CN111307336B - Spacecraft multichannel continuous rotary motion temperature measurement system and method - Google Patents
Spacecraft multichannel continuous rotary motion temperature measurement system and method Download PDFInfo
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- G01K13/00—Thermometers specially adapted for specific purposes
- G01K13/04—Thermometers specially adapted for specific purposes for measuring temperature of moving solid bodies
- G01K13/08—Thermometers specially adapted for specific purposes for measuring temperature of moving solid bodies in rotary movement
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Abstract
The invention discloses a spacecraft multichannel continuous rotary motion temperature measurement system in a vacuum cold and black environment, which mainly comprises a sensor interface unit, a normal pressure environment unit, a temperature measurement unit and a signal transmission unit, wherein the sensor interface unit is used for connecting a sensor on a test piece in a vacuum thermal test with the normal pressure environment unit; the normal pressure environment unit is used for providing a simulated atmospheric environment for the temperature measuring unit and the signal transmission unit; the temperature measuring unit is used for completing acquisition, conversion, storage and transmission of sensor signals on a vacuum thermal test piece and keeping synchronous motion with the test piece; the signal transmission unit is used for transmitting the temperature signal subjected to conversion processing from the rotating state to the fixed channel so as to be transmitted out of the space environment simulator. The invention carries out local automatic measurement and data storage under the condition of communication interruption of the slip ring, ensures continuous and reliable operation of temperature measurement and ensures that the spacecraft adopts the solar simulator to simulate external heat flow to carry out a vacuum thermal test and smoothly and reliably operate.
Description
Technical Field
The invention belongs to the technical field of space environment simulation, and particularly relates to a system and a method for carrying out multichannel temperature measurement on a continuously rotating motion test piece in a vacuum cold black environment.
Background
After the artificial satellite and other spacecrafts enter the orbit flight phase, the artificial satellite and other spacecrafts are in the ultra-high vacuum and ultra-low temperature environment for a long time and are influenced by the heat flow environment outside the space, and in order to verify the correctness of the thermal design of the satellite, ensure the long-term reliable work of the satellite and complete various preset tasks, in the process of developing the satellite, the artificial satellite and other spacecrafts must be subjected to test rulesIn the specification, the space is vacuum (negative pressure environment, the environmental pressure is 10)-5～10-7Pa) and cold and black (no illumination and low temperature environment, the temperature is-100 to-270 ℃) and vacuum thermal tests (including thermal vacuum and thermal balance tests) are carried out on the spacecraft system level and the component level under the simulation of heat flow outside the space.
When a solar simulation method with a solar simulator is adopted to simulate the heat flow outside the space, the system structure of the solar simulator determines that the spacecraft needs to perform 360-degree continuous rotation motion to simulate the attitude of the spacecraft when the spacecraft operates on the orbit. The method is different from an infrared simulation method for fixing a tested piece, the requirement for continuous rotation of the spacecraft is increased, a conventional temperature measurement system for a stationary spacecraft needs to be changed into a method capable of carrying out temperature data acquisition and signal processing on the tested piece of the continuously rotating spacecraft, signals are transmitted to the outside of a container to carry out spacecraft temperature data acquisition and detection, and the integrity and effectiveness of data acquired by a spacecraft vacuum thermal test are ensured.
In normal temperature and normal pressure environment, the rotary power supply and the conventional signal transmission can be realized by installing rotary transmission parts such as a slip ring, a rotary joint and the like in a transmission cable. However, in the process of a spacecraft thermal vacuum test, a tested piece is always in a vacuum black environment, and a temperature measurement sensor is a small signal sensor such as a multi-purpose thermocouple, and a millivolt-level voltage signal of the small signal sensor cannot pass through a slip ring to be accurately measured.
Therefore, it is very necessary to design a temperature measurement system and a temperature measurement method that can be used in a vacuum cold-black environment, and can realize reliable acquisition and data transmission of multi-channel high-precision temperature data required by a vacuum thermal experiment.
Disclosure of Invention
Based on the above, the invention aims to provide a temperature measurement system for multi-channel continuous rotary motion of a spacecraft in a vacuum cold and black environment, which can perform local automatic measurement and data storage, ensure continuous and reliable operation of temperature measurement, and ensure smooth and reliable operation of a vacuum thermal test performed by the spacecraft by adopting a solar simulator to simulate external heat flow.
The invention also aims to provide a method for measuring the temperature of the spacecraft in the vacuum cold and black environment through the multi-channel continuous rotary motion, which can still realize the reliable acquisition and data transmission of multi-channel high-precision temperature data required by a vacuum thermal test when the spacecraft performs the continuous rotary motion, so that the spacecraft can complete the vacuum thermal test for simulating the heat flow outside the space by adopting a solar simulation method of a solar simulator.
