84 results about "Solar sail" patented technology

The invention discloses a parameter self-adaptive control method for a low-orbit satellite with a solar panel driving device. Belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The satellite attitude control method solves the technical problems of overcoming the influence of the rotational inertia of the satellite and compensating the influence of the SADA vibration disturbance torque and the atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in a quaternion form, and constructing a satellite attitude kinematics and dynamics equation for designing an attitude controller on the basis of the satellite attitude kinematics and dynamics equation; defining an error attitude kinematics and dynamics equation of the satellite through the attitude control error and the angle control error; carrying out segmentation conversion on a to-be-estimated parameter and a known state variable, and expressing the negative influence of a rotational inertia error on the attitude by matrix operation; designing the control torque of the attitude controller by adopting a mode of combining nonlinear feedback control with self-adaptive control; and updating the self-adaptive estimated value in the control parameters of the attitude controller by adopting a parameter self-adaptive law.

The invention provides an in-orbit energy management method and system for a damaged satellite, network equipment, a computer readable storage medium and a computer program product. The method comprises the following steps: autonomously determining whether to execute the reasoning of the SADA rotation angle based on the current state data of a satellite, and selecting a reasoning model under the current energy limitation condition; simulating a regression relationship between the SADA rotation angle and the predicted current of the solar panel according to the selected reasoning model and the current state data of the satellite; and searching the corresponding SADA target rotation angle capable of enabling the predicted current of the solar panel to be maximum from the regression relationship based on a Bayesian optimization strategy, and converting the SADA target rotation angle into an SADA control instruction. According to the method, a damaged satellite energy management framework is constructed, real-time and accurate prediction of the optimal sun-pointing angle of the current peak value of the solar panel in the current damaged state is realized, energy fluctuation caused by satellite attitude and sun vector change is avoided, and an important basis is provided for autonomous maintenance of energy supply on a satellite.

An integrated spaceborne SAR satellite platform, belonging to the field of spacecraft technology, solves the problems of poor temperature uniformity of existing satellite platform mounting plates, large SAR antenna envelope, inability to meet high-precision single-unit installation requirements, and low structural strength. The platform comprises: a SAR antenna, a feed assembly, a single-unit mounting plate, an external single-unit device, an internal single-unit device, a platform main frame, and a solar sail; the platform main frame, single-unit mounting plate, and solar sail constitute the cabin; the solar sail is located below the platform main frame; the SAR antenna, feed assembly, and external single-unit device are mounted on the outer surface of the single-unit mounting plate; the internal single-unit device is mounted on the inner surface of the single-unit mounting plate; the single-unit mounting plate includes a honeycomb panel, metal gaskets, and pre-embedded heat pipes; the honeycomb panel includes a honeycomb core and a honeycomb panel skin; the metal gaskets are connected to the pre-embedded heat pipes. This invention is applicable to satellite platform spacecraft platforms, spacecraft configuration and layout design, spacecraft structural thermal control design, spacecraft on-orbit monitoring, and SAR satellite on-orbit active compensation.

The invention provides a thermal control device for a high-power wireless energy transmission driving mechanism. The thermal control device comprises a wireless driving mechanism, a heat dissipation plate, a heat pipe, a heat expansion plate, a heat conduction filler and a thermal control coating, the wireless driving mechanism comprises two high-heat-flux heat sources, namely a static coil and a rotating coil; the rotating coil is mounted on the insulating mounting plate, the insulating mounting plate is mounted on the magnetic conductive material, the magnetic conductive material is mounted on the rotating end cover, and the heat pipe is arranged on the rotating end cover to transmit heat to the heat dissipation plate on the reverse side of the solar panel; the static coil is mounted on the static end cover through an insulating mounting plate and a magnetic conductive material, and a heat pipe is arranged on the static end cover to transmit heat to the heat dissipation plate on the nightside; heat is mainly radiated to a cold space after reaching a radiating surface from the coil through the insulating mounting plate, the magnetic conductive material, the end cover and the heat pipe. The invention aims to solve the design problem of a high-power wireless energy driving mechanism of a satellite or a space station, and provides a design basis for solving the temperature control of a wireless energy transmission driving mechanism.

