46 results about "Solar sail" patented technology

This invention relates to a test device and verification method for the space-ground consistency of a flexible spacecraft control system based on uncertainty quantification, belonging to the fields of intelligent manufacturing equipment industry and space-ground consistency verification technology. To address the problem of space-ground consistency verification and evaluation for flexible spacecraft, this invention includes an air-floating robot and a flexible component. The air-floating robot simulates the central rigid body of the flexible spacecraft, and the flexible component simulates the deployed solar panels. One end of the flexible component is connected to the upper support column of the air-floating robot, while the other end is free and equipped with a weight. Fiber optic grating sensors are distributed on the flexible component to measure flexible vibrations. A piezoelectric ceramic plate is installed at the end of the flexible component connected to the air-floating robot for driving and controlling the vibration of the flexible component. This invention evaluates the space-ground consistency of the designed flexible spacecraft control test system and, based on uncertainty quantification analysis, provides the impact level of uncertainty sources on space-ground consistency.

The application discloses a solar sail installation rotating hinge device, which comprises a hinge body and a sail fixator. The hinge body is provided with hinge tables at both ends, the sail fixator is provided with connecting tables at both ends, the hinge tables and the connecting tables are connected through rotating joints, the sail fixator can rotate around the hinge tables, the rotating joints are sleeved with hinge torsional springs matched with the hinge tables and the connecting tables, and one connecting table is further provided with a time sequence torsional spring at the outer end. The application is fixedly installed on the side surface of a satellite structure frame and can be reliably connected with the sail when in use. Meanwhile, the rotating joints for rotating and connecting the sail are built in the hinge body, flexible rotating connection between the sail and the satellite body is realized, and the secondary expansion requirement of the sail is supported. The hinge is further provided with a time sequence control structure to prevent interference and collision in the expansion process and prevent over-expansion or under-expansion.

This invention relates to a method for the active deployment of solar panels based on path planning. After the solar panel clamping release mechanism is unlocked, the initial value of the rotation angle of the tilting mechanism in the clamped state is read, and rotation path planning is performed according to the actual state of the solar panel. When the tilting mechanism starts to drive the solar panel deployment, the solar panel deployment time is timed. If the deployment process time has not reached the maximum time, the solar panel deployment is determined sequentially by calculating the rotation angle and angular velocity of the tilting mechanism, the rotation angle calculated by the potentiometer output, and continuously judging the working mode of the tilting mechanism over multiple cycles. If the deployment process time reaches the maximum time and the solar panel has been fully deployed, the solar panel deployment process ends. If the deployment process time reaches the maximum time and the solar panel has not been fully deployed, a safety protection setting is implemented. This invention is applicable to the design of active solar panel deployment modes for satellites with solar panels featuring extension rod structures and tilting mechanisms, driving the solar panels to deploy smoothly, stably, safely, and reliably.

The invention relates to a multi-antenna SAR (Synthetic Aperture Radar) satellite configuration and a rocket, and the multi-antenna SAR satellite configuration comprises a satellite platform, a first antenna, a second antenna, a first solar panel and a second solar panel, the peripheral side face of the satellite platform comprises a first installation face, a second installation face, a third installation face and a fourth installation face which are sequentially arranged in the circumferential direction of the satellite platform, and the first antenna is assembled on the first installation face and can be unfolded in the direction parallel to the axis of the satellite platform. The second antenna is assembled on the second installation face and can be unfolded in the direction away from the first antenna, the first solar panel and the second solar panel are installed on the end faces of the two axial ends of the satellite platform respectively and can be stacked on the third installation face in a surrounding mode, and the fourth installation face serves as a rocket assembly face. The satellite is compact in overall configuration layout, adapts to a multi-load antenna, is high in integration level, is small in envelope size, is high in rocket adaptability, can be laterally hung and arranged at the top, improves the rocket plot ratio, and can achieve one rocket with multiple satellites.

