162 results about "Satellite bus" patented technology

The invention discloses a satellite-borne space-based measurement and control method based on a non-ground network and related equipment. The method is used for a satellite-borne space-based measurement and control device, the satellite-borne space-based measurement and control device is arranged on a low-orbit satellite platform, and the method comprises the following steps: respectively obtaining real-time orbit parameters of a low-orbit satellite, GNSS positioning information and a pre-stored high-orbit satellite coverage information table; based on the current orbit parameter, the GNSS positioning information and a high-orbit satellite coverage information table of a high-orbit satellite pre-stored or periodically updated from a satellite service computer, calculating a space-time relationship between the low-orbit satellite and a cell coverage range; and under the condition that it is predicted and determined that the cell coverage of the target high-orbit satellite is about to enter according to the time-space relationship obtained through calculation, waking up in advance and initiating cell search and a control flow of accessing at any time. The problems that an existing space-based measurement and control technology is high in construction cost, limited in multiple access capacity and poor in dynamic adaptability can be solved.

A relay satellite platform communication device for space-ground networking belongs to the field of relay satellites, and is formed by system-level integration of four core technologies including a structure-electrical integrated central bearing frame, an integrated three-surface pluggable modular system, a conformal antenna integrated multifunctional shell and an internal and external cooperative heat pipe heat dissipation network. A framework-module-skin three-layer deep coupling system framework is constructed, six faces of a center bearing structure are divided into a movable area and a static area, the movable area is used for achieving function reconfiguration and on-orbit maintenance, and the static area is used for deploying a high-reliability platform foundation system which is fixed once launched. According to the invention, the problems of difficulty in on-orbit maintenance, low reliability and efficiency bottleneck caused by discrete architecture and functional coupling of a traditional relay satellite are solved, and a high-reliability, high-flexibility and easy-to-maintain satellite platform foundation is provided for constructing a dynamic reconfigurable space-ground integrated information network.

The application provides a small satellite platform structure which carries a large amount of propellant and is connected and separated in point mode, belongs to the technical field of spacecraft structure, and comprises a frame, a bottom plate, a lower cabin partition plate, a storage tank support rod, a middle layer plate, an upper cabin partition plate, a top plate and a side plate, so that the side plate surrounds the parallelly arranged bottom plate, middle layer plate and top plate into a closed box body, the bottom plate is connected to the frame, the storage tanks are arranged diagonally in the round holes on the bottom plate and the middle layer plate, and the storage tank support rods are uniformly distributed around the round holes; the middle layer plate separates the box body into an upper cabin and a lower cabin, the upper cabin partition plate is connected in the upper cabin, and the lower cabin partition plate is connected in the lower cabin. The application proposes a configuration of "point connection + storage tank sinking", meets the demand of the small satellite platform for carrying a large amount of propellant, effectively reduces the whole satellite center of mass, and improves the satellite dynamic environment; meanwhile, the bottom main load-bearing frame cooperates with a detachable adapter block, can adapt to different mechanical interfaces of launch vehicles, and realizes point connection and separation.

The application discloses a space remote sensing scanning imaging system self-rotation eliminating turntable mechanism and a construction method thereof, and belongs to the technical field of space remote sensing. The mechanism comprises a first self-rotation eliminating mechanism and a second self-rotation eliminating mechanism, the first self-rotation eliminating mechanism and the second self-rotation eliminating mechanism are respectively used for driving the scanning movement and the swing of the space remote sensing scanning imaging system; the first self-rotation eliminating mechanism is connected with a satellite platform, and the second self-rotation eliminating mechanism is connected with the space remote sensing scanning imaging system; the first self-rotation eliminating mechanism comprises a first double-output-shaft motor, which is used for driving the coaxial reverse movement of the connected space remote sensing scanning imaging system and a first rotation eliminating device; the second self-rotation eliminating mechanism comprises a second double-output-shaft motor, which is used for driving the coaxial reverse movement of the connected space remote sensing scanning imaging system and a second rotation eliminating device. When the space remote sensing scanning imaging system is in scanning movement and swing, the satellite is provided with equal and opposite angular momentum, so as to offset the disturbance brought to the satellite by the movement of the space remote sensing scanning imaging system, and the satellite keeps a stable attitude.

