1371 results about "Astrophysics" patented technology

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The present disclosure relates to a solar plant with solar panels (2, 2′) which are prefabricated in a leporello folding (3) which can be unfolded in an unfolding direction (A), and a substructure on which the solar panels are arranged, wherein the substructure is designed to be flexible in the unfolding direction (A) and is connected to the solar panels in a positionally fixed manner in such a way that the leporello folding (3) of the solar panels can be unfolded with the substructure (8, 9, 10, 24) and determines angular positions of the solar panels (2, 2′) relative to one another.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapid accompanying flying, fly-around flying and sweeping flying of an elliptical orbit coplanar target, and relates to a fuel-saving control on-satellite autonomous solving method for rapid approaching and high-precision accompanying flying / fly-around flying / sweeping flying from a long-distance position on the same orbit surface. Comprising the following steps: providing a three-time planning and five-time maneuvering long-distance approaching accompanying flight / fly-around / grazing flight configuration forming method; the efficient iteration method is suitable for solving the initial value of a minimum fuel control analytical solution and the final value of an iterative precise solution of a high-precision orbit extrapolation model through a satellite autonomous solution two-body model. The method is suitable for a near-earth space random elliptical orbit satellite to quickly approach a long-distance target on an approximate same orbit to carry out accompanying flight / fly-around / grazing flight, and the control precision can reach a ten-meter level; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The invention belongs to the technical field of speed reducers, and particularly relates to a multi-point meshing compound planetary gear speed reducer, a driving wheel set and a robot driving module. Wherein the multi-point meshing compound planetary gear reducer is provided with a sun gear serving as power input; the duplex planet gears are uniformly distributed around the sun gear, the duplex planet gears are composed of first planet gears and second planet gears, and the sun gear is meshed with the first planet gears; an intermediate planet gear exists between any two adjacent second planet gears in the circumferential direction, and each intermediate planet gear is meshed with the two adjacent second planet gears at the same time; an inner gear ring is arranged on a machine shell of the speed reducer, and the inner gear ring is meshed with the intermediate planet wheel. An existing planetary reducer has defects in the torque transmission capacity and the bearing capacity, a combined structure of a duplex planet wheel and an intermediate planet wheel is adopted, the intermediate planet wheel forms a multi-point meshing relation in the transmission process, loads are evenly dispersed, and the torque transmission capacity and the bearing capacity of the reducer are improved.

A heat balance optimal satellite sun orientation control method comprises the following steps: firstly, determining sun orientation target parameters according to a sun vector direction calculated by an orbit, using preset sun orientation target parameters when regular preset parameters are set, and then calculating a sun orientation target attitude angle and angular velocity by using a formula provided by the method, and finally, the sun orientation of the satellite to the target attitude angle and angular velocity is realized according to the measurement of the terahertz gyroscope and the plane control of the jet phase.

The invention discloses a super-large plane array type space computing power center, and belongs to the field of space technologies and space infrastructures. The space computing power center is deployed in a morning and evening orbit and comprises a plurality of computing power satellites which are physically docked in orbit through a docking mechanism. Each computing power satellite comprises a satellite platform, and a computing power load, a condensation type energy system and a space pump drive fluid loop system which are arranged on the satellite platform, and a docking mechanism for inter-satellite in-orbit physical docking is also assembled on the satellite platform. A plurality of computing power satellites form a planar array type rigid whole body through the docking mechanism, and energy scheduling among the light condensation type energy systems is realized based on the docking mechanism. According to the space computing power center, through multi-satellite physical docking and energy collaborative scheduling, the energy utilization efficiency and the power supply reliability of the system are improved.

The invention provides a constellation phase keeping method and device and electronic equipment, and the method comprises the steps: obtaining the mean root semi-major axis of each satellite in a target constellation in a target epoch, and selecting a satellite with the maximum mean root semi-major axis as a reference satellite; and constellation phase keeping is carried out based on the reference satellite. According to the scheme, only a relatively small period difference needs to be adjusted, the overall orbit adjustment amplitude and frequency can be reduced, and then fuel consumption and the complexity of orbit control operation are reduced; the influence of atmospheric resistance on the reference satellite is relatively small, so that the condition that other satellites frequently perform orbit control due to frequent change of the reference satellite orbit can be reduced; constellation configuration confusion caused by instability of the reference satellite itself is reduced in the face of various perturbation force and environment changes; the operation period of the reference satellite is long, the orbit is relatively stable, and too complex orbit change conditions do not need to be considered when an orbit control strategy is formulated, so that an orbit control algorithm and operation are easier to implement.

