A space environment comprehensive monitoring system based on the last stage in orbit

Abstract

The present invention discloses a space environment comprehensive monitoring system based on a last stage of orbit retention. Aiming at the problems of the scarcity of opportunities and high costs for carrying space environment monitoring payloads on traditional space environment research satellites and general business satellites, the system uses the last stage of orbit retention of a preset orbit retention control platform and the space environment monitoring payload loaded on the last stage of orbit retention. Based on the large elliptical orbit formed after the last stage of orbit retention is passivated and discharged, high-energy proton measurement, single particle effect monitoring, radiation dose measurement and earth neutral atmosphere density monitoring at different orbital altitudes and perigee altitude of the large elliptical orbit within the range of the adjacent space orbit are performed. The acquired monitoring data is used for the fault analysis of the spacecraft and the precise orbit prediction of the spacecraft. A new type of space environment comprehensive monitoring system with high cost-effectiveness, short mission cycle, wide orbit coverage and distributed load is provided, which can obtain large-area, real-time, synchronous, multi-scale and multi-factor space environment information.

Description

Technical Field

[0001] The invention belongs to the technical field of space environment detection, and in particular relates to a space environment comprehensive monitoring system based on a last stage remaining in orbit. Background Art

[0002] With the vigorous development of the aerospace industry, the space for human activities continues to extend into outer space. The space environment is increasingly affecting various types of human production and life, becoming the fourth environment for human survival in addition to land, ocean and atmosphere. The space environment described in this patent refers to all environmental factors in the solar-terrestrial space that affect human activities, including high-energy protons, single-particle radiation, neutral atmospheric pressure, temperature, density, etc. At present, with the continuous deepening of human space activities, people have gradually realized that the space environment is one of the main factors affecting the high reliability and long life of spacecraft. Therefore, it is necessary to carry out comprehensive space environment monitoring and research, understand and grasp the general conditions and changes of the space environment, and ensure the safety of economic and scientific and technological activities based on spacecraft.

[0003] At present, the monitoring of space environment is mainly achieved by carrying space environment monitoring equipment on special space environment research satellites and general business satellites (such as the "Fengyun" series, "Beidou" series, and "Ziyuan" series satellites). Many space environment research satellites have formed certain business capabilities while achieving scientific goals, and play an important role in conventional space environment monitoring business. However, space environment monitoring for the purpose of guaranteeing services is different from space physical detection for the purpose of scientific research. The main difference is that space environment monitoring requires continuity and long-term stability. In addition, the space environment has a wide coverage, more monitoring elements, higher spatiotemporal resolution, and a longer time period. All of these determine that the space environment monitoring payload requires low-cost, multi-orbit, and long-term continuous flight tests. Due to the high cost and high risk of existing space environment research satellites and general business satellites, the opportunities for carrying traditional space environment monitoring payloads are still scarce and the cost is still high, which seriously restricts the development of my country's space environment monitoring system. Summary of the invention

[0004] The purpose of the present invention is to provide a space environment comprehensive monitoring system based on the last stage on orbit, which can solve the problems of scarce opportunities and high costs for carrying space environment monitoring payloads on traditional space environment research satellites and general business satellites. Moreover, based on the last stage on orbit, large-area, real-time, synchronous, multi-scale and multi-factor space environment information can be obtained, and it has the characteristics of high cost-effectiveness, short mission cycle, wide orbit coverage and distributed carrying.

[0005] To solve the above problems, the technical solution of the present invention is:

[0006] A space environment comprehensive monitoring system based on a terminal stage on orbit includes: a terminal stage on orbit of a preset control platform on orbit and a space environment monitoring payload. The terminal stage on orbit of the preset control platform on orbit is equipped with energy, attitude control and communication equipment to enable it to have the conditions and capabilities to carry out space environment monitoring, becoming an effective means and method for comprehensive space environment monitoring.