The invention provides a control method of a panel satellite large-size flexible solar wing driving device. The driving device comprises a satellite body, a double-shaft SADA, a swing rod, a hinge, a pressing and releasing mechanism and an ROSA solar wing. The control method comprises the steps that a release signal is sent to the pressing and releasing mechanism, the pressing and releasing mechanism releases the pressed solar wing after receiving the release signal, and the solar wing is driven by a hinge at one end of a swing rod to be passively unfolded by a first angle; according to the target angle and the first angle, calculating a second angle that the double-shaft SADA needs to rotate; a rotation signal is sent to the double-shaft SADA, the double-shaft SADA drives the solar wing to continue to rotate by a second angle after receiving the rotation signal, and a rotation in-place signal is sent out; and an unfolding signal is sent to the solar wing, and the solar wing starts to unfold the cell array plane until the cell array plane is completely unfolded. The control method improves the energy utilization efficiency of the whole satellite platform.

The invention discloses a biomedicine standardized cubesat platform and a working method, and relates to the technical field of space.The biomedicine standardized cubesat platform adopts a 3U cubesat framework, the top layer of the 3U cubesat framework is provided with a satellite bus module, the middle layer of the 3U cubesat framework is provided with an energy and environmental control life support module, and the bottom layer of the 3U cubesat framework is provided with a biological experiment module; the satellite bus module is in bidirectional communication with the energy and environment control life support module and the biological experiment module through data interfaces. The satellite bus module further communicates with a ground station; the energy and environment control and life support module supplies energy to the satellite bus module and the biological experiment module through electrical interfaces, and provides oxygen supply, water supply and environment regulation and control support for the biological experiment module through a fluid transmission channel. Through a standardized and modularized cubesat structure, the functions of biological sample culture, environment regulation and control, real-time monitoring, data return and the like are integrated, and a low-cost and quick-response technical support is provided for space life science experiments.

The application relates to an on-orbit maintenance method for a self-repairable satellite-borne DSP, which is realized by an on-orbit maintenance system and is used for maintaining and managing a DSP chip; the on-orbit maintenance method comprises the following steps: S1, powering on the on-orbit maintenance system; S2, performing three-mode redundancy comparison error correction and CRC check word checking to judge whether the checking is passed; S3, mapping the DSP chip to a target chip area address and performing DSP program data loading; S4, judging whether the monitoring function of the DSP chip is normal; and S5, downloading the chip area address of the DSP program data which is not passed in the telemetry and executing satellite platform instructions to repair the DSP program data. The application has simple architecture, low implementation difficulty, can effectively prolong the on-orbit life of a satellite and other space vehicles while keeping low cost.

This invention relates to a method for inverting remote sensing data from overlapping strong and weak absorption lines of greenhouse gases, belonging to the field of satellite remote sensing technology. Specifically, it includes the following steps: Step 1: Constructing an optical system, including a satellite platform, camera 1, and camera 2; camera 1 and camera 2 are mounted at the bottom of the satellite platform; camera 1 and camera 2 are mirrored; the imaging spectral band of camera 1 moves from the weak absorption spectral band to the strong absorption spectral band along the positive push-broom direction to the negative push-broom direction; the imaging spectral band of camera 2 is in the opposite direction; the optical system moves horizontally from left to right to achieve horizontal push-brooming of greenhouse gases on the ground, obtaining a dual-camera fused image; the dual-camera differential absorption spectroscopy algorithm is used to invert the dual-camera fused image to obtain the greenhouse gas column concentration; this invention, based on the spectral characteristics of strong and weak absorption lines of greenhouse gases, adopts a detector focal plane segmentation method, and the dual cameras achieve differential absorption and inversion in the same region, which can guide the design of greenhouse gas detection micro / nano satellite payloads.

