55 results about "Atmospheric drag" patented technology

The embodiment of the invention provides a simulation evaluation method and device for a multi-satellite cooperative task, and the method comprises the steps: building a unified time-space reference based on a standard time signal, and building a complete environment disturbance model comprising sunlight pressure, atmospheric resistance and geomagnetic field interference. A fault injection mechanism of load, attitude and orbit control is innovatively designed, and real simulation of multi-satellite cooperative motion is realized. According to the system, a system-level, subsystem-level and component-level multi-level evaluation index system is established, a dynamic weight calculation method based on a task execution state is adopted, and matching analysis of a three-dimensional error curve and a key event time sequence is combined, so that accurate evaluation of multi-satellite cooperative task performance is realized; and a reliable decision basis is provided for space task planning and system optimization.

The invention discloses an ultra-low orbit guarantee method, system and equipment for a near-earth satellite. Comprising the following steps: determining a satellite orbit element at the current moment based on an atmospheric resistance perturbation item, a conventional perturbation item and a small thrust perturbation item, and inputting the satellite orbit element at the current moment into an orbit mechanical model; based on the orbit mechanical model and the satellite orbit elements, determining a first velocity position parameter of the satellite at the prediction moment under the ICRS coordinate system, and converting the first velocity position parameter into an ECEF coordinate system to obtain a second velocity position parameter at the prediction moment; inputting the time characteristic, the satellite state characteristic, the space environment characteristic and the second speed position parameter of the current moment into an orbit correction model to obtain a target error prediction value; adjusting the second speed position parameter based on the target error prediction value to obtain a target speed position parameter of the satellite at the prediction moment; and performing orbit attenuation risk early warning based on the target speed and position parameters of the satellite at the prediction moment.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

The invention discloses a forecasting method and device for an atmospheric resistance coefficient of a low-orbit satellite in the technical field of satellite navigation. The forecasting method comprises the following steps: acquiring satellite orbit determination data and a space environment index in a preset time period before forecasting; according to the satellite orbit determination data, performing simplified dynamic precise orbit determination on a low-orbit satellite, and fitting an atmospheric resistance coefficient in the simplified dynamic precise orbit determination process; performing data preprocessing on the space environment index and the atmospheric resistance coefficient and generating a sample set; constructing an LSTM prediction model, and training, adjusting and evaluating the LSTM prediction model through the sample set; predicting an atmospheric resistance coefficient in a future preset time period according to the finally obtained LSTM prediction model; and applying the atmospheric resistance coefficient in the future preset time period to the atmospheric resistance perturbation model of the low-orbit satellite for orbit forecasting. According to the method, the atmospheric resistance coefficient in the LEO satellite orbit prediction process is predicted in a high-precision manner through the LSTM prediction model, so that the orbit prediction precision is effectively improved.

The invention belongs to the technical field of aerospace, and relates to an unscented batch processing orbit determination method and system for a continuous low-thrust maneuvering satellite. The method comprises the following steps: acquiring a radar observation vector according to radar measured data or satellite ephemeris data in combination with radar site coordinates; obtaining an initial orbit state at a to-be-estimated moment by adopting a Lambert algorithm according to the radar observation vector; generating an initial expansion state quantity according to the initial orbit state at the to-be-estimated moment; a simplified orbit kinetic equation containing earth non-spherical perturbation, atmospheric resistance perturbation and constant small thrust acceleration under an RTN coordinate system is constructed under the ECI coordinate system; based on the simplified orbit kinetic equation, obtaining an expansion state quantity precision value of the to-be-estimated moment through the initial expansion state quantity, the radar observation vector and an unscented batch processing filtering method; and substituting the precision value of the expansion state quantity of the to-be-estimated moment into the simplified orbit kinetic equation to carry out orbit forecasting at other moments. According to the method, the orbit determination precision is remarkably improved, and the influence of single-point observation errors is reduced.

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.

Tether-based reaction control systems for spacecraft undergoing atmospheric entry or reducing orbital energy, including systems that use atmospheric drag on one or more tethers to alter the angle of attack of the spacecraft during its atmospheric transit to reduce the thermal and mechanical loads on the spacecraft during the reentry process. Systems may include tethers that together generate a force that functionally acts on the spacecraft at a point that is offset from the spacecraft's center of mass. Such systems may be either actively or passively controlled.

The invention discloses a low earth orbit satellite orbit forecasting method based on feature screening and a related device. The method comprises the following steps: defining a pseudo resistance coefficient according to an atmospheric resistance perturbation formula; an SHAP method is adopted to analyze influence values of different space environment characteristics on a pseudo resistance coefficient prediction result of the target prediction model, and an influence bee colony graph is generated; screening out an optimal feature combination based on the influence bee colony graph, and predicting a pseudo resistance coefficient according to the optimal feature combination and the target prediction model to obtain a pseudo resistance prediction value; and performing prediction calculation on the orbit state of the low-orbit satellite according to the pseudo resistance prediction value to obtain an orbit prediction result. The process does not depend on orbit forecast errors and is not influenced by uncertainty of error divergence directions. Therefore, the technical problem that the prediction result is lack of accuracy and reliability due to the fact that the existing orbit prediction depends on the prediction error but the error is uncertain relative to the divergence direction of the real atmospheric resistance can be solved.