[0007] The space environment monitoring payload is loaded on the last stage of the orbit retention stage, and based on the large elliptical orbit formed after the passivation and discharge of the last stage of the orbit retention stage, high-energy proton measurement, single particle effect monitoring, radiation dose measurement and neutral atmosphere density monitoring of the earth at different orbital altitudes and the perigee altitude of the large elliptical orbit within the range of the near-space orbit are performed. The acquired monitoring data is used for spacecraft fault analysis and spacecraft precise orbit prediction.

[0008] According to one embodiment of the present invention, the orbit-retaining final stage draws on the attitude control method of early spin-stabilized satellites, maximizes the effectiveness of the remaining propellant by utilizing the remaining capacity of the carrier rocket's attitude control system after the launch mission, implements long-term on-orbit spin-stabilization control around the roll axis, and conducts multi-angle and multi-directional space environment monitoring.

[0009] According to an embodiment of the present invention, the orbit retention control platform includes an orbit retention control device, a solar array, a measurement and control antenna, and a cable network, which meet the launch vehicle mission requirements in terms of total weight, envelope, installation interface, test process, and environmental condition verification;

[0010] The orbit retention control device is installed on the orbit retention final stage support cabin and adopts shock absorption measures. The solar cell array is installed on the outer surface of the orbit retention final stage support cabin and adopts equipotential and heat insulation measures. The measurement and control antenna is installed on the surface of the transition section of the orbit retention final stage support cabin and meets the mechanical and thermal requirements of the active section of the launch vehicle and is used for uplink and downlink communication of data information of the orbit retention control device. The orbit retention control device is connected to the solar cell arrays installed at different positions on the support cabin through a cable network and is used for energy supply for long-term monitoring of the orbit retention platform and space environmental loads during the orbit retention process.

[0011] According to an embodiment of the present invention, the orbit retention control device includes an energy subsystem, a communication subsystem, an attitude control subsystem, a digital control subsystem and a thermal control subsystem, wherein the communication subsystem maintains a radio silent state, and after the orbit retention final stage is separated from the main satellite, it is turned on according to a predetermined program control program to carry out the orbit retention payload function application;

[0012] The energy subsystem uses a 28V DC bus to connect the solar array and the battery pack to supply power to the on-orbit control platform and the space environment monitoring payload;

[0013] The communication subsystem adopts measurement and control bidirectional omnidirectional link communication. The telemetry antenna and remote control antenna are wide-beam low-gain antennas. The telemetry antenna and remote control antenna in the same frequency band use different sources and different frequencies to avoid signal anomalies in the antenna interference area.

[0014] The attitude control subsystem adopts attitude control components including flywheels and magnetic rods to perform high-efficiency and low-cost three-axis stable control, so as to realize long-term on-orbit attitude stability control and attitude pointing control of the last stage of the large inertia cabin.

[0015] The digital control subsystem is based on the CPU, performs flight control and information detection, and realizes various digital signal collection, control and payload communication functions;

[0016] The thermal control subsystem adopts a combination of active and passive thermal control methods, with passive thermal control as the main method and active thermal control of electric heaters as the auxiliary method, to provide a suitable thermal environment for the on-orbit control platform and the space environment monitoring payload.

[0017] According to one embodiment of the present invention, the space environment monitoring payload mainly includes an environmental monitoring device, which realizes LET spectrum monitoring and radiation dose monitoring at different orbital altitudes, and utilizes the perigee height of the large elliptical orbit formed after the last stage passivation emission to be within the near-space orbit range, thereby realizing neutral atmosphere density monitoring and neutral atmosphere temperature monitoring.

[0018] According to one embodiment of the present invention, the space environment monitoring payload includes a space particle radiation effect comprehensive measuring instrument payload, which uses a silicon sensor to measure the linear deposition energy of different particles in the silicon material to obtain an LET spectrum; and uses silicon material as the target material for radiation dose rate measurement to perform radiation dose monitoring.