The invention relates to a satellite capturing and locking device, and belongs to the field of satellite docking. The satellite capturing and locking device comprises an active end and a passive end, the driving end comprises a driving end base, a guide mechanism, a driving mechanism and two locking mechanisms; the guide mechanism comprises a guide frame and a guide plate; the guide plate is in an inverted cone shape; the passive end comprises a passive end base and a caught rod; the guide plate is used for guiding the captured rod into the opening; the driving mechanism is used for driving the two locking mechanisms to move oppositely so as to clamp the captured rod, and therefore locking between the driven end and the driving end is achieved. The caught rod at the driven end is clamped left and right through the two locking mechanisms to achieve locking, the two locking mechanisms are arranged in a bilateral symmetry mode, namely, the two locking mechanisms are arranged at intervals in the transverse direction, the axial distance between two satellites after locking is reduced, and therefore the size of a satellite platform is reduced. The locking mechanism comprises a first-stage locking unit and a second-stage locking unit, and the ejection speed during separation can be guaranteed.

This invention relates to a lightweight, miniaturized 3S configuration triaxial integrated fiber-MEMS inertial measurement unit, belonging to the field of measurement devices. It includes a base cover, a power supply and detection board assembly located on the top of the base cover, a light source assembly located above the power supply and detection board assembly, a MEMS assembly mounted above the light source assembly, a base fixed to the base cover, fiber optic ring assemblies fixed to the base, and a reference mirror fixed to the top of the base. The power supply and detection board assembly, the light source assembly, and the MEMS assembly are all arranged inside the base. The three fiber optic ring assemblies are evenly distributed on a conical surface with an outer half-cone angle of 54.74° and a bottom projection of 120°. The light source assembly shares a light source with the three fiber optic ring assemblies via a 1*3 coupler. This invention eliminates the unnecessary space occupied by the traditional three independent light source configuration, reduces the overall volume, and improves optical path consistency and system stability. It can meet the engineering requirements of commercial satellite platforms for the miniaturization and mass production of inertial measurement components.

A first satellite includes a first host-side bridge and a first host, connected with a network via the first host-side bridge, mounted thereon. A second satellite includes a first I / O-side bridge and a first I / O device, connected with the network via the first I / O-side bridge, mounted thereon. A management device selects the first host mounted on at least one of the first satellites and the first I / O device mounted on at least one of the second satellites, according to a configuration request, and forms an information processing device in which the first host-side bridge connected with the selected first host and the first I / O-side bridge connected with the selected first I / O device constitute a virtual computer internal bus.

The invention relates to a connecting rod driving unfolding mechanism for a large-scale deployable array plane antenna, and belongs to the technical field of spaceborne antennas. Comprising two antenna multi-layer array surface structures, a satellite platform and two groups of driving unfolding assemblies, wherein the satellite platform is horizontally placed; the two antenna multi-layer array surface structures are arranged below the satellite platform in a vertical folded state; the two driving unfolding assemblies are symmetrically arranged on the front side wall and the rear side wall of the satellite platform. Each driving unfolding assembly comprises a fixed truss, two short connecting rods, two long connecting rods, two synchronous tooth holders, a driving synchronous belt, two follow-up synchronous belts, ten belt wheels and inter-plate hinges. Power is provided for the two antenna multi-layer array surface structures to be changed from a completely folded state to an unfolded state through the two groups of driving unfolding assemblies, and the unfolding speed of the antenna multi-layer array surface structures is controlled; the connecting rod driving unfolding mechanism for the large-scale deployable array plane antenna provides unfolding power and unfolding rigidity for the antenna, and has the characteristics of high storage ratio, high rigidity, high precision and controllable unfolding.

