CN112046787A - Domain-based microsatellite system - Google Patents
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Abstract
The invention discloses a domain-based microsatellite system which comprises a plurality of functional units which are arranged in a distributed mode, wherein the functional units comprise a structural unit, a communication and conduction unit, a comprehensive information unit, a GNC unit, a load unit, a thermal control unit and a power supply unit. The method is based on the design concept of a functional domain, adopts a miniaturized, standardized and modularized design method to perform technical analysis and design optimization of the whole satellite system architecture of the microsatellite, divides the satellite into different functional modules or functional areas according to task functions and support functions, breaks through the physical boundary of the system/subsystem/component in the traditional microsatellite, and solves the problem that the microsatellite is difficult to combine multiple functions, high density, functionalization and centralization.
Description
Technical Field
The invention relates to the technical field of space. And more particularly to a domain-based microsatellite system.
Background
A satellite is an unmanned spacecraft that orbits the earth in space and generally consists of a dedicated system and a support system. A dedicated system refers to a system directly related to the tasks performed by the satellite, also referred to as payload. The guarantee system is a system for guaranteeing the normal work of the satellite and the special system in the space, is also called as a service system, and mainly comprises a structural system, a power supply system, a thermal control system, an attitude control and orbit control system, a radio measurement and control system and the like. In order to fulfill the goals and requirements of users, satellite systems are generally used to develop comprehensive technologies such as electromagnetic compatibility technology, software engineering, comprehensive testing technology, environmental testing technology, and integrated design.
Since the 90 s of the 20 th century, with the successful operation of high-performance, low-cost, low-risk microsatellites, the field of application of microsatellites has continued to expand, and better economic and social benefits have been generated. Due to the increasing mission of microsatellites, the development period of the microsatellites is shorter and shorter. In order to complete the microsatellite task in a 'multi-, fast-, and provincial' manner, a general platform technology based on modularization, integration and serialization becomes a key. In addition, with the concern of low cost and the transparency of the industrial chain, users generally proceed to widely research the matching situation of industrial terminal development before project establishment, and grasp the first information of industrial matching. Therefore, in order to meet the demands of mass-produced and low-cost users, the microsatellite development model is gradually changed from platform-centered to load-centered. In general, modern microsatellites have high functional density and meet user requirements by developing platform-loaded integrated designs.
At present, from the perspective of system engineering, a microsatellite mainly comprises attitude and orbit control, a structure and a mechanism, thermal control, a power supply, information, communication, an antenna, a test load and other systems, wherein the subsystem comprises a plurality of subsystems, and the subsystems are further decomposed into various components, equipment and the like. That is, there is a physical limit for the system/subsystem/components in the conventional microsatellite, and it is difficult to combine the multifunction and high density, and the functionalization and centralization of the microsatellite.
Therefore, it is desirable to provide a microsatellite system which can break through the physical limitations of the system/subsystem/components of the conventional microsatellite and solve the problems of multifunction and high density of the microsatellite and difficulty in combining the functionalization and centralization.
Disclosure of Invention
In order to solve the above problems, the present invention provides a domain-based microsatellite system, which performs technical analysis and design optimization of the microsatellite whole satellite system architecture by using a centralized, expandable and reducible general architecture design and a miniaturized, standardized and modularized design method according to a domain design concept, and divides a satellite into different functional modules or functional areas according to task functions and support functions to form a highly integrated body, so as to break the physical boundary of the system/subsystem/component existing in the conventional microsatellite and solve the problem that the multifunction, high density, functionalization and centralization of the microsatellite are difficult to be combined.
In order to achieve the purpose, the invention adopts the following technical scheme:
a domain-based microsatellite system comprises a plurality of functional units which are arranged in a distributed mode, wherein the functional units comprise a structural unit, a communication and conduction unit, a comprehensive information unit, a GNC unit, a load unit, a thermal control unit and a power supply unit, and the system comprises a domain-based microsatellite system, a domain-based microsatellite system and a domain-based microsatellite system, wherein the domain-based microsatellite system comprises a plurality of functional units which
The structural unit is used for providing installation and support conditions for each functional unit of the microsatellite;
the communication and measurement unit is used for receiving a ground instruction and forwarding the ground instruction to the comprehensive information unit so as to realize the functions of ground measurement and control, ground data transmission and satellite navigation;
the comprehensive information unit is used for receiving the ground instruction forwarded by the communication and conducting unit and realizing the power distribution function, initiating explosive device control, heater control, temperature acquisition, bus remote measurement, star management and GNC software support tasks of the whole satellite based on the ground instruction;
the GNC unit comprises a GNC measuring unit and a GNC executing unit, wherein the GNC measuring unit is used for collecting attitude information of the whole star, and the GNC executing unit is used for finishing attitude and orbit control;
the thermal control unit is used for providing a working temperature environment for the microsatellite system in a task period;
the load unit is used for completing an on-orbit test task according to a preset flow based on the ground instruction; and
and the power supply unit is used for providing a direct current power supply bus for the microsatellite system in the service life of the whole satellite so as to finish power transmission of the whole satellite.
