124results about "Artificial satellites" patented technology

A reusable orbital vehicle (14) for a space transport system (10) comprises a fuselage (22) the shape of which widens in a direction extending from the front end (24) of the reusable orbital vehicle to a rear end (26) of said fuselage (22) and inside which there is defined a housing (32), and a crew-evacuation vehicle (34) housed in said housing (32) and oriented in such a way as to be extracted from the housing (32) in a direction (D1) directed from said rear end (26) of the fuselage (22) toward said front end (24) of the reusable orbital vehicle (14).

[Problem] The present invention makes it possible to provide, to an underwater apparatus located in an underwater area, an environment for communicating with an external communication apparatus or Internet connection. [Solution] The present invention is a communication system comprising: an information acquisition unit that acquires information relating to an underwater communication system for providing, to an underwater object located underwater, an environment in which a communication connection to an Internet connection or the like is possible; at least one unmanned watercraft that is capable of navigating on the water surface and is provided with a communication unit capable of communicating with the underwater object and / or the underwater communication system; a communication path determination unit that determines, on the basis of the information acquired by the information acquisition unit, a relay communication path for a communication line for establishing a relay connection between the underwater object and the Internet connection by using the unmanned watercraft and the underwater communication system; and an unmanned watercraft command determination unit that determines, on the basis of the determined relay communication path, an unmanned watercraft control command for configuring the relay communication path.

The application provides a satellite-borne sunlight reflection device and a satellite-borne sunlight reflection control method, and the satellite-borne sunlight reflection device comprises at least one on-satellite reflection plate and a satellite main body; the satellite main body is used for adjusting an attitude reflection angle between the on-satellite reflection plate and parallel sunlight; the on-satellite reflection plate is connected with the satellite main body, and in operation, a reflection surface faces a direction of irradiation of the parallel sunlight and is used for reflecting the parallel sunlight to a dark area on the earth surface; the satellite main body adjusts the attitude reflection angle of the on-satellite reflection plate by adjusting an attitude, and adjusts a ground reflection light intensity. The satellite-borne sunlight reflection device can adjust sunlight reflection to a shadow position on the earth, so as to accurately control the reflectivity, and the reflectivity and reflection position of a reflection area can be accurately controlled, and the influence of the satellite-borne sunlight reflection on other areas is reduced.

This application relates to the field of aerospace technology, and in particular to an irregularly shaped multi-satellite distributor. The lower end face of its irregularly shaped load-bearing structure is abutted against the upper end face of its conical load-bearing structure. The lower end face of the irregularly shaped load-bearing structure is circular, and the upper end face is elliptical. Its load-bearing beam is fixed within the mounting cavity of the irregularly shaped load-bearing structure. A top satellite mounting bracket is fixed to the upper end of the irregularly shaped load-bearing structure, and a side satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the major axis direction of the upper end face. An irregularly shaped satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the minor axis direction of the upper end face. Regularly shaped satellites are fixedly mounted on the top satellite mounting bracket and / or the side satellite mounting bracket, and irregularly shaped satellites are fixedly mounted on an irregularly shaped satellite mounting bracket. This application allows the protruding points of the satellites to be enveloped within a safe envelope range, and avoids increasing the size of the fairing.

This application relates to the field of intelligent aerospace technology and discloses a satellite autonomy method driven by embodied intelligent agents, applicable to subsatellites belonging to an in-orbit constellation. It constructs a multi-agent framework comprising multiple autonomous intelligent agents, including at least a resource assessment agent and a resource optimization agent. In response to a resource assessment command regarding the target mission, the method acquires the subsatellite's in-orbit operating parameters. The resource assessment agent is invoked to perform resource assessment based on the in-orbit operating parameters, obtaining the mission resource assessment result, which is then fed back to the in-orbit constellation. If mission feasibility indicates that the subsatellite can execute the target mission, the method acquires a single-satellite mission plan fed back by the in-orbit constellation. The resource optimization agent is invoked to optimize resource allocation based on the single-satellite mission plan, obtaining a mission operation plan. Its beneficial effects are that it improves the intelligence level of subsatellites in remote sensing observation missions and enhances the satellite's in-orbit autonomy.

