A stereoscopic imaging satellite attitude control method based on feedforward model prediction

Abstract

The present invention discloses a stereoscopic imaging satellite attitude control method based on feedforward model prediction, which is characterized by comprising the following steps: step 1, performing attitude trajectory planning according to the operation information of the stereoscopic imaging satellite, and judging whether the maneuverability meets the requirements through the angular momentum of the stereoscopic imaging satellite, wherein the operation information of the stereoscopic imaging satellite includes the maneuvering time planned by the imaging task, the attitude, angular velocity and angular acceleration of each imaging strip at the initial imaging moment and the terminal imaging moment; step 2, constructing a controller based on the feedforward model prediction, and calculating the three-axis command torque; step 3, calculating the frame angle command angular velocity of the control torque gyro group according to the three-axis command torque obtained in the previous step. The method solves the technical problems of long attitude stabilization time and poor attitude stability in the satellite rapid attitude maneuvering process and the subsequent stable imaging process.

Description

Technical Field

[0001] The invention relates to the technical field of remote sensing satellite attitude control, and in particular to a stereoscopic imaging satellite attitude control method based on feedforward model prediction. Background Art

[0002] Stereo imaging of remote sensing satellites can provide three-dimensional earth observation data, which helps to more accurately measure terrain height, depth and other information. It plays a very important role in surveying and mapping, urban planning, environmental monitoring and other fields. Agile satellites using control torque gyros can quickly adjust their attitude. However, due to the large torque noise and obvious vibration of large actuators, the transition time from attitude maneuvering to stable imaging using traditional proportional-integral-differential control (PID) is long and the stabilization speed is slow, which affects the satellite performance. Summary of the invention

[0003] In view of the deficiencies in the prior art, the present invention proposes a stereoscopic imaging satellite attitude control method based on feedforward model prediction, which solves the technical problem of long attitude stabilization time during the satellite's rapid attitude maneuvering process and the subsequent stable imaging process. This method greatly improves the stereoscopic imaging process, and the entire satellite can maneuver and stabilize quickly, thereby extending the effective imaging working time and obtaining longer stereoscopic imaging strips.

[0004] In order to solve the above technical problems, the technical solution of the present invention is:

[0005] A stereoscopic imaging satellite attitude control method based on feedforward model prediction comprises the following steps:

[0006] Step 1: Perform attitude trajectory planning according to the operation information of the stereo imaging satellite, and judge whether the maneuverability meets the requirements through the angular momentum of the stereo imaging satellite. The operation information of the stereo imaging satellite includes the maneuvering time planned for the imaging mission, the attitude, angular velocity and angular acceleration of each imaging strip at the initial and terminal imaging moments;

[0007] Step 2: Construct a predictive controller based on a feedforward model. The expression is as follows:

[0008]

[0009] M [k] It is the command torque of the stereo imaging satellite, used to output to the control torque gyro group; is the feedforward control quantity at the kth moment; Δu [k|k] is the model prediction control quantity at the kth moment; J is the inertia tensor of the whole star; ω=[ω x ω y ω z ] T is the three-axis angular velocity of the spacecraft;

[0010] Step 3: Calculate the frame angle command angular velocity of the control torque gyro group according to the three-axis command torque obtained in the previous step.

[0011] Preferably, the posture trajectory planning method:

[0012] Set trajectory planning coefficients a0, a1, a2, a3, a4, a5, a6, and a7. The trajectory planning coefficient expressions are as follows:

[0013]

[0014] a7=α0;

[0015] Among them, t m is the duration of attitude maneuver; a i are polynomial coefficients, where i = 0-7, α · Can be set as three-axis attitude angle The calculation formulas of θ and ψ are consistent. The subscript 0 indicates the starting attitude angle of the attitude maneuver, and the subscript m indicates the terminal attitude angle of the attitude maneuver; α0 is the three-axis attitude angle at the initial moment of the attitude maneuver, α m It is the three-axis attitude angle at the end of attitude maneuver.

[0016] As a preference, the target attitude α and angular velocity at time t are calculated based on the polynomial coefficients. Angular acceleration

[0017] α=a0t 7 +a1t 6 +a2t 5 +a3t 4 +a4t 3 +a5t 2 +a6t+a7;

[0018]

[0019] Where t is relative to the start time of the attitude maneuver, 0≤t≤tm, and α can be set as the attitude angle and roll angle of the three axes respectively. The calculation formulas for the pitch angle θ and the yaw angle ψ are the same.