[0019] According to an embodiment of the present invention, the space particle radiation effect comprehensive measurement instrument payload includes a plurality of sensor devices and a data processing device;

[0020] The sensor device is used to collect LET spectrum data of particles in the last stage orbit, total radiation dose in the support cabin and single particle upset data, and conduct joint observation;

[0021] The data processing device is used to process the data acquired by the sensor device and transmit the data to the on-orbit control platform.

[0022] According to an embodiment of the present invention, the space environment monitoring payload further includes a neutral atmosphere measuring instrument payload, and the neutral atmosphere measuring instrument payload is used to monitor 1×10 -14 ~5×10 -12 kg / m 3 Atmospheric density and temperature within a range.

[0023] According to an embodiment of the present invention, the neutral atmosphere measuring instrument load includes a multifunctional sensor, a power supply, a microcurrent detection unit and a communication unit;

[0024] The multifunctional sensor is used to collect the atmospheric temperature at the sensor inlet and the pressure and temperature inside the sensor;

[0025] The power supply is used to provide the high voltage electricity and electrode voltage required for the operation of the multifunctional sensor;

[0026] The micro-current detection unit is used to detect the micro-current signal generated by the multi-functional sensor;

[0027] The communication unit is used to transmit the data acquired by the multifunctional sensor and the data acquired by the microcurrent detection unit to the on-track control platform.

[0028] Due to the adoption of the above technical solution, the present invention has the following advantages and positive effects compared with the prior art:

[0029] The space environment comprehensive monitoring system based on the last stage of the orbit in one embodiment of the present invention aims at the problems of the scarcity of opportunities and high cost for carrying space environment monitoring payloads on traditional space environment research satellites and general business satellites. By utilizing the last stage of the orbit of the preset orbit control platform and the space environment monitoring payload loaded on the last stage of the orbit, based on the large elliptical orbit formed after the last stage of the orbit is passivated and discharged, high-energy proton measurement, single particle effect monitoring, radiation dose measurement and earth neutral atmosphere density monitoring at different orbital altitudes and the perigee altitude of the large elliptical orbit within the range of the near-space orbit are performed, and the acquired monitoring data are used for the fault analysis of the spacecraft and the precise orbit prediction of the spacecraft. A new type of space environment comprehensive monitoring system with high cost-effectiveness, short mission cycle, wide orbit coverage and distributed load is provided, which can obtain large-area, real-time, synchronous, multi-scale and multi-factor space environment information, becoming a beneficial supplementary way to build a space environment monitoring system, and having great engineering application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 This is a monitoring effect diagram of a space environment comprehensive monitoring system based on the last stage remaining in orbit in one embodiment of the present invention;

[0031] Figure 2 It is a block diagram of a space environment comprehensive monitoring system based on the last stage remaining in orbit in one embodiment of the present invention;

[0032] Figure 3 is a block diagram of a track-staying control device in one embodiment of the present invention;

[0033] Figure 4 is a block diagram of an environmental monitoring device in one embodiment of the present invention;

[0034] Figure 5 is a schematic diagram of the LET spectrum measurement principle in one embodiment of the present invention;

[0035] Figure 6 This is a block diagram of the load composition of a space particle radiation effect comprehensive measurement instrument in one embodiment of the present invention;

[0036] Figure 7 This is a schematic diagram of the neutral atmosphere density and temperature measurement principle in one embodiment of the present invention;

[0037] Figure 8 The figure is a block diagram of the neutral atmosphere measuring instrument load composition in one embodiment of the present invention.

[0038] Description of reference numerals:

[0039] 1: Orbit retention control device; 2: Solar cell array; 3: Measurement and control antenna; 4: Data transmission antenna; 5: Space environment monitoring payload; 6: Silicon sensor. DETAILED DESCRIPTION

[0040] The following is a further detailed description of a space environment comprehensive monitoring system based on the last stage of the orbiting sub-stage proposed by the present invention in conjunction with the accompanying drawings and specific embodiments. The advantages and features of the present invention will become more apparent from the following description and claims.

