52 results about "Orbital maneuver" patented technology

The invention provides a space target orbit maneuver identification method and system based on adaptive LSM, and the method comprises the steps: obtaining the orbit data of a space target, calculating the six orbit elements based on the orbit data, and coding the six orbit elements into a pulse sequence; and inputting the pulse sequence into an improved adaptive LSM for detection to obtain an orbit maneuver behavior identification result of the space target. According to the method provided by the invention, six orbital elements are calculated through the speed and the position of the orbital, then the orbital parameters are input into an improved LSM, and the improved LSM adopts an adaptive threshold neuron as a calculation unit, so that the threshold can be dynamically adjusted according to the current pulse response and attenuated along with time; meanwhile, an improved particle swarm optimization algorithm is combined to carry out optimization setting on an initial value of a neuron, so that the initial performance of the model is improved; and a learnable neuron connection weight based on an attention mechanism is introduced, so that the accuracy and reliability of orbit prediction are improved.

The invention relates to a spacecraft maneuver detection method based on TLE data and kernel density estimation. The invention relates to the technical field of spacecraft maneuver detection. The method comprises the following steps: acquiring and inputting satellite historical TLE data; gaussian filtering denoising is carried out on the input data; based on the de-noised data, propagation forecasting is carried out through an SGP4 model; constructing a joint residual sequence, and carrying out error distribution modeling and sample scoring; determining a maneuvering threshold value, and judging to obtain abnormal data; and aggregating abnormal data and carrying out maneuvering identification. The method can still realize accurate and stable track maneuver automatic detection under the constraint conditions of unstable TLE data, complex error distribution, no external auxiliary data and the like. The method has good adaptability, robustness and interpretability, and is a track anomaly detection technical scheme which is complete in structure, rigorous in logic and capable of achieving engineering landing.

The invention relates to an autonomous measurement and control orbit determination method and system for a high-orbit satellite. The method comprises the following steps: acquiring satellite-borne GNSS broadcast ephemeris data and single-frequency pseudo-range observation data of a high-orbit satellite; based on the GNSS broadcast ephemeris data and the single-frequency pseudo-range observation data, coarse orbit data of the high-orbit satellite are generated through a dynamic pseudo-range orbit determination algorithm; performing outlier elimination preprocessing on the coarse orbit data to obtain orbit data after outlier elimination; performing smooth fitting and equal-interval sampling on the orbit data after outlier elimination to obtain smooth orbit data; determining an orbit determination arc section according to the scheduling time and the orbit maneuvering time; and using the motion equation and the state transition matrix of the high-orbit satellite to process least square iteration in batches to obtain optimal orbit data, and executing orbit forecasting. The method provided by the invention does not depend on information transmission of ground measurement and control resources, realizes on-satellite autonomous orbit determination, reduces the complexity of data fusion calculation, and is small in calculation amount.

Method, electronic device, system, and computer-readable medium embodiments are disclosed. Some embodiments include a signal processing workflow incorporating a graphical user interface for displaying orbital information for satellites and other spacecraft. In some embodiments, a generative adversarial network (GAN) is employed for evaluating satellite orbital positions, for predicting future orbital movements, for detecting orbital maneuvers of a satellite, and for analyzing such maneuvers for potential nefarious intent.

The invention relates to the technical field of constellation satellite maintenance regulation and control, and discloses an intelligent regulation and control system and method for low-orbit Walker constellation configuration maintenance, and the method comprises the steps: obtaining the change trend of an in-plane phase angle and an inter-plane distance of an orbit of a satellite in real time through a high-precision orbit prediction model; inputting the in-plane phase angle and the inter-plane spacing into a pre-trained decision model, judging whether the parameters exceed a preset configuration maintenance threshold, and positioning an unstable satellite; if the parameters exceed the limit, the model outputs the optimal orbit maneuvering parameters of the satellite based on dual-target constraints of minimum fuel consumption and shortest service interruption; a phase maintenance demand and an on-satellite observation task are fused, an optimal control scheme is generated, and constellation configuration is maintained in combination with task planning and implementation; continuously monitoring the satellite constellation configuration state, and maintaining the stable operation of the constellation. According to the invention, automatic maintenance of Walker constellation configuration is realized, and the control precision and efficiency are improved.