Optionally, each functional unit uses a CAN bus as a backbone communication network.
Optionally, the integrated information unit is further configured to manage and schedule satellite information, monitor states of the subsystems, autonomously manage work and telemetry data and state information of the subsystems, and control the execution component according to data provided by the GNC unit, and the integrated information unit includes a host module, a standby module, and an interface function board, where the integrated information unit includes a host module, a standby module, and an interface function board card
The host module is integrated with a processor and is used for carrying out unified scheduling management on satellite resources so as to realize integration and processing of telemetering data, monitoring the satellite state so as to realize instruction control on satellite equipment, managing power distribution, thermal control, time correction and safety of all satellite subsystems and processing a part with a fault in time;
the standby module is configured as a hardware heterogeneous backup of the host module; and
the interface function board card is integrated with a data storage module and a data acquisition module and is used for providing electrical connection for the comprehensive information unit and providing a CAN bus interface, a heating drive control interface, an initiating explosive device line interface and an OC instruction interface.
Further optionally, the Combined information Unit employs a first level fault tolerance mechanism and a second level fault tolerance mechanism, wherein
The first-level fault-tolerant mechanism comprises a dual-core mutual detection technology and a rollback recovery technology; and
the second level fault tolerance mechanism includes software EDAC and triple modular redundancy techniques for data flow errors, and exception traps and watchdog techniques for control flow errors.
Optionally, the GNC measurement unit includes a star sensor, a sun sensor, a magnetometer, and a navigation receiving mechanism, and is configured to collect attitude information of a whole star, measure and calculate spatial position, velocity, attitude, and time information of a satellite, and transmit the information to the integrated information unit to generate an orbit maintenance and attitude stabilization control instruction; and
the GNC execution unit comprises a micro-propulsion unit, a micro flywheel and a micro magnetic torquer and is used for realizing the three-axis attitude stabilization and attitude control of the satellite.
Optionally, the measurement and communication unit includes a measurement and control terminal body, a pair of measurement and control transceiver antennas, a pair of data transmission transmitting antennas and a pair of satellite navigation receiving antennas, wherein
The measurement and control terminal body comprises a radio frequency module, a satellite navigation receiving unit, a signal processing module and a secondary power supply, wherein the radio frequency module, the satellite navigation receiving unit, the signal processing module and the secondary power supply are arranged on the inner wall of a satellite;
the pair of measurement and control transmitting and receiving antennas are symmetrically arranged on the sky surface and the ground surface of the satellite;
the satellite navigation receiving antenna is arranged on the opposite sky surface of the satellite; and
the secondary data transmission transmitting antenna is arranged on the ground of the satellite.
Optionally, the power supply unit comprises a solar cell array, a storage battery pack and a power supply controller, wherein
The solar cell array comprises a body-mounted three-junction gallium arsenide cell slice and is used for converting solar light energy into electric energy through photoelectric conversion;
the storage battery pack is used for providing energy for a load when the power provided by the shadow area or the peak power consumption solar cell array is insufficient; and
the power supply controller is used for controlling the working state of the solar cell array, the charge and discharge management of the storage battery pack and the power supply change.
Optionally, the functional units adopt standard electrical interfaces.
Optionally, the functional units are integrally designed and integrated in a three-dimensional heterogeneous manner.
Optionally, each functional unit is electrically and structurally interconnected in a synchronous manner by adopting board-level longitudinal plug-in.
The invention has the following beneficial effects:
based on a functional domain design concept, the invention adopts a miniaturized, standardized and modularized design method to perform technical analysis and design optimization of the whole satellite system architecture of the microsatellite, divides the satellite into different functional modules or functional areas according to task functions and support functions, breaks through the physical boundary of the system/subsystem/component in the traditional microsatellite, and solves the problem that the microsatellite is difficult to combine multiple functions with high density, functionalization and centralization.
The microsatellite system adopts a distributed network system structure and is set as a highly integrated system for uniformly scheduling and managing tasks, functions and resources, the system adopts a centralized, expandable and reducible general architecture design, the functions of the whole satellite (or part of electronic equipment) are integrated, and the tasks of information sharing, attitude control, task scheduling and the like of the whole satellite are realized.
Furthermore, the micro-satellite system in the invention takes micro technology and intelligent technology as the core, so that the satellite platform and the satellite effective load develop towards miniaturization and light weight on the basis of keeping the original various functions, the weight of the satellite is reduced, the development cost is reduced, the development period of the satellite is shortened, and the like. Due to the adoption of a system and subsystem integration technology, the integration of electronic equipment and functions of the whole satellite can be realized, the volume and the weight of the whole satellite are reduced by 60-70%, and the number of printed boards is reduced by 70-75%.