The invention provides a solar synchronous orbit quick response SAR satellite load platform integrated configuration, which comprises a satellite body, an SAR antenna, a solar array, a phased array data transmission antenna, a star sensor and a body-mounted solar cell, the satellite body is in a quadrangular prism shape, the cross section of the satellite body is an isosceles trapezoid, the satellite body comprises a group of opposite parallel surfaces and a group of opposite inclined surfaces, the phased array data transmission antenna and the star sensor are respectively arranged on two parallel surfaces, and the SAR antenna and the solar array are respectively arranged on two inclined surfaces; the body-mounted solar cell piece and the solar array are installed on the same inclined plane and located on the inner side of the solar array, and the solar array comprises two solar array plates and has an unfolded state and a folded state. According to the method, load platform integration is taken as a design thought, so that the AIT process of the SAR satellite and the time required by AIT are greatly shortened; meanwhile, the attitude requirements of energy and load work are met, and the reliability of the initial injection stage is improved through sufficient energy.

The invention discloses an earth surface downlink short-wave radiation space downscaling method based on a super-resolution reconstruction technology, belongs to the technical field of meteorological data analysis, and solves the problems of insufficient solar radiation data space resolution and weak earth surface heterogeneity characterization capability in the prior art. Carrying out super-resolution reconstruction on the radiation data set by adopting a super-resolution combination model, and carrying out weighted fusion on a reconstructed data set output by the super-resolution combination model; according to the method, the output results of the multiple models are subjected to weighted fusion, weight optimization of the super-resolution sub-models is performed in combination with the high-resolution terrain factors, spatial downscaling from 5km to 1km resolution is finally realized, the spatial precision of the solar radiation data is remarkably improved on the basis of keeping the hour-level time resolution of the original data, and the accuracy of the solar radiation data is improved. And more earth surface heterogeneity information can be captured, and meanwhile, the high-time-resolution application requirement is met.

The invention relates to an epicyclic gear with a housing, a sun gear, a planet carrier, at least one planet gear and a ring gear. The planet carrier is rotatably mounted in the housing. The housing defines an assembly direction for mounting the planet carrier. A bearing seat of the planet carrier, which serves for the rotatable mounting of the planet carrier in the housing and which is arranged adjacent to an insertion end of the planet carrier and in the use position of the planet carrier in the assembly direction behind the ring gear, has a diameter which is smaller than an inner diameter, in particular than a tip diameter, of the ring gear.

The invention discloses a simulation lunar dust excitation speed measurement system and method under a vacuum plume action condition, and relates to the technical field of aerospace engineering, and the system comprises a vacuum chamber and thruster system, an image collection assembly, a synchronous control collection system, a pulse laser light path, a high-speed spotlight, and a simulation lunar surface. In the invention, two optical systems of particle image velocity measurement and particle tracking velocity measurement are integrated in the same experiment, and the 532nm optical filter and the light trapping sheet are configured to realize compatibility of optical requirements, so that the problems that lunar dust particles with different scales and velocities cannot be measured synchronously and the optical systems interfere with each other in the prior art are solved; according to the method, velocity field data of micron-sized high-speed small particles and millimeter-sized low-speed large particles can be synchronously obtained and directly correspond to each other, the measurement integrity and reliability are improved in combination with simulation environment and synchronous control, and accurate data support is provided for spacecraft lunar dust prevention design.

The invention discloses a planetary gear train type speed reduction differential transmission mechanism and a transmission method thereof, and belongs to the technical field of speed reducers. A stator of the motor is connected with the shell, a rotor of the motor is connected with a first sun gear and arranged outside the first output shaft in a sleeving mode, the first sun gear is meshed with a first planet gear, the first planet gear is meshed with a first gear ring, the first planet gear is installed on a first planet carrier, and the first planet carrier is connected with the first output shaft. The second sun gear is meshed with the second planet gear, the second planet gear is meshed with the third planet gear, the third planet gear is meshed with the second gear ring, and the second planet gear and the third planet gear are both installed on the second planet carrier. The first gear ring is connected with the second planet carrier, and the second sun gear is connected with the second output shaft. Or the gear ring I is connected with the sun gear II, and the planet carrier II is connected with the second output shaft; the first output shaft, the second output shaft, the first planetary gear train, the second planetary gear train and the motor are coaxially arranged. The differential mechanism is compact in structure, large in transmission ratio and easy to manufacture, the differential function is achieved through the planetary gear train, and the technical bottleneck of a traditional transmission mechanism is broken through.

