A control system index evaluation method based on external fine measurement mode

Abstract

The present application relates to a kind of control system index evaluation method based on external fine measurement mode, first based on analog star body builds external fine measurement system, provides control system index evaluation system environment.The angle between the two pointing target angles before and after the theodolite calibration analog star body maneuver is recorded The total time of maneuver, verify attitude control maneuver speed index.Using autocollimator measure the angle change of the initial pointing position of load simulator and the return position after maneuver, calculate the angle difference between the two stable positions before and after maneuver, combined with system error and random error, verify the attitude control precision index of control system.Using dual-frequency laser interferometer and analog star body form stable interference measurement optical path, from the imaging start time to the imaging end time, the stability of the real-time data measured by dual-frequency laser interferometer Angle, verify the attitude control stability index of control system, realize the index evaluation of spacecraft control system through external fine measurement mode, reduce professional requirements, good universality.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of spacecraft control system design and verification, and particularly relates to a control system index evaluation method based on an external fine measurement mode. BACKGROUND

[0002] At present, a control system is an important component of a spacecraft, and has the characteristics of system complexity, great design difficulty and ground verification difficulty. For a major engineering spacecraft, the system is more complex, and in the ground verification stage, a full physical simulation test mode is often used to verify the key performance of the control system. In the ground verification development process, the control system index evaluation is usually achieved by statistically controlling the telemetry parameters of the control system, and calculating the control system index. The control system data is relatively complex, and if the data acquisition frequency requirement is high, the index result must be directly calculated by the control system controller. However, this mode requires that the index evaluation personnel must be the spacecraft model development personnel, and has a deep understanding of the control system of the spacecraft, so that the professional requirement is high. The system index evaluation method has a strong customization phenomenon, and cannot be quickly transplanted to other systems, resulting in resource waste and poor universality. SUMMARY

[0003] The purpose of the present application is to provide a control system index evaluation method based on an external fine measurement mode, which solves at least one technical problem of the spacecraft control system in the verification process in the background.

[0004] In order to achieve the above purpose, the present application provides a control system index evaluation method based on an external fine measurement mode, characterized by comprising the following steps:

[0005] S1, a control system index evaluation system environment is built based on a simulated star body, including an attitude control maneuvering speed test environment, an attitude control precision test environment and an attitude control stability test environment. The simulated star body includes a first attitude control platform, a second active pointing control platform, a tracking pointing mechanism and a first autocollimator. The second active pointing control platform can move synchronously with the first attitude control platform, and can rotate around the rotation center relative to the first attitude control platform. The tracking pointing mechanism is installed on the second active pointing control platform, and can move synchronously with the second active pointing control platform, and can rotate relative to the second active pointing control platform;

[0006] S2, based on the attitude control maneuvering speed test environment, the attitude control maneuvering speed evaluation is carried out: the control system performs maneuvering control on the tracking pointing mechanism, and moves from an initial pointing position to a specified pointing position. The included angle between the initial pointing position and the specified pointing position is calibrated by a theodolite as θ, the total time from the initial pointing position to the specified pointing position is recorded, and the attitude control maneuvering speed index is verified.

[0007] S3, based on the attitude control precision test environment, the attitude control precision evaluation: the first autocollimator measuring the initial pointing position of the secondary active pointing control platform and the angle change of the return position after the maneuver, the angle difference of the two stable positions before and after the maneuver is calculated, combined with the system error and the random error, the attitude control precision index is verified, wherein the random error is three times the standard deviation;

[0008] S4, based on the attitude control stability test environment, the attitude control stability evaluation: after the control system enters the directional mode, the position change of the secondary active pointing control platform is monitored in real time by the laser interferometer, the optical axis stability in the imaging period is calculated, and the attitude control stability index is verified, wherein the optical axis stability is three times the standard deviation.