[0020] Preferably, the method for determining whether the maneuverability of the stereoscopic imaging satellite meets the requirements is:

[0021] The three-axis attitude is iteratively calculated according to the set maneuvering process time, and the Euler angular velocity is calculated based on the three-axis attitude:

[0022]

[0023] Then calculate the angular momentum of the stereo imaging satellite:

[0024]

[0025] in is the value of the principal axis of inertia of the stereo imaging satellite;

[0026] The angular momentum during the maneuvering process is calculated to determine whether it exceeds the threshold set by the control torque gyro angular momentum envelope. If it exceeds the threshold, it reflects that the angular velocity maneuverability cannot meet the maneuvering time requirements and subsequent attitude maneuvers are stopped. If it does not exceed the threshold, subsequent attitude maneuvers can be performed.

[0027] Preferably, the feedforward model-based predictive controller includes a maneuvering process target posture determination module, a feedforward torque compensation calculation module, and a posture model predictive control module.

[0028] Preferably, the method for constructing the feedforward model predictive controller is:

[0029] Step 2.1: Establish the spacecraft attitude dynamics and kinematics model as follows:

[0030]

[0031] Where j is the inertia tensor of the entire star; ω = [ω x ω y ω z ] T is the three-axis angular velocity of the spacecraft; is the actuator angular momentum; The torque is generated by changing the angular momentum of the actuator; M c Represents the three-axis control torque of the satellite; q1, q2, q3 are the vector parts of the spacecraft attitude quaternion. Euler angles are often used in engineering A kinematic model describing the satellite's attitude, θ, ψ are the roll angle, pitch angle, and yaw angle respectively. When moving at a small angle, It can be considered that

[0032] Step 2.2: Define system state variables Generalized control quantity u=-ω×J ω +M c , the discretization model of the attitude system is established as follows:

[0033] x[k+1]=Ax[k]+Bu[k];

[0034] In the formula, k represents the current moment, k+1 represents the next moment;

[0035] Step 2.3: Use the maneuver process target attitude determination module to calculate the expected three-axis attitude angle α and angular velocity at each moment and angular acceleration

[0036] Step 2.4: Use the feedforward torque compensation calculation module to calculate the feedforward torque compensation and feedforward control quantity. ω r_last is the expected value of the angular velocity of the previous beat, Δt is the control period, m k is the adjustment factor;

[0037] Step 2.5: Use the attitude model prediction control module to establish the objective function:

[0038]

[0039] Where N p represents the prediction interval, S and Q are n×n semi-positive definite diagonal matrices, which are the terminal cost matrix and the running cost matrix respectively, x [k+j|k] Represents the state predicted at the k+j moment from the kth moment;

[0040] Step 2.6: Obtain the standard form of quadratic programming through transformation:

[0041]

[0042] In the formula, H=Γ T ΩΓ; in the formula, the superscript T represents the transpose of the matrix, the superscript ref represents the reference state trajectory, the subscript [k] represents the multi-step state quantity or control quantity at time k, the subscript [k|k] represents the state at time k, and the subscript [k+i|k] represents the state at time k+i predicted at time k.

[0043] Using the general method to obtain the optimal control sequence ΔU [k] ;

[0044] Step 2.7: Obtain a predictive controller based on a feedforward model through the above modules.

[0045] Preferably, when the stereoscopic imaging satellite moves at a small angle, in θ and ψ are the roll angle, pitch angle, and yaw angle, respectively.

[0046] Preferably, the conversion method of the objective function established by the posture model prediction control module is:

[0047] First, the discretized model is transformed into a multi-step form:

[0048] X [k] =Φx [k|k] +ΓU [k] ;

[0049]

[0050] The control quantity can be broken down into expressions:

[0051] The objective function is simplified as follows:

[0052]

[0053] In the formula,

[0054]

[0055] Except for the last two items in the above formula which are affected by the model prediction control quantity, the rest are not affected by the control quantity, so the performance index is simplified to:

[0056]

[0057] Finally, the objective function is transformed into the standard quadratic programming form.