The application discloses a kind of lightweight high-reliability rotor position detection method and device of large hollow permanent magnet motor, belong to permanent magnet synchronous motor control technical field.The method includes: S1 in stator slot opening embedded installation 3N linear hall sensor, wherein N is positive integer and N≥1, according to slot pitch angle α Calculation common offset, 3N original signal is time shift processing, obtain 3 (N+M) reconstructed hall signal, M is integer and M≥0;S2 executes amplitude range detection and phase relationship detection, reconstructs fault signal using adjacent leading phase normal sensor signal;S3 is calculated rotor position by vector space decoupling transformation and phase extraction method;S4 upper motor controller carries out fault detection and feedback channel switching instruction, realizes system level fault tolerance.The application realizes structure lightweight, high-precision detection, sensor level fault tolerance and system level fault tolerance, is suitable for space power station solar sail etc. Motor drive system high-reliability control demand.

The invention discloses a small space debris-oriented passive capturing and disposing spacecraft, cubesat and method, solves the problems that in the prior art, small space debris cannot be captured, and an existing capturing device performs active capturing, and has the beneficial effects that the space debris is passively captured and thoroughly eliminated. According to the specific scheme, the passive grabbing and handling spacecraft for the small space debris comprises a chip satellite and a capturing assembly, a plurality of solar sails are connected to the peripheral side of the chip satellite, the chip satellite controls the bending state of the solar sails so as to control flight of the spacecraft, and the capturing assembly comprises a folding storage wall plate, a triggering platform and a bistable capturing arm; the folding storage wall plate is connected to the peripheral side of the chip satellite through the hinge, the hinge is a bent hinge and has elasticity, the triggering platform is connected with the chip satellite, and after the bistable capturing arm passes through the triggering platform, the bistable capturing arm is in a straight state when a spacecraft is in a non-working state.

The application discloses a satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints, including the constraint conditions of comprehensively considering sun-pointing, inter-satellite communication pointing and the drastic change of sunlight conditions under low-inclination orbit, proposing a nominal expected attitude design method, a modified expected attitude design method, a transition expected attitude design method, a flip nominal expected attitude design method and a flip modified expected attitude design method, and giving an expected angular velocity calculation method of manipulation aiming at the corresponding expected attitude. The application can meet the basic condition that the satellite solar panel normal direction points to the sun, ensure the communication pointing requirements between space formation satellites at the same time, and avoid the influence of the polarity change of the satellite on the low-inclination orbit, thereby providing a stable and safe long-term sun-pointing attitude for the satellite.

The application provides a spacecraft solar panel power regulation system and relates to the technical field of spacecraft power supply.The system comprises a detection module for acquiring telemetry data of a photovoltaic array; a calculation module for determining target control parameters of power control according to the telemetry data and generating remote control instructions according to the target control parameters; a power supply lower computer module for acquiring the global energy state of a spacecraft solar panel; determining target working modes of power control according to the global energy state in combination with the telemetry data and generating mode instructions according to the target working modes; an MPPT master control module for adjusting control parameters of power control or switching working modes of power control according to the remote control instructions or the mode instructions; and performing power control according to the telemetry data of the photovoltaic array based on the working modes and the control parameters. The application improves the management effect of the power of the spacecraft solar panel.

The present disclosure relates to a method for calculating the area of a solar array shielded by a satellite based on the finite element method, which comprises: constructing a finite element calculation model of a satellite shielding a solar array, the model including a satellite body and a solar array; proving the convergence of the finite element calculation model of the satellite shielding the solar array; and calculating the area of the solar array shielded by the satellite body according to the proof result. The method for calculating the area of a solar array shielded by a satellite proposed by the present disclosure divides and calculates the first solar surface of the solar array by using the finite element method, which can effectively calculate the area of the solar array shielded by the satellite body in various attitudes, and the calculation result is basically consistent with the real data of the satellite in orbit.

