39 results about "Space technology" patented technology

The invention discloses a super-large plane array type space computing power center, and belongs to the field of space technologies and space infrastructures. The space computing power center is deployed in a morning and evening orbit and comprises a plurality of computing power satellites which are physically docked in orbit through a docking mechanism. Each computing power satellite comprises a satellite platform, and a computing power load, a condensation type energy system and a space pump drive fluid loop system which are arranged on the satellite platform, and a docking mechanism for inter-satellite in-orbit physical docking is also assembled on the satellite platform. A plurality of computing power satellites form a planar array type rigid whole body through the docking mechanism, and energy scheduling among the light condensation type energy systems is realized based on the docking mechanism. According to the space computing power center, through multi-satellite physical docking and energy collaborative scheduling, the energy utilization efficiency and the power supply reliability of the system are improved.

The invention relates to the technical field of manned spaceflight, in particular to an atmosphere adjusting device for a manned spaceflight scientific test, which comprises a mounting bracket, a main gas source gas cylinder, an adjusting gas source gas cylinder, a first control module and a second control module, a main air source output joint and an adjusting air source output joint are arranged at the front end of the side wall of the mounting bracket; the main gas source gas cylinder is arranged on one side of the top of the mounting bracket through the gas cylinder fixer, and the adjusting gas source gas cylinder is arranged on the other side of the top of the mounting bracket through the gas cylinder fixer; a main air source pipeline and an adjusting air source pipeline are arranged in the lower part of the mounting bracket; the first control module is arranged on the main air source pipeline, and the second control module is arranged on the adjusting air source pipeline. Various test gases can be provided for manned spaceflight scientific tests, high-pressure storage and accurate supply of the gases can be realized, and quick replacement and on-orbit disassembly and movement of the gas cylinders can also be realized.

The invention concerns a spacecraft propellant tank designed for storing and supplying fuel to the propulsion system, as well as its application method, which is related to space technology and particularly to the design of tanks for storing liquid propellants within rocket engines. The tank in question comprises a body featuring at least one convex edge that creates a corner capillary and an intake port positioned on one of these convex edges. Additionally, the tank's interior surface is treated to be hydrophilic. The objective of this invention is to develop a spacecraft propellant tank characterized by its lightweight and straightforward design. It facilitates the creation of various propellant tank shapes while minimizing the quantity of internal components that reduce useful volume. It also ensures consistent and reliable functionality of the propellant system with continuous supply of propellant to the intake port, even in the absence of gravity and during multiple activations of the propulsion system. The claimed invention achieves its objective by implementing technical results that include enhancing the efficiency of propellant supply within the tank to the intake port, improving reliability, augmenting the useful internal volume, and offering a variety of tank shapes. The stated objectives are met in particular through the hydrophilic properties of the propellant tank's inner surface, the design of the propellant tank shape featuring at least one convex edge that forms a corner capillary, and positioning of the intake port on one of the convex edges that forms the corner capillary. The application method for the spacecraft propellant tank involves initially filling the tank with liquid propellant and pressurizing gas through the intake port prior to spacecraft launch. Subsequently, once the spacecraft is in orbit, propellant is conveyed from the tank to the propulsion system through the intake port. The movement of liquid propellant towards the intake port is facilitated by the pressure of the pressurizing gas and a localized pressure reduction generated during propellant discharge. This reduction occurs in the corner capillary near the intake port relative to the pressure of the propellant in other sections of at least one corner capillary, thus establishing a flow of propellant through the corner capillaries towards the intake port area.

The application relates to the field of manned space technology, in particular to a quick disconnection type bottle mouth valve for a space station combustion test, which comprises a gas cylinder interface, a quick disconnector fixed end and a pointer type pressure gauge interface arranged on a ball valve body, wherein the gas cylinder interface, the quick disconnector fixed end and the pointer type pressure gauge interface form a tee structure in the interior of the ball valve body; one end of the gas cylinder interface is fixed on the ball valve body, the other end is connected with a gas cylinder to form a first passage; one end of the quick disconnector fixed end is fixed on the ball valve body, the other end is connected with a free end of a quick disconnector of a downstream pipeline to form a second passage; one end of the pointer type pressure gauge is fixed on the ball valve body, the other end is connected with a pointer type pressure gauge to form a third passage. The application has high integration, effectively reduces the number of external pipelines, joints and the like of the bottle mouth valve, has a more compact structure, is lighter in quality, and can realize quick disconnection and communication of the bottle mouth valve with the downstream pipeline under microgravity.

