1788 results about "Astrophysics" patented technology

The invention provides a self-adaptive anti-backlash planetary gear reducer and a working method thereof. The self-adaptive anti-backlash planetary gear reducer comprises an input shaft and a shell. An electromagnet coil, an electromagnet, a sliding block seat, a supporting seat, a sun gear and a second inner gear ring sequentially penetrate through one end of the input shaft from one end to the other end; the side, away from the input shaft, of the second inner gear ring is connected with an output shaft. A planet carrier is slidably arranged on the supporting seat in the radial direction, a first planet gear and a second planet gear are coaxially arranged on the planet carrier, a first inner gear ring is arranged on the inner side of the shell, and the sun gear, the first planet gear and the second planet gear are engaged with the first planet gear, the first inner gear ring and the second inner gear ring correspondingly; the electromagnet is fixed to the sliding block base, the sliding block base, the supporting base and the second inner gear ring are all rotationally connected with the input shaft, and the sliding block base can slide on the input shaft in the axial direction. And the sliding block seat is hinged with the planet carrier through a connecting rod. The speed reducer is simple in structure, has the self-adaptive anti-backlash capability, and can realize self-adaptive real-time elimination of the transmission backlash.

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 invention discloses an LEO on-orbit real-time precision orbit and clock error determination method and system based on PPP-B2b enhancement, and belongs to the technical field of low-orbit satellite high-precision autonomous orbit determination and clock error estimation. According to the method, enhanced information broadcasted by Beidou No.3 B2b signals is used for correcting a BDS-3 / GPS satellite orbit and clock correction, under the support of the regional BDS-3 / GPS enhanced information, when a new BDS-3 / GPS tracking satellite exists, before measurement updating, the inter-epoch clock correction variable quantity obtained through calculation by using a carrier phase time difference method is used as two orbit determination filtering clock correction parameters for constraint, so that the measurement accuracy is improved. In-orbit real-time separation of receiver clock error and non-difference floating point ambiguity parameters is carried out; and finally, substituting an LEO precision orbit output by orbit determination filtering and a non-difference floating point ambiguity parameter into a phase observation equation to calculate the precision clock error of the BDS-3 / GPS receiver. The method is suitable for the situation that only regional GNSS enhancement information support exists, and the influence of orbital maneuver is small.

The utility model discloses an actuator with a two-stage planetary reducing mechanism and a robot joint. The actuator with the two-stage planetary reducing mechanism comprises a shell and a motor, the shell comprises a side wall, a top wall and a bottom wall, an inner gear ring is installed in the shell, and an installation space is defined by the inner edge face of the inner gear ring. A first sun gear, a first planet gear, a first planet carrier, a second sun gear, a second planet gear and a second planet carrier are arranged in the installation space, the first planet gear is connected to the first planet carrier in a pivoted mode and meshed between the inner gear ring and the first sun gear, and the second sun gear is synchronously connected with the first planet carrier. The second planet gear is pivoted on the second planet carrier and is meshed between the inner gear ring and the second sun gear; the lower end of the first sun gear is connected with the output end of the motor, and the upper end of the second planet carrier is pivoted to the top wall of the shell. According to the actuator, the space in the shell can be reasonably utilized, so that the interior of the actuator has a relatively compact structure, and particularly, the axial size of the actuator is greatly reduced.

A coaxial electrified drive axle is provided for a vehicle. The coaxial electrified drive axle has an electric motor connected to a gear box housing containing a stepped planetary gear set operatively coupling a drive shaft of the electric motor to a linkshaft and an output shaft through a differential assembly. The differential assembly is contained within a carrier which also supports the planet gears through planet pins. The carrier is supported by a single carrier support bearing. The drive shaft, the linkshaft, and the output shaft are coaxial and share a common axis of rotation.

A drive assembly for a motor vehicle drivetrain includes a differential configured for driving a first output shaft and a second output shaft about a center axis and for allowing the first output shaft and the second output shaft to rotate about the center axis at different speeds, and planet gears each rotatably mounted on a respective planet carrier. The planet gears each include a radially extending surface and blades circumferentially spaced part from each other. The blades axially protrude from the radially extending surface for directing a fluid flow axially toward the differential during operation of the drive assembly.