In addition, the microsatellite system adopts a high-reliability dual-redundancy two-stage safety strategy, namely a satellite management and control dual-redundancy architecture and two-stage fault tolerance (the on-board management computer and the on-board control computer are mutually redundant, full-function backup is ensured when one computer fails, and a GNC bottom layer controller ensures the safety of the system when the two computers fail), which is close to the safety standard of a large satellite.
Drawings
In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings needed to be used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts.
FIG. 1 shows a domain-based microsatellite system architecture in an embodiment of the present invention.
Fig. 2 shows a schematic structural diagram of a power supply unit in the embodiment of the invention.
Fig. 3 shows a schematic diagram of the structure of the integrated information unit in the embodiment of the present invention.
FIG. 4 is a block diagram illustrating the overall satellite information flow of a domain-based microsatellite system in an embodiment of the present invention.
Detailed Description
In order to make those skilled in the art better understand the technical solution of the present invention, the technical solution in the embodiment of the present invention will be clearly and completely described below with reference to the drawings in the embodiment of the present invention. It is to be understood that the described embodiments are merely exemplary of a portion of the invention and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that the terms "first", "second", and the like in the description and claims of the present invention and in the above drawings are used for distinguishing objects, and are not necessarily used for describing a specific order or sequence. It is to be understood that the objects so used are interchangeable under appropriate circumstances such that the embodiments of the invention described herein are capable of operation in sequences other than those illustrated or otherwise described herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.
As shown in FIG. 1, the present invention provides a domain-based microsatellite system which comprises a plurality of functional units arranged in a distributed manner, wherein the plurality of functional units comprise a structural unit, a communication unit, a comprehensive information unit, a guidance, navigation and control (GNC) unit, a load unit, a thermal control unit and a power supply unit.
Specifically, the structural unit is used for providing installation and support conditions for each functional unit of the microsatellite; the communication and measurement unit is used for receiving a ground instruction and forwarding the ground instruction to the comprehensive information unit so as to realize the functions of ground measurement and control, ground data transmission and satellite navigation; the comprehensive information unit is used for receiving the ground instruction forwarded by the communication and conducting unit and realizing the power distribution function, initiating explosive device control, heater control, temperature acquisition, bus remote measurement, star management and GNC software support tasks of the whole satellite based on the ground instruction; the GNC unit comprises a GNC measuring unit and a GNC executing unit, wherein the GNC measuring unit is used for collecting attitude information of the whole star, and the GNC executing unit is used for finishing attitude and orbit control; the thermal control unit is used for providing a working temperature environment for the microsatellite system in a task period; the load unit is used for completing an on-orbit test task according to a preset flow based on the ground instruction; and the power supply unit is used for providing a direct current power supply bus for the microsatellite system in the service life of the whole satellite so as to complete the power transmission of the whole satellite.
The invention divides the microsatellite system into different functional units according to the task function and the supporting function, comprising: forming a comprehensive information unit by integrating the electronic control part; forming a power supply unit by integrating energy storage and control; the measurement and control and navigation are integrated to form a measurement and communication unit; the mechanical part and the sensor are integrated to form a GNC measuring and executing unit; and respectively forming a structural unit, a thermal control unit and a load unit by the structure, the thermal control and the load.
Based on a functional domain design concept, the invention adopts a miniaturized, standardized and modularized design method to perform technical analysis and design optimization of the whole satellite system architecture of the microsatellite, divides the satellite into different functional modules or functional areas according to task functions and support functions, breaks through the physical boundary of the system/subsystem/component in the traditional microsatellite, and solves the problem that the microsatellite is difficult to combine multiple functions with high density, functionalization and centralization.
In the embodiment of the invention, the structural unit provides installation and support conditions for each functional unit of the microsatellite, for example, the solar sailboard can be reliably unfolded and locked.
In order to obtain structural rigidity and strength meeting requirements, the design criteria of force flow continuity and shortest force transmission path can be taken as well as the overall thermal control design requirements of the platform and the installation requirements of all devices and effective loads. In one embodiment, the microsatellite platform adopts a traditional cubic star cabin scheme, the interior of the microsatellite platform is divided into different functional areas according to functions, and the unfolded solar wings are arranged on the outer sides of the side plates of the cabin.
In the embodiment of the invention, the GNC unit comprises a GNC measuring unit and a GNC executing unit, wherein the GNC measuring unit comprises a micro-magnetic sensor, a micro-star sensor and a micro-space sensor, and the GNC executing unit comprises a micro-propulsion device, a micro-flywheel and a micro-magnetic torquer.