The application discloses a kind of spacecraft ground microgravity simulation platform dynamics modeling and mass balancing method, belong to the technical field of space technology.The application includes the following steps: according to the component assembly relationship of spacecraft ground microgravity semi-physical simulation platform, distinguish each component and component group;According to the condition of each component and component group and whole, respectively, each component and component group are established coordinate system;Dynamics derivation is carried out based on coordinate system, and dynamics model is obtained;The complete dynamics model of ground simulation platform is obtained by analyzing dynamics model by Lagrange method;The complete dynamics model of ground simulation platform is simplified by hypothesis, and simplified dynamics model is obtained;Mass balancing is carried out to simplified dynamics model, and the spacecraft ground microgravity semi-physical simulation platform of approximate quasi-zero stiffness is obtained.The introduction of pivot system makes the system reach approximate quasi-zero stiffness, and then reconstructs microgravity environment on the ground.

A beamforming design method for intelligent metasurface arrays based on orthogonal time-frequency-space technology includes: aggregating multipath channels under direct illumination into cascaded channels with near-field direction vectors; randomly initializing the phase shift of intelligent metasurface units; establishing an effective channel matrix based on the cascaded channels; then establishing a multi-user and rate objective function; maximizing the objective function through phase shift optimization using a hierarchical search strategy; repeatedly executing the hierarchical search strategy to search and optimize subsequent blocks until the number of intelligent metasurface units in each block reaches a preset value; then performing an exhaustive search in the final block, thus completing one iteration; repeating the iteration until the maximum number of iterations is reached; outputting the final optimal phase shift; and implementing beamforming design based on the final optimal phase shift. This invention can reduce the complexity of beamforming design and ensure reliable transmission of high capacity during high-speed movement.

A connecting rod hook locking type locking separation mechanism for rocket stage separation relates to the field of space technology. The mechanism comprises an upper stage connecting plate, a lower stage fixing plate and a frame, the frame is fixed on the lower stage fixing plate, the mechanism further comprises a locking assembly, a transmission assembly and a driving assembly, the locking assembly comprises a hook lock and a first fixing shaft, the hook lock is hinged to the frame through the first fixing shaft; the transmission assembly is hinged between the frame and the hook lock; the driving assembly is connected with the transmission assembly, and is used for providing a driving force to drive the transmission assembly to rotate the hook lock around the first fixing shaft to realize locking and separation of the hook lock and the upper stage connecting plate. The mechanism realizes decoupling of locking force and unlocking force, has the advantages of compact structure, high reliability and strong environmental adaptability, and is suitable for stage connection and separation of a new generation of carrier rockets.

The invention discloses a space target long-time-sequence intelligent collision early warning method, and belongs to the technical field of space. The implementation method comprises the steps of constructing a double-branch multi-scale time convolution network long-time-sequence orbit forecasting model, and on the basis of a traditional time convolution network, establishing a relation between the model and perturbation of different time scales by establishing two branch expansion convolution structures and setting different expansion coefficients, so that the forecasting precision of the model is improved. On the basis, a voxel hash mapping rapid screening method is established, a Gaussian negative logarithm likelihood loss function is introduced, statistical characteristics of forecast errors are utilized, a voxel cube is constructed, only voxels where satellites are located and the distance between targets in neighborhoods of the voxels and the satellites need to be calculated during screening, the screening calculation scale is greatly reduced, and the screening speed is increased. The method does not depend on a high-precision dynamic model, can sense and extract the features of different time scales, improves the track forecasting precision and the collision risk screening efficiency, and accelerates the collision early warning speed.

In the current space navigation, communication, and defense technologies, it is difficult to integrate local control of time, gravity operation, use of dark matters, and communication with the outer space (the Buddhist monk Tenkai's gyrus), and intergalaxy communication, ultrahigh-speed movement, and creation of a time acceleration / deceleration space cannot be realized by modern technologies.SOLUTION: The present invention is a future note related to an ultra-future space technology capable of integrally operating time, gravity, and material behavior by performing local and large-scale dark matter control by a BHc / WHc gate based on a dark matter fluid space model and a space gate theory.SELECTED DRAWING: None

The application provides a longitudinal multiple determination method and device for earth coverage period calculation and equipment, relates to the field of space technology, and comprises the following steps: orbit recursion is carried out based on satellite parameters to form target longitude and latitude sequence data of the satellite under each regression orbit; sea and land identification processing is performed on the target longitude and latitude sequence data to obtain a target imaging arc segment timing diagram; the imaging time length of the load per orbit is taken as a constraint to determine the number of regression periods required for completing all the imageable arc segments in the target imaging arc segment timing diagram under each regression orbit, and the number of regression periods is taken as a longitudinal multiple for the calculation of the earth coverage period of the satellite. The application can adapt to the coverage period calculation requirements under the requirements of "only land imaging" or "only ocean imaging", improves the adaptability of the earth coverage period calculation, and meets more refined analysis requirements in actual work.