[0041] The last stage of the launch vehicle is the part that actually enters orbit. After the launch vehicle completes the space launch mission, its last stage of the launch vehicle is idle in the space orbit and does not fully exert its application efficiency. It becomes the largest volume of space debris generated by the space launch mission and occupies precious orbital resources for a long time. After the last stage of the launch vehicle is adaptively modified and equipped with energy, attitude control, communication and other equipment, it has the conditions and capabilities to carry out comprehensive monitoring of the space environment, becoming an effective means and method of space environment monitoring that "turns waste into treasure".

[0042] In the present embodiment, a space environment comprehensive monitoring system based on a staying-on-orbit final stage comprises: a staying-on-orbit final stage of a preset staying-on-orbit control platform, and a space environment monitoring payload loaded on the staying-on-orbit final stage. The space environment monitoring payload performs high-energy proton measurement, single particle effect monitoring, radiation dose measurement at different orbital altitudes, and monitoring of the earth's neutral atmosphere density at a perigee altitude of the large elliptical orbit within the range of a nearby space orbit based on a large elliptical orbit formed after the stay-on-orbit final stage is passivated and discharged. The acquired monitoring data is used for spacecraft fault analysis and spacecraft precise orbit prediction.

[0043] After the last stage of the launch vehicle is discharged, it forms a natural large elliptical orbit. Figure 1As shown in the figure, the apogee height is about 20,000 km. For details, please refer to the selection of different primary satellite mission orbits.) In different orbital areas below the altitude, high-energy proton measurement, single particle effect monitoring, and radiation dose measurement can be carried out. The monitoring data can effectively support ground-based work such as spacecraft fault analysis. At its perigee (such as Figure 1 As shown, the perigee height is about 200km. For details, please refer to the selection of different launch vehicles for passivation orbits). The neutral atmosphere density of the Earth can be monitored in this orbital space to obtain its distribution changes over time and space, thereby serving the precise orbit prediction of spacecraft.

[0044] For details, please see Figure 2 The orbit-retention platform installed on the orbit-retention terminal sub-stage support cabin in this embodiment includes an orbit-retention control device 1, a solar cell array 2, a measurement and control antenna 3 and a cable; the measurement and control antenna 3 is installed on the surface of the orbit-retention control device 1, and is used for uplink and downlink communication of data information of the orbit-retention control device 1; the solar cell array 2 is installed on the outer surface of the orbit-retention terminal sub-stage support cabin, and the solar cell array 2 is connected to the battery group in the orbit-retention control device 1 through a cable, and is used for energy supply of the orbit-retention platform during the orbit-retention process. In addition to supplying power to the orbit-retention platform, the solar cell array 2 also supplies power to the space environment monitoring payload 5.

[0045] The orbit control device in this embodiment is a comprehensive electronic device integrating energy management, measurement and control communication, digital control, attitude control and thermal control. The orbit control device is the core part of the orbit platform, and can also be equipped with a digital transmission antenna 4 to realize related data transmission.

[0046] The on-orbit control device is the core part of the on-orbit control platform. It is a high-function density integrated stand-alone machine that integrates energy management, measurement and control communications, digital control, attitude control, thermal control, etc. Figure 3 The orbit control device mainly includes a structural subsystem, an energy subsystem, a communication subsystem, an attitude control subsystem, a digital control subsystem and a thermal control subsystem.

[0047] Among them, the structural subsystem is one of the main subsystems of the space environment comprehensive monitoring system. The orbit control unit (OCU) adopts a polyhedron box panel configuration. The main structure of the OCU adopts an integrated and lightweight design of the orbit-class cabin based on 3D printing. The box panel adopts a lattice sandwich structure, the assembly components adopt a lightweight composite laminate structure, and the connecting parts adopt lightweight titanium alloy fasteners. The OCU structure studied in this patent has the characteristics of light weight and high strength.