In one embodiment, the GNC measurement unit is used for a measurement part responsible for performing measurement and calculation of information such as satellite spatial position, speed, attitude, time and the like, and transmitting the information to the comprehensive information unit for generating an orbit maintenance and attitude stabilization control instruction, and comprises a star sensor, a sun sensor, a magnetometer and a navigation receiver. The GNC unit execution part is responsible for completing the three-axis attitude stabilization and attitude control of the satellite and comprises a flywheel, a magnetic torquer and a micro-motor.
In the embodiment of the invention, the thermal control unit is responsible for providing a good working temperature environment for the satellite products in the whole task period, mainly adopts a passive thermal control mode, and has the function of assisting the active thermal control for key equipment such as a storage battery pack with narrow temperature indexes, calibration loads and the like.
In the embodiment of the invention, the communication and measurement unit consists of a baseband module, a radio frequency module and an antenna and is used for completing the functions of ground measurement and control, ground data transmission and satellite navigation. In one embodiment, the system comprises a measurement and control data transmission terminal, a remote control receiving antenna, a remote measurement data transmission antenna, a radio frequency cable and the like. Specifically, the communication measurement and conduction unit is composed of a pair of measurement and control receiving and transmitting antennas, a pair of data transmission transmitting antennas, a pair of satellite navigation receiving antennas, a satellite navigation receiving unit, a radio frequency module, a signal processing module and a secondary power supply, and a software radio equipment framework is selected for hardware scheme design. The radio frequency module, the satellite navigation receiving unit, the signal processing module and the secondary power supply form a body of the measurement and control terminal, and the body is installed on the inner wall of the aircraft. A pair of measurement and control receiving antennas are symmetrically arranged on the opposite sky surface and the opposite ground surface of the aircraft, a satellite navigation receiving antenna is arranged on the opposite sky surface of the aircraft, and a data transmission transmitting antenna is arranged on the opposite ground surface of the aircraft.
In the embodiment of the invention, the load cell is also called a test load cell and is mainly used for scientific tests.
As shown in fig. 2, in the embodiment of the present invention, the power supply unit is responsible for providing a reliable dc power supply bus for the whole satellite during the whole satellite life, and completes power transmission of the whole satellite by using the cable network through the integrated electronic interface unit, and is composed of a solar cell array, a storage battery pack, a power supply controller, a cable network, and the like.
In a specific embodiment, the power supply unit mainly provides sufficient electric energy for the whole satellite to work, and mainly comprises a solar cell array, a storage battery pack and a power supply controller. The solar cell array is used for converting solar light energy into electric energy through photoelectric conversion, and a body-mounted triple-junction gallium arsenide cell is adopted; the storage battery pack is used for storing energy and mainly provides energy for a load when the power provided by the shadow area or the peak power consumption solar cell array is insufficient; and the power supply controller is used for controlling the working state of the solar cell array, the charge and discharge management of the storage battery pack and the power supply change. Specifically, the battery pack may be a lithium battery pack. The power controller includes power regulation, power distribution, and battery management.
As shown in fig. 3, in the embodiment of the present invention, the integrated information unit is used to complete the tasks of the power distribution function of the whole satellite, initiating explosive device control, heater control, temperature acquisition, bus telemetry, star management, GNC software support, and the like.
In a specific embodiment, the integrated information unit is a core unit for controlling and managing the whole satellite equipment, and consists of a core module and an interface module, and the basic functions of the integrated information unit are to manage and schedule the whole satellite information, monitor the state of each subsystem, autonomously manage the work of each subsystem and telemetering data and state information, and provide a data control execution component according to the GNC system, so that the safe and reliable operation of the whole satellite system is realized.
Aiming at the application of limited power and volume in a space environment, in order to realize the construction of a comprehensive information unit system with low power consumption, small volume, high integration degree, high performance and high reliability, the invention adopts a sub-modular design idea to integrate specific functional components on different functional board cards. The connection relationship between the host processing function board card and the interface function board card and the internal structure module are shown in fig. 3. Three functional board cards are designed in the system: the host module, the standby module and the interface function board card are connected with each other.
The host function module is a core control part for data management of the integrated information unit, an ARM processor is integrated on the board card to carry out unified scheduling management on satellite resources, integration and processing of telemetering data are achieved, the state of the whole satellite is monitored, instruction control on satellite equipment is achieved, power distribution, thermal control, time correction and safety of all satellite subsystems are managed, timely and effective processing is carried out on a part with a fault, and safe and reliable operation of the satellite system is guaranteed. The board card is also integrated with a data storage module and a data acquisition module.
The interface function board card is a hub for information exchange between the comprehensive information unit and all the subsystems of the whole satellite and provides electrical connection for the comprehensive information unit and other parts on the satellite. The board card provides a CAN bus interface for the transmission of control instructions and telemetering data, and also provides a heating drive control interface, an initiating explosive device line interface, an OC instruction interface and a power supply module, so that the storage of satellite telemetering data and state information and the power supply of the whole comprehensive information unit are realized. Because the integrated information unit interface is more complex and more in number, the single integrated interface board card can effectively reduce the whole volume of the unit and improve the expansibility of the whole satellite system.