The invention discloses a planet generation method and system based on spherical display equipment. The method comprises the following steps: acquiring a planet material to be projected; based on preset UV coordinate mapping data, performing spherical projection on the planet material to generate a planet with two poles based on vertical projection and an equator based on cylindrical projection; the acquisition mode of the UV coordinate mapping data specifically comprises the following steps: generating a corresponding cylindrical projection through a cylindrical projection algorithm based on the auxiliary texture; corresponding spherical coordinates are obtained based on auxiliary texture mapping; converting the obtained spherical coordinates into 3D rectangular coordinates; the obtained 3D rectangular coordinates are mapped back to the 2D texture space again, and corresponding vertical projections are generated; determining an equator mask; and generating corresponding UV coordinate mapping data based on the equatorial mask, the cylindrical projection and the vertical projection. According to the invention, through the design of the UV coordinate mapping data, in the process of carrying out spherical projection on the planet material, the resource consumption is reduced, and the dynamic projection precision can be improved at the same time.

A gear speed change device. The gear speed change device comprises a first planetary gear train (100). The first planetary gear train (100) comprises a first ring gear (101), a first planetary carrier (102), first planetary gears (103), a solar idle gear (104), and a planetary idle gear (105). The planetary idle gear (105) and the first planetary gears (103) are all mounted on the first planetary carrier (102); each first planetary gear (103) comprises a pinion (103a) and a large gear (103b) coaxially connected to the pinion (103a); the planetary idle gear (105) and the pinion (103a) are both engaged with the inside of the first ring gear (101) and are both engaged with the outside of the solar idle gear (104); the pinion (103a) can float along the radial direction thereof relative to the first planetary carrier (102), and the maximum radial floating amount of the pinion (103a) is greater than the maximum radial floating amount of the planetary idle gear (105); and the large gears (103b) of at least two first planetary gears (103) are arranged in a staggered manner in the axial direction, and projections thereof on a plane perpendicular to the axial direction overlap with each other. The gear speed change device has a relatively high torque load density, and can give consideration to a small volume, a large transmission speed ratio and a high torque load capability.

The invention relates to a satellite-borne microwave radiometer cold air calibration attitude planning method based on multi-objective optimization, and belongs to the field of satellite microwave remote sensing, and the method comprises the steps: building a satellite attitude kinematics model and an attitude dynamics model through a geocentric equator inertial coordinate system, an orbital coordinate system and a satellite body coordinate system; establishing a brightness temperature sum model received by the microwave radiometer in the cold air calibration process; according to the method, a microwave radiometer receiving brightness temperature sum, brightness temperature fluctuation, platform maneuvering capability, energy and a satellite sensor field of view are comprehensively considered, a multi-target optimization model of microwave radiometer cold air calibration consideration constraints is established, solving is performed, constraint conditions are established, and a solution or a solution set is obtained. According to the method, the brightness temperature sum and the brightness temperature fluctuation in the observation area of the microwave radiometer in the cold air calibration process are optimized by adopting multi-objective optimization, and the on-orbit cold air calibration effect and the system reliability of the satellite-borne microwave radiometer are improved.

The invention relates to a step-by-step winding and unfolding ultra-large flexible solar cell array and an unfolding method, and belongs to the technical field of space solar cell arrays. The battery array comprises a battery array body structure, a pressing and releasing mechanism, an unfolding and locking mechanism and an extension arm mechanism, the cell array body structure has a folded state and an unfolded state; under folding, pressing force is provided through the pressing and releasing mechanism, so that the device is folded on the side walls of the two sides of the satellite body and can bear the mechanical load of the launching section; during unfolding, under the action of the unfolding locking mechanism, the folding mechanism and the stretching arm mechanism, the folding mechanism is unfolded in a distributed manner according to a time sequence; and after being unfolded, the antenna can rotate by 360 degrees to provide energy for satellites. The invention has the advantages of small folding volume, large unfolding area and light weight, and solves the contradiction that the power demand of a spacecraft power supply is large and the envelope size of a launching section is small.