[0009] Optionally, the attitude control maneuvering speed test environment in step S1 is built by the following method:

[0010] A first autocollimator is installed on the tracking pointing mechanism, a first pointing target is arranged at the starting position of the maneuver, a second pointing target is arranged at the end position of the maneuver, and a theodolite is arranged at the vertical intersection of the first pointing target and the second pointing target on the periphery of the simulated star body.

[0011] The first pointing target is provided with a first cubic reference prism, and the second pointing target is provided with a second cubic reference prism, and the first autocollimator is autocollimated with the first pointing target and the second pointing target when pointing to the first pointing target and the second pointing target.

[0012] Optionally, a second autocollimator is arranged on the periphery of the simulated star body, and a third cubic reference prism is installed on the tracking pointing mechanism. At the initial position, the second autocollimator is autocollimated with the third cubic reference prism.

[0013] Optionally, the attitude control stability test environment in step S1 is built by the following method:

[0014] A fourth cubic reference prism is installed at the front end of the secondary active pointing control platform, and the distance from the rotation center of the load simulator is not less than 2000mm;

[0015] A dual-frequency laser interferometer is arranged on the periphery of the simulated star body, and the dual-frequency laser interferometer forms a stable interference measurement light path with the fourth cubic reference prism.

[0016] Optionally, the calculation formula of the attitude control precision index in step S3 is:

[0017]

[0018] wherein, is the average value of the angle measured by the autocollimator before the maneuver, is the average value of the angle measured by the autocollimator after the maneuver, σ is the standard deviation of the random error, and δ is the systematic error.

[0019] Optionally, the formula for calculating the optical axis stability in step S4 is:

[0020]

[0021] wherein, θ i is the angle value collected by the laser interferometer in real time, is the average value of the collected angle, and n is the number of sampling points.

[0022] Optionally, the angle calibration error of the attitude control maneuver speed test environment is not greater than 4.3″.

[0023] Optionally, the autocollimator measurement error of the attitude control precision test environment is not greater than 0.1″.

[0024] Optionally, the angle test error of the attitude control stability test environment is not greater than 0.01″.

[0025] Optionally, the external precision measuring instrument includes a theodolite, a first autocollimator, a second autocollimator, and a dual-frequency laser interferometer, and the controlled object is a first attitude control platform, a second active pointing control platform, and a tracking designated mechanism, wherein the first attitude control platform is a single-axis air floatation table, and the second active pointing control platform is a load simulator.

[0026] The above technical solutions of the present application have the following advantages:

[0027] The application provides a control system index evaluation method based on an external fine measurement mode. Firstly, a control system attitude control maneuvering speed, attitude control precision and attitude control stability index evaluation system environment is built. The attitude control maneuvering speed evaluation process is as follows: the control system controls the simulated star body to maneuver, the included angle of the two pointing target angles is calibrated through the theodolite, the total time from the starting position to the target position is recorded, and the attitude control maneuvering speed index is verified. The attitude control precision evaluation process is as follows: the angle change of the initial pointing position of the load simulator and the position after maneuvering is measured by using the second autocollimator, the angle difference between the two stable positions before and after maneuvering is calculated, the system error and the random error are combined, and the attitude control precision index of the control system is verified. The attitude control stability evaluation process is as follows: after the simulated star body is pointed to the position, the control system enters the orientation mode to realize accurate and high-stable pointing to the target, the attitude is kept stable for a long time to meet the imaging requirement time, a good attitude environment is created for the load imaging, the stable interference measurement light path of the simulated star body is formed by using the dual-frequency laser interferometer, the real-time data stability of the angle measured by the dual-frequency laser interferometer from the imaging start time to the imaging end time is verified, and the attitude control stability index of the control system is verified. The above method realizes index evaluation of the spacecraft control system through the external fine measurement mode, does not require spacecraft model development personnel to perform index evaluation, reduces the professional requirement, and can be quickly transplanted to other systems and has good universality. BRIEF DESCRIPTION OF DRAWINGS

[0028] The proportions and quantities of the components in the drawings are not necessarily consistent with those of the actual product.

