32 results about "Atmospheric drag" patented technology

The invention provides a two-line element generation method and orbit prediction method adaptive to an analytic propagation model, and the method comprises the steps: obtaining the historical osculating orbit elements at each epoch moment, and carrying out the orbit element optimization step of a plurality of iteration rounds on each historical osculating orbit element through employing the analytic propagation model, obtaining a target osculating orbital element number at each epoch moment, enabling the historical osculating orbital element number to be the same as the target osculating orbital element number, and performing inversion to obtain an average orbital element number; and extracting each average semi-major axis of each average orbital element, fitting each average semi-major axis by adopting a least square method to obtain a semi-major axis attenuation rate, inverting the semi-major axis attenuation rate to obtain an atmospheric resistance parameter, and equivalently converting the atmospheric resistance parameter into a standard resistance term in two rows of elements to obtain the two rows of elements. According to the method, the problems of inconsistent models, insensitive residual errors, unable physical explanation of parameters and the like in the two-line element generation process can be solved, and the ephemeris extrapolation orbit prediction precision is improved.

The invention discloses a fuel-optimal side-by-side four-star configuration periodical maintenance control method. The method comprises the following steps: establishing a complete kinetic model containing J2 perturbation and atmospheric perturbation effects in an LVLH coordinate system; based on the model, the change of the phase along with time under the influence of atmospheric resistance is obtained, and the monotonous change rule of the center deviation of different satellite pairs in the formation satellites along with time is obtained; the baseline length of the satellite pair and the center deviation distance between the first and fourth tracks of the satellite and the second and third tracks of the satellite are used as core constraint conditions. In the optimization framework, weighted fuel cost is used as a target function, and differentiated fuel weight distribution is carried out on different satellite pairs. A hybrid optimization strategy of global search and local refining is carried out by adopting a genetic algorithm, an optimal control variable is efficiently solved, and finally, the fuel consumption is minimized on the premise of meeting all constraints.

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 discloses a satellite orbit forecasting method based on in-orbit calibration of model parameters. Belongs to the technical field of remote sensing satellite measurement, operation and control and particularly relates to the technical field of remote sensing satellite orbit forecasting. According to the method, the satellite windward side-to-mass ratio in an atmospheric resistance model is calibrated regularly according to actually measured orbit data, the resistance coefficient in the atmospheric resistance model is calibrated in real time, and more accurate model parameters are used for modeling a perturbative force model in the satellite orbit forecasting process, so that the error of the atmospheric resistance model is reduced, and the orbit forecasting precision is improved. In the time period that the space weather changes gently, the atmospheric resistance modeling precision is improved to 90% or above from about 60%-80% of a traditional method through the method.

The present disclosure provides a method and device for controlling the orbit configuration of a satellite cluster under J2 perturbation and a medium; the method comprises: constructing a component form expression of relative motion dynamics equation; constructing a space perturbation model based on J2 perturbation interference and atmospheric resistance perturbation interference; describing the communication topology of the satellite cluster through an undirected graph; constructing the relative motion error of each satellite by setting a virtual reference star based on the component form expression of the relative motion dynamics equation and the space perturbation model; constructing a finite-time extended state observer corresponding to the relative motion error of each satellite; designing a self-organizing synchronization controller according to the relative motion error of each satellite, the finite-time extended state observer, the corresponding designed nonlinear fast terminal sliding mode surface, and the communication topology of the satellite cluster.

The application relates to a configuration design method of a three-dimensional orbiting sail with posture stability, and belongs to the field of spacecraft attitude dynamics and structure design.The application is realized by the following method: considering the environmental perturbation of atmospheric resistance, atmospheric resistance moment and gravity gradient moment, an orbit attitude coupling dynamics model of a three-dimensional orbiting sail system is established based on a position vector and Euler angles; the three-dimensional orbiting sail and a spacecraft body are regarded as rigid bodies, and the windward area is considered to be blocked by airflow so as to more accurately describe the orbit movement and attitude movement of the system in the orbiting process; the attitude movement is simplified for the attitude dynamics of the system; the attitude stability of the system in the orbiting process is analyzed by using a phase plane method, a Lyapunov energy function method and other stability analysis methods of nonlinear dynamics systems; the relationship between important structure parameters such as a cone angle and a support rod length of the three-dimensional orbiting sail and the attitude stability is further constructed; and the three-dimensional orbiting sail satisfying the preset attitude stability requirement is obtained based on the relationship, so that the configuration design is realized.

