42 results about "Orbital period" patented technology

The invention relates to a method for eliminating a double-planetary-calendar forecast error based on post-event precise ephemeris. The method comprises the following steps: forecasting a satellite position of the day at an ephemeris reference moment based on double planetary calendars and an SDP4 model; calculating the satellite position of the ephemeris reference moment on the day based on the post-event precise ephemeris of the double planetary calendars on the ephemeris reference moment on the day, and taking the satellite position as a reference satellite position; acquiring a satellite position forecast error by comparing the difference between the forecast satellite position and the reference satellite position of the day at the ephemeris reference moment, and fitting the satellite position forecast error of the satellite in one orbit period to obtain an error function of the double-planetary-calendar forecast satellite position; and performing error elimination on the forecast satellite position of the target date according to the error function to obtain the forecast satellite position after error elimination. By adopting the method, the satellite position forecasting error of the double planetary calendars can be eliminated, and the precision of the double planetary calendars for medium and long term forecasting is improved.

The invention provides an IGSO satellite tracking method and device of an automatic satellite tracking antenna and a medium. The method comprises the steps that orbit parameters and epochs of a target IGSO satellite are acquired; extrapolating the orbit from the orbit epoch moment, setting the extrapolation time as an orbit period T and calculating the step length S; converting the orbit data into earth-fixed system coordinate system coordinates and sub-satellite point data at corresponding moments; and calculating the direction vector of the antenna pointing to the satellite according to the latitude and longitude of the tracking antenna and the earth-fixed system coordinate data at the corresponding moment, and converting the direction vector into an azimuth angle and a pitch angle, thereby generating an antenna guide data file. And then calculating a guide data sequence number at the current moment through a date difference value between the current system time and the track epoch moment. And finally, the antenna orientation and the pitching motor are controlled to move to the target position corresponding to the guide data serial number, and the technical scheme provided by the invention can realize the tracking of the traditional automatic satellite tracking antenna on the IGSO satellite on the premise of not changing the antenna form.

The invention discloses a rapid topology sampling method and device for a large-scale constellation in an unreliable space environment, and mainly solves the problem of low sampling efficiency in the prior art. According to the implementation scheme, satellite constellation configuration parameters are obtained; different orbit link establishment modes of all satellite nodes under different distortion factors are determined; determining distribution modes of all satellite nodes on a virtual plane under different constellation configurations; judging whether the satellites have synchronism in the critical latitude region crossing event or not, and judging whether topological dynamics can be caused when the first orbit satellite and the tail orbit satellite cross the critical latitude region or not; automatically calculating the number and length of topology snapshots of constellation distortion topology in an orbit period according to a different orbit link establishment mode, a distribution mode of satellite nodes on a virtual plane and a topology dynamic result, and determining a distribution rule of inter-satellite topology snapshots; and converting a calculation result into an inter-satellite topology sampling time point and a corresponding topology snapshot length according to the distribution rule. The method is high in sampling efficiency and low in use threshold, and can be used for design and optimization of a large-scale distorted topological constellation network protocol.

The application discloses a method and device for orbit transfer control of a lunar spacecraft, and realizes orbit transfer of the spacecraft into a full-parameter target orbit. The method comprises the following steps: a three-impulse control strategy of a far-lunar maneuver, phase control and period control is designed; based on an orbit plane, an arch line and a perigee height, control time and omnidirectional control impulses for orbit transfer from an initial orbit to a first intermediate orbit are determined; based on an orbit phase, a first tangential impulse for orbit transfer from the first intermediate orbit to a second intermediate orbit is determined; and based on an orbit period, a second tangential impulse for orbit transfer from the second intermediate orbit to the target orbit is determined. The first tangential impulse and the second tangential impulse are taken as variables, and a far-lunar maneuver position vector deviation is taken as a fitness value, and a genetic optimization algorithm is used to solve initial values of the impulses. The method solves the problem of orbit control solution caused by nonlinear enhancement of orbit evolution under perturbation influence and complexity of omnidirectional impulse control parameters.