The present invention relates to a method for calculating a long-distance rendezvous orbit maneuver based on multiple laps of latitude arguments, comprising: constructing a long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments; performing a perturbation orbit multi-pulse maneuver calculation based on multiple laps of latitude arguments for given maneuvering parameters to obtain the relative position and velocity of the long-distance rendezvous terminal; using the multi-lap maneuvering point latitude arguments and maneuvering pulses as design variables, and the terminal's relative position and velocity as constraints, employing a sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments, thereby obtaining the maneuvering pulse design variable values, the number of laps, and the maneuvering point latitude arguments for the given number of laps. Accordingly, the present invention proposes a long-distance rendezvous orbit maneuver calculation system. The present invention overcomes the convergence instability problem caused by the fixed number of laps in traditional long-distance rendezvous orbit maneuver calculation technology, and achieves stable calculation of long-distance rendezvous orbit maneuvers by performing a cross-lap search using multiple laps of latitude arguments.

The invention discloses a method and a system for planning an orbital task of a maneuvering satellite, and belongs to the field of spaceflight simulation analysis. The method comprises the following steps: creating a planning task and setting basic parameters of a maneuvering satellite; setting basic parameters of a target satellite and a general control center position point; maneuvering analysis is carried out with maneuvering satellite parameters, target satellite parameters and the position point of the master control center as constraints, and a task time window meeting the constraints is obtained; after the task time window is determined, an orbital maneuvering multi-strategy algorithm is called for calculation to obtain a maneuvering mode and a maneuvering parameter list of the maneuvering satellite arriving at the task point; and finally, entering a planning event processing module, sorting events according to time, judging whether impact events exist or not, and outputting a planning task success / failure state. The general control center evaluates the fuel remaining amount in real time based on a fuel formula and supports a user to dynamically formulate and issue an orbital transfer scheme, and integrated planning of maneuvering satellite task window calculation, maneuvering strategy selection and safety check is achieved.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

The invention discloses a satellite orbit forecasting method and device considering maneuvering acceleration, and belongs to the field of communication navigation remote sensing, and the method comprises the steps: carrying out the initial value calculation of a satellite orbit before orbit maneuvering, and obtaining an initial value calculation result; the initial value calculation result comprises six Kepler numbers at an initial moment and a sunlight pressure parameter; taking the orbit maneuvering acceleration as a random pulse, and introducing the random pulse and the initial value resolving result into a motion equation to obtain a simplified motion equation considering the maneuvering acceleration; wherein the starting moment, the ending moment and the numerical value of the orbit maneuvering acceleration are known parameters; the starting and ending moments determine the width of the random pulse; and performing numerical integration on the simplified motion equation under the constraint of a plurality of perturbative force models to obtain a satellite orbit forecasting result subjected to simplified dynamic constraint at any moment. According to the method, the track forecasting precision and the calculation load can be effectively balanced, and efficient and high-precision track forecasting is realized.

The invention provides a low-orbit constellation configuration maintenance method using perturbation force and adjacent position relative station deviation, and belongs to the technical field of low-orbit constellation configuration maintenance and control. The method comprises the following steps: acquiring orbit parameters of each satellite and adjacent satellites thereof in a low-orbit satellite constellation in a space, and calculating a relative drift distance and an allowable drift range of each satellite; on the basis of a low-orbit satellite constellation dynamic model and the relative drift distance, calculating theoretical retention time of each satellite in an allowable drift range, and further determining a constellation control sequence and an optimal control moment of each satellite; and when the satellite reaches the optimal control moment, determining the control quantity of the satellite orbit parameters and the propulsion system parameters of the satellite during orbit maneuver so as to maintain the constellation configuration. According to the method, the relative standing position relation of the adjacent satellites serves as control input, autonomous stable maintenance of the constellation relative configuration is achieved by controlling the orbit height and the inclination angle of the satellites, and unnecessary maintenance control and fuel consumption are avoided.