The purpose of the invention is realized by the following technical scheme:
the spacecraft multichannel continuous rotary motion temperature measurement system under the vacuum black environment mainly comprises a sensor interface unit (100), a normal pressure environment unit (200), a temperature measurement unit (300) and a signal transmission unit (400), wherein the sensor interface unit (100) is mainly used for connecting a sensor on a tested spacecraft or a test piece in a vacuum thermal test with the normal pressure environment unit, and connecting a sensor signal to the temperature measurement unit (300) for temperature measurement through the normal pressure environment unit (200); the normal-pressure environment unit (200) is mainly used for providing simulated atmospheric environment for the temperature measurement unit (300) and the signal transmission unit (400) and ensuring the normal work of the temperature measurement unit (300) and the signal transmission unit (400); the temperature measuring unit (300) is mainly used for collecting, converting, storing and transmitting sensor signals on a tested spacecraft and a test piece in a vacuum thermal test, and the temperature measuring unit (300) is installed on a rotating unit (210) in a normal-pressure environment unit (200), wherein the normal-temperature normal-pressure environment unit is provided with a rotating unit and a fixing unit, the fixing unit is connected with an environment simulator, the rotating unit realizes rotating motion and is rigidly connected with the tested spacecraft and keeps synchronous motion with the tested spacecraft, and the accuracy and precision of temperature measurement are ensured; and the device is in a normal pressure environment, so that the normal work of the instrument and equipment is ensured; the signal transmission unit (400) is mainly used for transmitting the temperature signals converted and processed by the temperature measurement unit (300) to a fixed channel from a rotating state so as to be transmitted to the outside of the space environment simulator for real-time monitoring and providing power supply necessary for operation for the temperature measurement unit (300)).
Wherein the sensor interface unit (100): the temperature sensor signal transmission device is mainly used for transmitting a temperature sensor signal on a spacecraft to a rotating unit (210), introducing the signal into a normal-temperature normal-pressure environment unit (200) through a signal through-wall sealing unit (223), and finally accessing a temperature measuring unit (300) for temperature data acquisition.
Wherein the ambient unit (200) is: the temperature measurement device is mainly used for providing a simulated atmosphere environment for the temperature measurement unit (300) and the signal transmission unit (400) and ensuring the working environment of the temperature measurement device, and mainly comprises: the device comprises a rotating unit (210), a fixing unit (220), a vacuum sealing unit (230) and an environment simulation unit (240), wherein the rotating unit (210) and the fixing unit (220) are main body parts of a normal-temperature normal-pressure environment unit structure and are used as normal-pressure environment simulation cabins to connect a satellite test piece and a vacuum environment simulator; the vacuum sealing unit (230) is mainly used for providing vacuum sealing for the simulation cabin in the normal-temperature and normal-pressure environment, and the environment simulation unit (240) is used for maintaining and controlling the normal-temperature and normal-pressure environment of the simulation cabin.
The rotating unit (210) and the spacecraft to be tested rotate synchronously, a signal transmission interface is provided for the sensor, and an installation interface is provided for the temperature measuring unit (300).
Wherein the fixed unit (220) is arranged in the space environment simulator, is kept still, provides an installation interface for the rotating unit and provides an interface for transmitting signals to the outside of the vacuum environment simulator.
The vacuum sealing unit (230) mainly comprises a dynamic sealing unit (231), a fixed sealing unit (232) and a signal through-wall sealing unit (233), wherein the dynamic sealing unit (231) provides vacuum sealing for a dynamic sealing surface connected with the rotating unit (210) and the fixed unit (220), and three sealing mechanisms are adopted to ensure that the leakage rate can meet the sealing requirement at the connecting part between the rotating unit and the fixed unit when the rotating unit moves and is static; the fixed sealing unit (232) is used for vacuum sealing of the rotating unit (210) and the cabin of the fixed unit (220); the signal through-wall sealing unit (233) is arranged on the rotating unit and used for introducing a sensor signal transmitted by the sensor interface unit (100) into the room temperature and normal pressure environment chamber, providing 1000 single-wire system measuring channels and 60 four-wire system measuring channels, and meeting the requirement of a large-data-volume sensor temperature measuring channel of a large-scale spacecraft vacuum thermal test.
The environment simulation unit (240) comprises a pressure simulation unit (241), a temperature simulation unit (242) and a humidity simulation unit (243), wherein the pressure simulation unit (241) obtains pressure information in a normal-temperature normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the air pressure of an air supply pipeline between the normal-pressure cabin and the space environment simulator, and realizes the function of simulating atmospheric pressure in the normal-pressure cabin; the temperature simulation unit (242) obtains temperature information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the gas temperature of an air supply pipe opening between the normal-pressure cabin and the space environment simulator, and realizes the function of simulating the atmospheric temperature in the normal-pressure cabin; the humidity simulation unit (243) obtains humidity information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the air humidity of an air supply pipe opening between the normal-pressure cabin and the space environment simulator, and finally realizes the function of simulating atmospheric humidity in the normal-pressure cabin.
The measuring unit (300) is integrally arranged in the rotating unit (210) and is used for carrying out data acquisition, signal data processing and data storage and transmission on signals of the temperature sensor on the spacecraft in the vacuum thermal test.
Further, the measuring unit (300) comprises a small signal interface unit (310), a data acquisition unit (320), a data processing unit (330), a transmission interface unit (340) and a data storage unit (350), wherein the small signal interface unit (310) is used for receiving signals of the temperature sensor on the spacecraft to be measured, which are transmitted by the sensor interface unit (100), and accessing the signals to the data acquisition unit (320) for data acquisition, and the data acquisition unit (320) is used for performing data acquisition on the accessed sensor signals and performing digital-to-analog conversion on small signal analog quantities generated by the sensor to obtain digital quantity signals; the data processing unit (330) is used for processing and fitting the sensor signals acquired by the data acquisition unit (320) to obtain real-time temperature data information of each part of the spacecraft; the transmission interface unit (340) is used for transmitting the sensor signal acquired by the data acquisition unit (320) and the temperature signal processed by the data processing unit (330), and converting the rotary motion signal into a fixed signal through the signal transmission unit (400); the data storage unit (350) is mainly used for locally storing the data generated by the data acquisition unit (320) and the data processing unit (330) in real time.