This invention proposes a novel spaceborne SAR antenna device and control method for simultaneous dual-side-looking imaging, thus meeting the application requirements of simultaneous dual-side-looking imaging of the Earth. The antenna system is a phased array system with a wide-range electronic scanning capability. It employs one transmit / receive link, one receive-only link, and one calibration link. Through an improved beam control code configuration and timing control method, simultaneous observation and imaging of both left and right sides is achieved. This invention achieves simultaneous dual-side-looking imaging of the Earth without requiring satellite platform attitude maneuvers, effectively improving system performance.

A low-stress temperature control device of a satellite-borne optical reflector comprises an optical reflector, a flexible core shaft, a reflector back plate, a radiation temperature control plate, a temperature control flexible supporting column, a heating assembly, a thermistor and a thermal control multi-layer heat insulation assembly. The optical reflector is installed in a suspended mode through the flexible mandrel, flexible glue bonding, radiation temperature control and the like, the temperature adaptability is improved, and the influence of a temperature control device on the surface type of the optical reflector is reduced. And the reflecting mirror surface adopts silver and a dielectric film, and the area except the reflecting surface is coated with a plurality of layers of heat insulation components, so that the adaptability of the device to a complicated and changeable space heat flow environment is improved. In a limited rotation angle range of the reflector, through asymmetric layout of a long axis and a short axis, the emission height size is reduced. The method can be applied to a space pointing mechanism, is suitable for occasions where out-cabin height direction resources are limited such as a flat plate type satellite platform, has the advantages of being good in space environment adaptability, low in launching envelope and flexible in application, and is suitable for various orbit application scenes.

A kind of micro-nano satellite module rapid integration method belongs to the field of relay satellite platform, there is standardization backplane and standardization module, standardization backplane is provided with bus area and slot area, bus area includes uniform specification power bus and data bus, slot area is arranged with multiple standardization slots, standardization module is inserted into standardization slot.The structure of standardization slot is uniform, and a uniform standard data electrical interface is provided inside, each standardization slot corresponds to fixed physical port and logical address / node number, or unified address configuration is realized through slot coding, according to functional task requirement, select the standardization module that meets functional task requirement, and build the satellite system that meets functional task requirement.The present application significantly shortens the development cycle, greatly reduces the development cost, greatly improves the system flexibility and scalability, improves the system reliability and maintainability, and promotes standardization and industrialization.

The invention provides a satellite-borne high-function-density comprehensive electronic computer based on an SM750 processor. Comprising a casing, and a first general processing and small solid memory module, a second general processing and small solid memory module, a first secondary power supply module, a second secondary power supply module, a channel gateway module, an I-type comprehensive interface module, an II-type comprehensive interface module, a serial port communication module, an initiating explosive device and power distribution module and a power driving module which are arranged on the casing, the system comprises a relay responder module, a measurement and transmission integrated responder module, a navigation module and an internal bus mother board. According to the invention, a satellite platform measurement and control function, a measurement and transmission integrated function, a satellite service management function, a time base function, an attitude and orbit control management function, a power supply and distribution management function, an initiating explosive device and hot knife driving function, a thermal control management function, a mechanism pointing trajectory planning function and a navigation function are integrated in one single machine, so that the integration level of core functions of the satellite platform can be improved; and the quality and power consumption overhead of basic functions of the platform are effectively reduced.

The invention relates to a large-aperture fixed surface radio telescope system and an on-orbit construction method, and the system comprises a radio telescope which is a large-aperture radio telescope; the mechanical arm is configured to perform in-orbit installation on the plurality of second antenna panel split plates so as to form the second antenna panel; and a satellite platform configured to carry the radio telescope and the robotic arm. The radio telescope includes: a first antenna panel; the second antenna panel comprises a plurality of second antenna panel sub-plates; and the butt joint device is configured to be in butt joint with the first antenna panel and the second antenna panel. The docking device can be conveniently docked with the second antenna panel on the outer side of the first antenna panel, and locking is completed through motor retraction after docking locking, so that the telescope system has two states of assembling and working, the requirement for the control precision of a mechanical arm is reduced, and the collision risk is reduced.

