Pseudo-closed-loop attitude control method for thin-film spacecraft

Abstract

The invention provides a pseudo-closed-loop attitude control method for a thin film spacecraft. The pseudo-closed-loop attitude control method comprises a plurality of open-loop control stages and a measurement stage, in the open-loop control stage, open-loop control is carried out by accelerating and decelerating at a constant speed; attitude control is not carried out in a measurement stage, an attitude error and an angular velocity error are evaluated through a least square method, the error is eliminated through a next open-loop control stage, and the control precision is improved through continuous iterative optimization; direct adoption of closed-loop control on a time domain is avoided, and thin film attitude jitter can be avoided; three-section control is set in an open-loop control stage, the thin film structure is restored to an undeformed shape under the non-thrust condition in a constant-speed stage, and structural deformation is further reduced; the attitude motion state of the thin-film spacecraft is accurately estimated in the measurement stage, the problem that a single-point sensor is greatly influenced by structural deformation is avoided, and the problem that sensor data change drastically along with time due to thin-film vibration is also avoided.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of space device control, and particularly relates to a pseudo-closed-loop attitude control method for a thin-film spacecraft. BACKGROUND

[0002] In recent years, with the substantial growth of spacecraft launch missions, the number of space debris is also rapidly increasing. In order to maintain the safety of the space environment and promote the sustainable use of space orbital resources, it is urgent to carry out space debris removal work. At present, various active space debris removal methods have been proposed at home and abroad, such as thin-film spacecraft. The thin-film spacecraft is an ultra-light and ultra-flexible two-dimensional spacecraft, which integrates thin-film solar cells, electric thrusters, communication, navigation, shape control and other complete functions. Due to the ultra-light mass and ultra-strong orbital maneuvering capability of the thin-film spacecraft, it provides a new idea for low-cost removal of small space debris. However, the thin-film spacecraft belongs to a new type of two-dimensional spacecraft. In the process of orbit and attitude control, the thrust will cause a large structural deformation of the thin-film, and the structural deformation will further affect the position and direction of the thruster, thereby affecting the orbit control force and the attitude control moment, so that the thin-film spacecraft appears a unique orbit-attitude-structure coupling effect. The orbit-attitude-structure coupling effect will bring great challenges to the control system design of the thin-film spacecraft.

[0003] The orbit, attitude and structure control problem of flexible spacecraft during on-orbit operation has always been a hot issue in the field of aerospace. The flexible vibration often has a significant impact on the stability and performance of the control system. Past researches are all aimed at the control problem of spacecraft containing flexible structures. Among them, the solar sail is widely studied for various space missions due to its advantages of not needing separate propellants and being able to obtain stable momentum through attitude control of the blades, movable mass and reflectivity control board. In active attitude control, the solar sail is usually configured with special mechanical devices, such as control blades, sliding mass and gimbal control arms, as actuators.

[0004] For spacecraft with flexible structures, the key to high-precision attitude control is the coupling disturbance caused by flexible accessory vibration. Since the thin-film spacecraft has no supporting structure, the thruster can only be directly configured on the flexible structure, and since the thruster placed at the corner point changes the attitude without passing through the center of mass, it will cause significant structural deformation. In addition, after the thin-film deforms, the attitude sensor data at different positions will change over time, and how to use the attitude sensor data of multiple thin-film vibrations for attitude control also faces great difficulties, and the existing technology does not have a good solution. SUMMARY

[0005] In view of the deficiencies of the prior art, the present application provides a pseudo-closed-loop attitude control method for a thin-film spacecraft to solve the problems in the background art.

[0006] The technical scheme of the present application is a pseudo-closed loop attitude control method for a thin film spacecraft, wherein the pseudo-closed loop comprises multiple open loop control stages and measurement stages; the method comprises the following steps: S1), designing multiple open loop control stages, and setting corresponding measurement stages not containing attitude control after each open loop control stage; S2), first performing three-stage control in the control stage, wherein the control parameters of the first open loop control stage are obtained from the integrated attitude angular velocity and attitude angle, then performing the measurement stage, and estimating the attitude angle and attitude angular velocity of the measurement stage by using the least square method; ; S3), updating the control parameters of the next open loop control stage based on the attitude angle and attitude angular velocity of the last measurement stage to perform the next open loop control stage control, and then entering the next measurement stage to estimate the attitude angle and attitude angular velocity; S4), continuously updating the open loop control stage and the measurement stage to form an iterative cycle of "open loop control, non-control measurement, open loop control, non-control measurement, …" until the attitude error and attitude angular velocity error reach the error limit given by the user.