In one embodiment, the integrated information unit may be divided into 5 basic function modules: the system comprises a central processing unit, a minimum system, an interface adaptation module, a data storage module, a power supply adaptation module and an inter-board interconnection module. The invention adopts ATSAMV71Q21 as the central processing unit of the system, and can realize the management and the dispatching of telemetering data, the monitoring of the whole star state and the star management of fault processing.
Specifically, the ARM data processing unit includes: the system comprises an on-board autonomous management control program, fault-tolerant satellite-borne satellite processing and attitude orbit control software, an embedded real-time operating system, an off-chip memory controller and an interface controller. The on-board autonomous management control program performs data processing and task scheduling on fault-tolerant satellite-borne satellite processing and attitude orbit control software, the fault-tolerant satellite-borne satellite processing sends data, state information and a control instruction to the attitude orbit control software to an embedded real-time operating system, the embedded real-time operating system returns the data and the state information to the attitude orbit control software to the fault-tolerant satellite-borne satellite processing, the embedded real-time operating system sends the data and the state information to an off-chip memory controller, the interface controller performs data interaction with the embedded real-time operating system, receives the control instruction from the embedded real-time operating system and sends the state information to the embedded real-time operating system. The data storage function module comprises an off-chip memory controller and an off-chip data storage unit, and the communication interface function module comprises an interface controller and a communication interface circuit.
It should be noted that the ATSAMV71Q21 CPU is a commercially available CPU, such as available from Shenzhen Dry age technology, Inc. The ARM processor based on the ARM Cortex-M732-bit kernel has the maximum operating frequency of 300MHz, adopts a low-power-consumption design, has the operating power consumption of 0.5W, is integrated with abundant I/O resources on a chip, and can be expanded into different types of interfaces. In order to realize normal startup and operation of the ATSAMV71Q21 microprocessor, a minimum system is required to be designed, and the minimum system comprises a power supply circuit, a clock circuit, a reset circuit and a Joint Test Action Group (JTAG) debugging circuit.
The interface adaptation module CAN realize the physical connection between the comprehensive management unit and all the devices of the whole satellite, and provides a CAN bus interface to realize the receiving and sending of the telemetering data and the control instruction on the on-satellite bus. In addition, an OC control instruction interface is provided for sending instructions to the whole satellite equipment, and a heating driving interface is provided for controlling the heating of the whole satellite equipment. And the power supply module is used for storing the satellite telemetering data and state information and supplying power to the whole comprehensive information unit.
The data storage module stores the telemetering data and the state information acquired by the whole satellite. The ATSAMV71Q21 is internally provided with an MMC interface, and can realize the reading and storage control of an off-chip SD card.
The power supply of the ATSAMV71Q21 minimum system is derived from a DC-DC conversion chip LMZ10503EXT of an interface board, and the LMZ10503EXT can convert the externally input 5V voltage into 3.3V for the system to use. It should be noted that the LMZ10503EXT chip is a commercially available chip, such as available from shenzhen shenshenshenshenshenshengsheng industry limited.
In the embodiment of the invention, a high-reliability dual-redundancy two-stage security strategy is adopted. Specifically, a multi-level heterogeneous backup mechanism is adopted on hardware. Aiming at satellite-borne software data flow errors and control flow errors caused by single-event upset, a dual-core mutual detection and rollback recovery technology is adopted as a first-level fault-tolerant mechanism, software EDAC and triple-modular redundancy technology aiming at the data flow errors, and an abnormal trap and watchdog technology aiming at the control flow errors are adopted as a second-level fault-tolerant mechanism, so that the detection and recovery of soft errors are realized.
In the embodiment of the invention, each functional unit takes a CAN bus as a backbone communication network.
As shown in fig. 4, the domain-based microsatellite system whole satellite information flow in the embodiment of the present invention mainly includes a whole satellite CAN bus and each functional unit performing information interaction with the CAN bus.
Specifically, the power supply unit supplies electric energy to each functional unit through the CAN bus.
The information synthesis unit comprises a master processing and a backup processing which are positioned on the mainboard and a monitoring system which is positioned on the interface board. Wherein the master processing system, the backup processing system and the monitoring system are respectively and electrically connected with the CAN bus. The monitoring system acquires status information of the master processing and/or backup processing and provides a reset signal to the master processing/backup processing. The information integration unit is also used for temperature acquisition and thermal control output.
The communication and conduction unit is electrically connected with the CAN bus and is electrically connected with the comprehensive information unit through the 485 bus so as to carry out instruction/signal interaction.
The load unit is electrically connected with the CAN bus, and is used for sending the working state of the load unit to the comprehensive information unit and receiving the instruction sent by the comprehensive information unit.