[0029] Figure 1 is a control system index evaluation method based on an external fine measurement mode in an embodiment of the application;

[0030] Figure 2 is a control system index evaluation system environment arrangement schematic diagram in an embodiment of the application;

[0031] Figure 3 is a raw data interface schematic diagram collected by the first autocollimator in an embodiment of the application.

[0032] In the drawings:

[0033] 1: simulated star body;

[0034] 11: secondary active pointing control platform;

[0035] 12: tracking pointing mechanism;

[0036] 13: first autocollimator;

[0037] 2: first pointing target;

[0038] 21: first cube reference prism;

[0039] 3: second pointing target;

[0040] 31: second cube reference prism;

[0041] 4: theodolite;

[0042] 5: second autocollimator;

[0043] 6: circular multi-scale indexing table;

[0044] 7: third cube reference prism;

[0045] 8: fourth cube reference prism;

[0046] 9: dual-frequency laser interferometer. DETAILED DESCRIPTION

[0047] In order to make the objects, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work belong to the scope of protection of the present application.

[0048] The control system is an important part of a spacecraft, and has the characteristics of complex system, great design difficulty, and difficult ground verification. For a major engineering spacecraft, the system is more complex, and in the ground verification stage, the key performance of the control system is often verified through full physical simulation test. In the ground verification development process, the control system index evaluation is usually achieved by collecting control system telemetry parameters and calculating control system indexes. The control system data is complex, and if the data acquisition frequency requirement is high, the index result must be calculated directly by the control system controller. However, this method requires that the index evaluation personnel must be the spacecraft model development personnel, and has a deep understanding of the control system of this spacecraft to complete this work. The system index evaluation method has strong customization, and cannot be quickly transplanted to other systems, causing resource waste.

[0049] To solve the above problems, the patent proposes a control system index evaluation method based on external fine measurement from the perspective of engineering application. The system index of the control system such as attitude control maneuvering speed, attitude control accuracy and attitude control stability is evaluated by building an external fine measurement system. First, the control system attitude control maneuvering speed, attitude control accuracy and attitude control stability index evaluation system environment is built. The attitude control maneuvering speed evaluation process: the control system controls the simulated star body to maneuver, the angle between the two pointing targets is calibrated by the theodolite, the total time from the starting position to the target position is recorded, and the attitude control maneuvering speed index is verified. The attitude control accuracy evaluation process: the angle change between the initial pointing position and the return position after maneuvering of the load simulator is measured by the second autocollimator, the angle difference between the two stable positions before and after maneuvering is calculated, the system error and random error are combined, and the attitude control accuracy index of the control system is verified. The attitude control stability evaluation process: after the simulated star body points to the position, the control system enters the orientation mode to realize accurate and high-stable pointing to the target, and keeps the attitude stable for a long time to meet the imaging requirement time, and creates a good attitude environment for load imaging. The stable interference measurement light path is formed by the dual-frequency laser interferometer, and the angle real-time data stability of the dual-frequency laser interferometer is measured from the imaging start time to the imaging end time, and the attitude control stability index of the control system is verified. The attitude control stability index of the control system is verified, and the control system index evaluation is realized by the above method.

[0050] The application is further described below through the specific embodiments:

[0051] The instruments for realizing external fine measurement in the embodiment include the theodolite, the autocollimator and the dual-frequency laser interferometer, and the controlled object is a single-axis air floating table, a load simulator and a tracking pointing mechanism.

[0052] As shown in Figure 1 , the control system index evaluation method based on external fine measurement provided by the embodiment of the application realizes control system index evaluation through the following four steps.