The invention discloses a space target area-to-mass ratio estimation method and system based on historical TLE data, and the method comprises the steps: obtaining the historical TLE data of a space target, and screening a segment of TLE data with the longest duration and the semi-major axis decreasing with time; for TLE data in a selected time period, extrapolating the TLE data into ephemeris for precise orbit determination, and synchronously solving an atmospheric resistance coefficient; inversely calculating a satellite surface-mass ratio according to the atmospheric resistance coefficient; and processing the satellite surface-to-mass ratio, and outputting the processed satellite surface-to-mass ratio. According to the method, the surface-to-mass ratio can be reversely deduced without target matching by screening effective data segments of semi-major axis decline and combining precise orbit determination and atmospheric resistance coefficient calculation by utilizing the disclosed historical TLE data, the problem that the surface-to-mass ratio of a non-cooperative target is difficult to obtain is solved, the feasibility of obtaining the surface-to-mass ratio is improved, and the application range of space target monitoring is expanded.

[Problem] To provide a spacecraft capable of properly controlling the attitude thereof, and a method for manufacturing same. [Solution] A spacecraft 2 comprises: a spacecraft body 10; and a movement control device 12 installed on the spacecraft body 10 and having a deployment membrane 14 that is deployed to act with atmospheric resistance, wherein the movement control device 12 is installed on the spacecraft body 10 such that the geometric center of gravity of the deployment membrane 14 in a deployed state is offset from the center of gravity of the spacecraft body 10 in the direction of a roll axis of the spacecraft. The moment of inertia of the spacecraft 2 around the roll axis is greater than the moment of inertia around a pitch axis and the moment of inertia around a yaw axis.

A method and system for optimizing on-board orbit mission sequences based on real-time windward area is disclosed. The system includes a data configuration module, an area mapping module, a high-precision prediction module, and an autonomous feedback optimization module. The data configuration module receives mission sequences and space environment characteristic parameters, and synchronizes the satellite's real-time orbit vector. The area mapping module analyzes the target attitude based on the attitude parameters in the mission sequence and calculates the equivalent windward area corresponding to each mission time period in real time by combining the velocity vector. The high-precision prediction module executes the attitude-coupled variable area HPOP algorithm, incorporating the dynamically calculated equivalent windward area as a key perturbation term into the atmospheric drag calculation, and outputs orbit prediction results reflecting the attitude adjustment impact. The autonomous feedback optimization module performs consistency judgment on the difference between two adjacent recursive results. If the difference does not meet the preset requirements, it autonomously adjusts the mission execution time and mission execution parameters in the mission sequence and feeds them back to the area mapping module for iterative calculation.

The application provides a satellite orbit prediction method and device based on orbit data of a spherical satellite. The method comprises determining a target atmospheric density model according to a target space environment condition of a satellite to be predicted, wherein the target atmospheric density model is obtained by correcting a preset atmospheric density model according to orbit data and a dynamic model of the spherical satellite in the target space environment condition; identifying an atmospheric drag coefficient of the satellite to be predicted based on the target atmospheric density model; and performing orbit prediction on the satellite to be predicted according to the target atmospheric density model and the atmospheric drag coefficient. According to the embodiment of the application, the spherical satellite maintains a stable and high-precision surface value ratio and atmospheric drag coefficient when operating in an orbit, the orbit data of the spherical satellite can be used to better achieve high-precision correction of the preset atmospheric density model, and the atmospheric drag coefficient of the satellite to be predicted can be more accurately identified based on the corrected atmospheric density model, thereby improving the orbit prediction accuracy of the satellite to be predicted.