The application discloses a kind of asymmetric configuration satellite attitude maneuvering process coupling compensation method, comprising: establishing the satellite attitude dynamics model including the non-diagonal item of moment of inertia matrix;Read current attitude sensor data, and calculate deviation quaternion and deviation angular velocity;According to the attitude dynamics model and current angular velocity data, calculate coupling compensation feedforward control torque;Based on the coupling compensation feedforward control torque, deviation quaternion and deviation angular velocity, the control torque after compensation is calculated;The control torque after compensation is limited and priority torque distribution, and the final control torque is obtained;Final control torque is sent to actuator drive satellite and carries out attitude maneuvering.Solve the problem that coupling effect caused by inertia product cannot be effectively compensated when asymmetric configuration satellite is in large-angle attitude maneuvering in prior art, and then affect control precision and stability.

The application discloses an ultra-low orbit adsorption type air-breathing pressurization system, which is used for an ultra-low orbit electric propulsion system and comprises an adsorption pressurizer on one side and a gas storage on the other side; the adsorption pressurizer is used for adsorbing rare thin gas molecules under low-temperature conditions and releasing the gas molecules under high-temperature conditions, and generates a pressurization effect; and the gas storage is used for storing the pressurized gas molecules and releasing the gas molecules on demand. The application discards the traditional mechanical compression mode and innovatively adopts an adsorption type pressurization principle. Through integrated design of a rotating staggered valve and an adsorption plate group, gas guiding, flow channel switching and adsorption pressurization functions are integrated into one, forming a highly compact overall structure. The application discards the traditional mechanical compression mode and innovatively adopts the adsorption type pressurization principle. The problems of a complex structure, a large size, extremely high power consumption and prominent energy contradiction of the prior art ultra-low orbit electric propulsion system are solved.

A kind of micro-nano satellite module rapid integration method belongs to the field of relay satellite platform, there is standardization backplane and standardization module, standardization backplane is provided with bus area and slot area, bus area includes uniform specification power bus and data bus, slot area is arranged with multiple standardization slots, standardization module is inserted into standardization slot.The structure of standardization slot is uniform, and a uniform standard data electrical interface is provided inside, each standardization slot corresponds to fixed physical port and logical address / node number, or unified address configuration is realized through slot coding, according to functional task requirement, select the standardization module that meets functional task requirement, and build the satellite system that meets functional task requirement.The present application significantly shortens the development cycle, greatly reduces the development cost, greatly improves the system flexibility and scalability, improves the system reliability and maintainability, and promotes standardization and industrialization.

This application provides a method, apparatus, electronic device, and storage medium for adjusting the solar array's attitude towards the sun, relating to the field of satellite control technology. The method for adjusting the solar array's attitude towards the sun includes: when the energy output corresponding to the target satellite's position at a preset orbital point is not within a preset energy output range, acquiring the target solar array's attitude information towards the sun and historical attitude information from the previous orbital period adjacent to the current orbital period; determining initial attitude adjustment information based on the target solar array's attitude information and the historical attitude information; determining multiple historical attitude adjustment information corresponding to the initial attitude adjustment information, and determining the target attitude adjustment information based on the multiple historical attitude adjustment information; determining the final solar array's attitude information based on the target attitude adjustment information and the target solar array's attitude information, and controlling the solar array to adjust its attitude based on the final solar array's attitude information.

[Problem] To efficiently determine the quality of acquired image information. [Solution] A processing device equipped with at least one processor that is used for imaging by a spacecraft capable of imaging the earth, wherein the at least one processor receives image information acquired by the spacecraft imaging an imaging target on the earth and / or attribute information associated with the image information, and executes processing for determining the quality of the image information on the basis of the received image information and / or attribute information.

Described are methods and systems for single-board satellite architectures.

The present invention provides an orbit design system for a global carbon inventory satellite, comprising: a long residence unit in the northern hemisphere configured to enable the global carbon inventory satellite operates in a mid-orbit elliptical orbit, and enable the global carbon inventory satellite to be located above the latitude of the human activity intensive region when it operates to its apogee.

The purpose of the present invention is to provide an image generation device that enables a user to easily acquire a satellite image. An image generation device according to the present disclosure comprises: an acquisition unit that acquires terminal information indicating an attitude of a terminal that images a prescribed region on a map; an estimation unit that estimates an attitude of a satellite that images the prescribed region on the basis of the attitude of the terminal; a generation unit that generates image data representing a case where three-dimensional data of the prescribed region is imaged from the satellite; and a communication unit that transmits the image data to the terminal.