The invention relates to the technical field of aerospace engineering and remote sensing, in particular to an extremely-low-orbit median-inclination-angle remote sensing satellite and a design method.The satellite runs on a median-inclination-angle orbit, and a spherical crown type windward side and smooth side plate transition structure is adopted to reduce aerodynamic resistance and homogenize heat flow; the detector and the refrigerator are connected with the main heat dissipation surface through a heat conduction structure to form a cold end heat dissipation channel; when a satellite is in a launching state, the solar panel is folded on the Zs side of a satellite body, the satellite is unfolded in the + Ys direction after entering an orbit, an included angle is formed between the normal of the solar panel and the-Zs direction, the SAR antenna is folded on the Ys side of the satellite body, the satellite is unfolded in the-Ys direction after entering the orbit, and the normal of the array plane of the SAR antenna points to the + Zs direction, so that a sunshade and counterweight structure is formed; the positive and negative Xs surfaces are provided with double Hall electric thrusters, the orbit solar angle is calculated through the GNSS and the sun vector, yaw switching is executed, and forward and reverse flight switching is achieved.

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 utility model discloses a solar panel unfolding mechanism which comprises a storage frame, first sliding groove plates are fixedly installed on the two sides of the storage frame, three moving plates are arranged on one sides of the first sliding groove plates from top to bottom, one side of the top moving plate is fixedly connected to one side of the first sliding groove plate, and the other side of the top moving plate is fixedly connected to the other side of the first sliding groove plate. By starting the motor, the motor drives the rotating rod to rotate, the rotating rod rotates to drive the rotating disc to rotate, the rotating disc rotates to drive the two traction rods to move towards the two sides respectively, the traction rods move to drive the bottom moving plate to move, and the moving plate moves to drive the solar panel to move. And meanwhile, a second strip-shaped frame moves to drive a pull rod to move, the pull rod moves to drive a supporting rod to move, the supporting rod moves to drive a pressing plate to move, and after the pressing plate moves, the effect of limiting a cover plate can be achieved.

The present application provides a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. The method comprises: monitoring the deployment status of the satellite's solar panels in real time based on a preset sensor, obtaining deployment status data of the solar panels, and determining the deployment status of the solar panels based on the deployment status data; in response to the abnormality in deployment of the solar panels, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panels, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. The present application achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment by obtaining the first inertia matrix of the satellite body and the second inertia matrix of the solar panels when an abnormality in deployment of the solar panels is detected in real time, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. This improves the stability of the satellite's attitude control.

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.

The invention relates to the field of spaceflight components, in particular to a high-rigidity high-torque two-degree-of-freedom sailboard driving mechanism which comprises a main support, 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 through the main support, one end of the double-axis connecting support is fixedly connected with the output end of the A-axis driving assembly, and the other end of the double-axis connecting support is fixedly connected with the output end of the B-axis driving assembly; the other end of the double-shaft connecting support is fixedly connected with the B-shaft driving assembly, one end of the sailboard connecting support is fixedly connected with the output end of the B-shaft driving assembly, and the other end of the sailboard connecting support is connected with the sailboard. The A-axis driving assembly and the B-axis driving assembly are each internally provided with a motor and a harmonic gear reducer so that the torsional rigidity of the output end can be larger than or equal to 12000 Nm / rad, and the output torque can be larger than 44 Nm. And the solar panel is driven to point to the sun and the power / signal transmission function is realized.

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 provides a zigzag mark hinge based on a shape memory material and a solar panel, and belongs to the technical field of space unfolding, the zigzag mark hinge comprises a framework unit, a driving unit and a rigid connecting part, and the framework unit comprises a fiber reinforced composite material panel; the butt-joint ends of the two fiber reinforced composite material panels are attached to each other and connected through polyimide films, the polyimide films are attached to the upper surfaces and the lower surfaces of the butt-joint ends, and preset zigzag marks are formed in the butt-joint positions; the shape memory polymer composite layers are attached to the surfaces of the framework units on the two sides of the zigzag marks, and the heating films are attached to the surfaces of the shape memory polymer composite layers so that the shape memory polymer composite layers can return to the straight state from the bent state and drive the framework units to be flattened; the zigzag mark design generates a self-adaptive nonlinear damping effect in the unfolding process, the damping effect is gradually enhanced along with the increase of the unfolding angle, the impact and vibration in the unfolding process are restrained, and the stability and controllability of the unfolding process are ensured.

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 application discloses a sail spacecraft underactuated three-axis control method based on a planar control moment mechanism and belongs to the technical field of spacecraft attitude control and maintenance. The method is based on an attitude dynamics equation, a kinematics equation and a space geometric relation to establish a sail surface normal control law, a sail surface roll axis angular velocity control law and a sail surface roll angle control law, and then orderly combines the three control laws, so that the underactuated three-axis control is realized only by using the control moment on the sail surface. The application solves the problem that sail spacecraft represented by a solar sail and an atmospheric sail is difficult to autonomously generate a control moment in the roll axis direction to realize three-axis attitude control. The application can reduce the overall mass of a platform of the sail spacecraft due to the use of an additional roll axis moment generating mechanism, reduce the failure risk caused by the use of an additional control mechanism, is favorable for deep space exploration of the solar sail, and prolongs the service life of the sail spacecraft.