[0048] The energy subsystem uses a 28V DC bus to connect the solar array and the battery pack to power the on-orbit control platform and the space environment monitoring payload. In actual applications, the energy subsystem adopts a solar array + battery pack combined power supply and a non-regulated 28V DC bus solution. The solar array design adopts a combination of a support cabin mounted solar array, a support cabin deployable solar array, an OCU mounted solar array, and an OCU deployable solar array. All solar cells use triple-junction gallium arsenide cloth sheets with an efficiency of more than 30%. After entering orbit, the solar array installed on the final stage support cabin and OCU is unfolded. During the illumination period, the solar array is used to generate electricity to power the platform and payload equipment and charge the battery; during the shadow period, the battery releases electricity to power the platform and payload equipment. This ensures stable and reliable power supply for the platform and payload equipment during normal operation during the operation of the space environment comprehensive monitoring system.

[0049] The communication subsystem adopts two-way omnidirectional link communication for measurement and control. The telemetry antenna and remote control antenna are wide-beam low-gain antennas. The telemetry antenna and remote control antenna in the same frequency band adopt different sources and different frequencies to avoid signal anomalies in the antenna interference area. In practical applications, the communication subsystem adopts an omnidirectional link solution to ensure that the space environment monitoring system can achieve stable and reliable ground measurement and control communication not only in a stable attitude state, but also in a spinning state when the last stage in orbit is in transit to monitor the ground station, and there is no communication blind spot when the last stage is in transit to monitor the ground station. That is, both the two-way measurement and control are designed for omnidirectional communication. Both the telemetry and remote control antennas are wide-beam low-gain antennas, and a single antenna can cover half the space. Telemetry and remote control in the same frequency band use two different frequencies, that is, different sources and different frequencies, to avoid signal anomalies in the antenna interference area.

[0050] The attitude control subsystem uses attitude control components including flywheels and magnetic bars to perform three-axis stable control, realizing long-term on-orbit attitude stability control and attitude pointing control of the last stage on orbit. In practical applications, the attitude control subsystem can carry out omnidirectional space environment monitoring according to the needs of space environment monitoring by using the special on-orbit mode of the attitude spin state of the last stage on orbit after passivation. At the same time, the OCU is equipped with attitude control components such as flywheels and magnetic bars, and has the ability to use a zero-momentum three-axis stable control solution to realize the long-term on-orbit attitude stability control and attitude pointing control of the last stage space environment comprehensive monitoring system on orbit.

[0051] The digital control subsystem is centered on the CPU, performs flight control and information detection, and realizes various digital signal acquisition and payload communication. In actual applications, the digital control subsystem is centered on the CPU, performs system flight control and information monitoring, includes attitude control computer functions, and completes various digital signal acquisition, control, payload communication and other functions. The digital control computer board adopts a dual-machine cold backup working mode, including two independent CPUs as the main and standby, attitude control drive modules, external storage modules, multi-way switches for external analog acquisition, and CAN / 485 / 422 transceivers, with high reliability and high integration.

[0052] The thermal control subsystem adopts a combination of active and passive thermal control methods, with passive thermal control as the main method and active thermal control of electric heaters as the auxiliary method, to provide a suitable thermal environment for the on-orbit control platform and the space environment monitoring payload. In practical applications, the thermal control subsystem, as a service subsystem of the space environment comprehensive monitoring system, has the main task of providing a suitable thermal environment for the platform and payload equipment, ensuring that the instruments, equipment, structural components, etc. on the space environment comprehensive monitoring system operate within the specified temperature index range during on-orbit operation. The thermal control subsystem adopts a combination of active and passive thermal control methods, with passive thermal control as the main method and active thermal control of electric heaters as the auxiliary method. By selecting a suitable heat dissipation surface, the temperature level can be controlled within a suitable range.