The embodiment of the present application discloses a kind of super low orbit satellite orbit height autonomous maintenance method.In a specific embodiment, the method comprises: the position and speed of satellite-borne GNSS receiver output as input, calculate the actual value of satellite's intersection period when satellite is in ascending node or descending node, according to the relationship between intersection period actual value and intersection period target value, execute orbit maneuver, to maintain orbit height.A kind of super low orbit satellite orbit height autonomous maintenance method of the present application, with the position and speed of satellite-borne GNSS receiver output as input, is realized by combination of electric propulsion and chemical propulsion, can comprehensively utilize the advantages of less electric propulsion working medium consumption and large chemical propulsion thrust, to realize orbit height on-orbit autonomous orbit maintenance.

The invention provides a common-view time transfer method based on multiple GEO satellites, and the method comprises the steps: selecting a plurality of GEO satellites in the visual range of two tracking stations, carrying out the common-view time transfer through employing each GEO satellite, and obtaining the common-view time transfer result of each GEO satellite; preprocessing the common-view time transfer result of each GEO satellite to obtain a preprocessed common-view time transfer result; traversing each GEO satellite, analyzing the stability of the link of the current GEO satellite, and meanwhile, referring to a broadcast ephemeris to judge whether the current GEO satellite is in an orbit maneuvering state; if not in the track maneuvering state, allocating weights to the track maneuvering state based on a static self-adaptive weight allocation method, otherwise, allocating weights to the track maneuvering state based on a dynamic self-adaptive weight allocation method; and based on the weight distributed for each GEO satellite, performing weighted average processing on each preprocessed common-view time transfer result, and finally obtaining a common-view time transfer result of the multiple GEO satellites.

The application discloses a low-fuel-consumption low-orbit satellite orbit recovery method, and belongs to the technical field of satellite orbit recovery.The application aims at the problem that the existing orbit maneuver method needs frequent orbit adjustment, which is not conducive to the long-term operation of the satellite in orbit with low fuel consumption.The application comprises the following steps: establishing an orbit dynamics equation considering the earth oblateness perturbation; using a semi-analytical method to solve the initial orbit elements of a low-fuel-consumption transition orbit according to the current orbit and the target orbit of the satellite; simultaneously, combining the orbit dynamics equation, calculating the drift time required for the satellite to drift on the transition orbit by using the earth oblateness perturbation without fuel consumption and the final orbit elements of the transition orbit at the end of the transition; and based on the Pontryagin maximum principle, calculating the optimal control rate sequence of the thruster fuel of the satellite from the current orbit to the transition orbit and from the transition orbit to the target orbit, and controlling the thruster to realize the orbit recovery of the satellite.The application is used for satellite orbit recovery.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The invention provides an angle-measurement-only multi-pulse self-adaptive closed-loop rendezvous guidance method based on second-order cone programming, and belongs to the relative navigation and rendezvous guidance technology of spacecrafts. According to the method, firstly, an angle-only-measurement relative navigation dynamical model and a measurement model based on a CW equation are constructed under LVLH coordinates, system state observability analysis is carried out, the orbit maneuvering condition enabling the state to be observable is obtained, and the condition is used for constructing an optimal guidance index. Secondly, analyzing and modeling the rendezvous guidance error source, and deducing to obtain a relative orbit guidance error covariance model through error covariance analysis; thirdly, on the basis of a second-order cone programming method, guidance fuel consumption and system state observability conditions are fused, and a multi-target weighted cost function of optimal rendezvous guidance is jointly constructed; meanwhile, based on the established guidance error covariance model, a cost function weight index system with a self-adaptive adjustment function is designed, and real-time self-adaptive adjustment of a guidance strategy along with change of system performance is achieved. The method solves the problems of high fuel consumption, poor flexibility and the like caused by separate design of the relative navigation and the rendezvous guidance model in the traditional only-sight-line measurement rendezvous guidance method, and has the ability of adaptively adjusting and optimizing the rendezvous fuel consumption and the relative navigation system observability.