[0053] S1, the control system index evaluation system environment is built based on the simulated star body, including: attitude control maneuvering speed test environment, attitude control accuracy test environment, attitude control stability test environment. Referring to Figure 2The simulation satellite 1 in the embodiment comprises a first attitude control platform (not shown in the figure) and a second active pointing control platform 11. The first attitude control platform provides coarse pointing control for the satellite, which is equivalent to the body of the satellite. In the embodiment, a single-axis air bearing table is used for simulation. The second active pointing control platform 11 performs fine pointing control on the basis of the coarse pointing control, and in the embodiment, a load simulator is used for simulation. The load simulator is installed on the single-axis air bearing table, can move synchronously with the single-axis air bearing table, and can rotate around the rotation center relative to the single-axis air bearing table. A tracking pointing mechanism 12 is installed on the load simulator, which can move synchronously with the load simulator and can rotate relative to the load simulator. Relative to the single-axis air bearing table and the load simulator, the tracking pointing mechanism 12 has a larger rotation angle, and can realize high-precision, high-stability, large-angle-range and high-agility control of the system. It should be noted that the single-axis air bearing table, the load simulator and the tracking pointing mechanism are all prior art, and will not be described here. It should also be noted that Figure 2 The marble platform in the above formula is a platform for bearing the simulation satellite and building the control system index evaluation system environment, and is not a limitation on the present application.

[0054] Attitude control maneuvering speed environment building:

[0055] Referring to Figure 2 A first autocollimator 13 is installed on the tracking pointing mechanism 12, and a first pointing target 2 is arranged at an initial pointing position of the simulation satellite 1. In the embodiment, the first pointing target is a parallel light pipe mirror, and a first cubic reference prism 21 is arranged on the parallel light pipe mirror. A second pointing target 3 is arranged at a specified pointing position of the simulation satellite, which is used to simulate a target star in the example, and a second cubic reference prism 31 is arranged on the second pointing target 3. When the pointing target angle is adjusted, the pointing target and the first autocollimator are autocollimated, and the pitch angle is consistent. When the angle is calculated, the pitch angle is not considered.

[0056] It should be noted that the dashed line at the theodolite 4 represents the position after the theodolite 4 is rotated by 90°, which is a position diagram of the theodolite 4 at different pointing positions. Similarly, the dashed line at the first autocollimator 13 represents the position after the first autocollimator 13 is rotated by 90°, which is a position diagram of the first autocollimator 13 at different pointing positions.

[0057] Taking α Maneuver_start = 0° as the initial pointing position (the position before maneuvering, which is the position of pointing to the first pointing target in the embodiment), and α Maneuver_end = 90° as the specified pointing position (the position after maneuvering, which is the position of pointing to the second pointing target in the embodiment), the specified pointing position α Maneuver_end is calibrated as follows:

[0058] a) The theodolite 4 is placed at the vertical intersection of the first pointing target 2 and the second pointing target 3 of the simulated star 1. The principle is shown in Figure 2 The first autocollimator 13 is collimated with the second pointing target 3, and the azimuth angle of the theodolite 4 is set to zero.

[0059] b) The control system controls the tracking pointing mechanism 12 to rotate clockwise by α Maneuver_end -α Maneuver_start , and the first pointing target is roughly pointed, so that the control system can identify the first pointing target.

[0060] c) The first cube reference prism 21 placed on the first pointing target 2 is aimed using the theodolite 4, and the azimuth angle α1 of the theodolite 4 at this time is recorded. The first cube reference prism 21 is adjusted based on the pointing of the first pointing target 2, so that the azimuth angle changes, for example, the angle increases by α, and the value of α can be arbitrarily set, for example, 5°, so that the theodolite 4 reading is (α Maneuver_end +α);

[0061] d) The control system controls the tracking pointing mechanism 12 to rotate counterclockwise by (α1-(α Maneuver_end +α)), and the second pointing target 3 is pointed, so that the control system can identify it.

[0062] e) The second pointing target 3 is finely pointed, so that the theodolite 4 reading is (α Maneuver_end +α), the first autocollimator 13 is collimated with the second cube reference prism 31, and the α Maneuver_end position calibration is completed.

[0063] The calibration method for the initial pointing position α Maneuver_start is basically the same as that for α Maneuver_end , which will not be described here.