The present invention proposes a method for modeling the ranging error between navigation satellites, belonging to the field of ranging between navigation satellites. First, precise orbit determination and precise clock error estimation are performed using the inter-satellite link observation data, and the inter-satellite orbit determination residuals and clock error estimation residuals are output. Secondly, the residual sequences are normalized to the same orbital period, and a high-order polynomial with a semi-orbital period term is used for fitting. Subsequently, through the F-test and t-test, the overall effectiveness of the model and the significance of each coefficient of the model are checked. Finally, the ranging error model and the ranging error model of the reverse-normalized one-way link are output. According to this process, the ranging error modeling work of the one-way observation value of any link can be completed. The present invention can effectively improve the inter-satellite link observation model, reduce the observation value residuals of orbit estimation and clock error solution, and meet the requirements of high-precision precise orbit determination and clock error estimation.

The application relates to a design method of a moon global positioning constellation based on a lunar space DRO reference orbit, wherein the initial orbit state and the orbit period of any DRO reference orbit in a DRO reference orbit to be deployed are determined based on the initial phase of the DRO reference orbit; the initial orbit state of any satellite to be deployed on the initial phase of the DRO reference orbit is determined based on the orbit state corresponding to the target phase; the orbit period of the any satellite is determined based on the orbit period of the any DRO reference orbit; the target orbit state corresponding to the any satellite is obtained based on the orbit period of the any satellite and the initial orbit state of the any satellite; and the moon global positioning constellation under a four-body model is constructed based on the target orbit state corresponding to each satellite. The moon global positioning constellation constructed by the application has higher stability and higher moon surface coverage capacity compared with the existing near-moon constellation.

The invention belongs to the technical field of spacecraft orbit design and operation management, and provides a satellite orbit life rapid calculation method, which combines a differential-integral method with Gaussian-Legendre integral, calculates the average change rate of orbit elements in an orbit period through the differential-integral method, and calculates the service life of the satellite orbit. The method combines the adaptive sampling characteristics of Gaussian-Legendre integral to realize large-step integral, greatly reduces the number of iterations and calculation amount, repeats the integral until the perigee height is less than the reentry height, the satellite falls down, and outputs the cumulative time of multiple integrals, namely the service life of the satellite orbit, thereby solving the technical problem that the calculation efficiency and precision cannot be considered at the same time. The method has important application value for satellite orbit maintenance, space debris management, reentry prediction and the like.

The invention provides a method for autonomously reducing power consumption in a shadow period of a low-orbit satellite, and the method comprises the steps: setting a first threshold configuration and a second threshold configuration which comprise different threshold ranges for each heater, and enabling the upper and lower limits of the second threshold configuration to be not lower than those of the first threshold configuration; establishing a mapping relation between a satellite orbital plane solar incident angle and an orbital period illumination duration, and setting a critical solar incident angle as a shadow period control starting condition; the method comprises the following steps: monitoring accumulation time of non-discharge of a storage battery pack in real time, triggering threshold switching condition judgment when a solar incident angle is within a certain range in combination with current illumination period duration and preset heater advanced temperature rise time, and switching from first threshold configuration to second threshold configuration for preheating if a specified condition is met. And after the storage battery is discharged, the first configuration is automatically recovered. Therefore, according to the invention, the discharge depth and the peak discharge current of the storage battery can be autonomously reduced, and the on-orbit energy safety management capability is improved.

The application discloses a satellite orbit transfer strategy planning method and device, equipment and a storage medium. The method comprises the following steps: initializing an initial time of a satellite and performing orbit period extrapolation to update a t time to an extrapolated time; initializing an orbit transfer flag as not completed, calculating a first orbit parameter of the satellite at the t time and a second orbit parameter of a target virtual satellite at the t time, and determining a relative orbit parameter of the satellite to the target virtual satellite; determining each orbit parameter correction flag according to the relative orbit parameter and a preset orbit threshold value, and determining a corresponding orbit control sub-mode based on the each orbit parameter correction flag; when the orbit transfer flag is completed, determining a firing parameter of the satellite based on the orbit control sub-mode, and performing orbit transfer maneuver and orbit transfer phase adjustment processing. The scheme can quickly select the corresponding orbit control sub-mode according to the current orbit transfer requirement, quickly generate the calculation and simulation verification of the electric propulsion orbit transfer parameter, and make the satellite accurately enter the predetermined orbit position.

