CN113495497B - Satellite simulation on-orbit working condition closed loop test system - Google Patents
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Abstract
The invention relates to a satellite closed-loop test system, which is used for modifying the existing test system; the system is suitable for all large-scale experiments or tests after satellite integration so as to carry out comprehensive examination of system functions and performances in a full-task stage and a real working state; the small dynamic simulator is adopted as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in the satellite integrated state, so that the transformation cost of the existing test system is reduced; accurately modeling the actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop; the satellite telemetry parameters comprise an actuator control instruction and a reaction wheel rotating speed; the reaction wheel replaces instruction modeling in a rotating speed modeling mode, so that the single machine performs closed loop testing in a real working state.
Description
Technical Field
The invention relates to the technical field of simulation and testing of space flight control systems, in particular to a satellite simulation on-orbit working condition closed-loop testing system.
Background
In order to ensure safety and reliability, satellites are subjected to a number of large-scale tests after integration, including thermal vacuum, mechanical, EMC, and aging tests. During the test, the real environment of the satellite in flight is simulated by using the ground test equipment as much as possible. The satellite functions and performance are checked by means of a power-up test before and after the test or during the test. When the power-up test, especially the hot and aging power-up test, is carried out for a long time, all functions and performances corresponding to each stage in the life cycle of the satellite are expected to be tested in the real in-orbit running state of the satellite, so that the satellite state can be more fully and comprehensively checked, and design defects are easier to expose so as to be solved in time when the satellite is on the ground.
For control systems, the in-orbit operation is a closed-loop control mode, while the post-ground satellite integration test is typically an open-loop mode, as the existing test systems can inevitably introduce external stimuli. 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 timely found in the ground test, and a certain risk is caused for the follow-up on-orbit operation.
Disclosure of Invention
The invention aims to provide a satellite on-orbit working condition simulation closed-loop test system, through which a sensor and an actuator of a satellite can be accurately simulated, so that the closed-loop performance of the test is ensured; in addition, the system is simple and low in cost, does not need to be greatly modified on the existing test system, and is suitable for being used in the existing test system.
In a first aspect of the invention, the object is achieved by a satellite simulation on-orbit operating mode closed loop test system comprising:
the system is used for modifying the existing test system;
The system is suitable for all large-scale experiments or tests after satellite integration so as to carry out comprehensive examination of system functions and performances in a full-task stage and a real working state;
The small dynamic simulator is adopted as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in the satellite integrated state, so that the transformation cost of the existing test system is reduced;
Accurately modeling the actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop;
The satellite telemetry parameters comprise an actuator control instruction and a reaction wheel rotating speed;
The reaction wheel replaces instruction modeling in a rotating speed modeling mode, so that the single machine performs closed loop testing in a real working state.
In another preferred embodiment of the present invention, the satellite simulation on-orbit condition closed loop test system comprises:
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 for the satellite sensor;
A pose and orbit dynamics model configured to simulate a real flight state of a satellite, wherein the pose and orbit dynamics model generates orbit and pose data of the satellite from thrust data and 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 status of satellites, wherein the telemetry status includes satellite control instructions and reaction wheel speeds, wherein the actuator model comprises:
a thruster model configured to model a thruster, wherein the modeling is an exponential model and simulates rising and falling edges of a thrust according to satellite control instructions to generate thrust data; and
A reaction wheel model configured to model the reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel from the reaction wheel rotational speed.
In a preferred embodiment of the invention, it is provided that the thruster model models a thruster by the following formula:
Wherein t n is the commanded jet time; t SR,tSD is the opening time delay and the closing time delay of the electromagnetic valve; t fr,tfd is the rise time constant and fall time constant of the thrust, and F i is the thrust amplitude.
By 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 accuracy of the test is ensured.
In a further preferred embodiment of the invention, it is provided that the reaction wheel model models the reaction wheel by the following formula:
Tc＝Tm+Td
Wherein T m is the output torque, z direction is the rotational direction of the reaction wheel, T mx＝Tmy＝0,Tmz＝Iw·(Ωt-Ωt-1),Iw is the rotational inertia of the reaction wheel, Ω is the rotational speed of the reaction wheel, and T d is the disturbance torque, wherein
Where U d is the dynamic unbalance and α 0 is the initial phase.