The invention provides an on-orbit self-connection flexible composite material stretching mechanism for a spacecraft. The on-orbit self-connection flexible composite material stretching mechanism comprises a flexible composite material stretching arm and a self-connection unit. And when the stretching arm units are unfolded and straightened to return to a partially overlapped C shape, the two parts of the self-connecting unit are in contact with each other and form self-connection through physical or chemical reaction. Self connection of the stretching mechanism in the space environment can be achieved, and the flexible mechanism has the advantages of being resistant to instability, high in rigidity and light in weight and is an important development direction of flexible mechanisms such as large-area-array winding type flexible solar wings, large-area-array deorbiting sails / solar sails, large-size deployable light shields and large-size deployable film antennas of spacecrafts.

The invention discloses a spacecraft solar panel thermal coupling dynamic compensation ground thermal control simulation system and method, and solves the problem that the existing spacecraft solar panel thermal simulation technology lacks dynamic feedback capability. The spacecraft solar panel thermal coupling dynamic compensation ground thermal control simulation system comprises a thermal coupling correction interface, a calculation unit, an environment simulation unit and a telemetering synchronization interface, and the calculation unit comprises a temperature difference trigger module, a dynamic calibration module, a heat flow compensation module and the like. The spacecraft solar panel thermal coupling dynamic compensation ground thermal control simulation method adopting the system comprises the steps of temperature collection, triggering calibration judgment, k value updating, thermal coupling compensation amount calculation, temperature updating, output temperature correction and the like. According to the method, the static limitation of thermal coupling modeling is broken through by constructing a temperature difference triggered closed-loop compensation framework.

The invention discloses a high-dynamic real-time simulation solar panel electric energy output closed-loop test method and system, and the method carries out the real-time dynamic simulation solar panel electric energy output closed-loop test through the downlink telemetry information of a spacecraft and the sun position information provided by ground equipment of a GNC subsystem. And the received uplink and downlink telemetering pose data of the spacecraft are calculated and converted in real time to obtain high-dynamic output voltage and current of the sun matrix for supplying power to the spacecraft in real time.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

An angle measurement system for a solar panel driving mechanism comprises a measurement and control device, a driving motor and an upper computer, the measurement and control device comprises an angle measurement module, a motor control module and a motor measurement module, and the rotation angle of the solar panel driving mechanism is directly measured through the angle measurement module; the motor measuring module acquires voltage and current data of the motor, whether the motor is abnormal or not can be diagnosed in an auxiliary manner, and the reliability of measurement and control is further guaranteed, so that the precision and reliability of a measuring system are improved, more accurate feedback data are provided for an attitude control system, and the reliability of the attitude control system is improved. Therefore, the attitude control precision and the operation stability of the solar panel are improved.

The present application relates to the field of aerospace technology, and particularly relates to a swing wire harness type sailboard driving mechanism. The present application provides a swing wire harness type sailboard driving mechanism, which comprises an A-axis driving assembly, a double-axis connecting support, a B-axis driving assembly and a sailboard connecting support, the A-axis driving assembly is connected with a star body through a main shell, one end of the double-axis connecting support is fixedly connected with an output end of the A-axis driving assembly, the other end of the double-axis connecting support is fixedly connected with the B-axis driving assembly, one end of the sailboard connecting support is fixedly connected with an output end of the B-axis driving assembly, the other end of the sailboard connecting support is connected with a solar sailboard, and the output end rotating shaft of the A-axis driving assembly and the output end rotating shaft of the B-axis driving assembly are perpendicular; a swing wire harness is connected with a low-speed transmission shaft in the A-axis driving assembly to transmit signals and power. The present application provides a swing wire harness type two-degree-of-freedom sailboard driving mechanism with high integration, low cost and batch production.

The invention discloses a low-orbit verification satellite, and relates to the technical field of satellites, and the satellite comprises a satellite main body which is of a flat structure, is provided with a front surface and a back surface, is provided with an attitude orbit control system, and is used for driving the satellite main body to overturn through the attitude orbit control system; the verification load is mounted on the front surface of the satellite main body; the back load is mounted on the back surface of the satellite main body; the solar sailboard is connected with the satellite main body through a sailboard driving device, and the sailboard driving device is used for driving the solar sailboard to turn over; according to the technical scheme, after the verification task is completed, the satellite main body performs attitude overturning through the attitude orbit control system, the solar panel is overturned through the panel driving device, and the back load of the satellite main body is started, so that the redundant value of the verification satellite can be fully utilized, and the service life of the satellite is effectively prolonged. The method not only reduces the cost of a single task, but also greatly improves the utilization efficiency and economic benefits of verification satellites.