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 present invention belongs to the field of aerospace technology, and relates to a method and system for orbit determination of a continuously small-thrust maneuvering satellite by unscented batch processing. The present invention obtains a radar observation vector according to radar measured data or satellite ephemeris data in combination with the radar station location coordinates; obtains the initial orbit state at the moment to be estimated according to the radar observation vector and by using the Lambert algorithm; generates an initial augmented state quantity according to the initial orbit state at the moment to be estimated; constructs a simplified orbit dynamics equation including the non-spherical perturbation of the earth, the atmospheric drag perturbation and the constant small-thrust acceleration in the RTN coordinate system in the ECI coordinate system; based on the simplified orbit dynamics equation, obtains the precise value of the augmented state quantity at the moment to be estimated through the initial augmented state quantity, the radar observation vector and the unscented batch processing filtering method; and substitutes the precise value of the augmented state quantity at the moment to be estimated into the simplified orbit dynamics equation for orbit prediction at other moments. The present invention significantly improves the orbit determination accuracy and reduces the influence of single-point observation errors.

The present invention belongs to the technical field of aerospace data decoupling processing. Specifically, it relates to a method and device for decoupling atmospheric drag and maneuvering force of a low-orbit continuously maneuvering target, including: Step 1) Select a certain low-orbit continuously maneuvering target and obtain continuous orbit tracking and monitoring data under different space environments; Step 2) Divide the continuous orbit tracking data in each space environment into n segments equally to obtain n segments of data; Step 3) Construct an orbit dynamics model of the maneuvering target; Step 4) For each equally divided segment of data, deduct external factors and only retain the factors of maneuvering force and atmospheric damping force to obtain the change in the semi-major axis of the orbit and the change in the argument of latitude of each segment of data; Step 5) Construct an optimization objective function for the semi-major axis of the orbit and an optimization objective function for the argument of latitude; Use the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the radial component of the maneuvering force for each segment of data; and integrate them respectively.

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.

The invention relates to a space debris cloud orbit evolution prediction method and system based on atmospheric resistance and group dynamics, and the method comprises the steps: obtaining an initial orbit parameter of a target debris cloud, carrying out the subclass division of a debris group according to the initial orbit parameter, and simulating and generating a random initial state of the debris group; and when the random initial state of the fragment group is simulated and generated, the orbit evolution trajectory of each subclass fragment is calculated, and a three-dimensional orbit evolution result is generated. According to the method, the crossing of debris cloud evolution from static homogeneous prediction to dynamic heterogeneous simulation is realized, and a high-credibility time sequence planning basis is provided for active removal of space debris.

The application discloses a method for adjusting and controlling the phase of a low-orbit micro-nano satellite constellation by means of atmospheric differential drag, and belongs to the field of on-orbit phase control of a constellation. The method aims at solving the problem of low phase control precision of the existing atmospheric differential drag control satellite constellation method. The method comprises the following steps: establishing a single-satellite atmospheric drag dynamics model to obtain the corresponding relationship between the single-satellite control input and the control command; determining the state constraint condition of other single satellites based on a reference satellite, combining the energy constraint and the task constraint, and adopting a simulated annealing algorithm to perform constellation slot allocation, and taking the shortest control time as the target to obtain the optimal result of the constellation slot allocation; calculating an initial control command matrix; taking the lowest adaptive function value and meeting the target control precision as the control target, optimizing the initial control command matrix to obtain an optimal control command matrix, so as to obtain the optimal satellite control input, and performing the angular acceleration control of the single satellite, thereby realizing the phase adjustment of the same-orbit constellation. The application is used for the phase adjustment control of the constellation.

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 present invention discloses a method for inverting the density of the Martian atmosphere, and its device and electronic equipment, which relate to the field of aviation measurement and control or other related technical fields. The method comprises: monitoring the operating status of a Mars probe in a Martian exploration orbit and collecting the operating data of the Mars probe; drawing an operation diagram of the Mars probe based on the operating data; utilizing the principle of height symmetry and calculating the energy attenuation of the Mars probe due to atmospheric resistance at the perigee and in each height symmetric interval on the operation diagram based on the operating data; inverting the average atmospheric density in each height symmetric interval of the Martian atmosphere based on the energy attenuation; and converting the average atmospheric density in each height symmetric interval to a specific altitude position to obtain the atmospheric density at each altitude position in the Martian atmosphere. The present invention solves the technical problem in the related art of inverting the Martian atmospheric density by expanding the inversion time interval, which results in low measurement and control accuracy.

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 invention discloses a Mars atmospheric density inversion method and device and electronic equipment, and relates to the field of aviation measurement and control or other related technical fields, and the method comprises the steps: monitoring the operation state of a Mars detector in a Mars detection orbit, and collecting the operation data of the Mars detector; drawing a running diagram of the Mars probe based on the running data; calculating the energy attenuation of the Mars probe on a near fire point on a working diagram and in each highly symmetric interval by using a highly symmetric principle and based on the operation data; inverting the average atmospheric density in each highly symmetric interval of the Mars atmosphere based on energy attenuation; and reducing the average atmospheric density in each height symmetric interval to a specific height position to obtain the atmospheric density of each height position of the Mars atmosphere. According to the invention, the technical problem that the measurement and control precision is low when the Mars atmospheric density is inverted by expanding the inversion time interval in the related technology is solved.

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 present application discloses a method, device, medium and product for determining the ultra-low orbit of a near-earth satellite. The method includes: when the thrust of the ion thruster of the satellite is constant, obtaining the thrust vector of the satellite engine in the satellite body coordinate system; obtaining the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; adding the small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain an input parameter, and inputting the input parameter into the orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term and the small thrust perturbation term.

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.