By this preferred solution, the reaction wheel can be modeled simply and accurately in a rotational speed modeling manner, rather than in 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. By means of the preferred scheme, the dynamic simulator can be directly installed on the satellite in an integrated state, and therefore 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 for achieving the satellite-to-ground time synchronization. By the aid of the optimal scheme, satellite-ground time synchronization can be achieved autonomously, and accuracy of a closed loop testing process is guaranteed.
In one embodiment of the invention, the dynamic simulator comprises one or more of the following: dynamic star image simulators, small solar simulators, and small infrared earth simulators. It should be noted herein that other optical or infrared detectors are also contemplated under the teachings of the present invention.
In one embodiment of the invention, it is provided that, depending on the configuration of the satellite sensors, a corresponding dynamic simulator is selected,
If the satellite is provided with a satellite sensor, a dynamic star map simulator is selected as an excitation source, and an optical head of the dynamic simulator is arranged at a lens hood of the satellite sensor, so that the focal plane position of the simulator coincides with the entrance pupil position of a lens of the sensor;
if the satellite is provided with an earth sensor, a small earth simulator is selected as an excitation source, and the infrared head of the simulator is arranged at the lens position of the earth sensor, so that the 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 solar simulator is selected, and the sun azimuth change is simulated by utilizing the intensity or azimuth change of the light source.
In a further preferred embodiment of the invention, the method for establishing an accurate satellite attitude dynamics model comprises: according to specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and performing the perturbation modeling, including light pressure, non-spherical or pneumatic, 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 an independent satellite sensor as a main attitude determination mode, and a thruster, a reaction wheel and a magnetic torquer are configured to complete the satellite attitude and orbit control;
The satellite transmits telemetry state data to a ground comprehensive testing system every second, and the comprehensive testing system reassembles corresponding data and forwards the data to a satellite on-orbit simulation working condition closed-loop testing system through a network;
The satellite simulation on-orbit working condition closed loop test system inquires forwarding packets in a period of 5ms, and after receiving forwarding data, the satellite simulation on-orbit working condition closed loop test system completes calculation of an actuator model, calculation of a dynamics model and calculation of simulator input 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 with a period of 100ms to generate a simulator excitation signal;
the satellite sensor collects the simulator signals, completes the calculation of the controller within 1s and issues telemetry.
In a further preferred embodiment of the invention, it is provided that the post-satellite-integration test or test comprises a thermal test and a burn-in test in order to carry out a full-mission phase, a full-scale functional, performance assessment of the system in the actual operating state.
The satellite thermal vacuum test is carried out by adopting a satellite simulation on-orbit working condition closed loop test system, which comprises the following steps:
Carrying out closed loop test of all task modes under each working condition of high temperature and low temperature;
all task modes comprise a sun oriented mode, a ground oriented mode, a normal working mode and a track control mode;
During the test, both the standalone and the software are operating in a truly on-track state.
The invention has the advantages that firstly, a small dynamic simulator is adopted as an excitation source of the sensor, and the simulator probe part has the characteristics of portability and easy installation, so that the simulator probe part can be directly installed on the surface of the sensor in a satellite integration state, thereby reducing the cost of the invention applied to the existing satellite test system; secondly, satellite telemetry parameters such as an actuator control command and a reaction wheel rotating speed are used as input to carry out accurate modeling of the actuator, so that the problem that the actuator enters a test closed loop is solved, and particularly, the reaction wheel replaces command modeling in a rotating speed modeling mode, and the closed loop test of a single machine in a real working state is ensured; thirdly, taking dynamic time as a reference, automatically feeding planetary ground time synchronization, and ensuring the correctness of a closed loop test process.
Drawings
The invention will be further elucidated with reference to a specific embodiment in conjunction with the drawings.
FIG. 1 illustrates a block diagram of a satellite simulated on-orbit operating condition closed loop test system according to the present invention;
FIG. 2 shows a schematic output curve of a thruster model;
FIG. 3 shows a graph of a relative-day mode attitude control curve when the scheme of the present invention is applied to a high orbit satellite;
FIG. 4 shows a plot of satellite versus mode attitude control for an application of the inventive solution to a high orbit satellite;
FIG. 5 shows a satellite normal operation mode attitude control graph when the scheme of the present invention is applied to a high orbit satellite;
FIG. 6 shows a satellite normal operating mode reaction wheel speed graph for a high orbit satellite with the scheme of the present invention applied; and
Fig. 7 shows a satellite orbit control mode attitude control graph when the scheme of the present invention is applied to a high orbit satellite.