The transmission interface unit (340) transmits small voltage below 100mV, does not interfere with signals of the small voltage, provides 1000 single-wire system measurement channels and 60 four-wire system measurement channels, and meets the requirement of a large data sensor temperature measurement channel for a large spacecraft vacuum thermal test.
The acquisition precision of the data acquisition unit (320) is not lower than 6 bits and half (or the AD conversion precision is not lower than 24 bits), and the acquisition module can complete data acquisition of the thermocouple, the thermistor and the platinum resistor.
Furthermore, the data acquisition unit (320) adopts a multi-way change-over switch to switch the acquisition channels, and sequentially accesses 1000 single-wire system acquisition channels and 60 four-wire system acquisition channel signals into the high-precision data acquisition module.
Wherein, high accuracy data acquisition module and multichannel conversion switch install on 3 instruments, and three instrument configurations are the same, can be each other for backup, and simultaneously, every instrument all is furnished with two high accuracy data acquisition modules, and acquisition modules are each other for backup.
The storage space of the data storage unit can store continuous data for not less than 2 years.
When the transmission signal is a digital signal, the main Ethernet transmission unit (341) transmits, and is also provided with a 485 transmission unit (342) as a backup transmission channel, the Ethernet transmission unit (341) serves as a main channel for data transmission, a TCP/IP protocol is adopted, the protocol has a packet loss continuous transmission function, and meanwhile, in a data transmission program, the Ethernet transmission unit has a breakpoint continuous transmission function, and the program can detect whether data is transmitted successfully in each period and support interruption continuous transmission for 30min at most; the 485 transmission unit (342) is mainly used as a backup means for data transmission, and when the Ethernet transmission unit (341) fails, the 485 transmission unit can access the FTP server of the temperature measurement unit (300) to read the temperature data information stored in the data storage unit (350).
The signal transmission unit (400) is mainly used for rotary transmission of signals, power supply and the like, converts fixed signals and power supply at a fixed end into continuously rotating signals, and accesses a transmission interface unit (340) in the temperature measurement unit (300), and mainly comprises an Ethernet slip ring unit (410), a low-frequency signal slip ring unit (420) and a power supply slip ring unit (430), wherein the Ethernet slip ring unit (410) adopts an Ethernet slip ring connection structure and is used for converting TCP/IP communication rotary signals of the Ethernet transmission unit (341) into fixed signals, the fixed signals are led out of a space environment simulator through the fixed unit (220), 3 paths of Ethernet slip ring channels are provided, independent communication channels are provided for three instruments in the temperature measurement unit (300), and the control and data monitoring requirements of a vacuum thermal test on the temperature measurement unit are met; the low-frequency signal slip ring unit (420) adopts a slip ring structure and is used for converting a low-frequency communication rotating signal of the 485 transmission unit (342) into a fixed signal, the fixed signal is led out of the space environment simulator through the fixed unit (220), 4 paths of low-frequency signal channels are provided, 2 paths of 485 transmission channels (one for one and one for standby) are provided for the temperature measurement unit (300), and the standby data reading requirement of a vacuum thermal test on the temperature measurement unit is met; the power supply slip ring unit (430) adopts a slip ring structure and is used for converting a fixed power supply channel outside the space environment simulator into a rotary power supply channel and providing 24 channels, the maximum voltage of 120Vac/dc and the maximum current of 5A for power supply, three data acquisition instruments in the temperature measurement unit (300) are supplied with power, the power supply voltage of the data acquisition instruments is 110Vac, the current of the data acquisition instruments is 10A, the conditions that current of each loop in actual engineering is possibly distributed unevenly and the power supply loop is accidentally failed are considered, each instrument supplies power through 4 power supply loops (8 channels), the loops are connected in parallel, the theoretical peak current is 2.5A, the actual power supply condition is considered, and when two power supply channels of the whole system can not work at all, the system can still supply power stably.
The atmospheric environment includes a pressure environment, a temperature environment and a humidity environment.
The invention relates to a method for measuring the temperature of spacecraft in multi-channel continuous rotary motion in a vacuum cold black environment, which comprises the following steps:
firstly, connecting a temperature measuring sensor signal to a system, connecting sensor signals including a thermocouple, a thermistor, a platinum resistor and the like in a tested part into the temperature measuring system, then carrying out channel configuration and sensor parameter configuration, and after establishing a vacuum black environment, electrifying the system, starting the system, and automatically completing functions of temperature data acquisition, recording, displaying and the like.