[0007] Preferably, in step S2), the open loop control stage comprises an acceleration segment, a uniform rotation segment and a deceleration segment, wherein the thrust of the electric thruster in the uniform rotation segment is 0.

[0008] Preferably, in step S2), in the open loop control stage, the attitude control moment of the thin film spacecraft is: In the formula, is the control moment amplitude of the acceleration segment; is the control moment amplitude of the deceleration segment; , , are the times of the acceleration segment, the uniform rotation segment and the deceleration segment, respectively; is the total time of one control period, wherein, = + is the total time of the open loop control stage, is the time of the measurement stage; is the time parameter in each control period, and is reset to 0 after each open loop control stage ends. to​​​​​​​

[0009] As preferred, in step S2), the attitude angular velocity and the attitude angle of the thin-film spacecraft in the acceleration segment of the first open-loop stage are respectively represented as: ; (2) ; (3) wherein, is the attitude angular velocity; is the attitude angle; is the rotational inertia of the controlled axis of the thin film.

[0010] When , the attitude angular velocity and the attitude angle at the end of the acceleration segment are respectively: ; (4) ; (5) For the first open-loop control stage, let , , the initial conditions of the attitude angle and the attitude angular velocity are both 0.

[0011] Therefore, for the first open-loop control stage, the time of the uniform rotation segment is: ; (6) wherein, is the rotation angle in the uniform rotation segment, is the total rotation angle.

[0012] As preferred, in step S2), in the measurement stage, the attitude angle and the angular velocity at the end of the measurement stage are estimated by using the least square method, i.e.: wherein, is the number of collected data points; represents the attitude angle corresponding to each collection point in the measurement stage; represents the time corresponding to each collection point in the measurement stage.

[0013] As preferred, in step S3), based on the attitude angle and the angular velocity of the last measurement stage, the control parameters of the next open-loop control stage are updated, i.e.: ; (9) wherein, is the rotation angle of the acceleration segment, the uniform rotation segment and the deceleration segment; For average angular velocity.

[0014] The beneficial effects of the present application are: 1、The present application uses a measurement stage without attitude control after each open-loop control stage, uses least square method to evaluate attitude error and angular velocity error, and then eliminates the errors according to the next open-loop control stage, thereby improving control precision by iteratively executing the open-loop control stage and the measurement stage; 2、The present application avoids directly using closed-loop control in time domain, thereby avoiding causing thin film attitude jitter; 3、The present application sets three-stage control through the open-loop control stage, and the thrust of the electric thruster is 0 in the uniform speed stage, so that the thin film structure is restored to the undeformed shape under no thrust, thereby further reducing structural deformation; 4、The present application accurately estimates the attitude motion state of the thin film spacecraft through the measurement stage, avoids the problem that a single-point sensor is greatly affected by structural deformation, and also avoids the problem that sensor data changes dramatically over time due to thin film vibration. BRIEF DESCRIPTION OF DRAWINGS

[0015] Figure 1 is a flowchart of the method of the present application; Figure 2 is a dynamic simulation result schematic diagram of the pseudo-closed-loop attitude control in embodiment 2 of the present application; Figure 3 is a schematic diagram of control forces of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 4 is a change schematic diagram of attitude angles of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 5 is a change schematic diagram of other attitude angles of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 6 is a schematic diagram of maximum structural deformation of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 7 is a change schematic diagram of attitude angles of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 8 is a change schematic diagram of attitude angular velocities of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 9 is a change schematic diagram of attitude angles of the pseudo-closed-loop control in embodiment 2 of the present application; Figure 10 is a schematic diagram of maximum deformation of the pseudo-closed-loop control in embodiment 2 of the present application. ​​​DETAILED DESCRIPTION

[0016] The specific embodiments of the present application will be further described with reference to the drawings as follows: Example 1 As shown in the figure, the embodiment provides a pseudo-closed loop attitude control method for a thin film spacecraft, the pseudo-closed loop includes multiple open loop control stages and measurement stages; the method includes the following steps: Figure 1 S1), design multiple open loop control stages, and set corresponding measurement stages without attitude control after each open loop control stage; S2), first perform three-stage control in the control stage, wherein the control parameters of the first open loop control stage are obtained from the integrated attitude angular velocity and attitude angle, then perform the measurement stage, and estimate the attitude angle and attitude angular velocity of the measurement stage by using the least square method; S3), update the control parameters of the next open loop control stage based on the attitude angle and attitude angular velocity of the last measurement stage to perform the next open loop control stage control, and then enter the next measurement stage to estimate the attitude angle and attitude angular velocity; S4), continuously update the open loop control stage and the measurement stage to form an iterative cycle of "open loop control, no control measurement, open loop control, no control measurement, …" until the attitude error and attitude angular velocity error reach the error limit given by the user. As preferred in the embodiment, in step S2), the open loop control stage includes an acceleration segment, a uniform rotation segment and a deceleration segment, wherein the thrust of the electric thruster in the uniform rotation segment is 0. As preferred in the embodiment, in step S4), in the open loop control stage, the attitude control moment of the thin film spacecraft is:

[0017] In the formula, Tacc is the control moment amplitude of the acceleration segment; Tdec is the control moment amplitude of the deceleration segment; T0 is the time of the acceleration segment; T1 is the time of the uniform rotation segment; T2 is the time of the deceleration segment; T is the total time of one control period, wherein T = T0 + T1 + T2.

[0018] In the formula, Tacc is the control moment amplitude of the acceleration segment; Tdec is the control moment amplitude of the deceleration segment; T0 is the time of the acceleration segment; T1 is the time of the uniform rotation segment; T2 is the time of the deceleration segment; T is the total time of one control period, wherein T = T0 + T1 + T2. , , ​​​​​​​​​​​​​​​the time for the measurement phase; the time parameter for each control cycle, the end of each open-loop control phase, to to is reset to 0.

[0019] In this embodiment, the attitude control moment of the open-loop control phase adopts a segmented trigonometric function form. This control moment has continuity, and can avoid the problem of large deformation of the membrane structure caused by sudden changes in the thrust of the thruster.

[0020] As preferred in this embodiment, in step S2), in the acceleration segment of the first open-loop control phase, the attitude angular velocity and attitude angle of the membrane spacecraft are respectively expressed as: ; (2) ; (3) In the formula, is the attitude angular velocity; is the attitude angle; is the rotational inertia of the controlled axis of the membrane.

[0021] When , the attitude angular velocity and attitude angle at the end of the acceleration segment are respectively: ; (4) ; (5) For the first open-loop control phase, let , , the initial condition attitude angle and attitude angular velocity are both 0, and the rotation angle in the acceleration segment is .

[0022] For the first open-loop control phase, the time for the uniform rotation segment is: ; (6) In the formula, is the rotation angle in the uniform rotation segment, is the total rotation angle.

[0023] As preferred in this embodiment, in step S2), in the measurement phase, the attitude angle and angular velocity at the end of the measurement phase are estimated by using the least square method, i.e.: In the formula, is the number of data points collected; θm represents the attitude angle corresponding to each collection point in the measurement phase; θm represents the time corresponding to each collection point in the measurement phase.

[0024] This method can accurately estimate the attitude motion state of the Brane Craft, avoid the problem that the single-point sensor is greatly affected by the structural deformation, and avoid the problem that the sensor data changes dramatically over time due to the vibration of the Brane Craft.

[0025] As preferred in this embodiment, in step S3), the control parameters of the next open-loop control phase are updated based on the attitude angle and the angular velocity of the last measurement phase, that is: ; (9) In the formula, is the rotation angle of the acceleration segment, the constant-velocity rotation segment and the deceleration segment; is the average angular velocity.

[0026] wherein the rotation angle is: ; (11) ; (12) In the formula, is the attitude angle error of the last measurement phase; is the estimated attitude angle of the last measurement phase.

[0027] The average angular velocity is represented as: ; (13) wherein and the angular velocity in the constant-velocity rotation segment is set to ;

[0028] In this embodiment, in order to avoid large deformation of the Brane Craft during attitude control, a smaller control torque and a longer control time are used, and according to the angle and time continuous conditions of the open-loop control phase, the time and the attitude angular velocity in the constant-velocity rotation segment are determined by the parameters and , without the need for additional settings, so that the only independent parameters to be determined are and . For the selection of the parameters and , the finite element analysis is selected to make the deformation of the Brane Craft not more than 25% of the side length, and is selected to be more than 3 times the maximum vibration period of the structure. In addition, this embodiment can also select and All other control parameters are calculated as free variables.

[0029] Example 2

[0030] This example simulates the pseudo-closed loop control of Example 1, as well as open loop control only, and closed loop control.