This application provides a method, apparatus, electronic device, and storage medium for adjusting the solar array's attitude towards the sun, relating to the field of satellite control technology. The method for adjusting the solar array's attitude towards the sun includes: when the energy output corresponding to the target satellite's position at a preset orbital point is not within a preset energy output range, acquiring the target solar array's attitude information towards the sun and historical attitude information from the previous orbital period adjacent to the current orbital period; determining initial attitude adjustment information based on the target solar array's attitude information and the historical attitude information; determining multiple historical attitude adjustment information corresponding to the initial attitude adjustment information, and determining the target attitude adjustment information based on the multiple historical attitude adjustment information; determining the final solar array's attitude information based on the target attitude adjustment information and the target solar array's attitude information, and controlling the solar array to adjust its attitude based on the final solar array's attitude information.

The application discloses a high-resolution isochronous stratigraphic framework construction method based on astronomical cycle and sedimentary interface verification, and comprises the following steps: obtaining an absolute astronomical age scale of a target reservoir unit; extracting an astronomical cycle signal from a natural gamma logging curve and identifying a characteristic orbit period; identifying natural sedimentary interfaces inside the reservoir; performing astronomical cycle phase analysis on the sedimentary interfaces by using the absolute astronomical age scale to verify and determine isochronous sedimentary interfaces; and performing isochronous stratigraphic division based on the verified isochronous interfaces and in combination with the periodicity of the astronomical cycle. The application combines the global isochronous constraint of the astronomical cycle with the verification of the internal physical sedimentary interfaces of the reservoir, effectively solves the problems that the division result is easy to be out of time and difficult to be compared across regions in the prior art, and the constructed isochronous stratigraphic framework has high resolution, isochronous and reasonable deposition, and provides a reliable stratigraphic correlation scale for fine description of deep reservoirs.

The invention discloses a low-thrust spacecraft asymmetric halo orbit design method, and belongs to the field of spacecraft orbits. The problem that the periodicity and stability of the orbit of a low-thrust spacecraft cannot be guaranteed under the condition that the plane of the halo orbit is asymmetric is solved. The method comprises the following steps: calculating a halo orbit satisfying a resonance relation with a second gravitational body orbit period; the eccentricity rate of the second gravitational body orbit is continued from zero to the actual eccentricity rate through numerical continuation in the first step, and a zero-thrust resonance halo orbit is obtained; through numerical continuation in the second step, the thrust acceleration amplitude is continued from zero to a preset upper limit while the thrust direction of the spacecraft is kept in the sun-spacecraft connecting line direction; through numerical continuation in the third step, the thrust taper angle is continued from zero to the preset upper limit under the gravity plane coordinate system; and through numerical continuation in the fourth step, the thrust azimuth angle is continued from zero to a preset value under the gravity plane coordinate system, and finally the asymmetric halo orbit is obtained. The method is suitable for circular and elliptical restrictive three-body problems.

The application discloses a scanning method and device of a non-cooperative target, electronic equipment and program product, and relates to the technical field of spaceflight measurement and control, wherein the scanning method comprises the following steps: acquiring initial orbit elements of the non-cooperative target, celestial body parameters of a target celestial body and antenna parameters of an antenna; determining the number of beam pointing positions required for the antenna to scan one orbit period of the non-cooperative target based on the initial orbit elements, the celestial body parameters and the antenna parameters; determining an acceleration multiple based on the orbit period of the non-cooperative target, the number of beam pointing positions and the antenna parameters; and generating guide data based on the initial orbit elements and the acceleration multiple, so as to control the antenna to scan the non-cooperative target through the guide data. The application solves the technical problem that the non-cooperative target cannot be efficiently scanned in the related art, thereby reducing the probability of capturing the non-cooperative target.

