CN113495497A - Satellite simulation in-orbit working condition closed-loop test system - Google Patents

Satellite simulation in-orbit working condition closed-loop test system Download PDF

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CN113495497A
CN113495497A CN202110638738.0A CN202110638738A CN113495497A CN 113495497 A CN113495497 A CN 113495497A CN 202110638738 A CN202110638738 A CN 202110638738A CN 113495497 A CN113495497 A CN 113495497A
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satellite
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simulator
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reaction wheel
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CN113495497B (en
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林宝军
熊淑杰
白涛
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Shanghai Engineering Center for Microsatellites
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Shanghai Engineering Center for Microsatellites
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    • G05BCONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
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Abstract

The invention relates to a satellite closed loop test system, which is used for reconstructing the existing test system; the system is suitable for all large-scale tests or tests after satellite integration so as to carry out comprehensive assessment on system functions and performances in a full-task stage and a real working state; the small dynamic simulator is used as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in a satellite integrated state, so that the modification cost of the conventional test system is reduced; accurately modeling an actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop; the satellite telemetry parameters include actuator control commands and reaction wheel rotational speed; the reaction wheel replaces the command modeling in a rotating speed modeling mode, so that the single machine carries out closed-loop test under a real working state.

Description

Satellite simulation in-orbit working condition closed-loop test system
Technical Field
The invention relates to the technical field of simulation and test of an aerospace control system, in particular to a satellite simulation on-orbit working condition closed-loop test system.
Background
In order to ensure safety and reliability, the satellite needs to perform a plurality of large tests after integration, including thermal vacuum, mechanics, EMC, aging tests and the like. During testing, the ground test equipment is utilized to simulate the real environment of the satellite during flying as much as possible. Before and after the test or during the test, the function and performance of the satellite are checked by means of a power-on test. When a power-up test, particularly a long-time power-up test such as hot and aging, is performed, it is expected that all functions and performances corresponding to each stage in the life cycle of the satellite are tested in a real in-orbit running state of the satellite, so that the state of the satellite can be more fully and comprehensively checked, and design defects are more easily exposed so as to be timely solved on the ground.
In the case of a control system, on-orbit operation is a closed-loop control mode, while ground satellite integrated testing is typically an open-loop mode, since the external excitation is inevitably introduced by existing test systems. Although the on-orbit working condition is simulated as much as possible by setting the working state of the single machine, the open-loop mode of the single machine cannot completely simulate the closed-loop state, and meanwhile, the system software cannot be fully verified, so that part of design defects cannot be found in time during ground testing, and certain risk is caused to subsequent on-orbit operation.
Disclosure of Invention
The invention aims to provide a satellite simulation on-orbit working condition closed-loop test system, through which a sensor and an actuator of a satellite can be accurately simulated, so that the test closed-loop performance is ensured; in addition, the system is simple and low in cost, does not need to be greatly modified for the existing test system, and is suitable for the existing test system.
In a first aspect of the invention, this task is solved by a satellite simulation in-orbit behavior closed-loop test system comprising:
the system is used for transforming the existing test system;
the system is suitable for all large-scale tests or tests after satellite integration so as to carry out comprehensive assessment on system functions and performances in a full-task stage and a real working state;
the small dynamic simulator is used as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in a satellite integrated state, so that the modification cost of the conventional test system is reduced;
accurately modeling an actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop;
the satellite telemetry parameters include actuator control commands and reaction wheel rotational speed;
the reaction wheel replaces the command modeling in a rotating speed modeling mode, so that the single machine carries out closed-loop test under a real working state.
In another preferred embodiment of the present invention, it is provided that the satellite simulated in-orbit working condition closed-loop test system includes:
a dynamic simulator configured to receive orbit and attitude data of the satellite from the attitude and orbit dynamics model to provide a dynamic excitation source to the satellite sensor;
an attitude and orbit dynamics model configured to simulate a true flight state of the satellite, wherein the attitude and orbit dynamics model generates orbit and attitude data of the satellite from the thrust data and the output torque data received from the actuator model; and
an actuator model configured to model an actuator, wherein the actuator model generates thrust data and output torque data from telemetry states of a satellite, wherein the telemetry states include satellite control commands and reaction wheel rotational speeds, wherein the actuator model comprises:
a thruster model configured to model a thruster, wherein the modeling is an exponential model and simulates a rising edge and a falling edge of thrust according to satellite control commands to generate thrust data; and
a reaction wheel model configured to model a reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel as a function of a reaction wheel rotational speed.