The invention discloses a satellite autonomous imaging task planning method and system, and the method comprises the steps: enabling a satellite service system to transmit a request data package to an AIS load according to a ground uploading instruction, enabling the AIS load to obtain the real-time state information of a target ship, and generating a detection data package; the satellite service system carries out on-satellite autonomous judgment according to the detection data packet sent by the AIS load, and an imaging target meeting the imaging constraint condition is screened out; and for the screened imaging target, the satellite service system generates a corresponding imaging task chain, and controls a satellite platform and an SAR imaging load to execute the imaging task chain so as to complete imaging data acquisition of the imaging target. According to the invention, full-autonomous closed-loop imaging task planning from target sensing to imaging execution can be realized on a satellite.

A rapid integration method for micro-nano satellite modules belongs to the field of relay satellite platforms, a standardized backboard and standardized modules are provided, the standardized backboard is provided with a bus area and a slot area, the bus area comprises a power bus and a data bus which are unified in specification, a plurality of standardized slots are arranged in the slot area, and the standardized modules are inserted into the standardized slots. The standardized slots are uniform in structure, data electric appliance interfaces with uniform standards are arranged in the standardized slots, each standardized slot corresponds to a fixed physical port and a logic address / node number, or uniform address configuration is achieved through slot position coding, and standardized modules meeting function task requirements are selected according to the function task requirements. And a satellite system meeting functional task requirements is constructed. According to the method, the research and development period is remarkably shortened, the research and development cost is greatly reduced, the flexibility and expandability of the system are greatly improved, the reliability and maintainability of the system are improved, and standardization and industrialization are promoted.

The application relates to a tracking radar decoupling combined test system, which comprises a scanning frame which is movable, a echo emission device which is connected with a signal source and fixed on the scanning frame to follow the scanning frame to move and used for simulating a target, a radar which is fixed on a rotating table to simulate disturbance of the radar fixed on a satellite through rotation of the rotating table, a host computer which is connected with the rotating table and used for controlling the rotating table, and a radar test device which is connected with the host computer at an input end, acquires angular velocity information sent by the host computer to the rotating table, and is connected with the radar at an output end to output the angular velocity information to the radar and used for simulating angular velocity information output by a gyroscope of a satellite platform; the radar executes decoupling processing according to the angular velocity information and outputs position information of the simulated target in real time for testing. The application firstly proposes to test the decoupling function of the radar in a system mode, simultaneously realizes combined testing of high-precision radar tracking targets and decoupling platform disturbance, and can simulate real use working conditions.

The invention relates to a multi-antenna SAR (Synthetic Aperture Radar) satellite configuration and a rocket, and the multi-antenna SAR satellite configuration comprises a satellite platform, a first antenna, a second antenna, a first solar panel and a second solar panel, the peripheral side face of the satellite platform comprises a first installation face, a second installation face, a third installation face and a fourth installation face which are sequentially arranged in the circumferential direction of the satellite platform, and the first antenna is assembled on the first installation face and can be unfolded in the direction parallel to the axis of the satellite platform. The second antenna is assembled on the second installation face and can be unfolded in the direction away from the first antenna, the first solar panel and the second solar panel are installed on the end faces of the two axial ends of the satellite platform respectively and can be stacked on the third installation face in a surrounding mode, and the fourth installation face serves as a rocket assembly face. The satellite is compact in overall configuration layout, adapts to a multi-load antenna, is high in integration level, is small in envelope size, is high in rocket adaptability, can be laterally hung and arranged at the top, improves the rocket plot ratio, and can achieve one rocket with multiple satellites.