The application provides a spacecraft pure magnetic control sun safety control method and system based on a magnetic moment device, which comprises the following steps: calculating a normal vector of a rotation shaft plane through a sun vector and a sun-pointing vector of the system, normalizing a magnetic field vector and constructing a unit orthogonal vector in the normal plane, solving an optimal rotation shaft vector in combination with a rotation inertia array, dynamically adjusting a control parameter according to a geometric included angle between the optimal rotation shaft and an inertial angular velocity of the system, superimposing a magnetic control sun pointing directional control amount and a damping control amount to obtain a total control amount. The system comprises vector calculation, optimal rotation shaft solving, control parameter adjustment and magnetic control amount output modules connected in sequence, and can be additionally provided with an execution driving module. The application fully considers the geometric relationship of multiple vectors, avoids attitude oscillation divergence, and can realize a sun pointing directional accuracy of ≤15° in less than one orbit period, thereby guaranteeing the charging of a solar cell array and providing a reliable pure magnetic control scheme for a spacecraft safety mode, and the engineering application value is remarkable.

The application relates to a space target local retrieval and identification method, and belongs to the cross field of space situation awareness, space dynamics and high-performance databases. The steps comprise the following steps: multi-source TLE orbit data and service attribute fusion pretreatment; multi-feature fusion dynamic decision tree adaptive time step allocation based on orbit period, eccentricity and task type; the HEALPix equal-area celestial sphere gridding algorithm is introduced to reduce the three-dimensional continuous observation space to a discrete one-dimensional grid identifier, and a lightweight space-time key-value pair is reconstructed in combination with discrete time slicing; a high-concurrency space-time inverted index is reconstructed by multi-process lock-free sandbox streaming writing and database kernel-level aggregation deduplication; a fast initial screening of candidate targets is realized by constructing an anti-time-space aliasing dynamic safety margin expansion search area; and high-precision target accurate matching identification. The application realizes fast retrieval and high-confidence identification of a large number of space targets with a constant-level time complexity, and significantly improves the real-time response capability of a space situation awareness system.

The application relates to the technical field of satellite task planning, and discloses a multi-constraint satellite task planning method and system based on U-Net. The method comprises the following steps: acquiring orbit parameter information of a target satellite; constructing multi-source constraint adjustment; constructing a unified gridded multi-channel feature set; constructing a three-dimensional sample feature vector based on an orbit period and a snake-shaped arrangement mode; generating a corresponding binary label feature vector; and outputting a satellite task planning prediction result. Compared with the prior art, the method can solve the technical problems that the prior art relies on manual rules to formulate tasks, is difficult to cope with multi-source heterogeneous constraint fusion and large-scale sample parallel processing, and especially under the conditions that task windows overlap densely or ground resources conflict seriously, the prior art cannot realize fast and efficient imaging task planning. Due to the three-dimensional space-time representation of multi-constraint information and the introduction of the U-Net model with a coding and decoding jump structure, the intelligent level of satellite imaging task planning is improved.

The present invention discloses a method and device for shale sequence stratigraphic division based on sediment noise simulation, including: (1) selecting various logging curves, performing preprocessing and then spectral analysis to identify astronomical orbital periods; (2) performing Gaussian band-pass filtering on the thickness corresponding to the astronomical orbital periods in the logging curves, selecting the filtered curves to establish a time-depth conversion model, and obtaining the time-domain sequences of various logging curves; (3) calculating the time-domain sequences of various logging curves through a sediment noise model to obtain several different relative sea / lake level change curves; (4) performing principal component analysis on the median curve of several different relative sea / lake level change curves to extract the maximum principal component PC1 curve; (5) identifying the maximum and minimum values at the 1200 kyr or 2400 kyr periods of the PC1 curve, and identifying the maximum sea / lake flooding surface and sequence stratigraphic interface. The present invention can solve the problem of constructing an isochronous sequence stratigraphic framework for mudstone shale in the prior art.

