errors to obtain the overall wavefront inclination degree and wavefront error data which are used for controlling a fine tracking system to correct the wavefront inclination and controlling a deformable mirror to correct the wavefront errors, and therefore a high-resolution image is obtained.

The invention discloses a method for compositely controlling attitudes and orbits of in-orbit dragging combination spacecrafts. The method includes steps of identifying mass property parameters of the combination spacecrafts in an in-orbit manner; determining composite control on the attitudes and the orbits of the combination spacecrafts in orbit maneuver periods. The method has the advantages that abandoned spatial targets can be deorbited and maneuvered under the conditions of high disturbance torque and uncertain parameters; the method can be used for cleaning abandoned satellite or space debris in the earth orbit, and accordingly clean running environments can be provided for satellites which run normally.

The invention discloses a relative positioning method for a spatial object. The method comprises the following steps of: releasing a micro satellite carrying a visible light camera by an on-track spacecraft; measuring the distance between the on-track spacecraft and the micro satellite by a microwave distance measurement method; measuring the azimuths and pitch angles of the on-track spacecraft and micro satellite in respective coordinate systems by the direction finding equipment; forming a double-star observation system from the on-track spacecraft and the micro satellite to observe the spatial object together; establishing a double-star inter-star coordinate system by taking the connection direction of the double stars as an X axis, wherein a Z axis is located in a plane formed by the double stars and the earth's core, is vertical to the X axis and points to the earth's core; calculating a transformation matrix from the coordinate system of the satellite to the double-star inter-star coordinate system; calculating the azimuth and pitch angle of the object in the double-star inter-star coordinate system through coordinate system transformation; and finally calculating the three-dimensional position of the spatial object by intersecting the sight lines of the double stars. In the invention, the relative relationship between the double stars is calculated firstly, a double-star inter-star coordinate system is established, then the azimuth and pitch angle measured by the coordinate system of the satellite are transformed to the double-star inter-star coordinate system through coordinate transformation, and finally the accurate position of the object is obtained by intersecting the sight lines of the double stars; the positioning speed is high, and the accuracy is high; and the method is particularly suitable for the real-time positioning for determining the space junk near the spacecraft.

A spacecraft for removing space debris is disclosed. The spacecraft includes a satellite bus, a shield member foldable on an outer side face of the satellite bus and disposed facing towards space debris to reduce a movement speed of the space debris, and a support member configured to support the shield member with respect to the satellite bus, in which the shield member includes a central panel configured to overlap one face of the satellite bus, a plurality of first panels connected to peripheral sides of the central panel and radially extended, and a plurality of second panels located between the first panels.

The invention discloses a space target high-precision on-orbit positioning method based on angle measurement and distance measurement information. An image of space debris is shot through a visible light camera; distance information of the space debris is obtained through a microwave distance measurement load; angle information of the space debris relative to a satellite is extracted through an optical recognition algorithm; and distance information of the space debris relative to the satellite is obtained through a radar pulse distance measurement model. The angle information and the distanceinformation of the space debris relative to the satellite are known, and a high-precision positioning result of the space debris can be obtained through a data fusion model. The arc segments of the satellite on-orbit tracking space debris can be reduced; the high-precision positioning of the space debris can be realized in combination with a calculation method convenient for software processing through the angle information and the distance information obtained through measurement in a relatively short tracking time; and the high-precision positioning of a space target can be realized withina space-based observation time from tens of seconds to a few minutes.

The invention relates to a low earth orbit satellite autonomous orbit determination method which comprises the following steps of selecting a plurality of high earth orbit objects (including large space debris) as calibrated satellites to acquire a precise ephemeris of the high earth orbit objects; when a low earth orbit satellite cannot acquire the precise ephemeris, utilizing a boarded space-based visible light camera to point to the high earth orbit calibrated satellites, performing tracking for a period of time to acquire an only angle measurement quantity; utilizing an on-board computer,combining rough ephemeris of the low earth orbit satellite, the precise ephemeris of the high earth orbit objects with the measurement quantity, performing reversed determination to obtain a precise orbit of the low earth orbit satellite. The invention provides a novel satellite autonomous orbit determination way; and when global positioning navigational satellite information cannot be utilized, the novel satellite autonomous orbit determination method can be used as an alternative method. The low earth orbit satellite autonomous orbit determination method provided by the invention is relatively simple in calculation model, fully utilizes space resources and meanwhile, is low in space-based visible light observation cost, small in energy consumption and strong in engineering applicability.

The invention discloses a space debris prevention structure of a spacecraft, which comprises a flat and vertical inner board, a corrugated outer board and a metal foam sandwich layer, wherein the metal foam sandwich layer is arranged between the flat and vertical inner board and the corrugated outer board. The space debris prevention structure of a spacecraft disclosed by the invention can effectively resolve the debris clouds generated by impacting the prevention structure at a high speed by the space debris of different incidence angles and scatter the damage energy of the debris clouds; the filled metal foam can further crash and melt the impacted space debris and reduce the speed of the debris clouds; when the debris clouds pass through an inner flat and vertical metal panel, the inner panel further scatters the damage energy of the debris clouds, reduces the speed of the debris clouds again and stops the residual space debris after passing through the metal foam layer to a certain extend, so as to effectively reduce the damages of the space debris to the external structure of a spacecraft and a sub system.

The invention relates to a measuring method for the real-time accurate direction of an astronomical telescope, and a control system works according to the following steps: (1) acquiring the observation time and the directing data; (2) acquiring images; (3) generating theoretical star atlases; (4) carrying out the full-frame scanning of the images; (5) carrying out the matching distribution of the theoretical star atlases and actually-measured star atlases; (6) measuring the direction of the telescope; and (7) giving out the accurate direction of the telescope in real time by a computer system according to the input data, and guiding the telescope to carry out the observation and the tracking of space debris. In the invention, the accurate direction of the telescope can be measured in real time according to an acquired CCD (Charge Coupled Device) image as well as the moment corresponding to the image and the direction with an error, the measured direction precision of the telescope is superior to one half of the image-element resolution of the CCD image, the differing degree of the error of the telescope is measured in time, and the actual direction of the telescope can also be accurately measured. The processing effect is good, and the measuring method can be widely applied to the fields of scientific researches and projects and is particularly favorable to the direction correction of movable-type observing equipment.

The invention discloses a remote sensing satellite space debris evasion maneuvering method based on ground-track keeping needs. According to an allowable ground track drifting range of remote sensing satellite on-orbit operation, a reference ground track keeping control ring is built, and is divided into different areas, and different maneuvering strategies are determined for a spacecraft according to the different areas, wherein, dangerous intersection possibly happens to the spacecraft, and the track maneuvering evasion space debris collision threat is needed to be carried out, so that the aim of avoiding the collision risk is achieved. According to the method, the ground-track keeping control ring is adopted to serves as the basis for generating the strategy, and the method is simple, efficient, easy to operate and particularly suitable for operation of space debris collision evasion with high timeliness; and the efficiency can be improved, time is saved, the valuable time is earned for space debris collision evasion, and satellite in-orbit operation safety is improved. Meanwhile, the influence of evasion maneuvering on a flying task of the spacecraft is reduced to the greatest extent, and the satellite space debris collision early warning and evasion capacity are improved.