Compared with the prior art, the invention has the following advantages: at present, the continuous rotating object is measured domestically, measuring channels are few, a single channel or a few channels are mainly used, measured signals are transmitted through a slip ring easily, and meanwhile, the system runs more and is in an atmospheric environment, and the continuous rotating temperature of the multi-channel small signals, which runs under a special vacuum black environment, is measured. The method can measure 1000-channel high-precision temperature, a temperature data acquisition instrument is arranged below a tested piece and rotates synchronously with a tested spacecraft, analog quantity signals of a sensor are subjected to local A/D conversion into digital signals, data processing is carried out simultaneously, processed data information is transmitted to a ground monitoring end through a communication slip ring, and real-time data monitoring is carried out. The temperature data acquisition instrument is provided with the local controller, so that under the condition of communication interruption of the slip ring, local automatic measurement and data storage can be performed, continuous and reliable operation of temperature measurement is ensured, and smooth and reliable operation of a vacuum thermal test performed by a spacecraft by adopting a solar simulator to simulate external heat flow is ensured.
Description of the drawings:
in order to more clearly describe the embodiments of the present invention, the drawings used in the embodiments will be briefly described below.
FIG. 1 is a block diagram of the structure of a multi-channel continuous rotary motion temperature measurement system of a spacecraft in a vacuum cold black environment;
FIG. 2 is a structural block diagram of a vacuum sealing unit in the multi-channel continuous rotary motion temperature measurement system of the spacecraft in the vacuum cold black environment;
FIG. 3 is a structural diagram of an environment simulation unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in a vacuum black environment according to the present invention;
FIG. 4 is a block diagram of a transmission interface unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in the vacuum cold black environment according to the present invention;
FIG. 5 is a block diagram of a signal transmission unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in a vacuum cold and black environment according to the present invention;
Detailed Description
The following is a description of the present invention, which is further illustrated by the following embodiments. The following detailed description, of course, is merely illustrative of various aspects of the invention and is not to be construed as limiting the scope of the invention.
As shown in fig. 1, it shows a multi-channel continuous rotary motion temperature measurement system of a spacecraft in a vacuum cold black environment. The measuring system comprises a sensor interface unit (100), a normal temperature and normal pressure environment unit (200), a temperature measuring unit (300) and a signal transmission unit (400). Sensor interface unit (100): the temperature sensor signal transmission device is mainly used for transmitting a temperature sensor signal on a spacecraft to a rotating unit (210), introducing the signal into a normal-temperature normal-pressure environment unit (200) through a signal through-wall sealing unit (223), and finally accessing a temperature measuring unit (300) for temperature data acquisition. Ambient unit at ambient temperature and pressure (200): the temperature measuring device is mainly used for providing simulated atmospheric environment (including pressure, temperature and humidity environment) for the temperature measuring unit (300) and the signal transmission unit (400) and ensuring the normal work of the temperature measuring unit and the signal transmission unit. The measuring unit (300) is integrally arranged in the rotating unit (210) and is used for carrying out data acquisition, signal data processing, data storage and transmission on signals of the temperature sensor on the spacecraft in the vacuum thermal test and the like. The signal transmission unit (400) is mainly used for rotary transmission of signals, power supply and the like, and converts the fixed signals and the power supply at the fixed end into signals capable of continuously rotating and then is connected into the temperature measuring unit (300). The normal temperature and normal pressure environment unit mainly comprises: the device comprises a rotating unit (210), a fixing unit (220), a vacuum sealing unit (230) and an environment simulation unit (240). The rotating unit (210) and the fixing unit (220) are unit structure main parts in normal temperature and normal pressure environments, are normal pressure environment simulation cabins and are connected with the satellite test piece and the vacuum environment simulator. The rotating unit (210) rotates synchronously with the spacecraft to be tested, provides a signal transmission interface for the sensor, and provides an installation interface for the temperature measuring unit (300). The fixed unit (220) is installed in the space environment simulator, keeps still and does not move, and provides an installation interface for the rotating unit. And providing an interface for signal transmission to the outside of the vacuum environment simulator.
Referring to fig. 2, fig. 2 is a block diagram of a vacuum sealing unit in a spacecraft multichannel continuous rotational motion temperature measurement system in a vacuum cold black environment, wherein the vacuum sealing unit (230) is mainly used for providing vacuum sealing for a simulation cabin in a normal temperature and normal pressure environment, and mainly comprises a dynamic sealing unit (231), a fixed sealing unit (232) and a signal through-wall sealing unit (233). And the dynamic sealing unit (231) provides vacuum sealing for a dynamic sealing surface connected with the rotating unit (210) and the fixed unit (220), and three sealing mechanisms are adopted to ensure that the joint between the rotating unit and the fixed unit can meet the use requirement on the leakage rate when the rotating unit moves and is static. And the fixed sealing unit (232) is used for vacuum sealing of the rotating unit (210) and the self cabin of the fixed unit (220). The sealing flange ensures the composite use requirement of the leakage rate of the normal-temperature and normal-pressure cabin, and the cabin body can achieve the sealed independent environment of normal temperature, normal pressure and constant humidity in a vacuum cold black environment. And the signal through-wall sealing unit (233) is arranged on the rotating unit and is used for introducing the sensor signal transmitted by the sensor interface unit (100) into the room temperature and normal pressure environment chamber, providing 1000 single-wire system measuring channels and 60 four-wire system measuring channels, and meeting the requirement of a large data sensor temperature measuring channel of a large-scale spacecraft vacuum thermal test.