[0058] Preferably, the control moment gyro group includes five control moment gyro group, which are installed in a pyramid configuration.

[0059] Preferably, the method for obtaining the three-axis command torque is:

[0060] The nominal angular momentum of each control moment gyro is h, and the combined angular momentum of the control moment gyro group is h CMGs :

[0061] h CMGs =h·(A cmg ·sinδ+B cmg ·cosδ)·I n×1 ;

[0062] in, is the CMGs frame angle vector matrix, δ i is the frame angle of the i-th control moment gyro; A cmg and B cmg is the matrix related to the installation position of CMGs, A cmg Each column is the angular momentum direction of each CMG when the frame angle is at 90°, B cmg Each column of is the direction of the angular momentum of the CMG when the frame angle of each CMG is at zero position. n×1 is an n×1 unit vector.

[0063] The synthetic torque M generated by the control torque gyro group is:

[0064]

[0065] Among them, A cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro when the frame angle of the control moment gyro is at 90°; B cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro rotating at high speed when the frame angle of the control moment gyro is at zero position; C is the Jacobi matrix of the torque output of the control moment gyro group;

[0066] The operating rate of the control moment gyro group is:

[0067]

[0068] in, is the frame angle command angular velocity; It is a singular robust inverse instruction operation law; It is the zero motion singularity avoidance operation law; is the nominal frame angle regression operation law;

[0069] Singular robust inverse instruction operation law:

[0070]

[0071] λ is the adjustment coefficient; according to the extreme value condition, we know that:

[0072]

[0073] The present invention has the following characteristics and beneficial effects:

[0074] By adopting the above technical scheme, the present invention introduces feedforward control information on the basis of model predictive control, which greatly improves the stereoscopic imaging process. The whole satellite can be quickly maneuvered and quickly stabilized. Compared with traditional PID control, the imaging working time is extended and longer stereoscopic imaging strips can be obtained. The method proposed in this article can achieve good attitude stability even in the active imaging process, and can be used for three-view stereoscopic imaging to obtain better stereoscopic imaging data than previous methods.

[0075] The present invention calculates whether the posture maneuverability meets the imaging requirements before executing the imaging posture maneuver, ensuring that the angular momentum is sufficient for imaging. There is no need to perform torque constraint calculation in the model predictive control algorithm, which reduces the calculation complexity and facilitates engineering implementation. BRIEF DESCRIPTION OF THE DRAWINGS

[0076] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative labor.

[0077] Figure 1 The figure is a flow chart of a stereoscopic imaging satellite attitude control method based on feedforward model prediction according to the present invention.

[0078] Figure 2 This is a block diagram of the attitude control of the stereoscopic imaging satellite of the present invention.

[0079] Figure 3 This is a curve showing the change of the three-axis attitude angle during satellite stereo imaging in an embodiment of the present invention.

[0080] Figure 4 1 is a curve showing the change of the three-axis attitude angular velocity during satellite stereo imaging in an embodiment of the present invention.

[0081] Figure 5 It is the MPC control attitude angle error curve in the embodiment of the present invention.

[0082] Figure 6 It is the PID control attitude angle error curve in the embodiment of the present invention.

[0083] Figure 7 It is a locally enlarged curve of the MPC control attitude angle error in an embodiment of the present invention.

[0084] Figure 8 It is a local enlarged graph curve of the PID control attitude angle error in an embodiment of the present invention.

[0085] Fig. 9 It is the MPC control attitude angular velocity error curve in the embodiment of the present invention.

[0086] Fig.10 It is the PID control attitude angular velocity error curve in the embodiment of the present invention.

[0087] Fig.11 It is a locally enlarged curve of the MPC control attitude angular velocity error in an embodiment of the present invention.

[0088] Fig.12 It is a locally enlarged curve of the PID control attitude angular velocity error in an embodiment of the present invention. DETAILED DESCRIPTION

[0089] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments may be combined with each other.

[0090] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first", "second", and the like are only used for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, features defined as "first", "second", and the like may explicitly or implicitly include one or more of the features. In the description of the present invention, unless otherwise specified, "multiple" means two or more.

[0091] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood by specific circumstances.

