337 results about "Aerospace technology" patented technology

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The invention relates to the technical field of aerospace, in particular to low-altitude traffic flow airspace-oriented real-time planning, and provides a centimeter-level precision detection network covering a low-altitude airspace by integrating multi-dimensional data sources such as radar electromagnetic feature recognition, ADS-B (Automatic Dependent Surveillance-Broadcast) automatic monitoring, Beidou or GPS (Global Positioning System) space-time reference positioning and the like. The three-dimensional trajectory and motion situation of the aircraft are solved in real time, intelligent reconstruction of an airspace sector and adaptive optimization of a flight corridor are realized by adopting a dynamic programming algorithm driven by reinforcement learning based on the real-time pose data of the aircraft, and a flight conflict prediction model constructed by combining a space-time convolutional neural network is used for predicting the flight conflict. According to the method, potential risks can be pre-judged, an optimal avoidance path can be generated, full-process digital management and control from identity verification to airspace authorization can be realized by constructing an aircraft digital identity authentication system and a dynamic access control mechanism, and modules such as a three-dimensional navigation information service module, a low-altitude digital communication private network module and an intelligent early warning and warning system module are integrated. And the guarantee of full-life-cycle service is provided for the low-altitude aircraft.

The invention discloses a predetermined time expansion state observer generation method and an observation error convergence method, which are applied to the technical field of aerospace, and comprise the following steps: constructing a controller for a second-order nonlinear system according to an all-drive system theory; converting the second-order nonlinear system according to the controller to obtain an all-drive system based on the main controller; wherein the main controller is a controller for controlling the all-wheel-drive system; designing a predetermined time expansion state observer based on the total disturbance and observation error of the all-wheel-drive system; wherein the predetermined time expansion state observer is an observer which limits the upper bound of convergence time of the observer based on a predetermined time parameter. Compared with most existing extended state observers which can only ensure asymptotic convergence and finite time or fixed time convergence, the observer generated based on the predetermined time extended state observer generation method can ensure that the observation error converges within the predetermined time, so that the converged predetermined time is more stable; and the user experience is improved.

The invention provides a calibration performance evaluation system and a calibration performance evaluation method suitable for a single performance index, and relates to the technical field of spaceflight. The method comprises the steps of receiving a remote control instruction, identifying a task intention and triggering an evaluation mode; acquiring a real-time physical state of a satellite, and calculating a dynamic calibration reference value of a single efficiency index based on a physical model; obtaining an actual measurement result of the index probe type measurement, and calculating the deviation between the actual measurement result and a reference value; and combining the task intention and the deviation adjustment evaluation parameter, and outputting a self-calibration calibration evaluation result. According to the method, the problems that dynamic response is lagged and errors are introduced by manual intervention in traditional multi-dimensional evaluation can be solved, rapid and reliable calibration of a single efficiency index in space-based micro-nano satellite networking is achieved, and safe switching of high-value services is guaranteed.

The invention discloses a spatial distributed observation method based on irregular region geometric feature sampling, and belongs to the technical field of aerospace. The implementation method comprises the following steps: establishing a coordinate system for describing the motion of the spacecraft; establishing a spacecraft dynamic model under the coordinate system; six initial orbit elements of a satellite can be obtained based on existing low-orbit constellation orbit data, the spatial position is calculated through the orbit elements, and then position components are converted into latitude and longitude coordinates, namely sub-satellite point coordinates; defining vertex coordinates of the polygon, and constructing a planar polygon by connecting the vertex coordinates end to end; discretizing the internal region of the closed polygon; on the basis of the kinetic model, on the basis of the coordinates of the sub-satellite point and the center coordinates of the discrete sampling circle, calculating the characteristic distance from the sampling point in the observation area to the sub-satellite point, and on the basis of the size relationship between the characteristic distance and the radius of the observation view field, judging whether the irregular area can be successfully observed in a preset time constraint interval; therefore, spatial distributed observation is realized.