Referring to fig. 3, fig. 3 shows a structural block diagram of an environment simulation unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in the vacuum cold black environment; the environment simulation unit (240) is used for maintaining and controlling the normal-temperature and normal-pressure environment of the simulation cabin, and comprises a pressure simulation unit (241), a temperature simulation unit (242) and a humidity simulation unit (243). The pressure simulation unit (241) obtains pressure information in a normal-temperature normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the air pressure of an air supply pipeline between the normal-pressure cabin and a space environment simulator, and finally realizes the function of simulating atmospheric pressure in the normal-pressure cabin; the temperature simulation unit (242) obtains temperature information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the gas temperature of an air supply pipe opening between the normal-pressure cabin and the space environment simulator, and finally achieves the function of simulating the atmospheric temperature in the normal-pressure cabin; the humidity simulation unit (243) obtains humidity information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the air humidity of an air supply pipe opening between the normal-pressure cabin and the space environment simulator, and finally realizes the function of simulating atmospheric humidity in the normal-pressure cabin. The measuring unit (300) comprises a small signal interface unit (310), a data acquisition unit (320), a data processing unit (330), a transmission interface unit (340) and a data storage unit (350), wherein the small signal interface unit (310) is used for receiving signals of the temperature sensor on the spacecraft to be measured, transmitted by the sensor interface unit (100), and accessing the signals to the data acquisition unit (320) for data acquisition. The interface unit transmits small voltage below 100mV, does not interfere with signals of the interface unit, can provide 1000 single-wire system measurement channels and 60 four-wire system measurement channels, and meets the requirement of a large data sensor temperature measurement channel for a large spacecraft vacuum thermal test. The data acquisition unit (320) is used for carrying out data acquisition on the signals accessed to the sensor and carrying out digital-to-analog conversion on the small signal analog quantity generated by the sensor to convert the small signal analog quantity into digital quantity signals. The core of the unit is a high-precision data acquisition module, the acquisition precision is not lower than 6 bits and a half (or the AD conversion precision is not lower than 24 bits), and the acquisition module can finish data acquisition of a thermocouple, a thermistor and a platinum resistor. Meanwhile, a multi-way change-over switch is adopted to switch the acquisition channels, and 1000 single-line acquisition channels and 60 four-line acquisition channel signals are sequentially accessed into the high-precision data acquisition module. In order to improve the reliability of the data acquisition unit, the high-precision data acquisition module and the multi-path conversion switch are installed on 3 instruments, and the three instruments are the same in configuration and can be used for mutual backup. Meanwhile, each instrument is provided with two high-precision data acquisition modules, and the acquisition modules are backup with each other. The influence of single module or single instrument fault on the whole temperature measurement system is reduced, and the reliability of the system is improved. The data processing unit (330) is used for processing and fitting the sensor signals acquired by the data acquisition unit (320) to obtain real-time temperature data information of each part of the spacecraft.
Referring to fig. 4, fig. 4 is a block diagram of a transmission interface unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in the vacuum cold black environment of the present invention; the transmission interface unit (340) is used for transmitting the sensor signal acquired by the data acquisition unit (320) and the temperature signal processed by the data processing unit (330), and the rotation motion signal is converted into a fixed signal through the signal transmission unit (400). The transmission signal is a digital signal, the main Ethernet transmission unit (341) transmits, and the 485 transmission unit (342) is also used as a backup transmission channel. The Ethernet transmission unit (341) is used as a main channel of data transmission, a TCP/IP protocol is adopted, the protocol has a packet loss continuous transmission function, the reliability of data transmission is improved, meanwhile, a breakpoint continuous transmission function is designed in a data transmission program, the program can detect whether data transmission of each period is successful, if a communication line or a transmission interface unit in the signal transmission unit (400) and the fixed unit (220) has a fault, transmission failure is caused, the program can record, when communication is recovered to be normal, all data packets during communication interruption can be automatically continuously transmitted, interruption continuous transmission of 30min at most can be supported, the reliability and integrity of data packet transmission are ensured, and the vacuum thermal test is ensured to be smoothly carried out. The 485 transmission unit (342) is mainly used as a backup means for data transmission, and when the Ethernet transmission unit (341) fails, the 485 transmission unit can access the FTP server of the temperature measurement unit (300) to read the temperature data information stored in the data storage unit (350). The data storage unit (350) is mainly used for locally storing data generated by the data acquisition unit (320) and the data processing unit (330) in real time, the storage space can store continuous data for not less than 2 years, and even if communication fails for a long time, the data cannot be lost, so that analysis and evaluation of vacuum thermal test data are influenced. Because the whole temperature measuring unit needs to synchronously rotate and swing with the spacecraft to be measured, vibration and shaking cannot be avoided, the storage medium adopts a solid hard disk, the problem that the magnetic disk cannot be normally read or damaged under the vibration or shaking condition of the traditional mechanical hard disk is avoided, and the reliability of local storage is improved.