[0064] Error analysis:

[0065] Prism angle processing error: 0.2", theodolite error (including human eye aiming error): 3", theodolite aiming twice can obtain the following formula:

[0066]

[0067] Rounding up, the angle calibration error of this method is 4.3".

[0068] Attitude control precision environment building:

[0069] Referring to Figure 2A second autocollimator 5 is arranged at the periphery of the simulated celestial body 1, and a third cube reference prism 7 is installed on the tracking pointing mechanism 12. In the initial position, the second autocollimator 5 is autocollimated with the third cube reference prism 7. The third cube reference prism 7 is aimed at the second autocollimator 5, and the third cube reference prism 7 is maneuvered with the load simulator. The angle change relationship between the stable positions before and after the maneuver of the load simulator is measured by the second autocollimator. In the embodiment, the second autocollimator 5 is installed on the circular multi-scale indexing table 6.

[0070] Error analysis:

[0071] Autocollimator error: 0.1″;

[0072] Therefore, the angle calibration error of the method is 0.1″.

[0073] Attitude control stability environment construction:

[0074] a) A fourth cube reference prism 8 is installed at the front end (the end far from the rotation center) of the load simulator. Since the farther the distance from the rotation center, the greater the position change, the fourth cube reference prism 8 is preferably installed at a position far from the rotation center. In the embodiment, the fourth cube reference prism 8 is installed at the front end of the load simulator, and the distance L from the rotation center of the load simulator is not less than 2000mm. A dual-frequency laser interferometer 9 is arranged at the periphery of the load simulator, and the dual-frequency laser interferometer 9 forms a stable interference measurement light path with the fourth cube reference prism 8.

[0075] b) During the rotation of the load simulator, the fourth cube reference prism 8 is aimed at the dual-frequency laser interferometer 9, and the position change of the fourth cube reference prism 8 is monitored.

[0076] Error analysis:

[0077] The test precision of the dual-frequency laser interferometer is 0.1μm, and the fourth cube reference prism 8 is calculated at a distance of 2000mm from the rotation center of the load simulator. The following formula is obtained:

[0078]

[0079] Therefore, the angle test error of the method is 0.01″.

[0080] S2, attitude control maneuvering speed evaluation

[0081] The attitude control maneuvering speed evaluation execution steps are described in detail in the following. Figure 2

[0082] a) The included angle between the initial pointing position and the specified pointing position is calibrated using the theodolite, so that the included angle is:

[0083] α Maneuver_end -α​Maneuver_start = 90°;

[0084] b) First, analog star pointing to α Maneuver_start = 0° position, to achieve the first pointing target alignment, start control system;

[0085] c) Control system control tracking pointing mechanism to yaw α Maneuver_end -α Maneuver_start = 90° rotation, to achieve the second pointing target alignment;

[0086] Attitude control maneuver speed evaluation data processing:

[0087] Calculate acceleration, uniform speed, deceleration and stable pointing time total T Maneuver_all .

[0088] The original data collected by the first autocollimator, see Figure 3 , where the horizontal coordinate is time, by Figure 3 It is known that the control system is yawed 90°, the starting time is 5.42s, and the ending time is 18.115s.

[0089] According to the attitude control maneuver speed evaluation method, the total time from the start to the end of the attitude maneuver is calculated, so the control system attitude maneuver 90° takes 12.695s.

[0090] S3, attitude control precision evaluation

[0091] Attitude control precision evaluation execution steps:

[0092] a) First, analog star pointing to yaw β accuracy_start = 0° position;

[0093] b) Start the control system and perform target pointing stability control;

[0094] c) After the system stabilizes, clear the initial pointing position and start the relative mode control to record the initial pointing position and start the maneuver;

[0095] d) The load simulator starts the maneuver from the initial pointing position, yawing clockwise β accuracy_end -β accuracy_start = 5°-0° = 5°, to the specified pointing position;

[0096] e) After the system stabilizes, maneuver back to the original position and stabilize the control;

[0097] f) After the system stabilizes, record the angle information of the second autocollimator, and the timing T accuracy_stat is 10 seconds, and the angle information after stabilization is recorded;

[0098] g) According to the second autocollimator measured load simulator before and after the maneuver to stabilize the position angle information, according to the following attitude control precision evaluation data processing part of the method to calculate the control system control precision.