[0031] In this example, the measurement time is set to about 30 seconds, and the simulation time is 150 seconds in the numerical simulation of the pseudo-closed loop control. The simulation results are shown in Figure 2 Figure 2 The membrane structure from 0 to 100 seconds is shown. After 100 seconds, the attitude angle stabilizes at 90 degrees.

[0032] It can be seen that there are two complete control periods within the 150 seconds. The maximum control force in the first control phase is Figure 3 , while the maximum control force in the second phase is . The attitude angle is shown in . It can be seen that the attitude angle is about 85.9 degrees at the end of the first control phase, and the angular velocity is Figure 4 . After that, at the end of the measurement phase, its value drops to 82.5 degrees. Then, the second control phase is carried out. At the end of the second control phase, the attitude angle reaches 90.02 degrees, and the angular velocity is . The final attitude angular velocity of the proposed pseudo-closed loop control method is two orders of magnitude lower than that of the open loop control method.

[0033] In Figure 5 , the changes of two other attitude angles are shown. The attitude angle changes in the range of about -0.017 degrees to 0.056 degrees, while the attitude angle changes in the range of 0 degrees to 0.35 degrees. These ranges are comparable to open loop control, but significantly smaller than closed loop control.

[0034] Figure 6 The maximum structural deformation during the control phase is shown. In the first control phase, the maximum deformation is 0.144 meters. In the subsequent control phase, the deformation of the membrane is less than 0.01 meters. In fact, the deformation of the structure can be further reduced by shortening the acceleration segment time. In summary, the proposed pseudo-closed loop control scheme is superior to open loop control and closed loop control in terms of accuracy of attitude angle and angular velocity. Through repeated correction of attitude error in multiple open loop control phases, this super flexible membrane spacecraft can ultimately achieve high-precision attitude control through this pseudo-closed loop attitude control method.

[0035] ​​The effect of the membrane flexibility on the attitude control results is investigated by using three attitude control algorithms, open-loop, closed-loop, and pseudo-closed-loop. The membrane flexibility can be adjusted by different parameters, such as the side length, thickness, or Young's modulus. However, changing the side length or thickness of the membrane will also change its moment of inertia and control performance. To avoid changing the mass properties, the membrane flexibility is changed by selecting different Young's moduli. The Young's modulus affects the bending, torsional, and tensile stiffness of the membrane and, ultimately, the deformation during the attitude control process. In the numerical simulations of this section, the Young's modulus is specified within a certain range, i.e., The attitude control simulation results are shown in Figure 7- Figure 10 The final attitude angles and angular velocities of the three controllers for different control strategies are shown in Figure 7 and 8 , respectively, which are measured at To further investigate its fluctuation characteristics, the fluctuation amplitudes of the closed-loop controller attitude angles are shown in Figure 9 In addition, Figure 10 shows the maximum deformation that occurs during the attitude control process.

[0036] For the open-loop controller, one of the significant features is that the final errors in the attitude angles and angular velocities are unavoidable and cannot be corrected. When the Young's modulus is small ( ), the final angular velocity reaches 4 deg / s, which means that the membrane spacecraft is actually rotating rapidly around the axis. Although the final error in the attitude angle is small, it will increase sharply afterwards. The reason for the large error in the angular velocity is that the deformation during the attitude control process is too large, and the direction of the thrusters is greatly affected. For example, at , the maximum deformation under the open-loop control is 40% of its size. For the case of , the final angular velocity is less than 0.25 deg / s. As the Young's modulus increases, the final attitude error, final angular velocity error, and maximum structural deformation of the membrane will decrease slowly.

[0037] In the closed-loop control mode, the error in the angular velocity is about 4 deg / s when In addition, the error in the attitude angle is greater than 3.5 deg when In fact, after , the attitude angle will fluctuate when using the closed-loop controller. In this case, an important feature that describes its control performance is the fluctuation amplitude, as shown in Figure 9 It can be seen that the fluctuation amplitude decreases from 33.59 deg at to 0.65 deg at The maximum deformation of the closed-loop control is similar to that of the open-loop controller.

[0038] By employing a pseudo-closed-loop attitude controller, the final attitude and angular velocity errors were significantly reduced within the selected Young's modulus range. Even when the Young's modulus was... The final attitude angle can still reach 89.77 degrees. When At this time, the attitude angular velocity remained below 0.019 degrees / second, which is approximately two orders of magnitude smaller than the other two controllers. Regarding structural deformation, the maximum deformation of the pseudo-closed-loop controller was also significantly smaller than that of the other controllers, because the acceleration time of the pseudo-closed-loop controller is shorter. In summary, the proposed pseudo-closed-loop controller enables high-precision attitude control of thin-film spacecraft with varying structural flexibility, exhibiting excellent control performance even under conditions of considerable deformation.