The invention provides a solar panel driving device control method and device, equipment and a medium. The method comprises the following steps: acquiring a current track illumination state; based on the current track illumination state, corresponding working modes are selected and entered, and the working modes at least comprise a sunshine area working mode and a shadow area working mode; according to the entered working mode, a corresponding driving instruction is generated and output, and the driving instruction is used for controlling movement of a driving mechanism of the solar panel driving device so that the angular acceleration of the driving mechanism can change continuously in the movement process. The scheme is used for solving the problem that the periodic motion of a slip-ring-free swing type SADA in the related technology can influence the attitude stability and the pointing precision of a satellite platform.

The invention belongs to the technical field of satellite navigation and satellite communication anti-interference, and provides a low-orbit satellite solar panel reflection GNSS interference modeling method and system, which is based on the orbit and attitude of a low-orbit satellite, and the attitude and geometric dimension of a solar panel, and depicts the geometric shielding effect of navigation signals such as Beidou in a satellite receiving link at the same time. And the additional interference effect of the ground power suppression type GNSS jammer signal on the satellite-borne receiver after the ground power suppression type GNSS jammer signal is subjected to mirror reflection and diffuse reflection through the satellite solar panel is avoided. According to the method, an interference link and a carrier-to-noise ratio C / N analysis model are established in a joint simulation environment formed by track and link simulation software and a numerical calculation platform, and data such as an attitude, a sailboard angle, a C / N time sequence and the like which are actually measured in a track can be directly received as input; the method is used for evaluating interference risks under different solar panel attitudes, interference source positions and power conditions, and provides a basis for solar panel control strategies, satellite operation scheduling and ground interference environment evaluation.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The application belongs to the technical field of solar sail and is specifically a heat-driven liquid crystal solar sail unfolding device. The liquid crystal solar sail unfolding device is composed of liquid crystal elastomers as a framework adhered to a solar sail film. The shape of the solar sail film includes a large equilateral triangle and three small squares outside the equilateral triangle, and each square film is symmetrically arranged with the triangular film. The liquid crystal elastomer is in the shape of a strip and is a double-layer film structure printed by a direct writing 3D printer. The bottom layer is horizontally printed, and the upper layer is vertically printed. When the temperature changes, the thermal expansion coefficients of the double-layer film are different, and the overall produces bending deformation. The application can achieve graded regulation of the unfolding degree by different temperatures to meet different requirements. Under the premise of realizing the intelligent storage and unfolding functions, the application has a high storage ratio. The structure is flexible and controllable, and the unfolding and contraction of a certain sail surface can be controlled to adjust the overall attitude and forward direction.

The utility model relates to the lighting equipment of sunlight introduction room, including the mounting seat, the drive assembly is installed on the right side upper side of mounting seat, the mounting seat is rotatably connected with the gear, the mounting seat is rotatably connected with the connecting piece, the connecting piece is connected with the gear, the connecting piece is connected with the light cover, the light cover is connected with the solar sail, the solar sail is installed with the photosensitive sensor, the light cover upside is connected with the light guide column, the solar sail is connected with the protective cover, the protective cover is connected with the reflector cover, and the battery is installed in the mounting seat. The utility model discloses through the hollow design of protective cover and the combination of reflector cover, guarantees the reflection efficiency while protecting the light cover from the outside environment interference, and the installation mode of reflector cover fixed on the protective cover makes the structure of reflector cover more stable in the outdoor use environment, and the protective plate and protective frame can protect the battery and main gear, gear, prolong the service life of equipment.