A method and system for suppressing leakage signal of continuous wave SAR by using secondary sampling, according to the characteristics that the leakage signal is strongly superimposed on the weak echo signal in the SAR system, after the first digital sampling of the SAR echo signal, secondary sampling is performed after the original signal cancellation (suppression of leakage signal) of the digital-to-analog conversion, and finally the effective echo data is obtained. The method can suppress the leakage signal without improving the antenna isolation, and can lay the foundation for the application of continuous wave SAR on satellite platform.

The invention discloses a satellite-ground integrated task planning method for multi-frequency tasks, and belongs to the field of satellite task planning. The method specifically comprises the following steps: firstly, receiving an aerospace observation grid element task set to form a grid element task pool; secondly, a ground task planning system comprehensively considers satellite electric quantity, storage and the like to preliminarily screen grid element tasks, and a single-satellite grid element task set is formed; secondly, uploading the single-satellite grid element tasks to each satellite platform through a measurement and control link; and finally, the satellite platform performs access calculation on the satellite based on an actual operation orbit to obtain accurate grid access time, namely load startup and shutdown time, and then single-satellite task planning is completed based on an improved genetic algorithm to form a final observation scheme. According to the invention, joint task planning is carried out by adopting satellite-ground resources, and a sectional task planning mode of obtaining a redundant observation scheme through ground rough planning and obtaining a final observation scheme through satellite fine planning is innovatively proposed, so that the observation coverage rate and task guarantee timeliness are effectively improved.

The invention discloses a power generation and radiation refrigeration integrated system for a high-power computing power satellite. The system comprises an integrated function panel, a photovoltaic power generation unit, an active radiation refrigeration unit and a closed pump drive heat transmission loop. The integrated function panel comprises a panel body, the panel body is provided with a first surface and a second surface which are opposite, and the panel body blocks heat conduction and heat radiation between the first surface and the second surface; the photovoltaic power generation unit is arranged on the first surface; the active radiation refrigeration unit is arranged on the second surface; the active radiation refrigeration unit is provided with an outer surface with spectral selective radiation refrigeration characteristics, and the active radiation refrigeration unit comprises an internal micro-channel structure thermally coupled with the outer surface; and the closed pump drive heat transmission loop is connected with a heating electronic element in the satellite body and the internal micro-channel structure. The power generation performance can be optimized while the satellite heat dissipation efficiency is improved, heat and electricity integrated management of a high-efficiency calculation satellite platform is achieved, the reliability and efficiency of the overall system are improved, and meanwhile the overall weight of an energy system is reduced.

This invention provides a method and system for autonomous calculation and solar avoidance by a laser communication terminal, comprising: before the start of a laser communication mission, the satellite platform sends platform information for the current mission time period of the satellite and the target satellite to the laser terminal via the onboard bus; based on the received platform information, the laser terminal autonomously calculates the angle between the communication optical axis of the satellite and the target satellite and the solar vector during the mission period; during the calculation process before the mission starts, the laser terminal determines whether there is a moment during the mission when the angle is less than a preset safety threshold; if so, the laser terminal sends a warning message to the satellite platform; the satellite platform decides whether to execute the laser communication mission based on the warning message. This invention achieves autonomous solar outage avoidance onboard without real-time ground-based telemetry and control support by completely migrating the complex solar vector calculation and avoidance decision-making logic to the satellite and having the laser terminal perform it autonomously.

The present application relates to a kind of motion coupling under spaceborne laser communication terminal fixed time disturbance control method, for the periscope type laser communication terminal fast pointing control problem under the influence of multi-source disturbance such as space environment interference, floating satellite platform coupling moment and rotation shaft friction moment, first, the deep coupling model of spaceborne communication terminal containing above-mentioned multi-source disturbance is established, and multi-source disturbance is regarded as total disturbance moment by analyzing interference structure and characteristics;Second, fixed time disturbance observer is designed to quickly estimate the total disturbance moment suffered by communication terminal;Third, nominal fixed time controller is designed and combined with disturbance observer to form fixed time compound anti-interference controller;Finally, combined with the pointing time demand of communication terminal, observer and controller parameter selection are analyzed and guided, and fixed time anti-interference control of spaceborne laser communication terminal under motion coupling is completed.The control method of the present application has the characteristics of strong robustness, fast pointing speed and guaranteeing pointing error fixed time convergence.