The present application relates to a method for designing a transfer orbit for a HALO orbiter to fly over a given ground point. The method includes: finding the moment when the lunar declination value is equal to the latitude value of the given ground point within the given mission time range, and selecting the time point closest to the initial moment as the initial value of the iteration for the orbit transfer moment; calculating the flight duration of the orbit transfer point relative to the starting point using the starting time and the initial value of the iteration for the orbit transfer moment; calculating the normalized time by dividing the flight duration by the HALO orbit period; calculating the magnitude of the orbit transfer velocity increment based on the normalized time using the polynomial fitting coefficients; establishing a mapping relationship between the magnitude of the orbit transfer velocity increment and the initial value of the iteration for the orbit transfer moment through numerical integration based on a high-precision force model; and completing the solution of the parameter to be solved using the Newton iteration method based on the mapping relationship. The method adopted has a short response time, high calculation result accuracy, and fast iteration convergence.

The application discloses a transfer orbit parameter optimization selection method of a lunar-circulating large-ellipse frozen orbit and belongs to the technical field of deep space exploration orbits. The method comprises the following steps: setting parameters that can be optimized in a transfer process of a lunar-circulating large-ellipse frozen orbit; analyzing influences of different near-lunar-point altitudes, inclination angles and periods of a capture orbit on velocity increments, measurement and control conditions and orbit safety; analyzing influences of different near-lunar-point altitudes, inclination angles and near-lunar-point amplitude angles of a transition orbit on velocity increments, measurement and control conditions and orbit safety; and analyzing influences of different mission orbit periods on velocity increments and orbit safety. If all are optimal, the optimization of the transfer orbit parameters is ended; if not, a mission orbit period is selected by means of bias parameter optimization. The application can obtain a complete capture-to-target orbit scheme and efficiently design orbit parameters from the moon capture to the large-ellipse frozen orbit.

The invention relates to a polar region sustainable observation point determination method based on a refraction type solar sail spacecraft, which comprises the following steps of: establishing a light pressure dynamic model of the refraction type solar sail spacecraft and based on a circular restrictive three-body problem (CR3BP) framework; generating an improved kinetic equation containing a refraction light pressure item, solving to obtain a space coordinate set of artificial translation points in the sun-earth system, and calculating an observation half-cone angle between each translation point in the space coordinate set and a target observation point in the polar observation geometric constraint model through coordinate transformation; the method comprises the following steps of: discretizing an earth revolution period, solving a continuity condition of observation coverage to obtain a feasible region meeting full-period polar region continuous observation, performing space matching on a space coordinate set of an artificial translation point and the feasible region, and screening out an optimal observation point position meeting requirements. According to the method, the sustainable observation area of the polar region can be enlarged by about two times, meanwhile, the sustainable observation position is determined based on a space geometry method without problem simplification, and compared with a traditional analytical solution method adopting a simplified problem, the calculation efficiency is high, and the precision is high.

The present invention discloses a method and device for identifying millennium-scale cycles based on bispectral analysis, including: (1) selecting paleoclimate proxy indicators, performing power spectrum analysis after preprocessing, and identifying thickness cycles corresponding to astronomical orbital cycles; (2) filtering the frequencies corresponding to the thickness cycles, selecting the filtered curve to establish a time-depth conversion model to obtain the time-domain sequence of the proxy indicator; (3) performing power spectrum analysis on the time-domain sequence to identify astronomical orbital cycles and millennium-scale cycles; (4) importing the time-domain sequence into the Matlab high-order spectral analysis toolbox, and calculating the non-linear phase coupling relationship between two sets of frequencies through bispectral analysis; (5) exporting the bispectral diagram of the entire time-domain curve in the time domain in step (4), observing and identifying strong response regions with high correlation, and further verifying the existence of the millennium-scale cycle. The present invention can solve the problem of identifying millennium-scale cycles and short-scale climate change cycles.