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.

A solar powered plane configured to operate in the stratosphere includes a hinged vertical stabiliser that that pivots about a hinge; the stabiliser extends both above the fuselage and also below the fuselage, and the base of the vertical stabiliser includes a skid. The hinged vertical stabiliser is configured such that when the skid contacts the ground on landing, the top of the vertical stabiliser pivots forwards around the hinge point, minimising damage to the aircraft on landing. the skid includes a lightweight, replaceable sacrificial layer.

The invention provides a photovoltaic system based on a sun-oriented trajectory, a rotation method and a storage medium. The photovoltaic system comprises a photovoltaic power generation panel, a motor, a data processing module, a motor control module, a sun-oriented tracking module and a photoelectric auxiliary tracking module. The motor is used for driving the photovoltaic power generation panel to rotate. The sun tracking module shoots a series of images of the sun and calculates the solar altitude and direction angle. The photoelectric auxiliary tracking module calculates the position of the sun through the angle of the sensor; and judging whether the illumination intensity reaches a threshold value or not, if so, selecting a sun tracking module for tracking, and if not, selecting photoelectric auxiliary tracking. The initial approximate position of the sun is tracked through an apparent sun view motion trail system, and the used tracking method is judged according to weather conditions and illumination intensity. And the sun can be stably and reliably tracked under most weather conditions.

The invention discloses an electric propulsion GEO satellite orbit transfer and point fixing method, and belongs to the technical field of satellite orbit dynamics, and the method comprises the following steps: S1, building a low-thrust satellite orbit dynamics model based on multi-perturbation and earth shadow constraints; s2, according to the low-thrust satellite orbit dynamic model, a Lyyunov function is constructed, and an RQ-Law feedback control model is constructed by using the Lyyunov function; s3, weight parameters of the RQ-Law feedback control model are optimized; and S4, using the optimized RQ-Law feedback control model to complete orbit transfer and point fixing. Aiming at the problem of long time of current satellite orbit transfer, the invention provides an innovative method, the traditional orbit transfer and fixed point staged optimization are integrated into one stage, and the integration of orbit transfer and fixed point is realized.

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.

The embodiment of the invention provides an ionization layer electron density inversion method, device, system and equipment and a medium. The method specifically comprises the following steps: determining a target area of electron density inversion; obtaining observation data corresponding to the signal link in the target area; the signal links comprise a first signal link from a GNSS satellite to a ground station, a second signal link from the GNSS satellite to a low earth orbit satellite, and a third signal link from the low earth orbit satellite to the ground station; determining the slant total electron content corresponding to the signal link according to the observation data corresponding to the signal link; solving an ionized layer chromatography function according to the oblique total electron content corresponding to the signal link so as to obtain an electron density value of a grid in the target area; the ionized layer chromatography function is used for representing a mapping relation between the oblique total electron content and the electron density value of the grid. According to the embodiment of the invention, the inversion precision of the ionosphere electron density can be 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.

A planetary wheel end includes a planetary gear set, a rotatable axle shaft, a rotationally fixed axle tube surrounding at least a portion of the axle shaft, and a planetary gear carrier arranged for fixing to a wheel. The planetary gear set includes a radially inner sun gear, a radially outer ring gear and a plurality of planet gears drivingly engaged with the sun gear and the ring gear. The axle shaft is rotationally fixed to the sun gear and the axle tube includes a spindle rotationally fixed to the ring gear. The planetary gear carrier includes a plurality of planetary shafts and each one of the planet gears is rotatable about a respective one of the planetary shafts. In an example embodiment, each one of the plurality of planet gears has a respective radial needle bearing disposed on the respective one of the plurality of planetary shafts.