In a preferred embodiment of the invention, it is provided that the thruster model models the thruster by means of the following formula:
Figure BDA0003106272770000021
wherein t isnTo command the air injection time; t is tSR,tSDOpening time delay and closing time delay of the electromagnetic valve; t is tfr,tfdIs the rise time constant and fall time constant of thrust, FiIs the thrust amplitude.
Through the preferred scheme, the rising edge and the falling edge of the thrust of the thruster can be simulated simply and accurately, so that the test accuracy is ensured.
In a further preferred embodiment of the invention, it is provided that the reaction wheel model models the reaction wheel by means of the following formula:
Tc=Tm+Td
wherein, TmFor output torque, z-direction is the direction of rotation of the reaction wheel, Tmx=Tmy=0,Tmz=Iw·(Ωtt-1),IwFor reaction wheel moment of inertia, Ω is reaction wheel rotational speed, and TdIs disturbance torque, wherein
Figure BDA0003106272770000031
Wherein U isdFor dynamic unbalance, α0Is the initial phase.
By this preferred solution, the reaction wheel can be simply and accurately simulated in a rotational speed modeling manner instead of a command modeling manner, so that the output torque is generated from the rotational speed difference in the satellite control command.
In a further preferred embodiment of the invention, it is provided that the dynamic simulator is arranged directly on the satellite sensor. Through the preferred scheme, the dynamic simulator can be directly installed on the satellite in an integrated state, so that the modification cost of the test system is reduced.
In a further preferred embodiment of the invention, it is provided that the system further comprises a time synchronization device, which is configured to perform a time calibration with reference to the ground dynamics time in order to achieve a satellite-to-ground time synchronization. Through the optimal scheme, satellite-ground time synchronization can be achieved autonomously, and the correctness of the closed-loop test process is guaranteed.
In one embodiment of the invention, it is provided that the dynamic simulator comprises one or more of the following: dynamic star maps simulators, small solar simulators, and small infrared earth simulators. It should be noted here that other optical or infrared simulators are also conceivable under the teaching of the present invention.
In a further development of the invention, provision is made for the corresponding dynamic simulator to be selected in dependence on the arrangement of the sensors of the satellites,
if the satellite is provided with the satellite sensor, selecting the dynamic star map simulator as an excitation source, and installing an optical head of the dynamic simulator at a light shield of the satellite sensor to enable the focal plane position of the simulator to be coincident with the entrance pupil position of a sensor lens;
if the satellite is provided with the earth sensor, selecting a small earth simulator as an excitation source, and installing the infrared head of the simulator at the position of a lens of the earth sensor, so that an infrared detector of the earth sensor can directly sense an infrared image generated by the simulator;
if the satellite is provided with a sun sensor, a small-sized sun simulator is selected, and the azimuth change of the sun is simulated by utilizing the intensity or the azimuth change of a light source.
In a further preferred embodiment of the invention, it is provided that the means for establishing the accurate satellite attitude dynamics model comprises: according to the specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and (4) carrying out perturbation force modeling, including light pressure, non-spherical or pneumatic, so as to obtain an accurate orbit dynamics model.
In a further preferred embodiment of the present invention, the method further comprises:
the satellite takes the attitude determination of the single satellite sensor as a main attitude determination mode, and is provided with a thruster, a reaction wheel and a magnetic torquer to complete the control of the satellite attitude and the orbit;
the satellite sends the telemetering state data to the ground integrated test system every second, and the integrated test system repackages the corresponding data and forwards the repackaged data to the satellite simulation on-orbit working condition closed-loop test system;
the satellite simulation on-orbit working condition closed-loop test system inquires the forwarding packet in a period of 5ms, and after receiving forwarding data, the calculation of an actuator model, the calculation of a dynamic model and the calculation of simulator input are completed within 5 ms;
the sensor model is used for simulating a simulator, and corresponding signals are received from the satellite attitude and orbit dynamics model according to the input of the simulator so as to generate a sensor simulation signal;
driving the simulator to generate a simulator excitation signal in a period of 100 ms;
the satellite sensor collects the simulator signal, completes the calculation of the controller within 1s and sends down the remote measurement.