The invention discloses a communication link automatic selection and concurrent transmission method under a helicopter limit condition, and relates to the technical field of aerospace. Comprising the following steps: firstly combing airborne communication link information, retaining four types of links including a satellite, an ultra-short wave, a short wave and a data transfer radio, and collecting parameters by means of a bus controller; filtering the data by using a Kalman filtering algorithm, and removing abnormal data through threshold verification; and finally outputting stabilized data. Determining a task scene and an environment scene based on the stabilization data in combination with airborne control screen task selection and environment sensor data; according to the technical key points, a fault precursor monitoring technology is adopted, a standby link pre-activation technology is combined, a seamless switching mechanism is matched, the defect that an existing system depends on manual switching is overcome, the automatic switching time of the link is shortened to be within 0.5 second, the communication interruption rate under the limiting condition is reduced to be lower than 5% from 30% or above, the real-time communication requirements of rescue, patrol and the like are met, and the real-time communication efficiency is improved. Good use prospects are realized.

The invention discloses a multi-spacecraft attitude and orbit coupling all-drive convergence time controller generation method, which is applied to the technical field of aerospace and comprises the following steps: establishing a multi-spacecraft attitude and orbit coupling error system all-drive model based on all-drive characteristics of a tracking error model; wherein the multi-spacecraft attitude and orbit coupling error system all-drive model is a pseudo-linear system model aiming at the stability problem of a multi-spacecraft attitude and orbit coupling error system; aiming at the multi-spacecraft attitude and orbit coupling error system all-drive model, constructing an all-drive convergence time controller based on an all-drive system method, a sliding mode control method and a preset time control method, and adjusting the convergence time of the system based on the all-drive convergence time controller; the all-drive convergence time controller is a controller designed based on a preset time parameter and an adjustable time parameter. According to the method for adjusting the preset time through attitude and orbit coupling, the multiple spacecrafts can maneuver to the expected attitude and orbit within the preset time, formation is completed, and the formation efficiency is improved.

The invention belongs to the technical field of aerospace, discloses a multi-gliding aircraft cooperative formation control method and system, and solves the problems that a multi-gliding aircraft is under-actuated in a high-speed reentry environment, a three-dimensional configuration and collision and obstacle avoidance are not considered in an existing method, and longitudinal errors are not converged. Firstly, a task scene is defined, a three-degree-of-freedom centroid dynamic model is established, and a three-dimensional configuration, collision and obstacle avoidance and control constraints are defined; then, preset time consistency control based on preset performance is designed, and lateral and height formation is achieved; respectively setting task condition correction for the same / different height scenes according to the longitudinal error, and leading a collaborative regulation strategy; setting a collision and obstacle avoidance strategy by using an improved artificial potential field method; and finally, constructing a control framework resolving instruction. According to the method, preset time error convergence, longitudinal configuration maintenance and collision avoidance safety are realized, and the aircraft clustering capability is improved.

The application discloses a spin-nutation collaborative inhibition method and system for a space tumbling target, and belongs to the technical field of aerospace. The method comprises the following steps: acquiring pose data of a space robot, and controlling the space robot to approach the space tumbling target according to the acquired pose data of the space robot; after the space robot approaches the space tumbling target, the space robot measures the space tumbling target, and acquires real-time state data of the space tumbling target by measuring the space tumbling target; a target stress position prediction model of spin-nutation collaborative inhibition is used to simulate and predict based on the acquired real-time state data of the space tumbling target, and a despinning condition required by a despinning tool in the space robot is determined; the space robot adjusts the pose according to the despinning condition, and controls the despinning tool to contact the space tumbling target, and implements contact despinning. The application can efficiently inhibit the nutation of the target in the initial stage of the despinning task, and effectively reduces the difficulty of the subsequent despinning task.

The invention discloses a coil excitation Hall thruster magnetic circuit structure suitable for a krypton working medium, relates to the technical field of aerospace technology and plasma propulsion, and solves the problems of low ionization efficiency and insufficient characteristic length of an existing Hall thruster after krypton is used. A magnetic conductive inner core is arranged at one end of a bottom plate, and a magnetic conductive cover is arranged at the other end; an inner magnetic pole is arranged at the upper end of the magnetic conductive inner core; an inner coil is arranged between the inner magnetic pole and the bottom plate and corresponds to the outer side of the magnetic conductive inner core; an outer coil is arranged between the outer magnetic pole and the bottom plate and corresponds to the inner side of the magnetic conductive cover; a ceramic channel is arranged between the inner coil and the outer coil; the bottom plate comprises an inner bottom plate and an outer bottom plate, the inner bottom plate and the outer bottom plate are coaxially arranged, and a bottom plate gap is formed between the inner bottom plate and the outer bottom plate. The characteristic length of a magnetic field in a discharge channel can be remarkably prolonged, the krypton ionization efficiency is improved, and meanwhile the structure of the thruster is simplified.