The invention relates to the technical field of spaceflight, and discloses an astronomical compass attitude drift correction method, device and equipment and a medium, the method comprises the following steps: determining a first measurement error corresponding to an accelerometer in an astronomical compass and a second measurement error corresponding to a sun sensor; determining an initial weight according to the first measurement error and the second measurement error; obtaining observation data of the astronomical compass in a preset time period; determining the attitude drift distance of the astronomical compass according to the initial weight and the observation data; determining whether the attitude drift amount meets a preset correction condition; if the attitude drift amount does not meet the preset correction condition, adjusting the initial weight to obtain an adjusted weight; and recalculating the attitude drift distance based on the adjusted weight until the recalculated attitude drift distance meets a preset correction condition. The correction precision of the astronomical compass can be improved.

The application relates to a satellite-ground ranging method suitable for a multi-satellite measurement and control situation, and relates to a satellite-ground ranging method. The application aims to solve the problem that a physical layer of advanced on-orbit system system service data stream and pseudo code data stream cannot be compatible in a multi-satellite measurement and control scene. The process is as follows: at a transmitting end: a downlink telemetry uplink remote control signal is directly sequence spread through a CA code at the transmitting end I path, so as to realize transmission of measurement and control information between a satellite and the ground; a pseudo code ranging sequence is directly sequence spread through the CA code at the transmitting end Q path, so as to realize a ranging function between the satellite and the ground; at a receiving end: 1, transmitting signals of different satellites are separated, and a pseudo code phase and a Doppler frequency shift are obtained; 2, a carrier tracking loop and a pseudo code tracking loop are realized to stably work; 3, I path coherent integration output is carried out to output continuous bits; Q path coherent integration output is carried out to output continuous bits, transmission time is solved, and then actual satellite-ground distance is obtained. The application is used in the field of space technology.

This invention relates to a power supply module for nanosatellite systems which will find application in the field of space technology and satellite communications, and in particular for powering nanosatellites. The created power supply module consists of at least one battery pack and at least one control and energy distribution module and provides maximum efficiency at a given illumination by adjusting the operating output voltage of the input stages (1.1, 1.2 and 1.3) according to the illumination of the panels. All nodes in the module are duplicated, which achieves complete redundancy of the module, which is activated after the main node is defective, or when the load is greater than the load which this main node can withstand. The use of power busbars, on the other hand, leads to a reduced voltage drop on the respective line, as well as to lower temperature losses. The input channels for the solar panels are transferred to the battery pack and it is possible to connect them in parallel when there are more than one.

The utility model discloses a laser inertial measurement unit test fixing support, and relates to the technical field of spaceflight, and the laser inertial measurement unit test fixing support comprises an adapter plate, a disc, a frame and at least three support leg assemblies used for adjusting height and levelness; the upper end of the adapter plate is provided with an identifier used for calibrating the installation direction of the tested device. Wherein the upper end of the adapter plate is connected with tested equipment, and the lower end of the adapter plate is connected with the upper end of the disc; the lower end of the disc is connected with the upper end of the frame; the at least three supporting leg assemblies are evenly arranged at the lower end of the frame at intervals. The laser inertial measurement unit can meet the requirements that the laser inertial measurement unit is stably placed according to the preset height and the preset direction and is not easily disturbed, vibrated and shaken when the electric performance of a rocket body is tested.

A spacesuit docking device based on a cam mechanism belongs to the field of space technology. The present invention can not only drive the first positioning pin and the second positioning pin to clamp the front pin of the spacesuit by rotating the first cam disk and the second cam disk to realize the six-degree-of-freedom constraint of the front pin of the spacesuit in the base, but also realize the self-locking of the two positioning pins through the friction force between the two cam disks and the two positioning pins, achieving the stability and reliability of docking; the first positioning pin and the second positioning pin are reset during the process of releasing the docking of the device of the present invention by using two springs, realizing the function of repeatable docking; the circumferential positioning flange is used to guide the docking and circumferential positioning of the front pin of the spacesuit and the base, with a large tolerance. The device of the present invention has a simple structure, high reliability, is suitable for the harsh environment in space, has a light weight, low processing, assembly and maintenance costs, and is suitable for high-intensity docking tasks.