The invention relates to the technical field of space biological tests, in particular to a reusable space life science test platform which comprises a service unit and a load cabin, and the service unit comprises a structural system, a service power system, a communication system and a spaceborne computer system; the structural system comprises a structural body, a service docking mechanism and a solar panel, the service power system, the communication system and the spaceborne computer system are all located in the structural body, the service docking mechanism is located on the end face of the structural body, and the solar panel is connected to the circumferential surface of the structural body; the load cabin comprises a cabin body structure, a cabin body power system, a cabin body docking mechanism and a cabin body clamping mechanism; the cabin body power system is positioned in the cabin body structure; the cabin docking mechanism is positioned in the center of the large bottom of the cabin structure; the cabin body clamping mechanism is located on the circumference of the large bottom of the cabin body structure. According to the invention, test results or products can be brought back to the ground in time, high-frequency and multi-sample attributes of the test are met, and the test cost can be reduced.

The invention discloses a reconfigurable modular comprehensive support system for a long-term lunar surface task, and particularly relates to the technical field of star catalogue detection. Comprising a lunar surface moving module which can move in all directions on the lunar surface and is provided with a module-level butt joint interface; the energy supply battery can provide a power supply for the lunar surface mobile module, and the energy supply battery is provided with an equipment-level butt joint interface and a module-level butt joint interface; the solar sail module is provided with a module-level butt joint interface, and the solar sail module is used for charging the energy supply battery; the lunar surface cabin section is provided with a cabin section level butt joint interface; the mechanical joint modules are provided with equipment-level butt joint interfaces, and the mechanical joint modules are connected through the equipment-level butt joint interfaces to form a mechanical arm; and the rotating base is arranged on the mechanical joint module, and a module-level butt joint interface is arranged on the rotating base. By the adoption of the technical scheme, the problem that an existing lunar surface facility cannot achieve modular replacement is solved, the maintainability and replaceability of a lunar surface task can be effectively improved, and the operation efficiency of the lunar surface task is improved.

The application provides a solar panel card stagnation under panel energy supply calculation method, device and electronic equipment, and the method comprises the following steps: obtaining the corresponding card stagnation angle γ when the solar panel is in the card stagnation state, the card stagnation angle γ is the included angle between the plane where the solar panel is in the card stagnation state and the plane where the solar panel is in the electrical zero position; determining the interval where the solar incidence angle of the solar panel is located in a period according to the solar panel in the card stagnation state, the interval comprises an illumination period and a shadow period, wherein the solar incidence angle is the included angle between the solar vector direction and the normal direction of the solar panel; calculating the output energy of the solar panel in the illumination period in a period, and the output energy is an expression related to the card stagnation angle, wherein a period of the solar panel corresponds to an orbit period of the satellite. The application can accurately characterize the solar panel energy supply under different card stagnation angles, and provides effective support for subsequent satellite on-orbit practical application.

The application discloses a spin-nutation coupling spin-down plume pointing method and system based on plume spin-down, and belongs to the technical field of spacecrafts. The method comprises the following steps: collecting the rotational inertia and angular velocity data in the motion state of a target; counting spin-nutation spin-down torque parameters that can be received by a solar panel of the target during spin-down; inputting the spin-nutation spin-down torque parameters, the rotational inertia and the angular velocity data into a coupling spin-down plume pointing model to obtain an optimal plume pointing point; inputting the optimal plume pointing point into a plume spin-down dynamics model to perform spin-down calculation and obtain a spin-down termination condition; and if the motion state of the target meets the spin-down termination condition, the spin-down task is completed. According to the motion characteristics, mass characteristics and shape characteristics of the target, the spin-nutation coupling spin-down of the target can be realized, the situation that a spin-down guidance law is insufficient can be solved, and the method is suitable for various spin-down modes of using the resistance moment generated by the solar panel of a failed satellite.

