1038 results about "Spaceflight" patented technology

The invention provides a PPO-based spacecraft attitude and orbit integrated game control method and device, and the method comprises the steps: building a spacecraft attitude and orbit integrated model, and carrying out the modeling of the interaction process of a tracking spacecraft and a target spacecraft as a non-cooperative zero-sum game; setting a multi-agent deep reinforcement learning framework, wherein the tracking spacecraft and the target spacecraft are respectively configured with an independent strategy network and an independent value network; training is carried out in a double-agent PPO mode, control instructions of orbit thrust and attitude torque are generated through the strategy network, and the state value is evaluated through the value network; and controlling and guiding the behavior of the spacecraft according to the trained strategy. According to the method, the autonomous decision-making capability and cooperative control performance of the spacecraft in complex tasks are effectively improved, and the method is suitable for high-precision aerospace tasks such as interception of non-cooperative targets, autonomous trajectory planning and attitude control.

A cabin virtual docking method based on actually measured point cloud data belongs to the technical field of spaceflight docking, and comprises the following steps: acquiring original point cloud data of a cabin structural member; inputting the initial point set into a feature extraction network; the key features comprise decoding features and coding features, the coding features are obtained through the feature processing block, and the decoding features are obtained through the feature recovery block; inputting the extracted key features into a feature segmentation network, segmenting the key features by the feature segmentation network according to categories, and performing parameter fitting on the segmented key features; optimizing matching indexes of the key characteristics after parameter fitting to obtain pose adjustment amounts of the upper cabin section and the lower cabin section so as to realize alignment of butt joint end surfaces and axis superposition of positioning pin holes; and virtual docking between the cabin sections is carried out by using the pose adjustment amount. According to the method, the cabin surface data are directly obtained through measurement, error accumulation is avoided, accurate calculation of the initial relative poses of the two cabins is achieved, and an efficient and reliable solution is provided for high-precision butt joint.

The invention relates to the technical field of spaceflight material forming, in particular to a tension control winding process of a spaceflight carbon fiber composite launch canister, which comprises pretreatment, winding forming, curing and post-treatment. The method comprises the following steps: S1, coating a release agent and a modified coating after processing a core mold, carrying out plasma treatment and eliminating stress; the carbon fiber filaments are soaked in resin containing nanometer components, and are pre-stretched after a hot melting reaction; s2, closed-loop tension control is conducted, winding is conducted according to segmented angles, and modified resin is coated between layers to be overlapped in a crossed mode; s3, stepwise heating and curing, dynamically regulating the pressure and introducing inert gas; and S4, polishing after hydraulic demolding, carrying out sand blasting, carrying out transition layer and ceramic coating, and carrying out multi-dimensional detection. According to the process, the mechanical property and environmental adaptability of the launch canister are improved, the process precision is high, defects are reduced, the forming time is shortened, the rework rate and cost are reduced, the spaceflight requirement is met, and the process is suitable for batch production.

The invention discloses an architecture system and method for large-scale task allocation, and belongs to the field of spaceflight measurement and control, and the system comprises a large-scale hardware device which is used for deploying processing units with different performances and various heterogeneous physical communication link resources; the hardware resource dynamic modeling and visualization framework is used for realizing a dynamic modeling system based on multi-source heterogeneous hardware topology perception; the resource topology awareness database captures and synchronizes system deployment changes by maintaining the dynamic topological relation of hardware resources, presents the life cycle state of the hardware resources, and supports system-level resource situation visualization and decision making; a versioning management engine of component arrangement is applied, a multi-version executable unit is packaged based on an atomization component, and an application instance is completed through a combined modeling component; and deploying a scheduling decision system for deploying a scheduling decision. The method can adapt to task deployment requirements of a current large-scale spaceflight measurement and control system.

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

The invention relates to the technical field of aerospace manufacturing and intelligent welding control, and solves the technical problems of insufficient single-mode perception, limited quality prediction precision, lagging parameter regulation and control and weak model adaptive capacity in traditional welding process monitoring. In particular to a spaceflight structure friction stir welding multi-modal sensing and intelligent quality control method, which combines multi-modal real-time sensing and reinforcement learning by introducing a body intelligent body thought, so that a system can autonomously understand a welding state, predict a quality trend and actively adjust process parameters in a welding process; and adaptive optimization control with quality as constraint is realized. According to the method, the limitation of traditional off-line modeling and manual adjustment is broken through, an intelligent control framework with self-learning and self-evolution capacity is provided for spaceflight-level friction stir welding, and a new technical approach is provided for achieving high-consistency and high-reliability welding seams. The intelligent level of friction stir welding of the spaceflight structural part and the welding seam consistency level are remarkably improved.