A planet carrier arrangement for a planetary gear system comprises a planet carrier (63) having a first support element (631) and a second support element (632) axially spaced from the first support element (631), a number of planet gears (62), and a coupling recess (634) in which a number of circumferentially spaced internal gear sectors (635) are provided. A number of support sections (636) are provided circumferentially between the internal gear sectors (635), the support sections (636) being located at least partially radially within an effective diameter of the internal gear sectors (635).Furthermore, an output element (71) is provided, to which a rotation of the planetary gear can be transmitted, wherein the output element (71) can be attached to the coupling recess (634) of the planet carrier (63) such that internal gear sectors (635) of the planet carrier (63) engage with external gear sectors (711) of the output element (71) to transmit a torque. The planet carrier arrangement is applicable in a drive system for a muscle-powered vehicle.

The present invention discloses a yaw planetary reduction motor applied to a wind turbine. It solves the problem that the existing yaw reducer has a large axial dimension, weak rigidity, and cannot be used in scenarios with strict spatial dimension requirements. The reducer uses the connecting flange of the motor to axially position the planetary gear of the first-stage planetary gear reduction mechanism, the current-stage planetary carrier axially positions the planetary gear bearing, and the upper-stage planetary carrier is axially positioned by the first external tooth of the next-stage sun gear, and a micro-gap is designed with the planetary gear of the next stage. At the same time, the planetary gear and the planetary carrier of the same stage are in direct contact to achieve mutual axial positioning; the planetary carrier of the Nth-stage planetary gear reduction mechanism is fixedly connected to the output gear to achieve power output. Compared with the existing reducer structure, the structure of the present invention not only shortens the overall axial dimension of the planetary gear reducer, but also reduces the characteristics and quantity of parts, thereby reducing manufacturing costs.

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 compact type planet cycloid joint module. The compact type planet cycloid joint module comprises a first-stage planet wheel speed reducing mechanism, a second-stage cycloid pin wheel speed reducing mechanism and a driving component. A hollow eccentric input shaft of the second-stage cycloidal-pin wheel speed reducing mechanism serves as an output shaft of the first-stage planet wheel speed reducing mechanism, and coaxial transmission is achieved. The hollow eccentric input shaft of the second-stage cycloidal-pin wheel speed reducing mechanism and the output shaft of the first-stage planet wheel speed reducing mechanism share one shaft, so that the axial size of the joint module is small, the module is compact, and the rigidity is improved; an output shaft of a first-stage planet wheel speed reducing mechanism is reduced, meanwhile, a large-diameter hollow eccentric input shaft is adopted, and the weight of the joint module is reduced.

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 maneuvering method considering sunlight avoidance based on a middle parking attitude belongs to the technical field of spacecraft attitude control, and comprises the following steps: calculating a minimum included angle between a camera optical axis and sunlight in a maneuvering process according to the pointing direction of the camera optical axis when maneuvering starts and ends; and if the minimum included angle is smaller than a set threshold value and maneuvering needs to be completed quickly, designing a satellite middle parking attitude, so that the included angle between the optical axis of the camera and the sunlight is always larger than the set threshold value in the maneuvering process of the satellite according to the initial attitude-middle parking attitude-tail end attitude. The method is realized by applying spherical geometric analysis design based on the assumption that the included angles of the rotating shaft, the sun vector and the camera optical axis are unchanged, and is small in calculation amount, simple and practical.

The invention relates to the field of electromagnetic calculation, in particular to a petal-shaped Bessel light-based distance measurement method in hypersonic plasma turbulence, which comprises the following steps of: simulating plasma sheath turbulence by establishing a multi-layer random phase screen, calculating diffraction transmission of petal-shaped Bessel light after the petal-shaped Bessel light passes through the turbulence, and capturing intensity distribution of reflected light beams; the rotation angle is determined based on the correlation with the ideal intensity distribution, and high-precision distance measurement is realized by using the linear relation between the rotation angle and the propagation distance. The petal-shaped Bessel light is formed by superposing two Bessel mode light beams with different orbital angular momentum orders, has a spatial structure strongly related to the propagation distance, and can keep higher stability and anti-interference capability in a turbulent environment. By combining the self-focusing and approximately non-diffraction characteristics of the Bessel beam, the target detection and distance measurement precision in hypersonic plasma turbulence is remarkably improved, and reliable technical support is provided for communication and target detection of hypersonic aircrafts.

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.