In a further preferred embodiment of the present invention, the tests or tests after the satellite integration include thermal tests and aging tests to perform a comprehensive assessment of system functions and performance in a full-mission stage and a real working state.
Adopt satellite simulation in orbit operating mode closed loop test system to carry out satellite thermal vacuum test, include:
carrying out closed loop tests of all task modes under high and low temperature working conditions;
all task modes comprise a sun orientation mode, a ground orientation mode, a normal working mode and a track control mode;
in the test process, the single machine and the software work in a real on-track state.
The invention has the following beneficial effects that firstly, a small dynamic simulator is adopted as an excitation source of the sensor, and the probe part of the simulator has the characteristics of portability and easy installation, so that the simulator can be directly installed on the surface of the sensor in a satellite integrated state, thereby reducing the cost of applying the invention to the existing satellite test system; secondly, the actuator is accurately modeled by using satellite telemetering parameters such as actuator control instructions and the rotating speed of a reaction wheel as input, so that the problem that the actuator enters a test closed loop is solved, particularly, the reaction wheel replaces instruction modeling in a rotating speed modeling mode, and a single machine is ensured to carry out closed loop test in a real working state; thirdly, the dynamics time is taken as a reference, the planet is automatically moved to the ground for time synchronization, and the correctness of the closed-loop test process is ensured.
Drawings
The invention is further elucidated with reference to specific embodiments in the following description, in conjunction with the appended drawings.
FIG. 1 shows a block diagram of a satellite simulated in-orbit behavior closed-loop test system according to the present invention;
figure 2 shows a schematic output curve of a thruster model;
FIG. 3 shows a graph of attitude control in a diurnal mode when the inventive arrangements are applied to high orbit satellites;
FIG. 4 shows a satellite-to-ground mode attitude control plot when the inventive arrangements are applied to high orbit satellites;
FIG. 5 is a graph illustrating attitude control in the normal operating mode of a satellite when the solution of the present invention is applied to an elevated orbit satellite;
FIG. 6 is a graph illustrating the normal operating mode reaction wheel speed for a satellite when the inventive arrangements are applied to an elevated orbit satellite; and
fig. 7 shows a satellite orbit control mode attitude control graph when the scheme of the present invention is applied to an high orbit satellite.
Detailed Description
It should be noted that the components in the figures may be exaggerated and not necessarily to scale for illustrative purposes. In the figures, identical or functionally identical components are provided with the same reference symbols.
In the present invention, "disposed on …", "disposed over …" and "disposed over …" do not exclude the presence of an intermediate therebetween, unless otherwise specified.
In the present invention, the embodiments are only intended to illustrate the aspects of the present invention, and should not be construed as limiting.
In the present invention, the terms "a" and "an" do not exclude the presence of a plurality of elements, unless otherwise specified.
It is further noted herein that in embodiments of the present invention, only a portion of the components or assemblies may be shown for clarity and simplicity, but those of ordinary skill in the art will appreciate that, given the teachings of the present invention, required components or assemblies may be added as needed in a particular scenario.
The numbering of the steps of the methods of the present invention does not limit the order of execution of the steps of the methods. Unless specifically stated, the method steps may be performed in a different order.
In the present invention, the various models refer to mathematical models simulating respective targets, which may be implemented by software, hardware, and/or firmware.
The invention aims to solve the following difficulties in performing closed-loop test of a control system after satellite integration: firstly, adding a dynamic excitation source of a sensor in an integrated state; secondly, acquiring the working state of the actuator, and obtaining the output force or torque of the actuator by using the acquired state; finally, time synchronization between the surface systems.
In order to solve the difficulties, the invention provides a satellite simulation on-orbit working condition closed-loop test system, which is used for reconstructing the existing test system; the system is suitable for all large-scale tests or tests after satellite integration so as to carry out comprehensive assessment on system functions and performances in a full-task stage and a real working state; the small dynamic simulator is used as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in a satellite integrated state, so that the modification cost of the conventional test system is reduced; accurately modeling an actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop; the satellite telemetry parameters include actuator control commands and reaction wheel rotational speed; the reaction wheel replaces the command modeling in a rotating speed modeling mode, so that the single machine carries out closed-loop test under a real working state.