The invention relates to a three-dimensional microgravity simulation system, and belongs to the field of spaceflight mechanism ground tests. According to the three-dimensional microgravity simulation system, ground six-degree-of-freedom microgravity simulation can be realized, the aircraft simulator and accessories thereof have a five-degree-of-freedom motion state on an air floating plane through the five-degree-of-freedom air floating platform, and the aircraft simulator can have a three-dimensional space six-degree-of-freedom motion state through the magnetic air suspension assembly. In the process of a ground microgravity experiment, the gravity of the movement of the aircraft and accessories thereof is unloaded by the magnetic suspension assembly, so that the movement interference of the aircraft simulator and the microgravity simulation tool is avoided, and the problem of dynamic high-precision gravity unloading is solved.

The invention discloses a multi-satellite cooperative operation method based on measurement and control link optimization, and relates to the technical field of spaceflight measurement and control and satellite communication optimization, and the method comprises the steps: obtaining the orbital elements of all satellites in a target constellation and the geographic coordinates of ground measurement and control stations, calculating the visible time windows of all satellites and all measurement and control stations, and generating a visible window matrix; constructing a link quality evaluation model to obtain a link quality index; distributing measurement and control time slots by using a weighted optimization algorithm, and generating a time slot distribution matrix; issuing a measurement and control instruction to each satellite and each measurement and control station, acquiring a link quality measured value, and comparing the link quality measured value with a link quality index to obtain a comprehensive link quality deviation value; and when the comprehensive link quality deviation value exceeds a preset threshold value, marking unexecuted measurement and control time slots in the time slot distribution matrix, and redistributing the unexecuted measurement and control time slots to generate an updated time slot distribution matrix, thereby realizing self-adaptive cooperative operation of the multi-satellite measurement and control link. According to the invention, dynamic optimal configuration of multi-satellite measurement and control resources is realized.

The invention provides a satellite power supply system topology architecture and a maximum power point tracking control method thereof, and aims to improve the energy utilization efficiency of a solar cell array of a spinning body spacecraft under a complex illumination condition. The architecture adopts a combined topology architecture of a preceding-stage sequence switch maximum power point shunt regulation (S3MPR) circuit and a backward-stage four-switch Buck-Boost (FSBB) converter, and is suitable for a special operation environment of a spinning body spacecraft power supply system. The system adopts a bus full adjustment mode, enables the DC bus voltage to be stabilized near the maximum power point voltage of the photovoltaic array through a double closed-loop control strategy, and carries out real-time shunting adjustment on an S3MPR circuit in combination with a multi-peak maximum power point tracking (MPPT) algorithm, thereby ensuring the efficient operation of the photovoltaic array. According to the invention, high-efficiency energy tracking and stable charging can be realized under a complex illumination condition, and the reliability and the energy utilization rate of a spinning body spaceflight power supply system are remarkably improved.

The invention relates to a satellite-borne electronic hardware design guarantee method and system. The method comprises the following steps: determining a satellite-borne hardware guarantee level (SHL) of satellite-borne electronic hardware; integrating a fault mode injection FM I into a design process of satellite-borne electronic hardware, and adjusting the execution depth of the FM I based on the SHL level; and providing a process cutting engine, and realizing adaptive process cutting based on a DO-254 standard according to the SHL level and in combination with the execution depth of the FM I. The system comprises an SHL evaluation module; an FM I management module; a space environment effect knowledge base; and a process cutting engine. According to the method, the problem that the DO-254 standard is insufficient in adaptability in the aerospace field is solved, the reliability design and verification efficiency of satellite-borne electronic hardware is remarkably improved, and the method is suitable for aerospace tasks such as various satellites and deep space probes.