The invention relates to the field of inter-satellite laser communication, and provides a satellite platform micro-vibration feature extraction method and device, and the method comprises the steps: obtaining the sampling data of a laser link related assembly in a laser terminal, the sampling data of the related components of the laser link at least comprises photoelectric encoder measurement value data, fast reflecting mirror driving value data and capturing and tracking camera miss distance data; extracting time domain spectrums of different granularity pointing compensation from sampling data of laser link related components; carrying out superposition processing on the time-domain spectrums of the granularity pointing compensation to obtain a micro-vibration signal time-domain spectrum; and determining micro-vibration characteristics according to the micro-vibration signal time domain spectrum. According to the invention, the extraction precision and efficiency of micro-vibration features can be improved.

The application provides a control method of a flat satellite large-size flexible solar wing driving device, the driving device comprising a satellite body, a double-shaft SADA, a swing rod, a hinge, a compression release mechanism and a ROSA solar wing; the control method comprising: sending a release signal to the compression release mechanism, releasing the compressed solar wing after the compression release mechanism receives the release signal, driving the solar wing to do passive unfolding of a first angle through the hinge at one end of the swing rod; calculating a second angle that the double-shaft SADA needs to rotate according to a target angle and the first angle; sending a rotating signal to the double-shaft SADA, driving the solar wing to continue rotating the second angle after the double-shaft SADA receives the rotating signal, and sending a rotating-in-place signal; sending an unfolding signal to the solar wing, and starting the battery array surface unfolding of the solar wing until completely unfolded. The control method improves the energy utilization efficiency of the whole satellite platform.

The invention provides a novel mechanical-thermal integrated satellite structure plate and a manufacturing method thereof. The novel mechanical-thermal integrated satellite structure plate comprises an upper skin (pretreated), a honeycomb core material, an embedded part, a heat pipe, a lower skin (pretreated) and an adhesive. The upper skin and the lower skin are respectively connected with the honeycomb core material through an adhesive, the heat pipe is arranged inside the honeycomb core material, the embedded part is adhered among the upper skin, the honeycomb core material and the lower skin through the adhesive, and the difference between the honeycomb core material and the honeycomb core material is that the upper skin and the lower skin are subjected to oxidation treatment before glue joint forming to obtain a thermal control oxidation layer. Therefore, integration of mechanical and thermal functions is achieved in the forming process, and an existing long-period structure and thermal control integration mode is improved. According to the method, multiple process links such as paint spraying and drying are reduced, the development period of a satellite structure platform is greatly shortened, and the satellite manufacturing requirement for quick response is met. Meanwhile, due to the simplification of the manufacturing process, the manufacturing cost of the satellite machine-hot platform is remarkably reduced, and the economical efficiency of satellite platform manufacturing is improved.

The present invention relates to an integrated satellite platform that will find application in the satellite industry, more particularly for reducing the time to integrate, test, and launch into orbit around the Earth various types of payloads and obtain data therefrom. The developed integrated satellite platform is composed of a power unit, a communication system, on-board computers, an attitude determination and control system, and a propulsion system. All components and systems of the platform are, on the one hand, connected to a single central bus (6) for power transmission, and, on the other hand, are bidirectionally interconnected via the single central bus (6) for telemetry and telecommand transmission within the platform. An embedded diagnostic and test module (7) is also connected to the single central bus (6) via interrupting mechanisms. The unified bus (6) allows data and power to be transferred between the components of the platform without the need for cabling between the components. The platform is equipped with standard power interfaces (8.1) and high-speed data communication interfaces (8.2) to the payload.