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 provides a solar panel tilt angle determination method, device and equipment based on solar panel illumination quantity. The method comprises the following steps: acquiring an orbital angle sequence and a solar panel tilt angle range corresponding to a solar panel of a satellite; for any target orbital angle in the orbital angle sequence, an initial sailboard swing angle value under the target orbital angle is determined, a target function is called in an optimizer, the maximum solar sailboard illumination quantity output by the target function for the sailboard swing angle value is taken as a target, the sailboard swing angle range is taken as a constraint, the sailboard swing angle value is subjected to iterative adjustment, and the optimal sailboard swing angle value is obtained. The optimal sailboard swing angle value under the target orbit angle is obtained; wherein the illumination quantity of the solar panel is the average value of the effective illumination quantity of the sun acting on the solar panel of the satellite in an orbital period of the satellite. The method has the characteristics of high precision, small calculation amount, high calculation speed and the like.

The application relates to the field of aerospace technology and discloses a thermal knife power supply method for satellite solar panel deployment based on a super capacitor, which comprises a super capacitor module for storing electric energy; a thermal knife assembly made of high-resistance material, which generates heat through instant power-on and is used for cutting the fixing device of a solar panel; and a control circuit comprising a charging module, a discharge control module and a temperature monitoring module; the super capacitor module comprises an MXene and COF composite electrode, and the composite electrode is prepared through the following steps: step one, etching TiVAlC powder to obtain TiVCTx MXene; and step two, mixing and dissolving tetraaminophenylene diamine and trialkyltrianiline and heating to react to obtain COF. In order to realize effective combination of COF particles and MXene, a hydrothermal assisted synthesis method is adopted in the research. The method utilizes chemical reactions under hydrothermal conditions to enable COF particles to grow in situ on the surface of MXene, thereby forming a novel MXene / COF (MCF) composite material.

The application provides a sun tracking control method and device of a solar panel, electronic equipment and a computer readable medium, wherein the method comprises: acquiring a thrust direction of a satellite; determining a yaw angle of the satellite based on the thrust direction of the satellite; determining a working mode of the satellite based on the yaw angle; acquiring a sun elevation angle and / or a sun azimuth angle of the satellite; in response to the working mode of the satellite, determining a roll angle, a pitch angle and a rotation angle of a solar panel driving mechanism of the satellite based on the sun elevation angle and / or the sun azimuth angle; and driving the solar panel to continuously rotate and realize sun tracking based on the roll angle, the pitch angle and the rotation angle. The application realizes attitude control and continuous and accurate sun tracking of the satellite during on-orbit operation, and has the advantages of simple attitude control algorithm and sun tracking algorithm.

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 invention relates to a solar panel rotation optimization method based on a measurement and control link availability metric, comprising: S100, obtaining a measurement and control tracking window set based on a measurement and control link budget; S200, pre-planning a solar panel rotation strategy based on the measurement and control tracking window set to obtain a solar panel shielding window set for the measurement and control link; S300, monitoring the measurement and control link signal quality in real time based on the solar panel shielding window set and in combination with a solar panel shielding signal historical trend atlas to obtain a solar panel signal trend atlas; S400, dynamically adjusting the solar panel rotation strategy if the solar panel signal trend is consistent or identical with a trend in the solar panel shielding signal historical trend atlas; S500, entering the next measurement and control tracking window after the current tracking arc ends based on the measurement and control tracking window set, and repeating steps S200 to S500. This method can significantly improve the long-term operational control efficiency of a space mission during its on-orbit period, achieving a balance between complexity and accuracy.

The invention relates to a method for processing a spacecraft solar sail and a solar sail, comprising the following steps: spin-coating a bottom layer on a silicon wafer, and then spin-coating a flexible base film on the bottom layer; forming a first deposition area on the flexible base film using a photolithography process, and depositing a thin film electrode in the first deposition area; forming a second deposition area on the overall structure after the thin film electrode deposition using a photolithography process, and then depositing an alloy beam in the second deposition area, wherein the alloy beam is deposited between two adjacent thin film electrodes; forming a third deposition area on the alloy beam using a photolithography process, and depositing a reflective beam in the third deposition area, removing photoresist and a sacrificial layer remaining between the alloy beam and the flexible base film and the thin film electrodes due to the photolithography process, and peeling the flexible base film from the bottom layer so that the flexible base film generates residual stress. The solar sail processed by the processing method of the invention has high bending rigidity and eliminates the need for a supporting mechanism.

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.