The GNC unit comprises a star sensor, a GNC drive control and an electric propulsion (including a controller) which are respectively electrically connected with the CAN bus. And the GNC drive control receives the signal of the comprehensive information unit and performs power on/off control on the star sensor and the electric propulsion. The GNC drive control also receives inertial navigation, micro flywheel actuating mechanism and space sensitive information, sends an instruction to the micro flywheel actuating mechanism and controls the magnetic torquer.
The load unit, the GNC unit and the propulsion module in the embodiment of the invention are independently designed, and adopt standardized interfaces, so that the customization and reconfiguration are easy, and the task adaptability of the system can be enhanced. In addition, a standard electrical interface is used, such as a PC104 connector.
In the embodiment of the invention, a structure and function integration technology is adopted, three-dimensional heterogeneous integration is utilized, a traditional satellite assembly integration method is overturned, and a main structure and a functional unit structure body are integrally designed. The mode of direct series connection by adopting a standardized unit electromechanical interface is adopted. The main characteristics include:
1) and bearing the system load by using the unit supporting structure.
2) The board-level longitudinal plug-in realizes the electrical and structural synchronous interconnection, simplifies the interconnection mode among units, improves the connection reliability and improves the assembly efficiency.
3) The circuit substrate for transmitting signals is manufactured by adopting multilayer composite dielectric materials in load design, and partial functional circuits are integrated in the multilayer substrate and can transmit microwave signals and digital signals; by adopting a three-dimensional vertical interconnection technology of a substrate, connecting wires are replaced by technical means such as fuzz buttons, micro-bumps and the like, and auxiliary structural support is arranged between boards, so that vertical interconnection between circuit boards in the aspects of machinery and electricity can be realized, interconnection between cables, connectors and the like among original subsystem components is reduced, the structure and the cable quality are reduced, and the system integration level is greatly improved.
4) The disassembly and the assembly are convenient. Can be assembled quickly and manufactured automatically in batch.
The microsatellite system simplifies interfaces and a development unit labor division interface, redefines units, adopts a distributed network architecture, is set as a highly integrated system for uniformly scheduling and managing tasks, functions and resources, adopts a centralized, expandable and reducible general architecture design, integrates the functions of the whole satellite (or part of) electronic equipment, and realizes the tasks of information sharing, attitude control, task scheduling and the like of the whole satellite. Due to the adoption of a multi-center and bus type network, the coupling among units can be reduced, and the working efficiency is improved.
Furthermore, the micro-satellite system in the invention takes micro technology and intelligent technology as the core, so that the satellite platform and the satellite effective load develop towards miniaturization and light weight on the basis of keeping the original various functions, the weight of the satellite is reduced, the development cost is reduced, the development period of the satellite is shortened, and the like. Due to the adoption of a system and subsystem integration technology, the integration of electronic equipment and functions of the whole satellite can be realized, the volume and the weight of the whole satellite are reduced by 60-70%, the number of printed boards is reduced by 70-75%, risks can be dispersed, and the safety is improved.
In addition, the microsatellite system adopts a high-reliability dual-redundancy two-stage safety strategy, namely a satellite management and control dual-redundancy architecture and two-stage fault tolerance (the on-board management computer and the on-board control computer are mutually redundant, full-function backup is ensured when one computer fails, and a GNC bottom layer controller ensures the safety of the system when the two computers fail), which is close to the safety standard of a large satellite.
It should be understood that the above-mentioned embodiments of the present invention are only examples for clearly illustrating the present invention, and are not intended to limit the embodiments of the present invention, and it will be obvious to those skilled in the art that other variations or modifications may be made on the basis of the above description, and all embodiments may not be exhaustive, and all obvious variations or modifications may be included within the scope of the present invention.
Claims (10)
1. A domain-based microsatellite system is characterized by comprising a plurality of functional units which are arranged in a distributed mode, wherein the functional units comprise a structural unit, a communication and conduction unit, a comprehensive information unit, a GNC unit, a load unit, a thermal control unit and a power supply unit, and the system is characterized in that
The structural unit is used for providing installation and support conditions for each functional unit of the microsatellite;
the communication and measurement unit is used for receiving a ground instruction and forwarding the ground instruction to the comprehensive information unit so as to realize the functions of ground measurement and control, ground data transmission and satellite navigation;
the comprehensive information unit is used for receiving the ground instruction forwarded by the communication and conducting unit and realizing the power distribution function, initiating explosive device control, heater control, temperature acquisition, bus remote measurement, star management and GNC software support tasks of the whole satellite based on the ground instruction;
the GNC unit comprises a GNC measuring unit and a GNC executing unit, wherein the GNC measuring unit is used for collecting attitude information of the whole star, and the GNC executing unit is used for finishing attitude and orbit control;
the thermal control unit is used for providing a working temperature environment for the microsatellite system in a task period;
the load unit is used for completing an on-orbit test task according to a preset flow based on the ground instruction; and
and the power supply unit is used for providing a direct current power supply bus for the microsatellite system in the service life of the whole satellite so as to finish power transmission of the whole satellite.