The invention provides a satellite launching / recycling system capable of moving back and forth freely and being loaded on orbit, and relates to the technical field of spaceflight. The system comprises a spacecraft which can be in on-orbit butt joint with a preset space station; the spacecraft comprises a three-section cabin body which comprises a front-section cabin body, a middle-section cabin body and a rear-section cabin body; a head fairing capable of being opened and closed is arranged at the end part of the front section cabin; a butt joint opening is formed in the head fairing. The top facing part of the middle cabin body is configured to be a cabin door capable of being opened and closed; an expandable folding solar wing and at least one satellite parking device are arranged on the inner side of the cabin door; a satellite parking point is formed after the cabin door is opened; a loading space used for loading a satellite to be launched and / or a recycled satellite, a mechanical arm device capable of grabbing the satellite and a satellite on-orbit launching device are arranged in the middle cabin body. The spacecraft is used as a carrier for satellite launching / recycling and / or a carrier for implementing orbital maneuver and returning to the ground in orbit. Satellite launching and recycling can be achieved with low cost and high efficiency.

This invention relates to a ground flight control and operation support system for a very low Earth orbit (UEO) satellite constellation. It incorporates predicted space environment information, satellite attitude change information, and orbit control information into the high-precision satellite orbit prediction process, effectively ensuring higher orbit prediction accuracy for the UEO constellation and maximizing the assistance to ground control stations during satellite transits in capturing and tracking the satellite. The satellite attitude and orbit control system simulation module performs real-time simulations of the impact of attitude control and orbit control on orbital maneuvers, thereby achieving higher orbit control accuracy and providing better assurance for the precise execution of missions by the UEO constellation. Automatic scheduling of Earth remote sensing and data transmission tasks, along with ground system fault identification and handling, and collision warning, provide more intelligent support for the operation of the UEO constellation.

The present invention provides a supplementary control method and device for the case of an orbital control thrust deficiency during deep space maneuvers, comprising the following steps: obtaining relevant parameters of the satellite and orbital control during normal orbital maneuvers, and determining whether an insufficient thrust failure occurs during orbital control; calculating the engine power-on duration required to achieve a predetermined speed increment under different thrust efficiencies, and determining the lower limit of thrust efficiency where supplementary control is not required; calculating the speed increment and power-on duration of the supplementary control based on the supplementary control interval; and determining whether the orbital control parameters of the supplementary control meet the thruster parameter requirements, and if so, completing the predetermined orbital control target. The present invention can promptly perform supplementary control when an engine thrust deficiency occurs during deep space maneuver orbital control, completing the orbital control action and reaching the predetermined target.

A reusable launch vehicle comprising auxiliary tanks for non-ascent burns, such as orbital maneuver burns, reorientation burns, deorbit burns, or landing burns, and / or power generation. The auxiliary tanks may provide propellants for an orbital maneuvering system, an orbital reaction control system, a reentry reaction control system, and / or a power generation system. The tanks are configured to maintain cryogenic properties of auxiliary fuel and auxiliary oxidizer for the duration of the mission. An auxiliary fuel tank, a main fuel tank, an auxiliary oxidizer tank, and a main oxidizer tank may have various relative positions with respect to each other within the launch vehicle. The auxiliary fuel tank and auxiliary oxidizer tank may comprise insulation to further insulate the cryogenic fluids.

The application provides a multi-GEO satellite-based common view time transfer method, which comprises the following steps: selecting multiple GEO satellites in the visible range of two tracking stations, respectively using each GEO satellite to perform common view time transfer, and obtaining common view time transfer results of the GEO satellites; preprocessing the common view time transfer results of the GEO satellites to obtain preprocessed common view time transfer results; traversing the GEO satellites, analyzing the stability of the link of the current GEO satellite, and simultaneously referring to broadcast ephemeris to determine whether the current GEO satellite is in an orbit maneuvering state; if not, assigning a weight to the GEO satellite based on a static adaptive weight distribution method, or assigning a weight to the GEO satellite based on a dynamic adaptive weight distribution method; and based on the weights assigned to the GEO satellites, performing weighted average processing on the preprocessed common view time transfer results to finally obtain multi-GEO satellite common view time transfer results.