Disclosed are systems and methods for a distributed space transportation network. Satellite launches to orbit are more efficiently performed by large rockets. Modern satellites are in smaller form factor, leaving the large launch rockets with excess capacity. Small satellite operators can use ride-shares, but do not have efficient options for delivering their satellites to their desired destination and may be forced to operate their satellites in compromise orbits. The disclosed distributed space transportation network maintains a fleet of space tugs, which can dock with satellites in space at an initial arrival destination and transport the satellites to their final destinations. In one embodiment, the space tugs can dock with satellite depots to obtain fuel and repairs.

Spacecraft servicing devices and related methods may include a propellant tank configured to store a propellant and to be placed into fluid communication with a portion of the target spacecraft.

The application provides a variable-configuration intercontinental high-speed transport system, which comprises a rocket booster and a cruise vehicle, and the cruise vehicle is arranged in series above the rocket booster; the transport system uses a mode of vertical launching by using a first-stage rocket booster+air-breathing cruise flight by using a second-stage rocket+vertical recovery, and has a variable-configuration capability, so that the vehicle meets different requirements of a vertical launching section, a hypersonic cruise flight section and a vertical landing section at the same time, and realizes efficient intercontinental transport capability.

Systems, methods, and devices in a satellite for on-orbit recovery are described. The device includes a plurality of circuits that are electrically connected to a backplane of the device, and a controller configured to monitor parameters of at least one of the plurality of circuits, and configured to store the parameters that are monitored with respective timestamps and respective orbit locations of the satellite. The controller is further configured to identify an error of a first circuit of the plurality of circuits based on the parameters that are monitored and the respective orbit locations of the satellite.

The application discloses a satellite flexible structure static deformation compensation method based on a piezoelectric actuator and belongs to the technical field of spacecraft control. The operation steps of the method comprise the following steps: installing p piezoelectric actuators at different positions of a flexible structure; measuring the static deformation of the flexible structure at m different measuring points; calculating the modal coordinates corresponding to the static deformation under the first n modes of the flexible structure, wherein p>n; taking the piezoelectric actuators as external force excitation sources of the flexible structure to construct a dynamic equation of the flexible structure; calculating the input voltage of each piezoelectric actuator according to the dynamic equation and the modal coordinates, thereby controlling the piezoelectric actuators to compensate the static deformation of the flexible structure. The method can quickly realize accurate and efficient compensation of the static deformation of the satellite flexible structure.

A rocket stage having a frame is disclosed. The frame includes a nose part and a mid-section part. The frame is configured to separate into two or more gliding parts after the stage has fulfilled its main purpose. After separation, the gliding parts glide down to the ground. The gliding part comprises a section of the nose part and a section of the mid-section part and a flat gliding surface alongside the section of the mid-section part. The slope of the nose part of the gliding part is steeper than the slope of the mid-section part of the gliding part. The air flows around the frame of the gliding part and a vortex forms above the upper edges of the gliding part, forcing the gliding part in the gliding position when it is in the atmosphere.

A small satellite docking adapter includes a probe adapter, the probe adapter including: one or more latches; one or more springs; and a cone adapter, the cone adapter including: one or more grooves, wherein: the one or more latches are configured to couple to the one or more grooves; and the one or more springs are configured to cause the one or more latches to engage with the one or more grooves to lock the probe adapter to the cone adapter, and to disengage with the one or more grooves to unlock the probe adapter and the cone adapter.

A satellite constellation forming system (100) forms a satellite constellation which is composed of a satellite group and in which the satellite group cooperatively provides a service. The satellite constellation has a plurality of orbital planes in each of which a plurality of satellites fly at the same nominal orbital altitude. A satellite constellation forming unit (110) continues providing the service while avoiding a collision between satellites by both or one of control of an orbital altitude and control of a passage timing of a satellite group flying in a region where the plurality of orbital planes intersect.