The application discloses a satellite adaptive discretization decomposition method and device for regional target observation planning, belongs to the field of space technology, and particularly relates to a strip decomposition of a satellite for a regional target. The method comprises the following steps: determining a discretization step of a regional target according to an observation range of a satellite load; and discretizing the regional target into a regional sub-node set based on the discretization step and the shape of the earth. The regional sub-node set comprises a plurality of sub-nodes. The satellite adaptive discretization decomposition method and device for regional target observation planning are suitable for single or multiple earth observation satellites to plan and calculate observation tasks for a single or multiple irregularly shaped regional targets.

The present application belongs to the field of space technology, and particularly relates to a spherical satellite locking and releasing device with shape maintaining function, which comprises a base, a spring actuating cylinder, an adapter frame and an unlocking and shape maintaining mechanism; the spring actuating cylinder is installed at the center of the base; the adapter frame is installed on the base; the unlocking and shape maintaining mechanism comprises a satellite locking assembly and a satellite shape maintaining assembly; the satellite locking assembly is installed on the adapter frame, and the satellite shape maintaining assembly is fixedly connected with the spherical satellite; when the spherical satellite is carried, the satellite locking assembly and the satellite shape maintaining assembly are rigidly connected to support and lock the spherical satellite; when the spherical satellite is unlocked and released, the satellite locking assembly and the satellite shape maintaining assembly are separated, and the satellite shape maintaining assembly maintains the surface shape of the satellite; the spring actuating cylinder works to push the spherical satellite to be released. The spherical satellite is separated and released from the adapter frame by cooperation of the satellite locking assembly and the satellite shape maintaining assembly, the separation surface is flat, and the requirements of reliable release separation of the spherical satellite and structure surface shape maintaining are effectively ensured.

The invention relates to the field of landslide, in particular to a method for predicting landslide mass motion. The invention relates to a method for predicting landslide mass motion by using a three-dimensional particle flow method. The method comprises the following steps: 1) macro and micro parameter calibration of particle flow calculation; 2) establishment of a three-dimensional geological sliding bed model and a real scene space model interaction and display technology; and 3) landslide particle flow simulation and landslide mass disaster-causing range inversion analysis and real scene result display. A new interdisciplinary method is adopted, a live-action space technology is applied to landslide mass disaster-causing range display, site information is obtained through unmanned aerial vehicle images, a landslide site live-action space model is generated by applying an oblique photography technology, and a disaster site is restored more visually in a three-dimensional form. Meanwhile, through online editing of a live-action model and coupling and interaction of a numerical calculation result and a live-action space model, a landslide mass disaster-causing range prediction result is displayed in a digital and three-dimensional manner, and a landslide disaster monitoring and prevention technical support is accurately and intuitively provided.

The utility model discloses a kind of multi-modal testing devices of simulated space environment, including vacuum chamber with vacuum module;Wherein, cooling module, heating module are respectively equipped in vacuum chamber, vacuum module, cooling module and heating module are respectively connected with the communication control module located outside;Vacuum chamber is also equipped with tray module, for placing sample to be measured;Cooling module can be relatively installed in vacuum chamber and displaced up and down, for adjusting the distance between it and sample to be measured;The utility model can greatly reduce the experimental time of cold and hot impact;While realizing the multi-modal testing effect of simulated space environment, reduce space solar cell array research and development test time and manufacturing cost of product control, help low-cost space technology development and application.

The present application discloses a method and device for planning spacecraft control events, a processor, and an electronic device, relating to the field of space technology. The method includes: determining a first execution duration and a first execution time point of on-board autonomous execution events, and determining a plurality of ground window execution events and corresponding second execution durations; determining a plurality of target TT&C arc segments according to the first execution duration and the first execution time point; determining a plurality of execution plans based on the target TT&C arc segments, the first execution time point, the first execution duration, and the second execution durations; scoring each execution plan based on a preset rule to obtain a plurality of target score values, and determining a target execution plan for the plurality of ground window execution events from the plurality of execution plans according to the target score values. Through the present application, the problem in the related art that the planning efficiency of the execution time of deep space exploration spacecraft control events is relatively low due to the planning of the execution time of deep space exploration spacecraft control events based on manual experience is solved.