2. A microsatellite system as recited in claim 1 wherein each of said functional units has a CAN bus as a backbone communication network.
3. The microsatellite system according to claim 1 wherein said integrated information unit is further used for managing and scheduling satellite information, monitoring the status of subsystems, autonomously managing the operation and telemetry data and status information of subsystems, and controlling the execution units based on the data provided by the GNC unit, said integrated information unit comprising a host module, a standby module and an interface function board, wherein said integrated information unit comprises a host module, a standby module and an interface function board
The host module is integrated with a processor and is used for carrying out unified scheduling management on satellite resources so as to realize integration and processing of telemetering data, monitoring the satellite state so as to realize instruction control on satellite equipment, managing power distribution, thermal control, time correction and safety of all satellite subsystems and processing a part with a fault in time;
the standby module is configured as a hardware heterogeneous backup of the host module; and
the interface function board card is integrated with a data storage module and a data acquisition module and is used for providing electrical connection for the comprehensive information unit and providing a CAN bus interface, a heating drive control interface, an initiating explosive device line interface and an OC instruction interface.
4. The microsatellite system according to claim 3 wherein said integrated information unit employs a first level fault tolerance mechanism and a second level fault tolerance mechanism wherein
The first-level fault-tolerant mechanism comprises a dual-core mutual detection technology and a rollback recovery technology; and
the second level fault tolerance mechanism includes software EDAC and triple modular redundancy techniques for data flow errors, and exception traps and watchdog techniques for control flow errors.
5. A microsatellite system according to claim 1,
the GNC measuring unit comprises a star sensor, a sun sensor, a magnetometer and a navigation receiving mechanism and is used for acquiring attitude information of the whole star, measuring and calculating spatial position, speed, attitude and time information of a satellite, and transmitting the information to the comprehensive information unit to generate an orbit maintaining and attitude stabilizing control instruction; and
the GNC execution unit comprises a micro-propulsion unit, a micro flywheel and a micro magnetic torquer and is used for realizing the three-axis attitude stabilization and attitude control of the satellite.
6. The microsatellite system according to claim 1 wherein the communication unit comprises a measurement and control terminal body, a pair of measurement and control transceiver antennas, a data transmission transmitting antenna and a satellite navigation receiving antenna, wherein
The measurement and control terminal body comprises a radio frequency module, a satellite navigation receiving unit, a signal processing module and a secondary power supply, wherein the radio frequency module, the satellite navigation receiving unit, the signal processing module and the secondary power supply are arranged on the inner wall of a satellite;
the pair of measurement and control transmitting and receiving antennas are symmetrically arranged on the sky surface and the ground surface of the satellite;
the satellite navigation receiving antenna is arranged on the opposite sky surface of the satellite; and
the secondary data transmission transmitting antenna is arranged on the ground of the satellite.
7. The microsatellite system according to claim 1 wherein said power supply unit includes a solar array, a battery pack and a power supply controller wherein
The solar cell array comprises a body-mounted three-junction gallium arsenide cell slice and is used for converting solar light energy into electric energy through photoelectric conversion;
the storage battery pack is used for providing energy for a load when the power provided by the shadow area or the peak power consumption solar cell array is insufficient; and
the power supply controller is used for controlling the working state of the solar cell array, the charge and discharge management of the storage battery pack and the power supply change.
8. A microsatellite system according to any one of claims 1 to 7 wherein each functional unit uses a standard electrical interface.
9. A microsatellite system according to any one of claims 1 to 7 wherein the functional units are integrated in a three dimensional heterogeneous integration using an integrated design.
10. A microsatellite system according to any one of claims 1 to 7 wherein the functional units are electrically and structurally interconnected in synchronism by board level vertical mating.