The utility model discloses a kind of aerospace single-shaft small mechanism measurement circle grating shafting devices, device includes mounting bracket, bearing mounting seat is fixed on it;Driving shaft, in bearing mounting seat, the outside of driving shaft is connected with bearing mounting seat through contact bearing;Circle grating mounting shaft, with driving shaft fixed connection, and located in mounting bracket, the end of circle grating mounting shaft is connected with circle grating;Reading head, symmetrically set on mounting bracket, and located in the opposite side of circle grating;Adapter shaft, set in the one end of driving shaft away from mounting bracket;Control device, set in the outside of bearing mounting seat, with reading head electric connection.After the device is connected with solar panel driving mechanism, the driving force of solar panel driving mechanism is used to drive the device synchronous rotation, in the process of rotation, the reading head located in the opposite side of circle grating can accurately read speed stability value and angle precision value in time, to improve the efficiency and accuracy of measurement.

This invention provides a solar panel assembly and its deployment method, belonging to the field of meteorological satellite technology. The solar panel assembly of this invention includes a solar panel composed of a left panel, a middle panel, and a right panel, a timing mechanism, a clamping and releasing mechanism, and a deployment locking mechanism. The timing mechanism is installed on the deployment shaft of the first deployed solar panel. Through the cooperation of a rotating pin and a semi-circular pin, the left and right panels are deployed step by step in a set sequence. The deployment locking mechanism drives a locking pin to slide along a guide slide and lock into a locking groove via a locking spring, triggering a microswitch to provide a feedback signal. After the clamping and releasing mechanism unlocks, the solar panel is deployed under the drive of the deployment locking mechanism. This invention achieves orderly deployment of the two panels through purely mechanical timing control, avoiding collisions and interference during the deployment process, and improving system reliability and safety.

The invention relates to a flexibly-connected three-dimensional solar panel mechanism, which comprises a panel main body which is configured to perform solar photoelectric conversion and provide power supply for in-orbit operation of a satellite, the panel main body is an inflatable shaping type panel main body, when the panel main body is unfolded in orbit, the panel main body is formed into a polygonal pyramid pair from a folded state, and the polygonal pyramid pair is formed into a rectangular pyramid pair from a folded state; the cone bottoms of the polygonal cone pairs are mutually attached and joined, and the cone angles are oppositely positioned at the two ends of the sailboard main body; a storage cabin; the booth forming assembly is configured to open the sailboard main body; the flexible connection assembly is configured to flexibly connect the sailboard main body with the satellite body and realize electric energy transmission and signal communication between the sailboard main body and the satellite body; and the electric propulsion assembly is configured to adjust the pose of the sailboard main body and provide orbit maintaining thrust for the satellite body. The interference of vibration, disturbance and the like of the solar panel on the satellite body is reduced to the maximum extent, and meanwhile the electric energy obtaining capacity of the solar panel is not reduced.

The utility model relates to a pressing and unfolding device of a satellite antenna, which does not need an independent locking and releasing mechanism. Comprising the steps that a flexible antenna is arranged on the side face, facing a solar panel, of a satellite body; when the solar panel is in a locked state, the flexible antenna is bent and arranged between the satellite body and the solar panel and is in a compressed state; when the solar panel is unfolded, the flexible antenna is unfolded along with the solar panel, the compressed state is converted into the natural state, the flexible antenna is smoothly unfolded, an additional locking and releasing mechanism is not needed in the whole process, the manufacturing and maintenance cost is reduced, and the risk of task failure caused by complex mechanism faults is reduced.

The invention relates to an angle-limited large elliptical orbit sailboard driving control method. The method comprises the following steps: establishing a geocentric orientation-sun orientation coordinate system according to a working mode and a mounting position of an angle-limited sailboard driving mechanism; selecting a sailboard tracking angle difference reference according to a sailboard simulation solar sensitive measurement value, a sailboard mechanism feedback angle and a solar azimuth; driving a sailboard to rapidly and roughly capture the sun according to the selected polarity of the tracking angle difference reference; different driving control laws are adopted to drive the sailboard to track the sun in a closed loop mode according to high and low orbit characteristics. The solar panel is guaranteed to face the sun to the maximum extent, and therefore sufficient energy is guaranteed.