The application discloses a design method of earth-moon periodic orbit robust maintenance point based on Lyapunov index, and particularly relates to the technical field of space, and establishes an equivalent position-velocity error model, converts the introduced initial velocity error into equivalent position error by using the equivalent position-velocity error model, and calculates the Lyapunov index of the initial velocity error of the earth-moon periodic orbit; according to the Lyapunov index of the initial position error of the earth-moon periodic orbit and the Lyapunov index of the initial velocity error of the earth-moon periodic orbit, a deep neural network is used to establish the convergence and divergence sphere of the initial position error and the initial velocity error of the earth-moon periodic orbit; according to the obtained convergence and divergence sphere of the initial position error and the initial velocity error of the earth-moon periodic orbit, a traversal method is used to design the robust periodic position maintenance point of the position error convergence, the velocity error convergence and the simultaneous convergence of the position and velocity error, and the error convergence axis is searched, so as to provide a reference for the control target of the orbit periodic position maintenance strategy.

The invention discloses an earth-moon periodic orbit robust retention point design method based on a Lyapunov exponent, and particularly relates to the technical field of space, an equivalent position speed error model is established, an introduced initial speed error is converted into an equivalent position error by using the equivalent position speed error model, and the equivalent position error is calculated. Calculating an initial velocity error Lyapunov index of the earth-moon periodic orbit; according to the earth-moon periodic orbit initial position error Lyapunov index and the earth-moon periodic orbit initial speed error Lyapunov index, establishing an earth-moon periodic orbit-oriented initial position error and initial speed error convergence ball by using a deep neural network; and according to the obtained earth-moon-oriented periodic orbit initial position error and initial speed error convergence ball, designing a robust periodic position holding point with position error convergence, speed error convergence and position and speed error simultaneous convergence by using a traversal method, and searching an error convergence axis, thereby providing a reference for a control target of an orbit periodic position holding strategy.

The invention discloses a method for evaluating the access capability of an artificial satellite to ground targets at different latitudes. Belongs to the field of aerospace dynamics and artificial satellite orbits. The method is based on mathematical modeling, main features of an orbit dynamics physical model are extracted, and the access capability of a satellite constellation to a ground specified latitude target is rapidly evaluated with low calculation cost, including the number of access windows and the total access duration in a given time period. According to the method, satellite motion is simplified into an elliptical orbit dominated by two-body gravitational force, and a static annulus model formed by projection of a satellite side-sway range on the earth surface is introduced; by analyzing the geometrical relationship and the time characteristic that a ground target passes through the annulus along with earth rotation, and combining a satellite orbit period and an earth rotation period, an analytical expression of the access probability, the window number and the time length is deduced. The method can effectively estimate the distribution of the access capability along with the latitude, does not need a large amount of numerical simulation, and remarkably improves the calculation efficiency.

The present application relates to a method for eliminating dual-planet ephemeris prediction errors based on post-precision ephemeris. The method comprises: predicting the satellite position for the ephemeris reference time based on the dual-planet ephemeris and the SDP4 model; calculating the satellite position for the ephemeris reference time based on the post-precision ephemeris of the dual-planet ephemeris for the ephemeris reference time and using it as the reference satellite position; obtaining the satellite position prediction error by comparing the difference between the predicted satellite position for the ephemeris reference time and the reference satellite position; and obtaining an error function for the dual-planet ephemeris predicted satellite position by fitting the satellite position prediction error within one satellite orbital period; and performing error elimination on the predicted satellite position for the target date based on the error function to obtain the predicted satellite position after error elimination. This method can eliminate the satellite position prediction error of the dual-planet ephemeris and improve the accuracy of the dual-planet ephemeris for medium- and long-term forecasts.

This invention relates to a satellite attitude determination method, system, and computer-readable medium based on a star sensor. The satellite attitude determination method includes: establishing a solar orientation coordinate system OX based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 Where O represents the satellite's center of mass, X i2 The axis is obtained by the cross product of the solar vector and the geocentric vector, Z. i2 The axis points in the negative direction of the solar vector, Y i2 The axis is X i2 Axis and Z i2 The axis is obtained according to the right-hand rule; the installation matrix of the star sensor is constructed according to the sun-oriented coordinate system; based on the installation matrix, the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector are calculated in the sun-oriented coordinate system; the satellite's attitude is controlled to keep the first angle within a first preset range and the second angle within a second preset range, thereby making the star sensor effective throughout the entire orbital period. This application can prevent the star sensor's field of view from being affected by light.