The embodiment of the invention provides a method and a system for accurately positioning an electric power operator based on a three-dimensional space technology, and belongs to the technical field of operator positioning. The method comprises the following steps: acquiring a field image of a power grid project; constructing a three-dimensional power grid project real scene model according to the field image of the power grid project; acquiring three-dimensional point cloud data of the power grid project; constructing a three-dimensional remote sensing space according to the three-dimensional point cloud data of the power grid project; performing joint mapping on the three-dimensional power grid project live-action model and the three-dimensional remote sensing space to obtain a three-dimensional point cloud map; a camera in a power grid project is mapped in a three-dimensional point cloud map, and a coordinate conversion relation is constructed, so that a real-time image of the camera can be mapped to the three-dimensional point cloud map in real time, the spatial position of an electric power worker can be determined conveniently, safe operation monitoring of the electric power worker is realized, and the working efficiency of the electric power worker is improved. The working efficiency and effect of the electric power operation personnel are guaranteed, and the operation convenience of the electric power operation personnel is improved.

The invention belongs to the technical field of space, and particularly relates to a micro turbojet power generation system capable of being used for primary energy of a pulse power supply, which comprises a micro turbojet engine and a generator, the micro turbojet engine comprises a gas compressor, a combustion chamber and a first turbine; the gas compressor, the combustion chamber and the first turbine are sequentially connected through gas paths in the gas inlet direction. The gas compressor is coaxially and fixedly connected with the first turbine through a mechanical shaft; the micro turbojet engine further comprises a second turbine, the second turbine is coaxially and fixedly connected with the generator through a mechanical shaft, and the second turbine is connected with the air outlet end of the first turbine through an air path. According to the invention, the micro turbojet engine is adopted to provide stable high-frequency and high-power alternating current for the pulse power supply, and is safer; the problems of small volume power, low power generation efficiency, poor safety and low space utilization rate of primary energy of the pulse power supply are solved.

The utility model relates to the technical field of spaceflight, in particular to a thermal vacuum test device. The thermal vacuum test device comprises a base, a temperature adjusting mechanism and an objective table, the temperature adjusting mechanism is arranged on the base, the objective table is arranged on the temperature adjusting mechanism and used for bearing a tested spacecraft, and the temperature adjusting mechanism can heat or refrigerate the objective table so as to heat or refrigerate the tested spacecraft. The objective table and the spacecraft on the objective table are heated or refrigerated through the temperature adjusting mechanism, so that the capacity of the tested spacecraft for resisting high temperature, low temperature and temperature alternation under the vacuum condition is tested, the problems that in the prior art, a heating structure and a refrigeration structure need to be independently arranged, and the size is large are solved, and the test cost is reduced; the objective table is in direct contact with the temperature adjusting mechanism, the tested spacecraft is in direct contact with the objective table, heat of the temperature adjusting mechanism is transferred in a heat conduction mode, and the heat transfer efficiency is improved.

The invention discloses a planetary probe vehicle, and belongs to the technical field of spaceflight. The planetary probe vehicle comprises a moving system, an intelligent sensing system, a control system and a power system, the moving system is used for moving the planetary probe vehicle and mounting and bearing other components of the planetary probe vehicle, and the intelligent sensing system is used for collecting visual information and depth information in front of the planetary probe vehicle; the control system is used for receiving the information data collected by the intelligent sensing system and sending a mobile control instruction to the mobile system after data processing, and the power system is used for supplying power to the mobile system and the control system. Wherein the control system comprises an embedded computer, a single chip microcomputer and an electronic speed regulator, an autonomous navigation module is arranged in the embedded computer, a navigation path can be dynamically generated by utilizing environmental data collected by the intelligent sensing system in real time, and the problem that an existing planetary probe vehicle depends on a multi-source sensor and a navigation pre-stored map is solved.

The present invention relates to the field of communications, and particularly to a covert communication method and device based on orthogonal time-frequency-space technology. The present invention simultaneously employs a conventional signal masking technique and a hopping code spread spectrum technique, and parasitizes a non-stationary covert signal using hopping code and direct spread spectrum on a stationary conventional signal with obvious characteristics using traditional direct spread spectrum in a manner of differential power combination. Furthermore, a joint design is carried out on the two. The covert signal adopts hopping code spread spectrum controlled by a random sequence, and directly uses the data information carried on the conventional signal (which can be approximately regarded as a random sequence) as the hopping code control sequence of the covert signal, thereby hiding the important service data of the covert signal in the conventional service data of the conventional signal. While ensuring the waveform covert performance, it can effectively improve the adaptability of the signal waveform to high-dynamic communication scenarios, and can be used to realize the transmission of non-stationary covert signals for high-mobility communications.