The invention is further illustrated by the following specific examples.
Fig. 1 shows a block diagram of a satellite closed loop test system 100 according to the present invention.
As shown in fig. 1, the system 100 includes a dynamic simulator 101. The dynamic simulator 101 is configured to receive orbit and attitude data of the satellite from the attitude and orbit dynamics model 102 (and optionally from the sensor model 104) to provide a dynamic excitation source to the satellite sensors of the integrated state satellite 105.
In the invention, the corresponding dynamic simulator can be selected according to the configuration condition of the sensor of the satellite. If the satellite is equipped with a satellite sensor, a dynamic constellation simulator may be selected as the excitation source. For example, the optical head of the dynamic simulator 101 may be installed at the light shield of the satellite sensor, so that the position of the simulator focal plane coincides with the position of the entrance pupil of the sensor lens; if the satellite is equipped with earth sensors, a small earth simulator can be selected as the source of excitation. For example, the infrared head of the simulator may be mounted at the lens position of the earth sensor so that the infrared detector of the earth sensor can directly sense the infrared image generated by the simulator. If the satellite is provided with a sun sensor, a small-sized sun simulator can be selected to simulate the azimuth change of the sun by utilizing the intensity or the azimuth change of a light source.
The system 100 also includes a pose and orbit dynamics model 102. The attitude and orbit dynamics model 102 is configured to simulate the true flight state of the satellite. The attitude and orbit dynamics model 102 generates orbit and attitude data for the satellite from the thrust data and output torque data received from the actuator model 103. The accurate satellite attitude dynamics model 102 is established, for example, by: according to the specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and (4) carrying out perturbation force modeling, including light pressure, non-spherical, pneumatic and the like, so as to obtain an accurate track dynamics model.
The system 100 also includes an actuator model 103. The actuator model 103 is configured to model the actuator, wherein the actuator model 103 generates thrust data and output torque data from telemetry states of the satellites. The telemetry state may be received from the ground integrated survey system 106, which includes satellite control commands and reaction wheel rotational speed. Telemetry status may also contain other satellite data.
The actuator model 103 includes:
a thruster model configured to model a thruster, wherein the modeling is an exponential model and simulates rising and falling edges of thrust according to satellite control commands to generate thrust data. One example of an exponential model of a thruster is:
Figure BDA0003106272770000071
wherein t isnTo command the air injection time; t is tSR,tSDOpening time delay and closing time delay of the electromagnetic valve; t is tfr,tfdIs the rise time constant and fall time constant of thrust, FiIs the thrust amplitude. Fig. 2 shows a comparison between the model and the actual thrust. As can be seen from fig. 2, by the exponential model, it is possible to accurately simulate the rising edge and the falling edge of the thrust, thereby simply and accurately simulating the thrust.
A reaction wheel model configured to model a reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel as a function of a reaction wheel rotational speed. One example of a model of a reaction wheel is:
the output torque of the reaction wheel is characterized by:
Tc=Tm+Td (2)
wherein, TmFor output torque, z-direction is the direction of rotation of the reaction wheel, Tmx=Tmy=0,Tmz=Iw·(Ωtt-1),IwFor reaction wheel moment of inertia, Ω is reaction wheel speed, TdIs disturbance torque, wherein
Figure BDA0003106272770000072
Wherein U isdFor dynamic unbalance, α0Is the initial phase.
The system 100 may optionally also include a sensor model 104. The sensor model 104 is used to simulate a simulator, which receives corresponding signals from the satellite attitude and orbit dynamics model to generate sensor simulation signals.
The following describes the flow of the satellite closed loop test according to the present invention.
First, thrust data and output torque data are generated by the actuator model 103 from satellite telemetry states received, for example, from the ground integrated test system 106, wherein the actuator model 103 is configured to model the actuator, wherein the telemetry states include satellite control commands and reaction wheel rotational speeds, and wherein the actuator model 103 includes a thruster model and a reaction wheel model.
Then, orbit and attitude data for the satellite is generated by the attitude and orbit dynamics model 102 from the thrust data and output torque data received from the actuator model, wherein the attitude and orbit dynamics model is configured to simulate the true flight state of the satellite.