Detailed Description
It should be noted that the components in the figures may be shown exaggerated for illustrative purposes and are not necessarily to scale. In the drawings, identical or functionally identical components are provided with the same reference numerals.
In the present invention, unless specifically indicated otherwise, "disposed on …", "disposed over …" and "disposed over …" do not preclude the presence of an intermediate therebetween.
In the present invention, the embodiments are merely intended to illustrate the scheme of the present invention, and should not be construed as limiting.
In the present invention, the adjectives "a" and "an" do not exclude a scenario of a plurality of elements, unless specifically indicated.
It should also be 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 the components or assemblies may be added as needed for a particular scenario under the teachings of the present invention.
The numbers of the steps of the respective methods of the present invention are not limited to the order of execution of the steps of the methods. The method steps may be performed in a different order unless otherwise indicated.
In the present invention, various models refer to mathematical models that simulate corresponding targets, which may be implemented by software, hardware, and/or firmware.
The invention aims to solve the following difficulties in closed loop testing of a control system after satellite integration: firstly, adding a dynamic excitation source of a sensor in an integrated state; secondly, collecting working states of the actuator, and obtaining output force or torque by utilizing the collected states; finally, time synchronization between the surface systems.
In order to solve the problems, the invention provides a satellite simulation on-orbit working condition closed loop test system which is used for modifying the existing test system; the system is suitable for all large-scale experiments or tests after satellite integration so as to carry out comprehensive examination of system functions and performances in a full-task stage and a real working state; the small dynamic simulator is adopted as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in the satellite integrated state, so that the transformation cost of the existing test system is reduced; accurately modeling the actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop; the satellite telemetry parameters comprise an actuator control instruction and a reaction wheel rotating speed; the reaction wheel replaces instruction modeling in a rotating speed modeling mode, so that the single machine performs closed loop testing in a real working state.
The invention is further illustrated below in connection with 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 satellite orbit and attitude data from the attitude and orbit dynamics model 102 (optionally from the sensor model 104) to provide a dynamic excitation source for 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 satellite sensor. If the satellite is configured with satellite sensors, a dynamic star map simulator may be selected as the excitation source. For example, the optical head of the dynamic simulator 101 may be mounted at the mask of the satellite sensor such that the simulator focal plane position coincides with the entrance pupil position of the sensor lens; if the satellite is configured with an earth sensor, a small earth simulator may be selected as the excitation source. 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 solar simulator can be selected, and the sun azimuth change is simulated by utilizing the intensity or azimuth change of the 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 real flight state of the satellite. The attitude and orbit dynamics model 102 generates orbit and attitude data for the satellites from the thrust data and output torque data received from the actuator model 103. The manner of establishing the accurate satellite attitude dynamics model 102 is, for example: according to specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and performing perturbation modeling including light pressure, non-spherical, aerodynamic and the like to obtain an accurate orbit 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 status of the satellites. Telemetry status may be received from the surface integrated system 106, which includes satellite control commands and reaction wheel speeds. Telemetry status may also include 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 a thrust according to satellite control instructions to generate thrust data. One example of an exponential model of a thruster is:
Wherein t n is the commanded jet time; t SR,tSD is the opening time delay and the closing time delay of the electromagnetic valve; t fr,tfd is the rise time constant and fall time constant of the thrust, and F i is the thrust amplitude. Fig. 2 shows the comparison between the model and the actual thrust. As can be seen from fig. 2, by this exponential model, the rising and falling edges of the thrust can be accurately simulated, thereby simply and accurately simulating the thrust.
A reaction wheel model configured to model the reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel from the reaction wheel rotational speed. One example of a model of a reaction wheel is:
The output torque of the reaction wheel is characterized by the following formula:
Tc＝Tm+Td (2)
Wherein T m is the output torque, the z direction is the rotation direction of the reaction wheel, T mx＝Tmy＝0,Tmz＝Iw·(Ωt-Ωt-1),Iw is the rotation inertia of the reaction wheel, Ω is the rotation speed of the reaction wheel, and T d is the disturbance torque, wherein
Where U d is the dynamic unbalance and α 0 is the initial phase.