This application relates to the field of space technology, and specifically discloses a method for screening space dangerous rendezvous targets, including obtaining the orbital elements of the target spacecraft and each space target in the set of targets to be screened; screening to obtain a first target set according to the prediction time and the perigee and apogee heights corresponding to the orbital elements of the space targets; respectively determining the orbital plane intersection lines, intersection positions, and times passing through the intersections of the target spacecraft and each space target in the first target set, and screening to obtain a second target set; setting an extrapolation interval and a step size and predicting the positions of the target spacecraft and each space target in the second target set at each step, so as to screen each space target in the second target set to obtain a third target set; correcting the rendezvous times of each space target in the third target set with the target spacecraft to obtain a dangerous rendezvous target set and a sequence of dangerous rendezvous time windows. Thereby, the screening accuracy and integrity of dangerous rendezvous targets are improved.

The present invention provides a method, device, equipment and medium for optimizing the orbital parameters of a Walker constellation, relating to the field of space technology, including: obtaining a plurality of first alternative parameter values of a first orbital parameter item to be optimized for the Walker constellation; for any first alternative parameter value of the first orbital parameter item, optimizing the second alternative parameter values of the second orbital parameter item to be optimized for the Walker constellation, and continuously calling an orbital simulation model during the optimization process to output an evaluation result corresponding to the second alternative parameter value until the second alternative parameter value corresponding to the optimal evaluation result is determined to obtain a parameter value combination; screening out a first target parameter value and a second target parameter value from a plurality of parameter value combinations according to the evaluation results. The present invention effectively reduces the workload of designers, and at the same time solves the problem of incomplete optimization caused by individual differences of designers, and effectively improves the design efficiency and design quality in the constellation demonstration stage.

A repeated locking and unlocking device for a flexible solar cell array aims to solve the problem that an existing solar cell array unfolding scheme cannot meet the energy requirement and the structural reliability requirement of a large spacecraft in a complex task scene, and comprises a power transmission assembly and a locking and unlocking execution assembly. The power transmission assembly is arranged on the outer side of the box body through a buffer spring and used for providing locking and unlocking power input. And the locking and releasing execution assembly is folded in the box body and is used for locking and releasing the solar cell array. According to the invention, multiple times of repeated locking and releasing operation can be performed on the solar cell array, and the fault-tolerant requirement of a large spacecraft for dealing with sudden faults in a complex task scene and the flexible requirement of dynamic energy management are effectively met. Meanwhile, on-orbit maintenance and subsequent upgrading of the solar cell array are supported, the on-orbit service cycle of the spacecraft is remarkably prolonged, and the operation cost of the whole life cycle is effectively reduced. The invention relates to the technical field of spaceflight.

The present invention relates to the field of space technology test technology, and provides an on-orbit microbial culture experimental payload and a cultivation method thereof. The on-orbit microbial culture experimental payload includes: a test module, which is composed of a plurality of culture units, each culture unit is provided with a culture pool and a heating element for heating the culture pool, and the culture pool is connected to an infusion pipeline with a pump body; a monitoring module, which is arranged relative to the test module; and a control module, which is respectively connected to the monitoring module and the water pump and the heating element in the test module. The on-orbit microbial culture experimental payload and the cultivation method thereof provided by the present invention can realize precise control and monitoring of the microbial culture process in a microgravity environment, thereby supporting space biology research, and through this integrated design, the experimental payload can provide high-throughput experimental capabilities while ensuring the safety and reliability of the experiment.

The invention discloses a large-scale symmetric star group internal collision assessment method, system and device and a medium, and belongs to the technical field of spaceflight. The method comprises the steps of obtaining an orbit error region of a target satellite in a satellite group; constructing a collision box of the target satellite based on the orbit error region of the target satellite; expanding the collision box of the target satellite to obtain an expanded collision box of the target satellite; according to the expanded collision box of the target satellite and the collision box of the slave satellite, an error-free position boundary of the two parties is constructed; calculating the time period of collision risk between the master satellite and the slave satellite through the error-free position boundaries of the two parties; and detecting the time period when the collision risk occurs between the master satellite and the slave satellite, and judging whether the collision occurs or not. According to the method, the symmetry and geometric characteristics of the satellite group are fully utilized, and efficient and accurate satellite collision evaluation is carried out on the large-scale satellite group composed of hundreds and even thousands of satellites.