This invention discloses a trajectory planning method for adjusting the visibility time and uncertainty of a spacecraft, relating to the field of spacecraft technology. The method includes: establishing a spacecraft dynamics model considering perturbations and a geometric visibility constraint model for a situational awareness device; defining a vertex detection angle as an intermediate variable relating the spacecraft's visibility time and state uncertainty; solving for the vertex detection angle of the spacecraft relative to the situational awareness device based on latitudinal argument matching and Newton's iteration method; constructing an optimization problem with the spacecraft's pulse velocity increment as the optimization variable and maximizing or minimizing the square of the vertex detection angle as the optimization objective, and numerically solving it using a sequential quadratic programming algorithm to obtain the optimal pulse velocity increment; and performing spacecraft orbital maneuvers based on the optimal pulse velocity increment. This invention can establish the relationship between the orbital maneuver vector and the detected characteristic indicators, enabling rapid optimization of the spacecraft maneuver vector.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The present invention discloses a satellite orbit prediction method and device that considers maneuvering acceleration, belonging to the field of communication, navigation, and remote sensing. The method comprises: performing an initial value calculation on the satellite orbit before the orbital maneuver to obtain an initial value calculation result; the initial value solution result includes six Kepler roots and a solar pressure parameter at the initial moment; introducing the orbital maneuvering acceleration as a random pulse, along with the initial value solution result, into the equation of motion to obtain a simplified equation of motion that considers the maneuvering acceleration; wherein the start and end times and values ​​of the orbital maneuvering acceleration are known parameters; the start and end times determine the width of the random pulse; and numerically integrating the simplified equation of motion under the constraints of multiple perturbation force models to obtain a satellite orbit prediction result under simplified dynamic constraints at any time. The present invention can effectively balance orbit prediction accuracy and computational load, achieving efficient and high-precision orbit prediction.

The application discloses an orbit maneuver compensation method based on constant acceleration and zero-sum speed pulse constraint, relates to the technical field of satellite orbit calculation, and comprises the following steps: acquiring GNSS observation data, satellite attitude data, ephemeris data and external source data, and performing orbit fitting to obtain a rough dynamic orbit; performing cycle slip detection and gross error elimination on the GNSS observation data to obtain clean observation values required for subsequent parameter estimation; using the clean observation values, satellite state, state transition matrix and parameter sensitivity matrix, adding a zero-sum speed pulse constraint equation to an original observation error equation to construct an augmented observation error equation; iteratively performing parameter estimation and orbit integration until convergence is achieved; and generating a high-precision dynamic orbit. The orbit maneuver compensation method based on constant acceleration and zero-sum speed pulse constraint improves the consistency between the dynamic orbit and the kinematic orbit and the KBR checking precision of the dynamic baseline.

The application provides a BDS satellite space signal availability evaluation method, system and electronic equipment, and the technical points are as follows: collecting all broadcast ephemeris data broadcasted in a limited time period of a BDS constellation; establishing a two-step method to detect orbit maneuver of IGSO satellites and GEO satellites of the BDS constellation, and calculating average fault interval time and average fault repair time of the orbit maneuver of the IGSO satellites and the GEO satellites; detecting space signal anomaly of BDS satellites, and calculating average fault interval time and average fault repair time of the space signal anomaly of the BDS satellites for IGSO satellites, GEO satellites and MEO satellites of the BDS constellation; and comprehensively establishing an availability comprehensive evaluation model of the space signal of the BDS satellites to evaluate and calculate the availability performance of the space signal of the BDS satellites. The application greatly improves the accuracy and reliability of the orbit maneuver detection of the BDS satellites, and provides a new method for BDS satellite space signal availability evaluation from the perspective of actual measurement data.

The application discloses a design method of a high-mobility stealth satellite communication system and relates to the technical field of satellite communication. The application comprises the following steps: step a: system silence and state presetting: a control center makes a satellite of a space segment enter an electromagnetic silence state, a satellite platform starts surface thermal management and retracts a protruding structure, and an orbit maneuver with no fixed period is planned and executed. The application improves the concealment and survivability of the satellite platform in orbit. By making the satellite enter a silence and survival state and implementing thermal management and shape control, combined with the orbit maneuver with no fixed period, the physical and orbit characteristics of the satellite are difficult to be observed and predicted for a long time, traditional detection and tracking means based on a fixed shape and a periodic orbit are invalid, and exposure risk of space assets is controlled. Meanwhile, communication adopts a "zero handshake" burst access mechanism based on one-time instructions, and open signaling interaction in a link establishment process is completely eliminated.