The invention belongs to the cross technical field of satellite gravity mechanics, geodetics, orbit perturbation, space science and the like. The invention particularly relates to a planet gravity field inversion method. The planet gravity field inversion method comprises the following steps: S1, acquiring shadow region duration of a planet in an orbital period through orbital device load data; s2, acquiring an included angle beta (t) between a sun-planet vector and an orbiter plane through a conical earth shadow model or a cylindrical earth shadow model according to the shadow region duration time tau (t), and further acquiring an ascending node right ascension advancing amount omega (t); s3, solving the ascending node right ascension advancement amount omega (t) to obtain a long-term item and a long-period item; and S4, obtaining a low-order even-number term harmonic coefficient and a low-order odd-number term harmonic coefficient of the gravity field by using a linear perturbation method, an average perturbation method or an analytical method. According to the invention, the low-order gravity field can be measured only by using the sun sensitive load carried by the orbiter without radio measurement, and long-distance measurement can be realized.

The invention relates to the technical field of spaceflight measurement and control and software testing crossing, in particular to a spacecraft measurement, operation and control full-link automatic testing method and system based on cloud native container arrangement. Aiming at the problems that a traditional spacecraft test, operation and control test system depends on physical equipment, the resource utilization rate is low, multi-satellite test capacity expansion is difficult, the fault reproduction capability is insufficient, and an existing Kubernetes orchestrator cannot adapt to a spaceflight scene, the method comprises the steps that a CCSDS spaceflight special protocol is converted into containerized micro-service, and TCP / IP time delay is avoided through shared memory communication; the extended Kubernetes scheduler introduces an orbit period constraint and an orbit attenuation factor to realize space-time perception scheduling; and full-link digital twinning is constructed, a satellite-ground-station collaborative test environment is dynamically constructed, and fault injection and snapshot playback are supported. According to the method and the system, the full-link test efficiency can be remarkably improved, the multi-satellite collaborative test capability is broken through, the reproduction accuracy of the jump fault is improved, and the method and the system can be widely applied to automatic test verification of spacecrafts.

The invention relates to a transfer orbit design method for a HALO orbital vehicle to fly to a given ground point. The method comprises the following steps: finding a moment when a lunar declination value is equal to a given ground point latitude value according to the lunar declination value in a given task time range, and selecting a time point closest to an initial moment as an orbit transfer moment iteration initial value; calculating the flight duration of the orbital transfer point relative to the starting point by using the starting point time and the iteration initial value of the orbital transfer time; calculating normalization time according to the quotient of the flight duration and the HALO orbital period; according to a polynomial fitting coefficient, calculating the increment of the orbital transfer speed for the normalization time; based on a high-precision force model, establishing a mapping relation between the increment of the orbital transfer speed and the iteration initial value at the orbital transfer time through numerical integration; and solving the to-be-solved parameters by adopting a Newton iteration method based on the mapping relation. The method is short in response time, high in calculation result precision and fast in iteration convergence.

The invention discloses a translation point interferometry constellation initial value design method based on phase optimization, and belongs to the field of aerospace. The implementation method comprises the following steps: establishing a dynamic model of a spacecraft in a translational point interferometry constellation; establishing a representation model of the configuration stability index of the translational point interferometry constellation; the method comprises the following steps: setting feature point states, orbital periods and expected terminal states of three spacecrafts of a translational point interferometry constellation, and generating periodic orbits of the three spacecrafts of the translational point interferometry constellation by using a single-step targeting method and numerical integration; establishing a mapping model of an initial phase parameter and a configuration stability index; and constructing an optimization objective function, establishing a translation point interferometry constellation global collaborative optimization model, and performing optimization by using a differential evolution algorithm to obtain a translation point interferometry constellation initial value. The translation point interferometry constellation meeting the long-term stability is obtained through recursion of the initial value, the long-term stability of the translation point interferometry constellation is achieved according to the constellation configuration, and the detection precision of the translation point interferometry constellation is improved.