Finally, the orbit and attitude data of the satellite is received by the dynamic simulator 101 from the attitude and orbit dynamics model 102 to provide a dynamic excitation source to the satellite sensors of the integrated state satellite 105.
An exemplary embodiment of the solution according to the invention on an elevated earth satellite is given below.
The satellite takes the attitude determination of an independent satellite sensor as a main attitude determination mode, and is provided with a thruster, a reaction wheel and a magnetic torquer to complete the control of the satellite attitude and the orbit. The satellite sends telemetering state data to the ground integrated test system 106 every second, the integrated test system repacks corresponding data and forwards the repacked data to the test system 100 through the network, the test system 100 conducts forwarding packet query in a period of 5ms, after receiving the forwarding data, the computation of the actuator model 103, the computation of the dynamic model 102 and the computation of simulator input (namely the sensor model 104) are completed within 5ms, and then the simulator is driven to generate a simulator excitation signal in a period of 100 ms. The satellite sensor collects the simulator signal, completes the calculation of the controller within 1s and sends down the remote measurement.
The large-scale tests after satellite integration, including integration tests, thermal vacuum tests, aging tests and the like, adopt the method to perform closed-loop tests in the whole task stage, and obtain good effects. The closed loop test of the satellite in the thermal vacuum test is taken as an example to illustrate the superiority of the invention in the application process.
The satellite thermal vacuum test adopts the method provided by the invention, and performs closed loop tests of all task modes (including a sun-oriented mode, a ground-oriented mode, a normal working mode and an orbit control mode) under various working conditions of high temperature and low temperature. The attitude control curves for the sun-oriented mode satellite are shown in fig. 3, the attitude control curves for the earth-oriented mode are shown in fig. 4, the attitude control curves for the normal operation mode are shown in fig. 5, the rotational speed curves for the normal operation mode reaction wheels are shown in fig. 6, and the attitude control curves for the orbit control mode are shown in fig. 7. In the test process, the single machine and the software work in a real on-orbit state, and the functions and the performances under each task mode are fully examined.
The invention has the following beneficial effects that firstly, a small dynamic simulator is adopted as an excitation source of the sensor, and the probe part of the simulator has the characteristics of portability and easy installation, so that the simulator can be directly installed on the surface of the sensor in a satellite integrated state, thereby reducing the cost of applying the invention to the existing satellite test system; secondly, the actuator is accurately modeled by using satellite telemetering parameters such as actuator control instructions and the rotating speed of a reaction wheel as input, so that the problem that the actuator enters a test closed loop is solved, particularly, the reaction wheel replaces instruction modeling in a rotating speed modeling mode, and a single machine is ensured to carry out closed loop test in a real working state; thirdly, the dynamics time is taken as a reference, the planet is automatically moved to the ground for time synchronization, and the correctness of the closed-loop test process is ensured.
Although some embodiments of the present invention have been described herein, those skilled in the art will appreciate that they have been presented by way of example only. Numerous variations, substitutions and modifications will occur to those skilled in the art in light of the teachings of the present invention without departing from the scope thereof. It is intended that the following claims define the scope of the invention and that methods and structures within the scope of these claims and their equivalents be covered thereby.

Claims (10)

1. A satellite simulation on-orbit working condition closed-loop test system is characterized in that,
the system is used for transforming the existing test system;
the system is suitable for all large-scale tests or tests after satellite integration so as to carry out comprehensive assessment on system functions and performances in a full-task stage and a real working state;
the small dynamic simulator is used as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in a satellite integrated state, so that the modification cost of the conventional test system is reduced;
accurately modeling an actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop;
the satellite telemetry parameters include actuator control commands and reaction wheel rotational speed;
the reaction wheel replaces the command modeling in a rotating speed modeling mode, so that the single machine carries out closed-loop test under a real working state.
2. The system of claim 1, the satellite simulated in-orbit behavior closed-loop test system comprising:
a dynamic simulator configured to receive orbit and attitude data of the satellite from the attitude and orbit dynamics model to provide a dynamic excitation source to the satellite sensor;
an attitude and orbit dynamics model configured to simulate a true flight state of the satellite, wherein the attitude and orbit dynamics model generates orbit and attitude data of the satellite from the thrust data and the output torque data received from the actuator model; and
an actuator model configured to model an actuator, wherein the actuator model generates thrust data and output torque data from telemetry states of a satellite, wherein the telemetry states include satellite control commands and reaction wheel rotational speeds, wherein the actuator model comprises:
a thruster model configured to model a thruster, wherein the modeling is an exponential model and simulates a rising edge and a falling edge of thrust according to satellite control commands to generate thrust data; and
a reaction wheel model configured to model a reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel as a function of a reaction wheel rotational speed.