The invention discloses a method and equipment for detecting a getter of a ring laser based on machine vision, and belongs to the technical field of laser gyroscope detection. According to the method, the thin film image of the getter is collected, the consumption condition of the getter is recognized by combining the color fading property of the getter after the getter reacts with the impurity gas, meanwhile, the getter parameter and the activation parameter of the ring laser are calculated according to the working condition data of the ring laser during standing and working, and the residual gas capacity of the getter is predicted based on the activation parameter. And combining the suction parameters to fit a suction characteristic function of the ring laser, and further predicting the remaining working time of the getter. According to the method, products with unqualified getter allowance can be effectively screened out, and the application reliability of the ring laser in the fields of aviation, spaceflight and the like is improved.

The invention relates to a system calling ptrace-based operating system micro-service security protection system and method, which are characterized in that the running state of micro-service is monitored in real time through calling of a ptrace system in a user state, and dynamic integrity verification is realized in combination with an SM3 hash algorithm. The method comprises the following steps: calculating an SM3 hash value of a code segment of a micro-service when the micro-service is started, storing the SM3 hash value as a static reference, periodically reading memory data of a target process through ptrace in a running process, recalculating a dynamic hash value, and comparing the dynamic hash value with a reference value; if it is detected that the hash values are inconsistent, it is judged that code injection attacks exist, and a defense mechanism is triggered. According to the method, the system stability risk caused by kernel module injection is avoided, meanwhile, real-time interception of tampering attacks during operation is achieved through high-frequency dynamic comparison, the lightweight design enables the method to be suitable for high-reliability scenes such as spaceflight with limited resources, the integrity of a micro-service code segment can be guaranteed, low performance loss of the system can be maintained, and the system reliability is improved. The balance blank between dynamic defense and system stability in the prior art is effectively filled.

The invention provides a vacuum laser welding method and system for spaceflight ultrahigh-strength steel, and the method comprises the following steps: a pretreatment step: pretreating the ultrahigh-strength steel, and assembling and fixing the ultrahigh-strength steel; a welding parameter setting step: setting welding parameters of a welding gun; the welding parameters comprise a welding angle, a welding speed and a welding path; a vacuum extraction step: extracting air in the vacuum cabin until the vacuum degree reaches a preset value; and a welding step: welding the ultrahigh-strength steel in the vacuum cabin based on the set welding parameters. According to the vacuum laser welding method, laser welding is carried out in the vacuum negative pressure environment, the plasma shielding effect is effectively restrained, the laser energy utilization rate and the welding penetration depth are remarkably increased, meanwhile, the defects such as air holes and cracks are avoided, and the vacuum laser welding method has the beneficial effects of being high in welding efficiency, small in heat input and excellent in weld joint mechanical property.

The invention discloses a space target monitoring constellation configuration optimization method and system considering complex constraints, and belongs to the technical field of aerospace mission design and optimization. The method comprises the following steps: firstly, constructing a discrete orbit modal library meeting regression orbit dynamic resonance conditions; constructing generalized hybrid design variables consisting of integer indexes and real number fine tuning quantities on the basis of the library; establishing an optimization model taking the space volume coverage rate, the maximum revisit time and the system construction cost as targets and taking illumination, dynamics and maneuvering capability as constraints; performing rapid analysis and coarse screening on invalid links through track geometry; and finally, performing global optimization by using an improved non-dominated sorting genetic algorithm to obtain a Pareto optimal constellation configuration set. According to the method, accurate evaluation of the three-dimensional space target coverage efficiency is realized, the physical feasibility and long-term stability of a configuration scheme are ensured, and a high-dimensional multi-target optimization problem is efficiently solved.

The invention provides a lunar rover polar region path planning method and device based on dynamic illumination condition analysis, and relates to the technical field of spaceflight measurement and control. The method comprises the following steps: generating a forward and reverse illumination safety corridor according to an illumination forecast file and preset information; determining a normalized light energy availability index parameter according to the forward and reverse illumination safety corridor and the illumination forecast file, and calculating a normalized light energy availability index according to the normalized light energy availability index parameter; determining shadow coverage cost according to the normalized light energy availability index, and superposing the shadow coverage cost to the forward and reverse illumination safety corridor to obtain an illumination cost graph; and planning the polar region path of the lunar rover through the cost map according to the illumination cost map and the terrain obtained in advance. The device executes the method. According to the method and the device provided by the embodiment of the invention, the global path planning capability of the lunar rover in extreme terrains, dynamic illumination and large-range unknown regions is remarkably improved.