The system 100 may optionally also include a sensor model 104. The sensor model 104 is used to simulate a simulator that receives corresponding signals from the satellite attitude and orbit dynamics models to generate sensor simulation signals.
The flow of satellite closed loop testing according to the present invention is set forth below.
First, thrust data and output torque data are generated by the actuator model 103 from satellite telemetry conditions, e.g., received from the ground surface comprehensive survey system 106, wherein the actuator model 103 is configured to model the actuator, wherein the telemetry conditions include satellite control commands and reaction wheel speeds, and wherein the actuator model 103 includes a thruster model and a reaction wheel model.
Orbit and attitude data for the satellite is then 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 real flight state of the satellite.
Finally, the orbit and attitude data of the satellites is received by the dynamic simulator 101 from the attitude and orbit dynamics model 102 to provide a dynamic excitation source for the satellite sensors of the integrated state satellites 105.
Exemplary embodiments on high orbit satellites according to the scheme of the present invention are given below.
The satellite takes the attitude determination of an independent satellite sensor as a main attitude determination mode, and a thruster, a reaction wheel and a magnetic torquer are configured to complete the satellite attitude and orbit control. The satellite transmits telemetry status data to the ground comprehensive testing system 106 every second, the comprehensive testing system reassembles corresponding data into packets and forwards the packets to the testing system 100 through a network, the testing system 100 inquires the forwarded packets in a period of 5ms, after receiving the forwarded data, the calculation of the actuator model 103, the calculation of the dynamics model 102 and the calculation of simulator input (namely the sensor model 104) are completed in 5ms, and then the simulator is driven to generate simulator excitation signals in a period of 100 ms. The satellite sensor collects the simulator signals, completes the calculation of the controller within 1s and issues telemetry.
The large-scale test after satellite integration, including the integration test, the thermal vacuum test, the aging test and the like, adopts the method to carry out the closed loop test at the full task stage, and achieves good effects. The superiority of the invention in the application process will be described below by taking a closed loop test as an example when a satellite is subjected to a thermal vacuum test.
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 sun-to-earth directional mode satellite attitude control curve is shown in fig. 3, the earth directional mode attitude control curve is shown in fig. 4, the normal operating mode attitude control curve is shown in fig. 5, the normal operating mode reaction wheel speed curve is shown in fig. 6, and the orbit control mode attitude control curve is 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 in each task mode are fully checked.
The invention has the advantages that firstly, a small dynamic simulator is adopted as an excitation source of the sensor, and the simulator probe part has the characteristics of portability and easy installation, so that the simulator probe part can be directly installed on the surface of the sensor in a satellite integration state, thereby reducing the cost of the invention applied to the existing satellite test system; secondly, satellite telemetry parameters such as an actuator control command and a reaction wheel rotating speed are used as input to carry out accurate modeling of the actuator, so that the problem that the actuator enters a test closed loop is solved, and particularly, the reaction wheel replaces command modeling in a rotating speed modeling mode, and the closed loop test of a single machine in a real working state is ensured; thirdly, taking dynamic time as a reference, automatically feeding planetary ground time synchronization, and ensuring the correctness of a closed loop test process.
Although some embodiments of the present application have been described in the present document, those skilled in the art will appreciate that these embodiments are shown by way of example only. Numerous variations, substitutions and modifications will occur to those skilled in the art in light of the present teachings without departing from the scope of the application. The appended claims are intended to define the scope of the application and to cover such methods and structures within the scope of these claims themselves and their equivalents.