Referring to fig. 5, fig. 5 is a block diagram of a signal transmission unit in the multi-channel continuous rotational motion temperature measurement system of the spacecraft in the vacuum cold black environment of the present invention; the signal transmission unit (400) is mainly used for rotary transmission of signals, power supply and the like, and converts fixed signals and power supply at the fixed end into continuously rotatable signals to be connected into the temperature measuring unit (300). The system mainly comprises an Ethernet slip ring unit (410), a low-frequency signal slip ring unit (420) and a power supply slip ring unit (430). The Ethernet slip ring unit (410) adopts an Ethernet slip ring connection structure and is used for converting a TCP/IP communication rotating signal of the Ethernet transmission unit (341) into a fixed signal, the fixed signal is led out of the space environment simulator through the fixed unit (220), 3 paths of Ethernet slip ring channels can be provided, independent communication channels are respectively provided for three instruments in the temperature measurement unit (300), and the requirements of a vacuum thermal test on the control and data monitoring of the temperature measurement unit are met. Low frequency signal sliding ring unit (420), adopt the sliding ring structure for change the low frequency communication rotating signal of 485 transmission unit (342) into fixed signal, draw forth outside spatial environment simulator via fixed unit (220), can provide 4 way low frequency signal channels, for temperature measuring unit (300) provide 2 way 485 transmission channels (one is used one and is equipped with), satisfy the stand-by data reading requirement of vacuum heat test to temperature measuring unit. The power supply slip ring unit (430) adopts a slip ring structure, is used for converting a fixed power supply channel outside the space environment simulator into a rotary power supply channel, can provide 24 channels, supplies power with the maximum voltage of 120Vac/dc and the maximum current of 5A, supplies power for three data acquisition instruments in the temperature measurement unit (300), supplies power with the data acquisition instruments with the voltage of 110Vac and the current of 10A, considers the conditions that current of each loop in actual engineering is possibly distributed unevenly and the power supply loop is accidentally invalid, each instrument supplies power through 4 power supply loops (8 channels), each loop is connected in parallel, theoretical peak current is 2.5A, considers the actual power supply condition, and the whole system can still stably supply power when two power supply channels can not work at all.
In a specific implementation mode, the temperature measuring instrument needs to be installed at the end of the motion simulator, keeps consistent with the motion of the spacecraft, changes a rotating temperature measuring signal into a relatively static measuring signal, is provided with a corresponding controller at the same time, and can be controlled locally, so that the temperature measuring instrument does not need a remote control computer to automatically complete temperature signal data measurement and acquisition and data processing. The temperature data measured by the temperature measuring instrument and the controller is transmitted to the master control system through the signal transmission Ethernet slip ring to complete the measurement, display and storage of the outer surface temperature of the spacecraft.
Because the whole temperature measuring unit needs to synchronously rotate and swing with the spacecraft to be measured, vibration and shaking cannot be avoided, the storage medium adopts a solid hard disk, the problem that the magnetic disk cannot be normally read or damaged under the vibration or shaking condition of the traditional mechanical hard disk is avoided, and the reliability of local storage is improved.
The invention relates to a spacecraft multichannel continuous rotary motion temperature measurement method for a vacuum cold black environment, which redesigns a whole set of temperature measurement system, can carry out high-precision measurement on a spacecraft temperature sensor in continuous rotary motion in the vacuum cold black environment, and provides complete and reliable spacecraft temperature data for a vacuum thermal test for carrying out space external heat flow simulation by adopting a solar simulation method of a solar simulator. The innovative data acquisition equipment is placed in a vacuum black environment and synchronously rotates with the spacecraft to be measured, and the technical problems that thousands of channel slip ring projects are difficult to realize and small signals cannot be accurately measured through slip rings are solved. On the basis, the data acquisition equipment is placed in a vacuum black and cold environment, communication and the like need to pass through a series of motion equipment such as a slip ring, local data processing and storage are carried out on the data processing and data storage functions, data loss caused by instability of a communication loop is prevented, the reliability of the system is further improved, and a foundation is laid for finally completing a vacuum thermal test by adopting a solar simulator.
Although particular embodiments of the invention have been described and illustrated in detail, it should be understood that various equivalent changes and modifications can be made to the above-described embodiments according to the inventive concept, and that it is intended to cover such modifications as would come within the spirit of the appended claims and their equivalents.
Claims (16)
1. The multi-channel continuous rotary motion temperature measurement system of the spacecraft in the vacuum black and cold environment mainly comprises a sensor interface unit, an atmospheric environment unit, a temperature measurement unit and a signal transmission unit, wherein the sensor interface unit is mainly used for connecting a sensor on the spacecraft to be measured or a test piece in a vacuum thermal test with the atmospheric environment unit, and connecting a sensor signal to the temperature measurement unit for temperature measurement through the atmospheric environment unit; the normal-pressure environment unit is mainly used for providing a simulated atmospheric environment for the temperature measuring unit and the signal transmission unit and ensuring the normal work of the temperature measuring unit and the signal transmission unit; the temperature measuring unit is mainly used for collecting, converting, storing and transmitting sensor signals on a tested spacecraft and a test piece in a vacuum thermal test, and is arranged on a rotating unit in a normal-pressure environment unit, wherein the normal-temperature normal-pressure environment unit is provided with a rotating unit and a fixing unit, the fixing unit is connected with an environment simulator, the rotating unit realizes rotating motion and is rigidly connected with the tested spacecraft and keeps synchronous motion with the tested spacecraft, and the accuracy and precision of temperature measurement are ensured; and the device is in a normal pressure environment, so that the normal work of the instrument and equipment is ensured; the signal transmission unit is mainly used for transmitting the temperature signals converted and processed by the temperature measurement unit to the fixed signal channel from the rotating state so as to be transmitted to the outside of the space environment simulator for real-time monitoring and providing power supply necessary for operation for the temperature measurement unit.