The present application belongs to the field of aerospace technology, and discloses a kind of multi-glider aircraft cooperative formation control method and system, solve the problem of underactuated of multi-glider aircraft in high-speed reentry environment, the problem that three-dimensional configuration and collision avoidance and obstacle avoidance are not considered in existing method and longitudinal error does not converge.Firstly, the task scenario is defined, a three-degree-of-freedom mass center dynamics model is established, and three-dimensional configuration, collision avoidance and obstacle avoidance and control constraints are defined;Then, a preset time consistency control based on preset performance is designed to realize lateral and height formation;For longitudinal error, task condition correction and leader-follower cooperative control strategy are set for same / different height scenarios respectively;Improved artificial potential field method is used to set collision avoidance and obstacle avoidance strategy;Finally, a control framework is constructed to calculate the instruction.The present application realizes preset time error convergence, longitudinal configuration maintenance and collision avoidance safety, and improves the cluster capability of aircraft.

The invention discloses a construction method and system for a low-thrust transfer orbit family between earth-moon translation point orbits, and belongs to the technical field of aerospace. The method comprises the following steps: firstly, solving a time optimal transfer orbit based on a continuous low-thrust spacecraft motion equation; then taking the orbit as a seed, performing forward and reverse phase continuation along a departure orbit and an arrival orbit by adopting a self-adaptive step length numerical continuation method, and generating a time optimal transfer orbit family covering different phases; secondly, mapping the time optimal orbit family into an energy optimal orbit family based on control law consistency, and expanding the transfer time range of the energy optimal orbit family through time continuation; and finally, gradually reducing homotopy parameters by using a homotopy method, detecting and processing thrust switching discontinuity in combination with a switching function, and converting the energy optimal orbit family into a fuel optimal transfer orbit family. According to the method, a continuous and complete orbit family is efficiently generated from a single solution, and systematicness and engineering applicability of orbit design are improved.

This invention provides a spacecraft escort control method based on zero-sum differential games, belonging to the field of aerospace technology. The control method includes the following steps: obtaining spacecraft escort problem parameters for both the pursuing and defending satellites; establishing a spacecraft escort model based on zero-sum differential games and inputting initial parameters into the model; transforming the spacecraft escort model into a two-point boundary value problem model; using a genetic algorithm to solve the two-point boundary value problem model to obtain the state trajectory within the game time, thus completing the spacecraft escort control. This spacecraft escort control method can effectively solve the spacecraft escort problem under continuous thrust, and is an effective extension of existing continuous thrust escort problem models and methods. The model employs a differential game method described by a zero-sum cost function, thus simplifying the solution of the difficult-to-solve weighted performance index escort problem and improving the solution efficiency of the spacecraft escort game problem.

The utility model discloses a suspension simulator, which belongs to the technical field of aerospace and comprises a shell, a front joint and a rear joint are arranged at the top of the shell, the shell is connected with an aircraft through the front joint and the rear joint, and a first accommodating bin and a second accommodating bin for accommodating internal equipment are sequentially arranged on the shell. A first cover plate matched with the first containing bin is arranged at the top of the first containing bin, and a second cover plate matched with the second containing bin is arranged at the top of the second containing bin. On the basis that the light weight is met, the structural strength is guaranteed, the whole body is formed through assembly, the size accuracy is considered, and a tester can be supported to complete training tasks such as disassembly, maintenance and testing of a hanging object.

The present application relates to the field of aerospace technology, in particular to a kind of spacecraft autonomous control system, including autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, by setting autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, to realize the full-automatic intelligent navigation of spacecraft in the process of space flight and the autonomous docking of spacecraft and space station, further not need manual intervention, by the autonomous control of spacecraft in the process of space flight, to greatly shorten the working time of astronauts in the small space of spacecraft, improve the work efficiency of astronauts, further solve the technical problems that the control system of existing spacecraft cannot realize automatic control, need manual intervention, to increase the working time of astronauts in the small space of spacecraft, reduce the work efficiency of astronauts.