Satellite constellation limited time cooperative orbit insertion control method belongs to the technical field of deep space exploration. The method of the present application is as follows: a satellite dynamics model is established with extended orbit elements as state variables, and the model is incorporated with J2 perturbation, atmospheric resistance, third body gravity and other modelable bounded perturbations; according to the orbit dynamics model, a state space expression in the reconstruction process is constructed, and a Lyapunov type artificial potential function about the target orbit state is constructed; in order to ensure collision safety in the reconstruction process, the artificial potential field method is used, the gravitational potential function and the repulsive potential function are constructed under the target orbit and constellation configuration constraint conditions, and the time control part is added to the constructed potential function, which is used to adjust the phase change rate, so that the target phase is reached at the specified time, and then the target convergence potential function is obtained; a multi-target potential function is constructed; and a control law is solved to enable the satellite to complete the orbit insertion task under multiple constraint conditions.

The application discloses a space target surface-mass ratio estimation method and system based on historical TLE data, and the method comprises the following steps: obtaining historical TLE data of a space target, and screening TLE data with the longest semi-major axis descending over time within a time period; extrapolating ephemeris from the TLE data within the selected time period to perform precise orbit determination, and simultaneously solving an atmospheric resistance coefficient; inversely calculating a satellite surface-mass ratio according to the atmospheric resistance coefficient; and processing the satellite surface-mass ratio, and outputting the processed satellite surface-mass ratio. The application inversely deduces the surface-mass ratio without the cooperation of the target by screening the effective data section with the descending semi-major axis, combining the precise orbit determination and the solution of the atmospheric resistance coefficient, solves the problem that it is difficult to obtain the surface-mass ratio of a non-cooperative target, improves the feasibility of obtaining the surface-mass ratio, and expands the application range of space target monitoring.

The application discloses a low-orbit satellite orbit prediction method based on feature screening and a related device, and the method comprises the following steps: defining a pseudo-drag coefficient according to an atmospheric drag perturbation formula; using a SHAP method to analyze the influence value of different space environment features on the pseudo-drag coefficient prediction result of a target prediction model, and generating an influence bee colony diagram; screening out an optimal feature combination based on the influence bee colony diagram, and predicting the pseudo-drag coefficient according to the optimal feature combination and the target prediction model to obtain a pseudo-drag prediction value; and predicting and calculating the low-orbit satellite orbit state according to the pseudo-drag prediction value to obtain an orbit prediction result. This process does not depend on the orbit prediction error and is not affected by the uncertainty of the divergence direction of the error. Therefore, the application can solve the technical problem that the existing orbit prediction depends on the prediction error, but the divergence direction of the error relative to the true atmospheric drag is uncertain, resulting in the lack of accuracy and reliability of the prediction result.

The application discloses a spacecraft of double-configuration atmosphere sail based on cross-section formation, which comprises front and rear atmosphere sail units (1, 6) and atmosphere sail units (2, 3, 4, 5) installed between the front and rear atmosphere sail units; a center joint (9) is arranged in the middle of the double-configuration atmosphere sail spacecraft; a right end folding shaft assembly (7) and a left end folding shaft assembly (8) are arranged at two ends of the center joint (9) respectively. The frame of the double-configuration atmosphere sail spacecraft is a rectangular frame unit formed by support rods and angle connectors. In order to solve the problems of large volume, large mass and large fuel consumption for stable control of a traditional spacecraft attitude and orbit control system, the double-configuration atmosphere sail spacecraft based on cross-section formation does not need to consume fuel, and only relies on atmospheric resistance to realize attitude and formation configuration stability, so that the light weight of the spacecraft is realized.

An autonomous mission planning method for ultra-low sun-synchronous orbit optical remote sensing satellites includes the following steps: extrapolating the satellite orbit considering atmospheric drag and thrust effects; planning orbit maintenance arcs according to an autonomous orbit maintenance strategy; planning imaging and data transmission arcs respectively according to imaging and data transmission mission requirements; eliminating overlapping orbit maintenance, imaging, and data transmission arcs based on priority; estimating whether the planned arcs meet the energy requirements; if not, eliminating the planned arcs; updating the surface-to-mass ratio based on the planned arcs, and replanning using the updated surface-to-mass ratio until the planned arcs remain unchanged compared to the previous one; completing the autonomous mission planning for the next few orbits. This invention uses electric propulsion for autonomous orbit maintenance as a key constraint in mission planning, prioritizing satellite energy for orbit maintenance, and on this basis, maximizing the performance of imaging and data transmission missions.