Claims (7)
1. A satellite simulation on-orbit working condition closed loop test system is characterized in that,
The system is used for modifying the existing test system;
The system is suitable for all large-scale experiments or tests after satellite integration so as to carry out comprehensive examination of system functions and performances in a full-task stage and a real working state;
The small dynamic simulator is adopted as an excitation source of the sensor and is directly arranged on the surface of the satellite sensor in the satellite integrated state, so that the transformation cost of the existing test system is reduced;
Accurately modeling the actuator by using satellite telemetry parameters as input so as to enable the actuator to enter a test closed loop;
The satellite telemetry parameters comprise an actuator control instruction and a reaction wheel rotating speed;
the reaction wheel replaces instruction modeling in a rotating speed modeling mode, so that a single machine performs closed-loop test in a real working state;
the satellite simulation on-orbit working condition closed loop test system comprises:
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 for the satellite sensor;
A pose and orbit dynamics model configured to simulate a real flight state of a satellite, wherein the pose and orbit dynamics model generates orbit and pose data of the satellite from thrust data and 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 status of satellites, wherein the telemetry status includes satellite control instructions and reaction wheel speeds, wherein the actuator model comprises:
a thruster model configured to model a thruster, wherein the modeling is an exponential model and simulates rising and falling edges of a thrust according to satellite control instructions to generate thrust data; and
A reaction wheel model configured to model the reaction wheel, wherein the reaction wheel model calculates an output torque of the reaction wheel from the reaction wheel rotational speed to ensure that the stand-alone performs a closed-loop test in a true operating state;
The thruster model models the thruster through the following formulas to simulate the rising edge and the falling edge of the thrust of the thruster, so that the accuracy of the test is ensured:
Wherein t n is the commanded jet time; t SR,tSD is the opening time delay and the closing time delay of the electromagnetic valve; t fr,tfd is the rising time constant and the falling time constant of the thrust, and F i is the thrust amplitude;
the reaction wheel model models the reaction wheel through the following formula, so that the reaction wheel is simulated in a mode of realizing rotational speed modeling instead of a mode of instruction modeling, and output torque is generated according to rotational speed difference in satellite control instructions:
Tc＝Tm+Td
Wherein T m is the output torque, z direction is the rotational direction of the reaction wheel, T mx＝Tmy＝0,Tmz＝Iw·(Ωt-Ωt-1),Iw is the rotational inertia of the reaction wheel, Ω is the rotational speed of the reaction wheel, and T d is the disturbance torque, wherein
Where U d is the dynamic unbalance and α 0 is the initial phase.
2. The system of claim 1, further comprising a time synchronization device configured to perform time calibration based on 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 procedure.
3. The system of claim 1, wherein the dynamic simulator comprises one or more of the following: dynamic star image simulators, small solar simulators, and small infrared earth simulators.
4. A system according to claim 3, wherein the corresponding dynamic simulator is selected in dependence on the configuration of the satellite's sensors,
If the satellite is provided with a satellite sensor, a dynamic star map simulator is selected as an excitation source, and an optical head of the dynamic simulator is arranged at a lens hood of the satellite sensor, so that the focal plane position of the simulator coincides with the entrance pupil position of a lens of the sensor;
if the satellite is provided with an earth sensor, a small earth simulator is selected as an excitation source, and the infrared head of the simulator is arranged at the lens position of the earth sensor, so that the 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 solar simulator is selected, and the sun azimuth change is simulated by utilizing the intensity or azimuth change of the light source.
5. The system of claim 1, the means for establishing an accurate satellite attitude dynamics model comprising: according to specific characteristics of the satellite, a sailboard flexible or liquid shaking accessory is selected to be added; and performing the perturbation modeling, including light pressure, non-spherical or pneumatic, to obtain an accurate orbit dynamics model.
6. The system of claim 1, further comprising:
The satellite takes the attitude determination of an independent satellite sensor as a main attitude determination mode, and a thruster, a reaction wheel and a magnetic torquer are configured to complete the satellite attitude and orbit control;
The satellite transmits telemetry state data to a ground comprehensive testing system every second, and the comprehensive testing system reassembles corresponding data and forwards the data to a satellite on-orbit simulation working condition closed-loop testing system through a network;
The satellite simulation on-orbit working condition closed loop test system inquires forwarding packets in a period of 5ms, and after receiving forwarding data, the satellite simulation on-orbit working condition closed loop test system completes calculation of an actuator model, calculation of a dynamics model and calculation of simulator input 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 with a period of 100ms to generate a simulator excitation signal;
the satellite sensor collects the simulator signals, completes the calculation of the controller within 1s and issues telemetry.
7. The system of claim 6, wherein the post-satellite integration test or test comprises a thermal test and a burn-in test;
The satellite thermal vacuum test is carried out by adopting a satellite simulation on-orbit working condition closed loop test system, which comprises the following steps:
Carrying out closed loop test of all task modes under each working condition of high temperature and low temperature;
all task modes comprise a sun oriented mode, a ground oriented mode, a normal working mode and a track control mode;
During the test, both the standalone and the software are operating in a truly on-track state.
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| CN109143280B (en) | 2021-05-07 |
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