[0099] The initial pointing position data is recorded in the following table, for brevity, only the first and last part of the data is shown, and the content of the table is a line of ellipsis, which represents the omission of part of the data in the middle:

[0100] Number Time [s] Absolute angle ["] X-axis angle ["] Y-axis angle ["] 0 0.000 0.857 -0.145 -0.845 1 0.049 0.908 -0.198 -0.886 2 0.099 0.447 -0.330 -0.302 3 0.154 0.172 0.052 -0.164 4 0.209 0.306 0.301 -0.057 5 0.253 0.601 0.432 -0.418 6 0.298 0.594 0.563 -0.191 7 0.358 0.605 0.294 -0.529 8 0.411 0.338 -0.035 -0.336 9 0.455 0.647 -0.233 -0.604 … … … … … 190 9.499 1.496 -1.386 -0.563 191 9.554 1.658 -1.440 -0.823 192 9.608 1.281 -1.206 -0.433 193 9.650 1.013 -0.845 -0.558 194 9.699 0.838 -0.771 -0.329 195 9.753 1.149 -1.017 -0.535 196 9.808 1.394 -1.338 -0.390 197 9.853 1.570 -1.481 -0.521 198 9.899 1.501 -1.404 -0.533 199 9.959 1.473 -1.350 -0.588

[0101] The data after the maneuver is recorded in the following table, for brevity, only the first and last part of the data is shown, and the content of the table is a line of ellipsis, which represents the omission of part of the data in the middle:

[0102]

[0103]

[0104] Attitude control precision evaluation data processing:

[0105] Control system attitude control precision evaluation method: control system before maneuvering to stabilize pointing and control system after maneuvering to stabilize pointing, the average value difference of the two second autocollimator measurement angles is the system error, 3σ accuracy is the random error, where σ accuracy is the standard deviation, and the system error plus the random error is the system control precision.

[0106]

[0107] where, is the average value of the first second autocollimator measurement angle, and the calculation can obtain -0.028″, is the average value of the second second autocollimator measurement angle, and the calculation can obtain -0.146″, ψ accuracy_ε is the average value difference of the two second autocollimator measurement angles, that is, the system error, N is the sampling point number, and in an exemplary embodiment, the sampling point number N is 200.

[0108] Standard deviation σ accuracy is calculated by the following formula:

[0109]

[0110] 3σ accuracy is the random error, and the calculation can obtain σ accuracy = 0.557″, therefore, the control system control precision is 3σ accuracy + ψ accuracy_ε0.146″-0.028″+3x0.557″=1.789″≈1.8″, so the control system control accuracy is 1.8″.

[0111] S4, attitude control stability evaluation

[0112] Attitude control stability evaluation execution steps:

[0113] a) using a dual-frequency laser interferometer and a fourth cubic reference prism to form a stable interference measurement light path;

[0114] b) the control system performs optical axis compensation movement through attitude control. It should be noted that the optical axis compensation movement through attitude control is prior art, which will not be described here;

[0115] c) the interference measurement light path starts to collect the shaking angle of the fourth cubic reference prism and records the data;

[0116] d) get T stability_stat is the optical axis stability (3σ) within 10 seconds, where σ is the standard deviation;

[0117] The data frequency of the dual-frequency laser interferometer is 1000HZ, and there are 10000 groups of data for recording 10s original data. The data at the beginning and end of the data recording are as follows: for the sake of brevity, only the first and last parts of the data are shown, and one row of ellipsis in the table represents that part of the data is omitted:

[0118]

[0119]

[0120] Attitude control stability evaluation data processing method:

[0121] Control system attitude control stability evaluation method: the real-time data stability (3σ) of the dual-frequency laser interferometer measured angle from the beginning of imaging to the end of imaging is the control system attitude control stability.