[0053] The space environment monitoring payload in this embodiment mainly includes an environment monitoring device. Figure 4 The environmental monitoring device realizes four major functions: LET spectrum monitoring, radiation dose monitoring, neutral atmosphere density monitoring and neutral atmosphere temperature monitoring. The environmental monitoring device is the core part of the space environment monitoring payload. It makes full use of the unique advantages of the distributed load-bearing of the final stage support cabin and can install various types of space environment monitoring payloads at the same time, so as to collect multi-orbit, multi-element and multi-region continuous space environment comprehensive monitoring data. After the final stage in orbit is passivated and enters the large elliptical orbit, various types of space environment monitoring payloads are powered on and begin to collect space environment data of different orbital altitudes and different environmental elements.

[0054] The space environment monitoring payload in this embodiment includes a space particle radiation effect comprehensive measuring instrument payload, which uses a silicon sensor to measure the linear deposition energy of different particles in the silicon material to obtain an LET spectrum; and uses silicon material as the target material for radiation dose rate measurement to perform radiation dose monitoring.

[0055] In practical applications, the space particle radiation effect comprehensive measurement instrument has the ability to carry out 0.1-37 MeV / (mg / cm2) and greater than 37 MeV / (mg / cm 2 ) range and 0-10 6 Radiation dose measurement within the Rad(Si) range can also monitor the single particle upset frequency of typical devices. The monitoring principle is to use the measurement method of silicon sensors to measure the linear deposition energy of different particles in silicon materials, namely the LET spectrum; at the same time, silicon materials are used as the target material for radiation dose rate measurement.

[0056] LET spectrum monitoring uses particles incident on a silicon sensor, where they lose energy ΔE (keV). If the thickness of the silicon sensor is known to be d (um), then the LET value of the particle in the silicon material (keV / um) = ΔE / d, which is the energy loss value of the particle per unit length in the silicon material. High-energy charged particles penetrate the cabin of the environmental monitoring unit and are incident on the sensor. Some particles penetrate all sensors and deposit part of their energy. The LET value of these particles can be obtained based on the definition of LET. Some particles deposit all of their energy and are blocked from entering the first sensor. The actual path length of the particles is less than the thickness of the sensor. The LET value of these particles is obtained by inverting the ground data based on the energy deposition of the incident particles. The principle of LET spectrum measurement is as follows: Figure 5 As shown, a typical component to be tested is added between two silicon sensors 6, and the information of single particle events occurring in the component to be tested can be monitored through the circuit system. In this way, the particle incident and the single particle information of the device can be monitored simultaneously to find the relationship between the two. The radiation dose is defined as dose D = ΔE / ΔM, that is, the total energy deposition of the particle per unit mass, and the radiation dose rate H = D / t, that is, the cumulative dose per unit time. The design and measurement of the radiation dose rate probe are completely based on the definition. The particle generates energy loss in the silicon detection material, and the energy loss value ΔE of the particle is recorded electronically. The total energy loss ΣΔE per unit time is accumulated and divided by the mass M of the silicon detection material to obtain the radiation dose rate.

[0057] Please see Figure 6 From the measurement principle of the payload of the space particle radiation effect comprehensive measurement instrument, it can be seen that the payload is mainly composed of a sensor system (or sensor device), an electronic circuit (or a data processing device) and a mechanical structure. The sensor system includes a semiconductor measurement instrument system and a typical chip for monitoring single-particle upsets. The semiconductor measurement instrument system is mainly used to complete the particle LET spectrum measurement of the last stage orbit LET value greater than 0.1MeV / (mg / cm2) and the total radiation dose measurement in the cabin, and at the same time conduct joint observations with the single-particle upset data obtained by the chip under test; the electronics part mainly pre-processes the electrical signals output by the above-mentioned measurement instrument system and the chip under test, collects and processes data, and communicates with the last stage platform bus to realize data transmission. The sensor part of the payload of the space particle radiation effect comprehensive measurement instrument is required to be installed facing the sky, and the sensor normal is perpendicular to the flight direction and facing the sky.