[0039] The embodiments and descriptions above are merely illustrative of the principles and preferred embodiments of the present invention. Various changes and modifications may be made to the present invention without departing from its spirit and scope, and all such changes and modifications fall within the scope of the present invention as claimed.

Claims

1. A method for pseudo-closed loop attitude control of a thin-film spacecraft, the method comprising: The pseudo-closed loop comprises a plurality of open-loop control stages and measurement stages; the open-loop control stages comprise the following steps: ​ S1), designing a plurality of open-loop control stages, and setting a corresponding measurement stage without attitude control after each open-loop control stage; S2), first, a three-stage control is performed in the control phase, wherein the control parameters of the first open-loop control phase are obtained from the integrated attitude angular velocity and attitude angle; then, a measurement phase is performed, and the attitude angle at the end of the measurement phase is estimated using the least square method and attitude angular velocity ; S3) the attitude angle based on the previous measurement phase and the attitude angular velocity updating the control parameters for the next open loop control phase for the next open loop control phase control, before entering the next measurement phase for the attitude angle and the attitude angular velocity estimation; S4), continuously updating the open-loop control stages and the measurement stages to form an iterative cycle of "open-loop control, non-control measurement, open-loop control, non-control measurement, …" until the attitude error and the attitude angular velocity error reach the error limit given by a user.

2. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 1, characterized in that: In step S2), the open-loop control stage comprises an acceleration segment, a uniform rotation segment, and a deceleration segment, wherein the electric thruster of the uniform rotation segment has a thrust of 0.

3. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 2, characterized in that: In step S2), during the open-loop control phase, the attitude control torque of the thin-film spacecraft... for: In the formula, The control torque amplitude during the acceleration phase; This refers to the control torque amplitude during the deceleration phase. , , These represent the times for the acceleration phase, the uniform rotation phase, and the deceleration phase, respectively. The total time of one control cycle. ,in, = + This represents the total time for the open-loop control phase. This refers to the timeframe for the measurement phase. The time parameter is defined for each control cycle, and the time parameter is defined at the end of each open-loop control phase. arrive Then reset to 0.

4. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 2, characterized in that: In step S2), in the acceleration segment of the first open-loop control stage, the attitude angular velocity and the attitude angle of the thin-film spacecraft are respectively represented as: ； （2） ； （3） wherein is the attitude angular velocity; is the attitude angle; is the rotational inertia of the film controlled axis; And when the attitude angular velocity and the attitude angle at the end of the acceleration section are respectively ； （4） ； （5） For the first open-loop control phase, let , , the initial conditions are 0 for both the attitude angle and the attitude angular velocity.

5. The method of claim 4, wherein: In step S2), for the first open-loop control stage, the time of the uniform rotation segment is: In step S2), for the first open-loop control stage, the time of the uniform rotation segment is: ； （6） In the formula, is the rotation angle of the uniform rotation section, is the total rotation angle.

6. The method of claim 1, wherein: In step S2), in the measurement phase, the attitude angles are estimated at the end of the measurement phase using a least squares method and the angular velocity i.e.: In the formula, is the number of data points collected; represents the attitude angle corresponding to each collection point in the measurement stage; represents the time corresponding to each collection point in the measurement stage.

7. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 6, characterized in that: In step S3) the control parameters for the next open loop control phase are updated based on the attitude angles and angular velocities of the last measurement phase, i.e. ； （9） In the formula, is the rotation angle of the acceleration section, the constant speed rotation section, and the deceleration section; is the average angular velocity.

8. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 7, characterized in that: In step S3, the rotation angle of the acceleration section, the constant speed rotation section, and the deceleration section is set to be: is: ； （11） ； （12） wherein is the attitude angle error of the previous measurement phase; is the estimated attitude angle of the previous measurement phase.

9. The pseudo-closed loop attitude control method of a membrane spacecraft according to claim 8, characterized in that: In step S3), the average angular velocity is expressed as: 。 10. The method of claim 3, wherein: According to the angle and time continuous conditions of the open-loop control phase, the time and attitude angular velocity of the uniform rotation segment are determined by parameters and Without additional settings, therefore, the only independent pending parameters are and For the selection of parameters and , through finite element analysis, the deformation generated by the Brane Craft does not exceed 25% of the edge length, and the selection is more than 3 times the maximum vibration period of the structure.

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