2. The spacecraft multichannel continuous rotary motion temperature measurement system of claim 1, wherein said sensor interface unit: the temperature sensor signal transmission device is mainly used for transmitting a temperature sensor signal on a spacecraft to a rotating unit, introducing the signal into a normal-temperature normal-pressure environment unit through a signal through-wall sealing unit, and finally accessing a temperature measuring unit for temperature data acquisition.
3. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 1, wherein said ambient environment unit: the main use provides simulation atmospheric environment for temperature measurement unit and signal transmission unit, guarantees its operational environment, mainly includes: the device comprises a rotating unit, a fixing unit, a vacuum sealing unit and an environment simulation unit, wherein the rotating unit and the fixing unit are main parts of a normal-temperature normal-pressure environment unit structure and are used as normal-pressure environment simulation cabins to connect a satellite test piece and a vacuum environment simulator; the vacuum sealing unit is mainly used for providing vacuum sealing for the simulation cabin in the normal-temperature and normal-pressure environment, and the environment simulation unit is used for maintaining and controlling the normal-temperature and normal-pressure environment of the simulation cabin.
4. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 1, wherein said rotary unit rotates synchronously with the spacecraft to be measured, providing a signal transmission interface for the sensor, and a mounting interface for the temperature measurement unit.
5. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 1, wherein said stationary unit is mounted in a space environment simulator, remains stationary, provides a mounting interface for the rotary unit, and provides an interface for signal transmission out of the vacuum environment simulator.
6. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 3, wherein the vacuum seal unit mainly comprises a dynamic seal unit, a fixed seal unit and a signal through-wall seal unit, the dynamic seal unit provides vacuum seal for a dynamic seal surface connected with the rotary unit and the fixed unit, and three seal mechanisms are adopted to ensure that the leak rate can meet the seal requirement at the connection part between the rotary unit and the fixed unit when the rotary unit moves and is static; the fixed sealing unit is used for vacuum sealing of the rotating unit and the cabin body of the fixed unit; the signal through-wall sealing unit is arranged on the rotating unit and used for introducing a sensor signal transmitted by the sensor interface unit into the room temperature and normal pressure environment cabin, providing 1000 single-wire system measuring channels and 60 four-wire system measuring channels, and meeting the requirement of a large-data-volume sensor temperature measuring channel for a large-scale spacecraft vacuum thermal test.
7. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 3, wherein the environment simulation unit comprises a pressure simulation unit, a temperature simulation unit and a humidity simulation unit, the pressure simulation unit obtains pressure information in a normal-temperature normal-pressure cabin through an environment detection sensor installed in the normal-pressure cabin, and adjusts the air pressure of an air supply pipeline between the normal-pressure cabin and the space environment simulator, so as to realize the function of simulating atmospheric pressure in the normal-pressure cabin; the temperature simulation unit obtains temperature information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, and adjusts the gas temperature of an air supply pipe opening between the normal-pressure cabin and the space environment simulator to realize the function of simulating the atmospheric temperature in the normal-pressure cabin; the humidity simulation unit obtains humidity information in the normal-temperature and normal-pressure cabin through an environment detection sensor arranged in the normal-pressure cabin, adjusts the gas humidity of an air supply pipe opening between the normal-pressure cabin and the space environment simulator, and finally realizes the function of simulating atmospheric humidity in the normal-pressure cabin.
8. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in any of claims 1 to 7, wherein the temperature measurement unit is integrally installed in the rotary unit and is used for carrying out data acquisition, signal data processing and data storage and transmission on the temperature sensor signals on the spacecraft in the vacuum thermal test.
9. The spacecraft multichannel continuous rotary motion temperature measurement system according to any one of claims 1 to 7, wherein the temperature measurement unit comprises a small signal interface unit, a data acquisition unit, a data processing unit, a transmission interface unit and a data storage unit, wherein the small signal interface unit is used for receiving signals of a temperature sensor on the spacecraft to be measured, which are transmitted by the sensor interface unit, and connecting the signals to the data acquisition unit for data acquisition, and the data acquisition unit is used for carrying out data acquisition on the signals connected to the sensor, carrying out digital-to-analog conversion on analog quantities of the small signals generated by the sensor and converting the analog quantities into digital quantity signals; the data processing unit is used for processing and fitting the sensor signals acquired by the data acquisition unit to obtain real-time temperature data information of each part of the spacecraft; the transmission interface unit is used for transmitting the sensor signal acquired by the data acquisition unit and the temperature signal processed by the data processing unit, and converting the rotary motion signal into a fixed signal through the signal transmission unit; the data storage unit is mainly used for locally storing the data generated by the data acquisition unit and the data processing unit in real time.
10. The spacecraft multichannel continuous rotational motion temperature measurement system of claim 9, wherein the transmission interface unit transmits a small voltage of less than 100mV without interfering signals thereof, and 1000 single-wire system measurement channels and 60 four-wire system measurement channels are provided to meet the requirement of a large data sensor temperature measurement channel for a large-scale spacecraft vacuum thermal test.
11. A spacecraft multichannel continuous rotary motion temperature measurement system as claimed in claim 9, wherein the data acquisition accuracy of the data acquisition unit is not less than 6 bits and a half, and the acquisition module can complete data acquisition of a thermocouple, a thermistor, and a platinum resistor.