The application discloses a kind of space target ultra-short arc initial orbit intelligent determination method, belong to aerospace technology field.The application is based on artificial intelligence technology to realize space target ultra-short arc initial orbit determination, there is no strict requirement to complex nonlinear problem, and can develop more complex perturbation condition and initial orbit determination under the operation law of multiple body, coverage is wider;The application generates multiple groups of target orbits randomly and calculates the observation and state quantity of each visible arc segment, constructs the sample data set with ultra-large capacity, which is helpful for neural network to learn the nonlinear relationship between observation and state quantity using rich samples;The application builds and trains long short-term memory network model, determines the optimal network model weight parameter by designing and adjusting hyperparameter, can obtain the fast prediction solution of unknown space target ultra-short arc initial orbit according to observation;The application can meet more space target ultra-short arc initial orbit determination tasks by considering different inputs and outputs.

The invention relates to the technical field of aerospace, in particular to a method for calculating the sea-launch take-off drift distance of a carrier rocket. The method comprises the steps that a launching platform shaking model is established under a launching platform body coordinate system, and a launching platform rolling angle and a launching platform pitching angle are obtained; establishing a carrier rocket vector form dynamical model under a carrier rocket body coordinate system to obtain the attitude angular velocity of the carrier rocket and the displacement of the carrier rocket mass center; obtaining relative displacement of a launch vehicle tail observation point and a launch platform motion reference point under a launch point gravity coordinate system according to the launch platform rolling angle, the launch platform pitching angle, the attitude angular velocity of the launch vehicle and the displacement of the launch vehicle mass center; and according to the relative displacement, obtaining the synthetic drift distance of the carrier rocket tail observation point relative to the launching platform under the body coordinate system of the launching platform. The method can reduce the prediction error of the sea-launch take-off drift distance of the carrier rocket, and meets the sea-launch high safety margin requirement.

The invention provides a spatial domain operation method and device, and relates to the technical field of aerospace. A spatial domain operation method comprises the steps that a space-time 4D occupation plan is generated according to a to-be-executed task of spatial mobile equipment, the space-time 4D occupation plan is sent to a scheduling approval end, and the space-time 4D occupation plan comprises occupation description of space-time galleries, 4D space-time blocks and / or 4D tracks; receiving first approval information from the scheduling approval end; and controlling the space mobile device according to the first approval information and the space-time 4D occupation plan. According to the method, the space domain used by the to-be-executed task is described by using the space-time gallery, the 4D space-time block and / or the 4D track based on the equipment operation end, so that the utilization rate and the safety of the space domain are improved; through a spatial mobile device-device operation end-scheduling approval end three-party interaction mechanism, rapid approval of the space-time 4D occupancy plan is realized, the conflict risk between spatial mobile devices is reduced, and the flexibility is improved.

The application discloses a kind of deep space survey path planning method and system, it is related to aerospace technology field, method includes: S1, ground control center obtains strategy network initial parameter of strategy network and value network initial parameter of value network according to expert experience, and upload to deep space survey vehicle;S2, the perception data obtained by deep space survey vehicle is input into the strategy network of strategy network initial parameter configuration and the value network of value network initial parameter configuration, path planning is carried out, and the path planned is obtained;S3, deep space survey vehicle moves according to the path planned, obtains actual movement data in the process of moving, and strategy network initial parameter and value network initial parameter are updated according to actual movement data.The present application can exert the ability of heaven and earth big system, integrates spacecraft intelligent agent, ground planning system and the intelligence of expert experience.

The invention belongs to the field of aero-engine rotor design, and provides an aero-engine two-dimensional thrust vectoring nozzle heat shield fixing and restraining structure which adopts a high-flexibility space curve notch design to actively control the rigidity of two dimensions of a bolt so as to release thermal deformation and avoid thermal damage to the structure caused by local overlarge thermal stress. And meanwhile, the material characteristics of metal rubber are used for improving the vibration absorption capacity of the aero-engine fixing and restraining structure, the problem that the vibration reduction capacity of an existing connecting structure is insufficient under airflow pulsation excitation is solved, and meanwhile the pulsation frequency band in the engine working condition is widened.