[0058] The space environment monitoring payload in this embodiment also includes a neutral atmosphere measuring instrument payload, which is used to monitor 1×10 -14 ~5×10 -12 kg / m 3 The atmospheric density and temperature within the range are measured with an accuracy better than 10%. Figure 7The measurement principle is to install thermometers at the sensor entrance and inside the sensor to measure the atmospheric temperature and the gas temperature data in the sensor, and then directly measure the sampling chamber pressure (sensing air pressure in the sensor). According to the gas molecular dynamics theory, the thermosphere atmospheric density can be obtained by pressure (sensing air pressure in the sensor) and temperature (gas temperature in the sensor). Therefore, when the gas molecules (atoms) enter the sampling chamber with the adjustment plate through the standard measuring hole, the incoming gas molecules (atoms) experience collisions with the sampling chamber wall and the adjustment plate, and its temperature is adjusted to approximate the wall temperature. The sampling chamber wall temperature (Tg) is directly measured by setting a thermometer, and the BA gauge ionization meter is used to measure the air pressure in the sensor. Since the charged ions or electrons entering with the neutral gas molecules are captured, the adjusted neutral molecules reach the sensing area of ​​the sensor and are ionized by the electrons emitted by the hot cathode and accelerated by the gate to form positive ions. The size of the collected positive ion flow can directly reflect the pressure in the sensor.

[0059] Please see Figure 8 The neutral atmosphere measuring instrument load in this embodiment includes a multifunctional sensor, a power supply, a microcurrent detection unit and a communication unit, wherein the multifunctional sensor is used to collect the atmospheric temperature at the sensor entrance (i.e., the standard measuring hole) and the pressure and temperature inside the sensor; the power supply is used to provide the high voltage and electrode voltage required for the operation of the multifunctional sensor; the microcurrent detection unit is used to detect the microcurrent signal generated by the multifunctional sensor; and the communication unit is used to transmit the data obtained by the multifunctional sensor and the data obtained by the microcurrent detection unit to the on-orbit control platform.

[0060] In actual applications, the neutral atmosphere measuring instrument load is installed on the windward side of the support cabin, the sensor extends out of the cabin, the measuring hole is parallel to the flight direction, and the field of view angle is not less than 120 degrees.

[0061] The space environment comprehensive monitoring system based on the last stage of the orbit in this embodiment has an independent power supply and no electrical interface with the original control system on the launch vehicle. Its working state has no effect on the control system on the rocket. In the future, it can be adaptively modified according to the demand of the space environment monitoring payload and the remaining carrying capacity of the launch vehicle, and can be expanded to the last stage of my country's main launch vehicle models (CZ-4B / C, CZ-2D, CZ-6, etc.).

[0062] The embodiments of the present invention are described in detail above in conjunction with the accompanying drawings, but the present invention is not limited to the above embodiments. Even if various changes are made to the present invention, if these changes fall within the scope of the claims of the present invention and their equivalents, they still fall within the protection scope of the present invention.

Claims

1. A space environment comprehensive monitoring system based on the last stage in orbit, characterized in that: include: The preset orbit-retention terminal stage and space environment monitoring payload of the orbit-retention control platform are provided. The orbit-retention terminal stage of the preset orbit-retention control platform is equipped with energy, attitude control and communication equipment to enable it to have the conditions and capabilities to carry out space environment monitoring. The space environment monitoring payload is loaded on the last stage of the orbit, and based on the large elliptical orbit formed after the last stage of the orbit is passivated and discharged, high-energy proton measurement, single particle effect monitoring, radiation dose measurement at different orbital altitudes and earth neutral atmosphere density monitoring at the perigee altitude of the large elliptical orbit within the range of the near-space orbit are performed, and the acquired monitoring data are used for spacecraft fault analysis and spacecraft precise orbit prediction; The space environment monitoring payload includes a neutral atmosphere measuring instrument payload, which is used to monitor 1×10 -14 ~5×10 -12 kg / m 3 The neutral atmosphere measuring instrument load installs thermometers at the sensor entrance and inside the sensor to measure the atmospheric temperature and the gas temperature data in the sensor, and then directly measures the sensed air pressure in the sensor. According to the gas molecular dynamics theory, the thermosphere atmospheric density is calculated through the sensed air pressure in the sensor and the gas temperature in the sensor.