We provide a transmissive optical repeater. [Solution] The transmissive optical repeater of the present disclosure is configured to be incorporated into a satellite and comprises a set of optical fibers (FO), a first optical head (OHU1), a low-noise optical amplifier (LNOA), a polarization separation controller (PCS) having a channel polarization separation controller (C-PCS) for each output channel of an optical bandpass filter (BPF), a high-power optical amplifier (HPOA) for each output channel of the polarization separation controller (PCS), a quadrature polarization coupler (PBC) configured to recombine the two output channels of each of two high-power optical multiplexers (HPWDM), and a second optical head (OHU2) configured to transmit the output optical signal of the quadrature polarization coupler (PBC).

A frame for a satellite is provided. The frame includes a body and posts that are constructed of aluminum. The body can receive spacecraft components therein. The posts extend from the body and are operable to releasably couple with posts from another frame for another satellite and / or with a payload adaptor fitting.

According to the present invention, the effectiveness of search missions and other tasks when searching for an object using multiple systems can be improved. [Solution] The present invention is a control system that searches for an object in a predetermined first area using a plurality of moving bodies, and comprises: an object information interpretation unit that interprets object information relating to an object present in the second area based on acquired information obtained from a cooperative system capable of measuring a second area which includes at least a part of an area outside the first area; an object motion prediction unit that generates entry prediction information relating to the predicted content of an object entering the first area from an area outside the first area based on the object information; and an information output unit that displays or transmits to the outside information relating to the entry prediction information generated by the object motion prediction unit, or information relating to a portion or all of the plurality of moving bodies generated by the moving body motion determination unit based on the entry prediction information.

The objective is to safely recover the rocket that has descended into orbit after satellite separation, and to reuse the recovered rocket. [Solution] A rocket is launched from a rocket launch site located between 30 and 50 degrees latitude, and a rocket recovery method is used to recover the rocket, which has descended into orbit after satellite separation, in an area different from the rocket launch site. The recoverable rocket 202a is equipped with an atmospheric deceleration device, and the recoverable rocket 202a recovered by the above rocket recovery method is transported to the rocket launch site by a reusable rocket transport means. The rocket launch site then performs a rocket launch method to launch the recoverable rocket again.

This invention discloses a time-division multiplexing system for power generation and controllable space electromagnetic pulse radiation based on a conductive tether, belonging to the field of aerospace electric tether technology. The invention uses a single retractable conductive tether as the core carrier and adopts a time-division multiplexing topology. A fully integrated radiation-resistant solid-state switch array within the cabin achieves physical-level electrical isolation and nanosecond-level disconnect-before-make switching between power generation and pulse radiation modes. The controllable space electromagnetic pulse radiation is actively excited high-energy pulse radiation, with high-energy nanosecond-level electromagnetic pulses actively injected through a pulse emission module, significantly increasing radiation power and possessing practical value for electromagnetic passivation of space debris. In power generation mode, the tether and space plasma form a closed loop to cut the geomagnetic field and generate electricity. In pulse radiation mode, the tether acts as an end-fed directional long-line antenna radiating electromagnetic pulses. Active sheath control stabilizes the sheath potential and suppresses losses. Radio frequency isolation achieves electrical decoupling between the cabin section and the radiation section, and the mode switching timing does not depend on the plasma's millisecond-level relaxation response, avoiding the risks of mutual exclusion and self-interference. This invention resolves the fundamental conflict between single-strand conductive tether power generation and pulsed radiation. It exhibits strong dynamic stability and meets electromagnetic compatibility standards. The core components are based on on-orbit verification technologies such as TSS-1R, PMG, and T-Rex, making it highly feasible for engineering applications. It can be applied to aerospace scenarios such as satellite on-orbit operation and maintenance and space debris management.

The present invention relates to a spacecraft 1 for supplying electric energy to a client ship 2 located in free space, in orbit and / or on a celestial body, the spacecraft 1 comprising a main structure 10 having an electric thruster 10A, a chemical thruster 10B and a solar electric energy generator 10C, a first fuel container 11 for fuel for the electric thruster 10A and a second fuel container 12 for fuel for the chemical thruster 10B, the spacecraft 1 being such that the main structure 10 can alternatively be connected or separated to the first fuel container 11 or the second fuel container 12, the first fuel container 11 and the second fuel container 12 can be connected to or separated from each other, and the solar electric energy generator 10C is intended to be deployed or stored.