The present application relates to a flexible variable camber wing control system and method driven by a rudder and a piezoelectric hybrid, belongs to the technical field of aerospace, solves the problem of narrow dynamic response band and insufficient high-frequency disturbance suppression capability in the prior art, comprising: step S1, setting a real-time shape control module based on aeroelastic coupling and a multi-modal sensing feedback module; step S2, obtaining the flight state parameters of the aircraft in real time through the flight control, calculating the optimal wing trailing edge camber curve under the current flight state, and converting it into a rudder driving instruction; step S3, the flexible structure of the wing trailing edge is pulled by the rudder to realize the deflection of the rudder surface, and the actual deformation of the wing trailing edge is monitored in real time by the strain gauge to generate an error signal; step S4, providing a driving voltage for the piezoelectric sheet; step S5, the strain of the wing is detected in real time by the strain gauge, and the driving voltage of the piezoelectric sheet is dynamically calculated; step S6, the piezoelectric sheet generates vibration output reverse strain based on the received driving voltage to drive the wing trailing edge to vibrate.

The invention discloses a large-scale constellation multi-type observation task high and low orbit collaborative planning method, and belongs to the technical field of aerospace. The implementation method comprises the steps of defining feature information of each type of tasks, and laying a foundation for evaluating the overall observation benefits of the multiple types of tasks. Designing a weighted overall observation benefit evaluation function which comprises global and local optimization targets and weights thereof; the earliest task completion time is a global optimization standard, and the ship target trajectory coverage and the regional target coverage rate triggered by the type identifier are used as local optimization targets. The constellation completes meta-task time window screening in a distributed mode, the planning process is based on an overall observation benefit evaluation function, observation satellites and time windows are determined through inter-satellite cooperation, and then large-scale constellation multi-type observation task high and low orbit collaborative planning is achieved. The method can deal with the problems of complex constraint and resource coupling caused by multiple types of observation tasks in a large-scale constellation, realizes multi-type task income evaluation at the same time, and has the advantages of high efficiency and high precision.

The invention discloses a satellite orbit simulation method and system based on a quad-rotor unmanned aerial vehicle, and belongs to the field of spaceflight technology and unmanned aerial vehicle systems.The method comprises the following steps that a satellite orbit and speed information are generated based on six elements of the satellite orbit, the satellite orbit is mapped according to a preset scaling, and the speed information of the satellite orbit is obtained; generating a reference orbit of flight of the four-rotor unmanned aerial vehicle; establishing a four-rotor unmanned aerial vehicle kinetic model; controlling the quadrotor unmanned aerial vehicle by using a proportional-integral-differential (PID) control algorithm for the obtained kinetic model; and filtering observation noise and external interference by using extended Kalman filtering (EKF), so that the quadrotor unmanned aerial vehicle can stably track and simulate a satellite orbit in a ground environment. The problems that satellite orbit in-orbit verification is high in cost, large in risk and non-repeatable can be effectively solved, ground rehearsal and advanced layout can be achieved on the strategic level, and the method has important significance on space task deployment and space safety in the future.

The invention belongs to the technical field of aerospace, and discloses a large-bevel-angle honeycomb sandwich structure manufacturing method and a large-bevel-angle honeycomb sandwich structure. The manufacturing method comprises the steps that prepreg is laid on the surfaces of the upper edge and the lower edge of a honeycomb core crude product, and pressure-bearing barrier strips are arranged on the four side edges and cured; removing the pressure-bearing barrier strip, and processing the honeycomb core crude product to form a honeycomb core semi-finished product; filling the hole grids at each side edge of the honeycomb core semi-finished product with a pouring sealant in rows, and curing to form a strip-shaped reinforcing structure; paving prepreg on the surface of the side edge of the honeycomb core semi-finished product, and curing; and paving the adhesive film, the inner skin and the outer skin, and curing. According to the manufacturing method, the problems of hole collapse, deformation and the like of the honeycomb core during production can be avoided, and then the phenomena of wrinkling, sinking and bending of the skin are avoided, so that the whole honeycomb sandwich structure has good appearance and mechanical properties, and formation of the high-thickness large-bevel-angle honeycomb sandwich structure is facilitated.