2. The space environment comprehensive monitoring system based on the last stage in orbit as claimed in claim 1 is characterized in that: The orbit-retaining final stage draws on the attitude control method of early spin-stabilized satellites. It maximizes the effectiveness of the remaining propellant by utilizing the remaining capacity of the carrier rocket's attitude control system after the launch mission, implements long-term on-orbit spin-stabilization control around the roll axis, and conducts multi-angle and multi-directional space environment monitoring.

3. The space environment comprehensive monitoring system based on the last stage in orbit as claimed in claim 1 is characterized in that: The orbit retention control platform includes an orbit retention control device, a solar array, a measurement and control antenna, and a cable network, and meets the launch vehicle mission requirements in terms of total weight, envelope, installation interface, test process, and environmental condition verification; The orbit retention control device is installed on the orbit retention final stage support cabin and adopts shock absorption measures. The solar cell array is installed on the outer surface of the orbit retention final stage support cabin and adopts equipotential and heat insulation measures. The measurement and control antenna is installed on the surface of the transition section of the orbit retention final stage support cabin and meets the mechanical and thermal requirements of the active section of the launch vehicle and is used for uplink and downlink communication of data information of the orbit retention control device. The orbit retention control device is connected to the solar cell arrays installed at different positions on the support cabin through a cable network and is used for energy supply for long-term monitoring of the orbit retention platform and space environmental loads during the orbit retention process.

4. The space environment comprehensive monitoring system based on the last stage in orbit as claimed in claim 3 is characterized in that: The orbit retention control device includes an energy subsystem, a communication subsystem, an attitude control subsystem, a digital control subsystem and a thermal control subsystem, wherein the communication subsystem maintains a radio silent state, and after the orbit retention final stage is separated from the main satellite, it is turned on according to a predetermined program control program to carry out the orbit retention payload function application; The energy subsystem uses a 28V DC bus to connect the solar array and the battery pack to supply power to the on-orbit control platform and the space environment monitoring payload; The communication subsystem adopts measurement and control bidirectional omnidirectional link communication. The telemetry antenna and remote control antenna are wide-beam low-gain antennas. The telemetry antenna and remote control antenna in the same frequency band use different sources and different frequencies to avoid signal anomalies in the antenna interference area. The attitude control subsystem adopts attitude control components including flywheels and magnetic rods to perform high-efficiency and low-cost three-axis stable control, so as to realize long-term on-orbit attitude stability control and attitude pointing control of the last stage of the large inertia cabin. The digital control subsystem is based on the CPU, performs flight control and information detection, and realizes various digital signal collection, control and payload communication functions; The thermal control subsystem adopts a combination of active and passive thermal control methods, with passive thermal control as the main method and active thermal control of electric heaters as the auxiliary method, to provide a suitable thermal environment for the on-orbit control platform and the space environment monitoring payload.

5. The space environment comprehensive monitoring system based on the last stage in orbit as claimed in claim 1 is characterized in that: The space environment monitoring payload mainly includes an environment monitoring device, which realizes LET spectrum monitoring and radiation dose monitoring at different orbital altitudes, and utilizes the perigee altitude of the large elliptical orbit formed after the last stage passivation emission to be within the near-space orbit range to realize neutral atmosphere density monitoring and neutral atmosphere temperature monitoring.

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