The invention provides an in-orbit star-observation calibration and correction method and system for geometric thermal stability of a space camera, and the method comprises the steps: calculating the positioning error of each star point in an along-orbit direction and a vertical-orbit direction through the star-observation imaging of a satellite under different initial temperature conditions, recording the satellite temperature and imaging time, and carrying out the calibration of the geometric thermal stability of the space camera. Fitting to obtain a temperature variation coefficient and a time variation coefficient of the exterior orientation element thermal stability of the space camera, and fitting to obtain a space camera interior orientation element zoom coefficient of each star observation task according to the positioning error of each star point after correction of the exterior orientation element thermal stability and the distribution condition of the star points in the camera field of view; the temperature change coefficient of the space camera interior orientation element thermal stability is obtained, and in-orbit calibration and correction of the space camera interior and exterior orientation element thermal stability are achieved. The method does not need a ground geometric field, has the advantages of high precision, high timeliness and low cost, and has good adaptability.

The utility model relates to the technical field of automobiles, and provides a wheel and an electric automobile, the wheel comprises a hub, a hub motor, a planetary reduction mechanism and a mandrel, the hub comprises a rim and a spoke, and the hub motor and the planetary reduction mechanism are both located on the same side of the spoke; the hub motor comprises a motor shell, a motor stator and a rotor assembly, the motor stator is fixed in the motor shell, and the radial outer end of the rotor assembly corresponds to the motor stator; the planetary reducing mechanism comprises a planet carrier and a sun gear, the radial inner end of the rotor assembly is fixedly connected with the sun gear, and the planet carrier is fixedly connected with the spoke; the mandrel is fixed on a vehicle body, the motor shell is fixed on the mandrel, and the planet carrier, the sun gear and the rotor assembly are all rotationally installed on the mandrel. The acting force borne by the hub in the wheel is transmitted to the central spindle through the planetary speed reducing mechanism and then transmitted to a vehicle body through the central spindle, and the problem that a shell of a hub motor is prone to deformation is solved.

A drone deployable modular system for remote solar energy generation utilizes remote-controlled multi-rotor drones configured to deliver and maintain modular heliostat units for assembly into heliostat fields in remote locations. The drone deployable modular heliostat unit may include a frame that supports at least one deployable, mirrored surface controlled by a heliostat driver. In one embodiment, the deployable, mirrored surface is a MYLAR film tensioned over a mirror frame portion supported and controlled by the heliostat driver to orient a working face of the mirrored surface in various positions. The drone deployable modular heliostat unit may include a deployable stand arrangement and / or an anchor system. The modular heliostat units may be configured to be deployed in heliostat fields in a generally peripheral arrangement to provide reflected solar energy to a centrally located concentrating solar collector / energy utilization system.

The present invention provides a riveting tooling for a planetary gear shaft, belonging to the field of planet carrier processing. The present invention includes a tooling body and a planet carrier placing base arranged above the tooling body. The planet carrier placing base can move vertically relative to the tooling body. The planet carrier placing base is further provided with a riveting hole corresponding to the position of the riveting position of the planetary gear shaft. A first punch is arranged on the tooling body below the riveting hole. The riveting tooling for the planetary gear shaft further includes a press for driving the planet carrier placing base to move vertically downward. A second punch corresponding to the first punch is arranged below the press. The first punch and the second punch cooperate to complete riveting. The beneficial effects of the present invention are: simple operation, simple structure and low cost.

A steering drive includes an electric motor with a motor shaft and a steering lever rotatable about a pivot axis. A first planetary stage has a first sun gear, first planet gears, a first ring gear, and a first planet carrier. A final planetary stage has a final sun gear, final planet gears, a final ring gear, and a final planet carrier. A force flow runs from the motor shaft to the first sun gear, and from the first sun gear via the first planet gears to the first planet carrier and to the first ring gear. The force flow runs via the final sun gear and via the final planet gears to the final ring gear, runs from the first ring gear and from the final ring gear to the steering lever, and runs via a first sun gear shaft from the electric motor to the first sun gear.