The utility model discloses a kind of mould of high-temperature-resistant composite material stator blade in aerospace technology field, including bottom plate, the bottom plate is buckled with enclosure, the enclosure is provided with pressing plate, the enclosure is formed with cavity, installation slot is opened in the bottom plate, the installation slot and cavity jointly form installation area, lower mould is arranged in the cavity, the upper mould is buckled on the lower mould, side mould assembly is arranged in the cavity.The utility model is formed by stage pressure and sequential die control forming process: first, main profile such as leaf back, leaf basin is compacted alone, ensure that fiber is evenly laid, avoid distortion wrinkle, then closed cavity by side mould assembly, make resin fully soak fiber in low pressure stage, resin has been gelled when last whole die, high pressure is used for compaction and does not cause migration, to synchronously solve fiber distribution uneven and resin migration problem, guarantee the consistency of high-temperature-resistant blade's compactness and mechanical property.

The present application relates to the field of aerospace technology, specifically to a kind of multi-pulse parallel construction method of gravitational wave detection formation considering timing optimization, each spacecraft of formation simultaneously departs from virtual center, each spacecraft is moved to the preset specified position by applying multiple pulses, so as to form a kind of spacecraft formation that can be used for space gravitational wave detection, by giving the steps of calculating the initial position and velocity of each spacecraft and applying initial control to reach the initial position and velocity, the state quantity of each spacecraft at each time in the formation construction process is calculated, so as to provide reference for the design of satellite formation task in actual engineering application, and the accuracy of the linearized relative motion dynamics model established by the multi-pulse parallel space gravitational wave detection formation is improved by differential correction, the error in the relative motion dynamics model is reduced, the initial state set in advance can be effectively reached, and the working efficiency of formation component is improved.

This application relates to the field of intelligent aerospace technology and discloses a satellite autonomy method driven by embodied intelligent agents, applicable to subsatellites belonging to an in-orbit constellation. It constructs a multi-agent framework comprising multiple autonomous intelligent agents, including at least a resource assessment agent and a resource optimization agent. In response to a resource assessment command regarding the target mission, the method acquires the subsatellite's in-orbit operating parameters. The resource assessment agent is invoked to perform resource assessment based on the in-orbit operating parameters, obtaining the mission resource assessment result, which is then fed back to the in-orbit constellation. If mission feasibility indicates that the subsatellite can execute the target mission, the method acquires a single-satellite mission plan fed back by the in-orbit constellation. The resource optimization agent is invoked to optimize resource allocation based on the single-satellite mission plan, obtaining a mission operation plan. Its beneficial effects are that it improves the intelligence level of subsatellites in remote sensing observation missions and enhances the satellite's in-orbit autonomy.

The invention relates to the technical field of spaceflight, in particular to a rocket engine, namely a real instant PD-6S manned rocket electromagnetic pulse protection chip mainboard, which comprises a mainboard body, a protection assembly, a mounting assembly and a chip assembly, the chip assembly comprises a detection module, a processing module and a generation module, the detection module is fixedly connected with the mainboard body, and the processing module is fixedly connected with the mainboard body. The processing module is fixedly connected with the mainboard body and is connected with the detection module, the generation module is fixedly connected with the mainboard body and is connected with the processing module, electromagnetic pulse signals are timely and accurately detected through the detection module, and then the processing module processes the detected electromagnetic pulse signals and extracts effective feature information. And finally, the generation module generates a reverse pulse signal according to the feature information, thereby effectively resisting interference of electromagnetic pulses on the rocket and ensuring stable operation of the rocket.

The application belongs to the technical field of aerospace, and relates to a distance-keeping game control framework and method for a maneuvering non-cooperative target. The framework comprises a rolling control window, an event perception window arranged in each rolling control window, and an event triggering mechanism arranged in each event perception window. The event perception window predicts an orbit state through a two-body dynamics model, the event triggering mechanism calculates a relative distance between an active satellite and a target satellite at each discrete point according to an orbit prediction result, and judges whether the relative distance constraint is met. If the relative distance constraint is met, an impulse maneuver strategy of the active satellite as a decision output is calculated through a game strategy optimization algorithm. Simulation results show that the application can realize long-time distance keeping, and the success rate reaches 98% in a 12-hour maintenance task, which is increased by 308.3% compared with a traditional algorithm.