The utility model discloses a spaceflight anti-shaking plate part forming device, which relates to the field of storage box anti-shaking plates, and comprises an upper template and a lower template which are oppositely arranged, the upper template is provided with a material pressing assembly, and the material pressing assembly is used for pressing a non-bent part of a part to be processed between the upper template and the lower template. A second forming block is fixedly arranged on the lower die plate, a first forming block matched with the second forming block is fixedly arranged on the upper die plate, and the first forming block and the second forming block are used for bending a to-be-bent part of a part into a specified size. And the surface quality of the product is effectively improved. And meanwhile, by means of the first forming block and the second forming block, forming of the bent part of the product is achieved.

The invention relates to a large-bearing memory alloy driven unlocking separation nut mechanism, and belongs to the technical field of spaceflight structures. The memory alloy separation nut comprises a shell assembly, a split nut assembly, a bearing force transmission assembly, a memory alloy driving assembly and a nut separation assembly, a modular design concept is adopted, configuration design is carried out on the memory alloy separation nut according to three functional parts of driving, transmission and execution, the bearing efficiency and the unlocking reliability are improved in a targeted mode, and the memory alloy separation nut is suitable for being used for a large-scale industrial production. The unlocking impact is reduced, and the separation nut is more convenient to use.

The invention relates to the technical field of thermal control of commercial spaceflight unfolding mechanisms, in particular to a deformable liquid metal thermal control commercial spaceflight flexible unfolding mechanism which comprises a flexible unfolding mechanism body and a flexible fluid circulation pipeline integrated inside or on the surface of the flexible unfolding mechanism body and synchronously folded and unfolded along with the flexible unfolding mechanism body. The fluid driving and heat exchange unit is arranged in the non-flexible area of the spacecraft platform or the body and is connected with the flexible fluid circulation pipeline to form a closed circulation loop; wherein the flexible fluid circulation pipeline is filled with a liquid metal working medium. The liquid metal working medium has high thermal conductivity and flow characteristics, and is matched with the vein-shaped / grid-shaped microtube array, so that the temperature of each region of the flexible unfolding mechanism can be quickly balanced, and structural stress deformation caused by temperature gradient is effectively eliminated; and the adaptive capacity of the mechanism in extreme space environments such as high and low temperature alternation and intense radiation is greatly improved.

The invention belongs to the technical field of spaceflight electronic integrated circuits, and particularly relates to a satellite load FPGA configuration data error correction method and system capable of resisting single event upset, in particular to a bit-symbol double-layer error correction method of RS decoding operator error correction based on finite field algebra cascading on the basis of two-out-of-three majority voting logic. Single-particle independent upset and multi-copy common-mode failure can be corrected at the same time, and the problem of traditional TMR error locking is solved. Meanwhile, when configuration data are written into a nonvolatile memory, a cross-sector space staggered and dispersed storage mechanism is introduced, symbols in the same RS code word are separated as far as possible in physical positions, and therefore aggregated physical damage is converted into dispersed symbol-level errors.

The utility model relates to the technical field of pneumatic heatproof ground test equipment for aerospace crafts, in particular to a conical spray pipe which comprises an inner shell, a transition shell layer and an outer shell, the transition shell layer is sleeved on the outer side of the inner shell, and the outer shell is sleeved on the outer side of the transition shell layer. The inner shell is divided into three sections in the axial direction, and a plurality of reinforcing ribs are evenly arranged on the outer side conical surfaces of the first section, the second section and the third section at intervals in the circumferential direction. The number of the reinforcing ribs at the second section is smaller than that of the reinforcing ribs at the first section and the third section, and a cooling liquid channel is formed between every two adjacent reinforcing ribs in the circumferential direction. According to the conical spray pipe, the reinforcing ribs are arranged in the axial direction of the spray pipe in a segmented mode, the number of the reinforcing ribs in the circumferential direction of the throat is smaller than that of the reinforcing ribs arranged on the contraction section and the expansion section, and the reinforcing ribs serve as structural reinforcing structures of the inner shell and serve as wall bodies of the cooling channels at the same time, so that passing of cooling liquid at the throat is guaranteed; the cooling effect at the throat is improved, and sufficient strength of the inner wall of the spray pipe is ensured.