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113815904A (en) * | 2021-09-06 | 2021-12-21 | 中国科学院微小卫星创新研究院 | Modularized energy system capable of on-orbit maintenance and replacement |
| CN114513244A (en) * | 2022-01-14 | 2022-05-17 | 上海卫星工程研究所 | Spacecraft integrated intelligent information system |
| CN115857586A (en) * | 2023-02-20 | 2023-03-28 | 银河航天(北京)网络技术有限公司 | Initiating explosive device temperature control system and temperature control method |
| CN116192809A (en) * | 2023-03-02 | 2023-05-30 | 苏州泰富晶宇科技有限公司 | Dual-machine cold-standby highly-reliable star service and attitude control interaction architecture system |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| BR9813884A (en) * | 1997-10-27 | 2000-09-26 | Discovery Semiconductors Inc | Microsatellite with integrated circuit, chip packaging system with integrated circuit and photolithographic spiral antenna |
| CN102431659A (en) * | 2011-10-27 | 2012-05-02 | 北京航空航天大学 | Halo orbit detector structure of sun-earth system and attitude directing thereof |
| CN102795350A (en) * | 2012-07-06 | 2012-11-28 | 中国航天科技集团公司第五研究院第五一三研究所 | Physical discrete type satellite system structure |
| CN103699069A (en) * | 2013-12-06 | 2014-04-02 | 上海卫星工程研究所 | Advanced electronic integrated system for microsatellite |
| CN104015939A (en) * | 2014-05-26 | 2014-09-03 | 中国科学院长春光学精密机械与物理研究所 | Comprehensive management system for platform and load integrated satellite |
| CN105045672A (en) * | 2015-07-24 | 2015-11-11 | 哈尔滨工业大学 | Multilevel fault tolerance reinforcement satellite information processing system based on SRAM FPGA |
| CN105501471A (en) * | 2015-12-16 | 2016-04-20 | 上海卫星工程研究所 | Configuration of satellite loaded with large deployable antenna with double reflecting surfaces |
| CN106542117A (en) * | 2016-10-17 | 2017-03-29 | 中国运载火箭技术研究院 | A kind of micro- integrated aircraft untethered electrical system |
| CN109131939A (en) * | 2018-07-20 | 2019-01-04 | 中国空间技术研究院 | A kind of satellite equipment layout method and system based on electromagnetic radiation environment |
-
2020
- 2020-07-31 CN CN202010769230.XA patent/CN112046787A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| BR9813884A (en) * | 1997-10-27 | 2000-09-26 | Discovery Semiconductors Inc | Microsatellite with integrated circuit, chip packaging system with integrated circuit and photolithographic spiral antenna |
| CN102431659A (en) * | 2011-10-27 | 2012-05-02 | 北京航空航天大学 | Halo orbit detector structure of sun-earth system and attitude directing thereof |
| CN102795350A (en) * | 2012-07-06 | 2012-11-28 | 中国航天科技集团公司第五研究院第五一三研究所 | Physical discrete type satellite system structure |
| CN103699069A (en) * | 2013-12-06 | 2014-04-02 | 上海卫星工程研究所 | Advanced electronic integrated system for microsatellite |
| CN104015939A (en) * | 2014-05-26 | 2014-09-03 | 中国科学院长春光学精密机械与物理研究所 | Comprehensive management system for platform and load integrated satellite |
| CN105045672A (en) * | 2015-07-24 | 2015-11-11 | 哈尔滨工业大学 | Multilevel fault tolerance reinforcement satellite information processing system based on SRAM FPGA |
| CN105501471A (en) * | 2015-12-16 | 2016-04-20 | 上海卫星工程研究所 | Configuration of satellite loaded with large deployable antenna with double reflecting surfaces |
| CN106542117A (en) * | 2016-10-17 | 2017-03-29 | 中国运载火箭技术研究院 | A kind of micro- integrated aircraft untethered electrical system |
| CN109131939A (en) * | 2018-07-20 | 2019-01-04 | 中国空间技术研究院 | A kind of satellite equipment layout method and system based on electromagnetic radiation environment |
Non-Patent Citations (3)
| Title |
|---|
| 张宏德等: "小卫星对上面级定向的姿控设计与验证", 《计算机测量与控制》 * |
| 徐喆垚等: "LH-2 微小卫星综合电子系统设计及姿态控制研究", 《载人航天》 * |
| 詹盼盼等: "卫星高功能密度综合电子系统设计", 《中国空间科学技术》 * |
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113815904A (en) * | 2021-09-06 | 2021-12-21 | 中国科学院微小卫星创新研究院 | Modularized energy system capable of on-orbit maintenance and replacement |
| CN113815904B (en) * | 2021-09-06 | 2023-11-10 | 中国科学院微小卫星创新研究院 | Modularized energy system capable of being maintained and replaced on orbit |
| CN114513244A (en) * | 2022-01-14 | 2022-05-17 | 上海卫星工程研究所 | Spacecraft integrated intelligent information system |
| CN114513244B (en) * | 2022-01-14 | 2023-09-26 | 上海卫星工程研究所 | Integrated intelligent information system for spacecraft |
| CN115857586A (en) * | 2023-02-20 | 2023-03-28 | 银河航天(北京)网络技术有限公司 | Initiating explosive device temperature control system and temperature control method |
| CN116192809A (en) * | 2023-03-02 | 2023-05-30 | 苏州泰富晶宇科技有限公司 | Dual-machine cold-standby highly-reliable star service and attitude control interaction architecture system |
| CN116192809B (en) * | 2023-03-02 | 2024-06-04 | 苏州泰富晶宇科技有限公司 | Dual-machine cold-standby highly-reliable star service and attitude control interaction architecture system |
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