The invention relates to the technical field of satellite orbits, in particular to a sun-synchronous orbit dip angle offset calculation method and device, equipment and a medium. According to the method, an orbit perturbation equation containing earth non-spherical gravity and atmospheric damping perturbation is constructed, and the influence of semi-major axis attenuation caused by atmospheric resistance on the inclination angle offset is considered. According to the method, the influence of atmospheric damping on the orbit height and different orbit height maintaining strategies are comprehensively considered, more accurate orbit injection parameters are calculated in combination with the prior art, and a more reliable guarantee is provided for subsequent on-orbit normal operation of a spacecraft. In the initial stage of orbit design, more accurate spacecraft orbit injection parameters can be calculated by introducing more perturbation models. The orbit height maintaining strategy is also considered, which means that the later-stage spacecraft in-orbit operation and maintenance cost is also considered, the engineering implementation is more reliable, the result is more accurate, and meanwhile, flexible adjustment can be realized by fully combining actual conditions.

A method and system for maintaining the coplanar flying formation configuration under the influence of space perturbations is disclosed, relating to the field of space satellite formation and configuration maintenance technology. Addressing the problem that traditional satellite formation configuration maintenance methods struggle to maintain the configuration of coplanar flying formation satellites under the influence of J2 term perturbations and atmospheric drag perturbations, this application directly compensates for the perturbation force in the control channel. Furthermore, the non-singular terminal sliding mode can handle nonlinear state-space equations, thus eliminating the need for further linearization of the dynamic equations incorporating the perturbation force, avoiding inherent errors caused by linearization. Therefore, the technical solution of this application can effectively maintain the configuration of coplanar flying formation satellites under the influence of J2 term perturbations and atmospheric drag perturbations.

This invention discloses a method for calculating the sub-satellite point based on the SGP4 orbital model, relating to the field of satellite navigation and orbit prediction technology. The polar spherical projector of this invention eliminates longitude jumps in latitude >75°, ensuring continuous trajectory output for scenarios such as polar scientific expeditions and glacier monitoring. By directly mapping three-dimensional radial information to generate latitude, it avoids the iterative convergence problem of the ellipsoidal model, ensuring natural symmetry of the North and South Poles coordinates and continuous longitude domain. Second-order differential compensation of the mean anomaly angle, combined with dynamic correction of the Earth's rotational angular velocity, strictly aligns the UTC timestamp with the orbital position. It suppresses the cumulative error generated by traditional linear interpolation at perigee. It enhances adaptability to complex perturbation environments; atmospheric drag vectorization compensation captures real-time aerodynamic direction through relative velocity cross terms, combined with continuous processing of the density index model, significantly improving the orbital attenuation prediction accuracy of low-Earth orbit satellites during geomagnetic storms.

Provided is an aerodynamic device that reduces the influence of atmospheric drag on a satellite. The aerodynamic device comprises an aerodynamic shaping part provided at the tip of a satellite body. The aerodynamic shaping part is configured to be deformable between a first state in which the aerodynamic shaping unit has a conical shape and a second state in which the aerodynamic shaping unit is folded from the first state.

The invention discloses a space target deorbiting method and device based on a diffraction light sail and electronic equipment, and relates to the technical field of spaceflight measurement and control, and the space target deorbiting method comprises the steps: configuring the diffraction light sail for a space target, determining the current orbit position of the space target, and determining the position of the space target on the basis of the current orbit position; and determining a first included angle between the velocity vector of the space target and the sun direction vector, determining a diffraction light sail strategy based on an included angle range to which the first included angle belongs, and adjusting a diffraction angle and a sailboard angle of the diffraction light sail by adopting the diffraction light sail strategy so as to drive the space target to deorbit through resistance generated by the adjusted diffraction light sail. The technical problem that the speed of natural deorbit of a space target depending on atmospheric resistance is low in the prior art is solved.