The invention provides a low-energy transfer window rapid design method and system for earth-moon DRO, and relates to the technical field of aerospace. Based on a double-circle restrictive four-body model, an earth-moon transfer orbit database is constructed, the earth-moon transfer orbit database comprises a plurality of transfer orbits from a target near-earth orbit to a target long-distance retrograde orbit; determining a target track point from the track distribution diagram corresponding to each track leveraging category according to the track information track distribution diagram of each transfer track and the input near-earth track information and epoch time of the target spacecraft, and correcting the track information of the transfer track corresponding to the target track point based on an ephemeris model, and obtaining a target transfer orbit of the target spacecraft. According to the method provided by the invention, rapid and accurate planning of the earth-moon DRO low-energy transfer window can be realized, the processing time is reduced, and the planning efficiency is improved.

The invention relates to a light-weight type planetary line gear reducer which comprises a sun gear, a planetary gear, an inner gear ring and a planetary carrier which are all light-weight type line gears, the sun gear is located in the center of the inner gear ring, the sun gear and the planetary carrier are both coaxially and rotationally installed on the inner gear ring, and the planetary gear is installed on the planetary carrier in a heteroaxial rotation mode and can rotate around the axis of the planetary gear. And the tooth surface normal of any contact point of the planet gear is collinear with the tooth surface normal of the corresponding contact point of the sun gear and the inner gear ring at the same time, the contact line is a cylindrical spiral line, and the transmission ratio and the spiral radius are decoupled. Through the space curve meshing theory, the reduction ratio exceeding 30 can be achieved through a single stage, the stage number of the planetary reducer is reduced, the force flow transmission path is remarkably shortened, the overall weight is reduced, and meanwhile the bearing capacity and reliability are improved; the technical problems that in the prior art, when the large reduction ratio is achieved, the multi-stage series connection structure is complex, the weight is increased, the transmission and positioning precision is reduced, and the load dispersing capacity is limited are systematically solved.

The invention provides a solar wing device which is characterized in that one end of a flexible solar panel (6) is wound on a winding drum (3), the other end of the flexible solar panel (6) is connected with a uniform load rod (5), and the uniform load rod (5) is always parallel to a rotating shaft of the winding drum (3); the scissor mechanism (4) is formed by cascading four equal-length rods which form parallelograms when unfolded, the flexible solar panel is unfolded and folded in the bottom end-top end direction, elastic components are arranged at the bending positions of the left end and the right end, opposite to each other, of each parallelogram, and the scissor mechanism is driven by the elastic components to be unfolded. The elastic component enables the scissor mechanism (4) to always keep a force for pushing away the uniform load rod (5), and in the unfolding state or the folding state of the solar blanket, the torque generated by the motor adapts to the elastic force generated by the elastic component. The flexible solar wing can be repeatedly unfolded and folded in orbit for many times.

The invention relates to the technical field of laser terminal pointing control, and discloses a laser terminal installation error on-orbit calibration method which comprises the following steps: S1, defining multiple coordinate systems including an ideal installation coordinate system of a laser terminal during ground calibration and an actual installation coordinate system caused by stress release or deformation after emission, and a real-time pointing coordinate system of in-orbit operation of the laser terminal. And S2, establishing a laser terminal installation error model, wherein the model defines an installation error angle through a direction relation of a coordinate system before and after terminal emission. According to the method, the installation error angle is dynamically estimated through the recursive least square method, high-precision dynamic correction of the laser terminal pointing error is achieved in combination with the rotation matrix pointing correction model, meanwhile, error feedback and time sequence control are introduced, and the problems of nonlinear error accumulation, insufficient fixed gain response and long-term error accumulation in the prior art are solved.

The invention provides a helicopter planetary gear bearing ring lower lubrication test bed. The helicopter planetary gear bearing ring lower lubrication test bed comprises a base provided with a mounting plate, a power transmission mechanism, an oil guide mechanism, a flight attitude simulation platform mounted on the lower surface of the base through a connecting assembly, a driving mechanism arranged on the base and a test main body, the test main body comprises a planetary gear train gearbox mounted on the base and a planetary gear train mounted in an inner cavity of the planetary gear train gearbox, and the planetary gear train comprises a sun gear, an inner gear ring, a planet carrier and three planetary gear mechanisms arranged on the planet carrier; the planetary gear mechanism comprises a hollow planetary shaft installed on the planetary carrier and a planetary gear embedded with a test bearing and an auxiliary bearing, the planetary gear is sleeved on the hollow planetary shaft, the planetary gear is meshed with the sun gear, and the planetary gear is meshed with the inner gear ring. According to the invention, the test bearing can rotate and revolve at the same time while simulating the real operation condition of the helicopter, so that the lubrication effect of the helicopter planetary gear bearing can be effectively evaluated.

