15 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 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.

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 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.

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.

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 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.

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.