The invention discloses a low-temperature upper-stage power system, and relates to the field of spaceflight. Comprising a gas cylinder, a pressure supplementing electromagnetic valve, a pressure supplementing pore plate, a self-generation pressurizing electromagnetic valve, a self-generation pressurizing pore plate, a storage tank, an overflow valve, a storage tank pressure sensor, a pressure stabilizing valve, a circulating pump, a circulating control valve, a pump front valve, a Joule-Thomson throttling expansion valve, a heat exchanger and a branch supply control valve; the main engine subsystem comprises a pump, a turbine, an auxiliary valve, a fuel gas generator, a main valve, a precooling backflow valve, a main thrust chamber, an evaporator, a starting gas cylinder, a starting gas cylinder control valve, a main thrust chamber igniter and a fuel gas generator igniter. The attitude control engine subsystem comprises a plurality of attitude control engines; each attitude control engine comprises an attitude control thrust chamber, and an oxidant control valve, a fuel control valve and an attitude control thrust chamber igniter are arranged on each attitude control thrust chamber. Propellant supply of the main engine and the attitude control engine is achieved through the unified pressurization conveying system.

The invention relates to the technical field of aerospace engineering, space robots and intelligent control, and particularly discloses a space debris intelligent capture and treatment method and system based on an SAC algorithm, and the method comprises the steps: obtaining the motion data of a space non-cooperative target and the system parameters of a space robot; constructing a hybrid filtering optimization framework based on the motion data and the system parameters; adjusting a hybrid weight and a noise forgetting factor in the hybrid filtering optimization framework by using an SAC algorithm to obtain self-adaptive estimation of a target attitude; a state space and an action space are defined based on a base of the space robot and a mechanical arm of the space robot respectively, and based on the state space and the action space, a mechanical arm movement path of the space robot is generated through an SAC algorithm in combination with a multi-target reward function; and the self-adaptive estimation of the target attitude is fed back to the motion path of the mechanical arm, dynamic cooperation of filtering and control is achieved, and stable capture of the space non-cooperative target is completed.

The invention provides an aerospace craft discrete trajectory precision reconstruction method based on a variable order, and the method comprises the following steps: calculating the order of a reconstruction model based on a given highest-order dynamic parameter of an aerospace craft discrete trajectory; calculating a coefficient of the reconstruction model; based on a given aerospace craft discrete trajectory, constructing a continuous time model of a flight trajectory; and substituting any moment into the continuous time model of the flight trajectory to obtain trajectory data of the aerospace craft at any moment. According to the method, reconstruction can be completed with higher precision based on given discrete trajectory data, and trajectory continuity is considered; order configuration can be flexibly carried out, and the adaptability to discrete trajectories with different dynamic change rules is improved.

The invention discloses a calculation method for determining a visible time period between a satellite and a ground station through TLE data, and relates to the technical field of spaceflight measurement and control and satellite communication scheduling. Through a dynamic sampling mechanism of cluster perception, the time resolution is automatically improved in a dense transit period, the sampling density is reduced in a non-cluster period, the redundancy calculation is remarkably reduced, and the calculation efficiency is improved. The calculation power peak congestion is avoided; a boundary detection algorithm based on elevation angle change rate symbol switching can capture a rapid jump point of a satellite elevation angle, the error merging phenomenon of short interval windows is eliminated, and the precision of an output timestamp reaches a sub-second level; the confidence identification dynamically reflects TLE data timeliness and cluster complexity, a user can intuitively evaluate the window time credibility, and the communication interruption risk caused by orbit prediction errors is reduced.

The invention relates to a carrier rocket secondary reusable electrical system based on a modular TTE-TSN bus, and aims to solve the problems of reliability, light weight and maintenance efficiency in high-speed reentry and precise recovery of a secondary rocket. The system comprises a modular comprehensive control unit, a redundant TTE-TSN high-speed bus, a distributed thermal protection monitoring network and an intelligent health assessment module. Through the modular design, electrical equipment is greatly reduced, and the mass is obviously reduced; the redundant TTE-TSN high-speed bus realizes high-bandwidth and high-reliability transmission of messages on the rocket; the thermal protection monitoring network and the attitude coordination system dynamically optimize reentry thermal load distribution; the intelligent health assessment model predicts the structural life and reuse feasibility based on machine learning. According to the method, the reuse economy and reliability of the two-stage rocket are remarkably improved, and the method is suitable for commercial spaceflight and hypersonic flight vehicle tasks.