[0122]

[0123] n is the number of sampling points, in this embodiment, the value of n is 10000; θ i is the dual-frequency laser interferometer measured angle acquisition value, is the mean value of the dual-frequency laser interferometer measured angle acquisition value, σ stability is the standard deviation, so the control system attitude control stability is 3σ stability = 0.067578″.

[0124] In summary, the control system index evaluation method based on the external fine measurement mode provided by the application can realize the evaluation of the control system attitude control maneuvering speed, attitude control precision and attitude control stability system indexes by building an external fine measurement system. First, the control system attitude control maneuvering speed, attitude control precision and attitude control stability index evaluation system environment is built. The attitude control maneuvering speed evaluation process is as follows: the control system controls the simulated star body to maneuver, the included angle between the two pointing target angles is calibrated by the theodolite, the total time from the starting position to the target position is recorded, and the attitude control maneuvering speed index is verified. The attitude control precision evaluation process is as follows: the angle change of the initial pointing position of the load simulator and the return position after maneuvering is measured by the second autocollimator, the angle difference between the two stable positions before and after maneuvering is calculated, the system error and the random error are combined, and the attitude control precision index is verified. The attitude control precision index of the control system is verified. The attitude control stability evaluation process is as follows: after the simulated star body points to the position, the control system enters the orientation mode to realize the accurate and high-stability pointing to the target, the attitude is kept stable for a long time to meet the imaging requirement time, a good attitude environment is created for the load imaging, the stable interference measurement light path is formed by the dual-frequency laser interferometer and the fourth cubic reference prism, the dual-frequency laser interferometer measures the angle real-time data stability from the imaging start time to the imaging end time, and the attitude control stability index of the control system is verified. The above-mentioned method can realize the index evaluation of the control system of the spacecraft through the external fine measurement mode, and can solve the problems that the index evaluation personnel must be the spacecraft model development personnel and have a deep understanding of the spacecraft control system to complete the work, that is, the professional requirement is high, and that the system index evaluation method has a strong customization phenomenon and cannot be quickly transplanted to other systems, resulting in resource waste.

[0125] Compared with the prior art:

[0126] The application is aimed at the key index evaluation conditions of maneuvering speed, control precision and control stability, the method does not depend on the control system design details, the index evaluation can be completed through the external fine instrument measurement method and instrument data processing, and the threshold for the understanding of the system by the evaluation personnel is reduced. The method steps are clear, and the operation and implementation of the design personnel and the test personnel are facilitated. In addition, the method has been successfully applied in ground tests and tests, and the effectiveness and feasibility of the method are verified.

[0127] It should be noted that, in the embodiment, the pointing movement of the first autocollimator can be realized by adjusting one or more levels in the multi-level structure, for example, by directly adjusting the first attitude control platform and / or the second active pointing control platform, or by adjusting the first attitude control platform and / or the second active pointing control platform first and then adjusting the tracking pointing mechanism. In the embodiment, it is limited.

[0128] What is not described in detail in the present application is the common knowledge in the art or the prior art.

[0129] Finally, it should be noted that the above examples are only used to illustrate the technical solutions of the present application, and are not limiting; although the present application has been described in detail with reference to the foregoing examples, those skilled in the art should understand that not every example contains only one independent technical solution, and in the absence of solution conflicts, the various technical features mentioned in each example can be combined in any way to form other embodiments that can be understood by those skilled in the art.