3. The system of claim 2, wherein the thruster model models the thruster to simulate a rising edge and a falling edge of thrust of the thruster by the following formula to ensure the accuracy of the test:
Figure FDA0003106272760000011
wherein t isnTo command the air injection time; t is tSR,tSDOpening time delay and closing time delay of the electromagnetic valve; t is tfr,tfdIs the rise time constant and fall time constant of thrust, FiIs the thrust amplitude.
4. The system of claim 3, wherein the reaction wheel model models the reaction wheel by modeling the reaction wheel by implementing a speed modeling approach instead of a command modeling approach, generating an output torque based on a speed difference in the satellite control commands by:
Tc=Tm+Td
wherein, TmFor output torque, z-direction is the direction of rotation of the reaction wheel, Tmx=Tmy=0,Tmz=Iw·(Ωtt-1),IwFor reaction wheel moment of inertia, Ω is reaction wheel rotational speed, and TdIs disturbance torque, wherein
Figure FDA0003106272760000021
Wherein U isdFor dynamic unbalance, α0Is the initial phase.
5. The system of claim 2, further comprising a time synchronization device configured to perform time calibration with reference to ground dynamics time to achieve satellite-to-ground time synchronization, autonomously achieve satellite-to-ground time synchronization, and ensure correctness of the closed loop test process.
6. The system of claim 2, wherein the dynamic simulator comprises one or more of: dynamic star maps simulators, small solar simulators, and small infrared earth simulators.
7. The system of claim 6, wherein the dynamic simulator is selected according to the configuration of the sensors of the satellites,
if the satellite is provided with the satellite sensor, selecting the dynamic star map simulator as an excitation source, and installing an optical head of the dynamic simulator at a light shield of the satellite sensor to enable the focal plane position of the simulator to be coincident with the entrance pupil position of a sensor lens;
if the satellite is provided with the earth sensor, selecting a small earth simulator as an excitation source, and installing the infrared head of the simulator at the position of a lens of the earth sensor, so that an infrared detector of the earth sensor can directly sense an infrared image generated by the simulator;
if the satellite is provided with a sun sensor, a small-sized sun simulator is selected, and the azimuth change of the sun is simulated by utilizing the intensity or the azimuth change of a light source.
8. The system of claim 2, the means for establishing an accurate satellite attitude dynamics model comprising: according to the specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and (4) carrying out perturbation force modeling, including light pressure, non-spherical or pneumatic, so as to obtain an accurate orbit dynamics model.
9. The system of claim 2, further comprising:
the satellite takes the attitude determination of the single satellite sensor as a main attitude determination mode, and is provided with a thruster, a reaction wheel and a magnetic torquer to complete the control of the satellite attitude and the orbit;
the satellite sends the telemetering state data to the ground integrated test system every second, and the integrated test system repackages the corresponding data and forwards the repackaged data to the satellite simulation on-orbit working condition closed-loop test system;
the satellite simulation on-orbit working condition closed-loop test system inquires the forwarding packet in a period of 5ms, and after receiving forwarding data, the calculation of an actuator model, the calculation of a dynamic model and the calculation of simulator input are completed within 5 ms;
the sensor model is used for simulating a simulator, and corresponding signals are received from the satellite attitude and orbit dynamics model according to the input of the simulator so as to generate a sensor simulation signal;
driving the simulator to generate a simulator excitation signal in a period of 100 ms;
the satellite sensor collects the simulator signal, completes the calculation of the controller within 1s and sends down the remote measurement.
10. The system of claim 9, wherein the post-satellite integration trials or tests include thermal trials and burn-in trials;
adopt satellite simulation in orbit operating mode closed loop test system to carry out satellite thermal vacuum test, include:
carrying out closed loop tests of all task modes under high and low temperature working conditions;
all task modes comprise a sun orientation mode, a ground orientation mode, a normal working mode and a track control mode;
in the test process, the single machine and the software work in a real on-track state.
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