12. The spacecraft multichannel continuous rotational motion temperature measurement system of claim 9, wherein the data acquisition unit uses a multi-way switch to switch the acquisition channels, and sequentially connects 1000 single-wire acquisition channels and 60 four-wire acquisition channels to the high-precision data acquisition module.
13. A spacecraft multichannel continuous rotary motion temperature measurement system according to claim 12, wherein the high precision data acquisition modules and the multiplexing switches are mounted on 3 instruments, the three instruments are configured identically and can be backed up each other, and each instrument is provided with two high precision data acquisition modules, which are backed up each other.
14. The spacecraft multichannel continuous rotational motion temperature measurement system according to claim 9, wherein when the transmission signal is a digital signal, the primary ethernet transmission unit transmits, and is further provided with a 485 transmission unit as a backup transmission channel, the ethernet transmission unit serves as a primary channel for data transmission, and a TCP/IP protocol is adopted, and the protocol has a packet loss continuous transmission function, and simultaneously has a breakpoint continuous transmission function in a data transmission program, and the program can detect whether data is transmitted successfully in each period, and supports interruption continuous transmission for up to 30 min; the 485 transmission unit is mainly used as a backup means of data transmission, and when the Ethernet transmission unit fails, the FTP server of the temperature measurement unit can be accessed through the 485 transmission unit to read the temperature data information stored in the data storage unit.
15. The spacecraft multichannel continuous rotary motion temperature measurement system of claim 9, wherein the signal transmission unit is mainly used for signal and power supply rotary transmission, converts fixed signals and power supply at a fixed end into continuously rotating signals, and accesses the continuously rotating signals into the transmission interface unit in the measurement unit, and mainly comprises an ethernet slip ring unit, a low-frequency signal slip ring unit and a power supply slip ring unit, wherein the ethernet slip ring unit adopts an ethernet slip ring connection structure, is used for converting TCP/IP communication rotary signals of the ethernet transmission unit into fixed signals, is led out of the space environment simulator through the fixed unit, provides 3 paths of ethernet slip ring channels, provides independent communication channels for three instruments in the temperature measurement unit respectively, and meets the control and data monitoring requirements of a vacuum thermal test on the temperature measurement unit; the low-frequency signal slip ring unit adopts a slip ring structure and is used for converting a low-frequency communication rotating signal of the 485 transmission unit into a fixed signal, the fixed signal is led out of the space environment simulator through the fixed unit, 4 low-frequency signal channels are provided, 2 485 transmission channels are provided for the temperature measurement unit, and the requirement of a vacuum thermal test on reading standby data of the temperature measurement unit is met; the power supply slip ring unit adopts a slip ring structure and is used for converting a fixed power supply channel outside the space environment simulator into a rotary power supply channel and can provide 24 channels, the maximum voltage is 120Vac/dc, the maximum current is 5A, power is supplied to three data acquisition instruments in the temperature measurement unit, the power supply voltage of the data acquisition instruments is 110Vac, the current is 10A, the conditions that current of each loop is possibly distributed unevenly and the power supply loop is accidentally failed in actual engineering are considered, each instrument supplies power through 4 power supply loops, the loops are connected in parallel, the theoretical peak current is 2.5A, the actual power supply condition is considered, and when two power supply channels of the whole system can not work at all, the system can still supply power stably.
16. A spacecraft multichannel continuous rotary motion temperature measurement system according to any of claims 1 to 7, wherein the atmospheric environment comprises a pressure, temperature and humidity environment.
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104085542A (en) * | 2013-11-27 | 2014-10-08 | 北京卫星环境工程研究所 | Transmission system for spacecraft capable of moving rotationally |
| CN106467174A (en) * | 2015-08-10 | 2017-03-01 | 北京卫星环境工程研究所 | Large Spacecraft mechanism normal pressure thermograde pilot system |
| CN109018456A (en) * | 2018-06-28 | 2018-12-18 | 上海卫星工程研究所 | A kind of thermal vacuum test Orbital heat flux analogy method of high rail remote sensing satellite platform |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR100513207B1 (en) * | 2002-07-24 | 2005-09-08 | 한국전기연구원 | Superconducting Rotor With Conduction Cooling System |
-
2020
- 2020-03-11 CN CN202010164914.7A patent/CN111307336B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104085542A (en) * | 2013-11-27 | 2014-10-08 | 北京卫星环境工程研究所 | Transmission system for spacecraft capable of moving rotationally |
| CN106467174A (en) * | 2015-08-10 | 2017-03-01 | 北京卫星环境工程研究所 | Large Spacecraft mechanism normal pressure thermograde pilot system |
| CN109018456A (en) * | 2018-06-28 | 2018-12-18 | 上海卫星工程研究所 | A kind of thermal vacuum test Orbital heat flux analogy method of high rail remote sensing satellite platform |
Non-Patent Citations (2)
| Title |
|---|
| 基于PXIe高速同步温度数据采集系统在真空热试验中的运用;邵静怡 等;《真空与低温》;20181231;第24卷(第6期);419-422 * |
| 旋转受照状态下天线温度场仿真分析;邵钰杰 等;《航天器环境工程》;20151231;第32卷(第6期);583-588 * |
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