The application relates to the field of spacecraft propulsion technology, in particular to a continuous adjustable electric propulsion testing method and device, the method comprising setting a test instruction change period synchronized with a satellite orbit period; constructing an atmospheric resistance interference model changing with an orbit phase and converting the atmospheric resistance interference model into a dynamic thrust demand benchmark; generating at least one dynamic thrust instruction in a sine wave, a triangular wave, a step signal or a composite waveform; using a three-dimensional interpolation matrix to solve power and gas supply parameters in real time and driving a thruster to perform wide-range dynamic adjustment; and evaluating dynamic tracking performance based on a root mean square error and a frequency response characteristic index. The application realizes high-fidelity ground reproduction of a complex interference environment of a satellite in orbit, overcomes the limitations of existing static tests and single working conditions, and provides a scientific and effective testing scheme for dynamic performance verification of a continuous adjustable electric propulsion system.

The application discloses a low-orbit satellite parameter adaptive control method with a solar sail driving device, belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The method solves the technical problem of overcoming the influence of satellite moment of inertia and compensating the influence of SADA disturbance torque and atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in the form of a quaternion, and on the basis, constructing a satellite attitude kinematics and dynamics equation for attitude controller design; defining the error attitude kinematics and dynamics equation of the satellite through an attitude control error and an angle control error; dividing and converting the parameters to be estimated and known state variables, and representing the negative influence of moment of inertia error on the attitude by matrix operation; designing an attitude controller control torque in a manner of combining nonlinear feedback control with adaptive control; and updating the adaptive estimation value in the attitude controller control parameter by using a parameter adaptive law.

A satellite system operates at altitudes between 180 km and 350 km relying on vehicles including an engine to counteract atmospheric drag to maintain near-constant orbit dynamics. The system operates at altitudes that are substantially lower than traditional satellites, reducing size, weight and cost of the vehicles and their constituent subsystems such as optical imagers, radars, and radio links. The system can include a large number of lower cost, mass, and altitude vehicles, enabling revisit times substantially shorter than previous satellite systems. The vehicles spend their orbit at low altitude, high atmospheric density conditions that have heretofore been virtually impossible to consider for stable orbits. Short revisit times at low altitudes enable near-real time imaging at high resolution and low cost. At such altitudes, the system has no impact on space junk issues of traditional LEO orbits, and is self-cleaning in that space junk or disabled craft will de-orbit.

The application discloses a spacecraft mass estimation method and device, a storage medium and an electronic device. The method comprises the following steps: obtaining a measured comprehensive drag coefficient of a target orbit in which a spacecraft is located and an atmospheric drag coefficient initial value of the target orbit, wherein the measured comprehensive drag coefficient is an actually measured comprehensive drag coefficient; determining a nominal comprehensive drag coefficient of the target orbit based on the atmospheric drag coefficient initial value and a preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is a theoretically calculated comprehensive drag coefficient; determining an atmospheric drag coefficient target value based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and estimating a target mass of the spacecraft based on the atmospheric drag coefficient initial value, the preset mass and the atmospheric drag coefficient target value. The application solves the technical problem of low efficiency in estimating the mass of the spacecraft in the related art.

The invention discloses a low earth orbit satellite orbit forecasting method and device based on Doppler dynamic constraint, and belongs to the technical field of spaceflight measurement and control and satellite navigation. According to the method, bidirectional coherent Doppler observation data are obtained through an inter-satellite laser link, and a Doppler dynamic constraint mechanism is constructed in combination with GNSS precise orbit determination information, so that high-precision prediction of a low-orbit satellite orbit is realized. The method comprises the following steps: adaptively adjusting a process noise matrix and dynamically correcting modeling errors of non-conservative forces such as atmospheric resistance and solar radiation pressure by using the energy change of a Doppler observation residual error in a tangential sensitive frequency band; when the GNSS signal is weak or lost, stable and reliable orbit constraint can still be provided; and outputting the orbit state and uncertainty evaluation in the future 30-120 minutes through filtering updating and dynamic extrapolation. According to the method, the precision and robustness of orbit forecasting are remarkably improved, and the method is suitable for low-orbit satellite task scheduling and safety management and control.