[0130] In addition, modifications to the technical solutions described in the foregoing examples, or equivalent replacements of some technical features, do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A control system index evaluation method based on external fine measurement mode, characterized in that, The method comprises the following steps: S1, based on the simulated star body, a control system index evaluation system environment is established, including: attitude control maneuvering speed test environment, attitude control precision test environment, attitude control stability test environment, the simulated star body includes a first attitude control platform, a second active pointing control platform, a tracking pointing mechanism and a first autocollimator, the second active pointing control platform can move synchronously with the first attitude control platform, and can rotate around the rotation center relative to the first attitude control platform, the tracking pointing mechanism is installed on the second active pointing control platform and can move synchronously with the second active pointing control platform, and can rotate relative to the second active pointing control platform; S2, based on the attitude control maneuvering speed test environment, the attitude control maneuvering speed is evaluated: the control system performs maneuvering control on the tracking pointing mechanism, and moves from an initial pointing position to a specified pointing position, the angle between the initial pointing position and the specified pointing position is calibrated by a theodolite to be θ, the total time from the initial pointing position to the specified pointing position is recorded, and the attitude control maneuvering speed index is verified; S3, based on the attitude control precision test environment, the attitude control precision is evaluated: the first autocollimator is used to measure the angle change of the initial pointing position and the returned position after maneuvering of the second active pointing control platform, the angle difference between the two stable positions before and after maneuvering is calculated, the system error and the random error are combined, and the attitude control precision index is verified, wherein the random error is three times the standard deviation; S4, based on the attitude control stability test environment, the attitude control stability is evaluated: after the control system enters the directional mode, the position change of the second active pointing control platform is monitored in real time by a laser interferometer, the optical axis stability in the imaging period is calculated, and the attitude control stability index is verified, wherein the optical axis stability is three times the standard deviation.

2. The method of claim 1, wherein: The attitude control maneuvering speed test environment in step S1 is established by the following method: A first autocollimator is installed on the tracking pointing mechanism, a first pointing target is arranged at the starting position of maneuvering, a second pointing target is arranged at the end position of maneuvering, and a theodolite is arranged at the vertical intersection of the first pointing target and the second pointing target on the outer periphery of the simulated star body; The first pointing target is provided with a first cubic reference prism, the second pointing target is provided with a second cubic reference prism, and the first autocollimator is autocollimated with the first pointing target and the second pointing target when pointing to the first pointing target and the second pointing target.

3. The method of claim 1, wherein: The attitude control precision test environment in step S1 is established by the following method: A second autocollimator is arranged on the outer periphery of the simulated star body, and a third cubic reference prism is installed on the tracking pointing mechanism, and the second autocollimator is autocollimated with the third cubic reference prism at the initial position.

4. The method of claim 1, wherein: The attitude control stability test environment in step S1 is established by the following method: A fourth cubic reference prism is installed at the front end of the second active pointing control platform, and the distance from the rotation center of the load simulator is not less than 2000mm. A dual-frequency laser interferometer is arranged at the periphery of the simulated star body, and forms a stable interferometric measurement light path with the fourth cubic reference prism.

5. The method of claim 1, wherein: The calculation formula of the attitude control precision index in step S3 is: wherein is the mean value of the angles measured by the motorized autocollimator before the movement, is the mean value of the angles measured by the motorized autocollimator after the movement, σ is the standard deviation of the random error, and δ is the systematic error.

6. The method of claim 1, wherein: The calculation formula of the optical axis stability in step S4 is: Wherein, θ i is the angle value collected by the laser interferometer in real time, is the average value of the angle collected by the laser interferometer, and n is the number of sampling points.

7. The method of claim 2, wherein: The angle calibration error of the attitude control maneuvering speed test environment is not more than 4.3″.

8. The method of claim 3, wherein: The autocollimator measurement error of the attitude control precision test environment is not more than 0.1″.

9. The method of claim 4, wherein: The angle test error of the attitude control stability test environment is not more than 0.01″.

10. The method of claim 1, wherein: The external fine measurement instruments include the theodolite, the first autocollimator, the second autocollimator and the dual-frequency laser interferometer, the controlled object is a first-level attitude control platform, a second-level active pointing control platform and a tracking designation mechanism, wherein the first-level attitude control platform is a single-axis air floating table, and the second-level active pointing control platform is a load simulator.

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