The invention relates to a repeatable unfolding and folding mechanism for a large solar wing, which aims to solve the problem that the requirements of the large solar wing on high rigidity, high reliability and repeatable unfolding and folding are difficult to meet in the prior art, and comprises a motor, two double-layer thin-wall elastic rod unfolding and folding mechanisms, a solar wing assembly and a gear transmission assembly, the two double-layer thin-wall elastic rod unfolding and folding mechanisms are arranged in parallel, the solar wing assembly is located between the two double-layer thin-wall elastic rod unfolding and folding mechanisms, and the motor drives the two double-layer thin-wall elastic rod unfolding and folding mechanisms to be unfolded or folded synchronously through the gear transmission assembly. The solar wing assembly is unfolded or folded through the two double-layer thin-wall elastic rod unfolding and folding mechanisms. The invention provides an unfolding and folding mechanism which is light in weight, simple in structure, large in unfolding and folding ratio, higher in rigidity and capable of being repeatedly unfolded and folded by utilizing a motor to drive a wound and folded double-layer thin-wall elastic rod to stretch out to recover rod-shaped unfolding and support a solar wing, and relates to the technical field of spaceflight.

The invention provides a spacecraft state control method and device based on a large language model, and belongs to the technical field of cosmic sailing, and the method employs a spacecraft knowledge graph to add constraint conditions for a fuzzy instruction in a natural language form, and generates a feasible spacecraft instruction in time through the large language model based on the constraint conditions, thereby achieving the control of the state of the spacecraft. The control effect of the state of the spacecraft can be improved, so that the spacecraft can better complete a spaceflight task.

Flexible hole forming equipment suitable for a rocket tube section end frame comprises an axial hole forming device, a radial hole forming device, a rotary table, a flexible tool and a control system. Wherein the rotary table is fixed on a foundation through an adjusting sizing block and a foundation bolt; the flexible tool is mounted on the rotary table through a positioning pin and a screw; the axial hole forming device and the radial hole forming device are installed on the two sides of the rotary table at the interval of 90-degree azimuth angles. The device is mainly used for machining axial butt joint holes and radial riveting holes of multi-model cylinder section end frames of spaceflight carrier rockets and axial milling, flexible hole forming of the rocket cylinder section end frames with the diameter of phi 3350 mm and phi 3800 mm can be met, and machining of all inner L-shaped and outer L-shaped end frames with the diameter of phi 3350 mm and phi 3800 mm is considered; the problem that a multi-drill jig tool corresponding to multi-type end frames is tedious to replace is solved, the size precision and the position precision of end frame drilling are greatly improved, and the machining period is shortened.

The invention discloses a satellite constellation autonomous deployment method and system, belongs to the technical field of aerospace orbit control, and solves the problems of long deployment period, high personnel investment, slow collision avoidance response, increased orbit safety risk and the like caused by a traditional satellite deployment method. The method comprises the following steps: constructing a constellation network according to initial orbit parameters of a satellite and orbit injection time and orbit parameters of different stages; outputting an optimal deployment strategy of each stage of the constellation through a genetic algorithm; the right ascension of the ascending node and the phase are adjusted to a target position together while the height is deployed in place through a coupling equation, and ignition constraints are constructed; generating an orbital transfer plan according to the thrust and a satellite initial orbit needing to be deployed; an orbital transfer instruction is automatically generated according to the orbital transfer plan; according to on-orbit telemetering, the orbital transfer progress condition is judged, and strategy closed-loop adjustment is carried out; cooperative deployment of multi-orbit-plane satellites is achieved, and fuel and time cost are saved. The method is particularly suitable for multi-orbit-plane collaborative deployment and autonomous management of large-scale satellite constellations.

The invention discloses a multi-source data fusion aerospace electronic system reliability evaluation method based on design and manufacturing process constraints, and belongs to the technical field of electronic system reliability evaluation. The problem that the reliability evaluation accuracy of the spaceflight electronic system is poor under the condition of incomplete information is solved. The method comprises the following steps: firstly, establishing a digital prototype model of the aerospace electronic system, and establishing a parameter degradation model of key components of the electronic system in combination with a failure physical model so as to obtain a performance degradation reliability function of each key electrical performance parameter of the electronic system; establishing an electronic system function failure propagation network model based on a logical relationship between an electronic system function structure and basic function network nodes, performing state updating on each function network node of the electronic system by adopting a Bayesian inference method, and obtaining a function failure reliability function of each function network node of the electronic system; and then based on the two reliability functions, constructing a system-level reliability model to obtain an overall reliability curve of the electronic system.