The invention discloses a moon shadow time calculation method based on an SGP4 model, and relates to the technical field of spacecraft orbit dynamics and energy management.The moon shadow time calculation method comprises the steps that a closed feedback loop is formed by extending a Kalman filter state variable and taking the instantaneous noise offset of an atmospheric resistance parameter B * as an independent estimation target; orbital perturbation changes caused by sun activity peaks, sudden geomagnetic disturbance and the like can be tracked in real time, and the stability of penumbra entering and exiting time of the near-earth satellite is remarkably improved; in order to solve the problem of frequent switching of shadow intervals of high-orbit satellites, a time period merging strategy based on an error ellipsoid and shadow area overlapping probability is provided. By quantizing the contribution of the position uncertainty to the shadow state, a pseudo switching event generated by model jitter is filtered out, a continuous shadow time period conforming to engineering practice is output, and invalid operation of an energy system is reduced.

The application discloses a low-orbit satellite orbit determination precision improvement method and device under geomagnetic storm conditions and belongs to the technical field of satellite navigation. The method proposes an adaptive orbit determination framework of dynamic model and stochastic model optimization. At the dynamic level, the segmented estimation length of empirical force and atmospheric resistance coefficient is dynamically adjusted according to the geomagnetic activity level, and the resolution capability for rapid changes of non-conservative force is improved. At the stochastic model level, an external physical driving weight function based on the SYM-H index and an internal data diagnosis weight function based on observation residuals are constructed, and real-time adaptive adjustment of observation value variance is realized. Through the "geomagnetic index-observation residual" double driving mechanism, the low-orbit satellite orbit determination error during the strong magnetic storm is reduced from the decimeter level to the centimeter level, and the precision and robustness of the low-orbit satellite orbit determination under the space weather event are significantly improved.

The application discloses a method for calculating atmospheric resistance of LEO spacecraft based on posture rotation mode, and relates to the field of wind tunnel test, and comprises the following steps: S1, selecting a corresponding posture rotation mode of the spacecraft according to a scene; S2, calculating the windward area and atmospheric resistance of the posture rotation mode of the spacecraft selected in S1 in a period; and S3, based on the calculation result of S2, calculating the period average atmospheric resistance of the spacecraft in the corresponding posture rotation mode; the method for calculating atmospheric resistance of LEO spacecraft based on posture rotation mode is a fast and efficient algorithm for atmospheric resistance under the operation condition of dynamic change of the in-orbit spacecraft posture, the period average atmospheric resistance is obtained by precisely calculating the atmospheric resistance in the rotation period, and thus an important basic prerequisite is provided for the medium and long term analysis and prediction of the orbit parameters of the LEO spacecraft.

The invention discloses a satellite atmospheric resistance modeling method based on a geometric effective area and a relative flow direction. The method belongs to the technical field of aerospace dynamics and satellite orbit determination. The defects that only resistance coefficient modeling is carried out in an existing satellite atmospheric resistance model, and the reference area is fixed to be a constant are overcome. Comprising the following steps: S1, defining a relative speed and a coordinate system; s2, effective windward area calculation based on a geometric model: constructing a satellite geometric model, and calculating a geometric effective windward area of a satellite; s3, constructing a resistance coefficient angle model: constructing the resistance coefficient angle model through the azimuth angle and the pitch angle of the relative flow direction under the body-fixed coordinate system; s4, defining an effective windward area-coefficient combination quantity: combining the resistance coefficient angle model with the geometric effective windward area of the satellite, and defining the effective windward area-coefficient combination quantity; and S5, constructing an improved atmospheric resistance acceleration model: constructing the improved atmospheric resistance acceleration model through the effective windward area-coefficient combination quantity.

The application discloses a kind of low-orbit satellite orbit prediction method and device based on Doppler dynamic constraint, belong to spaceflight measurement and control and satellite navigation technical field.The method obtains two-way coherent Doppler observation data by inter-satellite laser link, combined with GNSS precise orbit determination information, constructs Doppler dynamic constraint mechanism, realizes high-precision prediction to low-orbit satellite orbit.Core includes: using the energy change of Doppler observation residual in tangential sensitive band, self-adapting adjustment process noise matrix, dynamically revising the modeling error of non-conservative force such as atmospheric resistance and solar radiation pressure;When GNSS signal is weak or lost, still can provide stable and reliable orbit constraint;By filtering update and dynamics extrapolation, the orbit state and uncertainty evaluation of future 30-120 minutes are output.The application significantly improves the precision and robustness of orbit prediction, and is suitable for low-orbit satellite task scheduling and safety control.