Method for managing a satellite receiving CDM messages relating to a secondary object (2) that could collide with the satellite at a TCA date. At each orbit, an autonomous orbit control module establishes a safe current plan on board, defining station-keeping maneuvers (14, 15). A preliminary filtering of the received CDMs is performed on board based on geometric and / or temporal criteria. For each selected risky CDM, a collision risk level is estimated on board based on onboard navigation data propagated at the TCA date of said risky CDM. In the event of a confirmed risk, an ACA collision risk management module develops a new plan on board by removing or replacing at least one maneuver from the safe current plan with a maneuver (16, 17) that allows both station-keeping within a mission window (301-302) and avoidance. Figure for the abstract: Fig. 4

The invention relates to the technical field of satellite attitude control and orbital maneuver, in particular to a Brayton cycle-based satellite composite propulsion system and method. The composite propulsion system comprises a propellant storage unit which comprises a gaseous oxygen storage tank and a gaseous hydrogen storage tank; the turbocharging unit comprises an oxygen first-stage gas compressor and an oxygen second-stage gas compressor which are coaxially connected in series, a hydrogen first-stage gas compressor and a hydrogen second-stage gas compressor which are coaxially connected in series, and a turbine; the oxygen first-stage gas compressor, the oxygen second-stage gas compressor, the hydrogen first-stage gas compressor and the hydrogen second-stage gas compressor are coaxially connected with the turbine; an oxidant inlet of the fuel gas generator is communicated with an outlet of the oxygen secondary gas compressor, a fuel inlet of the fuel gas generator is communicated with an outlet of the hydrogen secondary gas compressor, and a fuel gas outlet of the fuel gas generator is communicated with an inlet of the turbine. A two-stage coaxial series-connection gas compressor structure is adopted, exhaust gas generated after a turbine does work is still kept to be higher than the pressure of a main thrust chamber, the exhaust gas can smoothly flow back to the main thrust chamber for afterburning, and full utilization of energy in a system is achieved.

The invention belongs to the technical field of satellite orbit maneuver detection, and particularly discloses an orbit maneuver detection method based on kinematic orbit epoch difference. According to the method, the kinematic orbit data is used for solving the orbit machine power through a non-difference method, so that the dependence on a prior dynamic model is eliminated, and a high-precision satellite-borne GNSS kinematic orbit product is directly utilized. The kinematic orbit is a pure geometric solution obtained through PPP (Precision Point Positioning), and is not smoothed through any dynamics, so that the position jump of the satellite during maneuvering is truly recorded. The method is simple and reliable, kinematics orbit data are easy to obtain, extra deviation products do not need to be provided externally, the algorithm is simple and easy to implement, and real-time application is convenient. In addition, the method also fills up the data blank during the shutdown period of the accelerometer, and can provide purer background field information for the inversion of the earth gravitational field.

The invention discloses a design method of a high-maneuverability stealth satellite communication system, and relates to the technical field of satellite communication. The method comprises the following steps: step a, system silent latent and state presetting: a control center enables a satellite in a space section to enter an electromagnetic silent state, a satellite platform starts surface thermal management and retracts a protruding structure, and meanwhile, orbital maneuver without a fixed period is planned and executed; according to the method, the concealment and survivability of the satellite platform existing in orbit are improved, the satellite enters the silent latent state, thermal management and appearance control are implemented, and in combination with orbit maneuver without a fixed period, the physical and orbit characteristics of the satellite are difficult to be stably observed and predicted for a long time; a traditional detection tracking means based on a fixed shape and a periodic track loses efficacy, the exposure risk of space assets is controlled, meanwhile, communication adopts a'zero handshake 'burst access mechanism based on a one-time instruction, and public signaling interaction in the link establishment process is thoroughly eliminated.