The invention discloses an unmanned aerial vehicle low-altitude target radar detection and recognition system, and belongs to the technical field of radar detection and target recognition. Comprising a quantum entanglement laser emission module, a quantum multi-dimensional receiving module, a multi-source heterogeneous data fusion module, a quantum-classical hybrid calculation module, a self-adaptive anti-interference module and an environment self-adaptive correction module, the quantum entanglement laser emission module is used for generating entanglement photon pairs carrying orbital angular momentum (OAM) and emitting the entanglement photon pairs to a detection airspace; and the quantum multi-dimensional receiving module demodulates the distance, the angle, the OAM attenuation spectrum, the polarization state and the reflection spectrum of the echo signal through coincidence measurement. The quantum entanglement-OAM combined modulation technology is adopted, and the non-locality of entanglement photon pairs and the multi-dimensional information bearing capacity of a 9-path orthogonal OAM mode are improved by more than three times compared with a traditional microwave radar; single-photon-level coincidence measurement is combined with quantum Fourier transform signal processing, the signal-to-noise ratio is increased by 3dB, and stable detection of a weak target is achieved.

The invention belongs to the technical field of gear transmission, and relates to a planetary gear train structure, in particular to a herringbone tooth planetary transmission speed reduction structure. Comprising a sun gear, nP planetary gears, a first inner gear ring, a second inner gear ring, a planetary gear carrier, a planetary gear mounting disc, a planetary gear pin shaft, a planetary gear bearing, a locking nut, a connecting bolt, a spline housing, a pressing plate, a spacer sleeve, a positioning bush and a screw. Through the spline floating design of the sun gear and the inner gear ring, the uniform load adjusting effect when load distribution is not uniform due to error accumulation and the like in the working process of the herringbone tooth planetary mechanism can be achieved. The inner gear ring is designed into a split design and combination form of the first inner gear ring, the second inner gear ring, the planetary gear mounting disc, the planetary gear pin shaft and the planetary gear carrier, and the problem of assembly interference of herringbone tooth planetary transmission can be solved.

The invention belongs to the technical field of aerospace asteroid defense, and relates to a variable mass gravitational traction deflection asteroid orbit full-flow design method based on real working condition constraint, which comprises the following steps of: obtaining orbit elements corresponding to a target asteroid according to initial conditions; substituting into a high-precision numerical integration model, and calculating perigee moments and perigee distances under the condition that the gravitational traction action is not applied; according to the early warning time requirement, the carrying capacity and the real engineering constraint, a genetic algorithm and sequential quadratic programming are used, the maximum perigee deflection distance serves as a target, and a transmitting window is optimized and solved; and substituting the optimized emission window into the high-precision deflection efficiency numerical integration model, and outputting the asteroid perigee deflection distance. According to the method, comprehensive analysis is carried out in combination with early warning time, emission capability and actual working conditions from the perspective of a whole-process task, so that the authenticity and engineering applicability of a calculation result are ensured.

The invention provides a dual-motor coupling driving robot joint module based on a planetary reducing mechanism, which comprises a first motor, a second motor, a first planetary reducing mechanism, a second planetary reducing mechanism and a gear ring, the first motor comprises a first casing and a first output shaft, and the second motor comprises a second casing and a second output shaft. The first planetary reducing mechanism comprises a first sun gear and a first planet gear, the first sun gear is connected with the first output shaft, the first planet gear is rotationally arranged on the first machine shell, and the first planet gear is meshed with the first sun gear; the second planetary reducing mechanism comprises a second sun gear and a second planet gear, the second sun gear is connected with the second output shaft, the second planet gear is rotationally arranged on the second machine shell, and the second planet gear is meshed with the second sun gear; and the gear rings are annularly arranged outside the first planet wheel and the second planet wheel correspondingly, the gear rings are engaged with the first planet wheel and the second planet wheel correspondingly, the gear rings are used for power output, and the compact design of the joint module is achieved.