[0092] The present invention provides a stereoscopic imaging satellite attitude control method based on feedforward model prediction, such as Figure 1 As shown, the attitude trajectory is planned according to the maneuvering time planned for the imaging task and the attitude, angular velocity and angular acceleration information at the initial and terminal moments of imaging of each imaging strip, and it is determined whether the maneuvering capability meets the requirements.

[0093] Specifically, in this embodiment, the maneuvering time planned by the task and the attitude, angular velocity and angular acceleration information at the initial and terminal moments of imaging of each imaging strip are as follows:

[0094]

[0095]

[0096] The posture trajectory planning involved in this embodiment is a method based on seventh-order polynomial trajectory planning.

[0097] Trajectory planning coefficients a0, a1, a2, a3, a4, a5, a6, a7, the trajectory planning coefficient expressions are as follows:

[0098]

[0099] Among them, t m is the duration of attitude maneuver; a i are polynomial coefficients, where i = 0-7, α · Can be set as three-axis attitude angle The calculation formulas of θ and ψ are consistent. The subscript 0 indicates the starting attitude angle of the attitude maneuver, and the subscript m indicates the terminal attitude angle of the attitude maneuver; α0 is the three-axis attitude angle at the initial moment of the attitude maneuver, α m It is the three-axis attitude angle at the end of attitude maneuver.

[0100] According to the polynomial coefficients, calculate the target attitude α and angular velocity at time t Angular acceleration

[0101] α=a0t 7 +a1r 6 +a2t 5 +a3t 4 +a4t 3 +a5t 2 +a6t+a7;

[0102]

[0103] Where t is relative to the start time of the attitude maneuver, 0≤t≤tm, and α can be set as the attitude angle and roll angle of the three axes respectively. The calculation formulas for pitch angle θ and yaw angle ψ are the same. It can be understood that

[0104] The three-axis attitude (θ is the pitch angle, is the roll angle, ψ is the yaw angle) and the Euler angular velocity:

[0105]

[0106] Then calculate the angular momentum of the stereo imaging satellite:

[0107]

[0108] in is the value of the main axis of inertia of the whole satellite, and H is the value of the angular momentum of the satellite body during the maneuver. The angular momentum during the maneuver is calculated to determine whether it exceeds the threshold set by the control moment gyro angular momentum envelope. In this embodiment, the threshold is set to 10Nms. If it exceeds, it reflects that the angular velocity maneuverability cannot meet the requirements of the maneuvering time and the subsequent attitude maneuvering action is stopped. If it does not exceed, the subsequent attitude maneuvering can be carried out.

[0109] Step 2: Design a predictive controller based on a feedforward model, such as Figure 2As shown in the figure, it mainly includes the target attitude determination module during maneuvering, the feedforward torque compensation calculation module, and the attitude model prediction control module. The target attitude determination module during maneuvering gives the real-time target attitude angle, angular velocity and angular acceleration according to the imaging requirements; the feedforward torque compensation calculation module calculates the feedforward torque; the satellite attitude model prediction control module calculates the optimal control sequence based on the model prediction control principle, uses the first control quantity of the optimal control sequence to superimpose the feedforward torque, and then outputs the control torque to the satellite body through the control torque gyro group for attitude control; the attitude measurement and determination module calculates the satellite attitude and angular velocity in real time through the onboard gyro and star sensor.

[0110] Specifically,

[0111] 1) The spacecraft attitude dynamics and kinematics model is established as follows:

[0112]

[0113] It can also be written as:

[0114]

[0115] Where J is the inertia tensor of the entire star; ω = [ω x ω y ω z ] T is the three-axis angular velocity of the spacecraft; is the actuator angular momentum; The torque is generated by changing the angular momentum of the actuator. Here, the actuator is a group of control torque gyroscopes, which can be simplified as M c Represents the three-axis control torque of the satellite. Euler angle is often used in engineering A kinematic model describing the satellite's attitude, They are roll angle, pitch angle, and yaw angle respectively.

[0116] Understandably, when moving at small angles, It can be considered that

[0117] 2) Define system state variables Generalized control quantity u=-ω×Jω+M c , the discretization model of the attitude system is established as follows:

[0118] x[k+1]=Ax[k]+Bu[k] (7);

[0119] In the formula, , k represents the current moment, k+1 represents the next moment;

[0120] 3) The target attitude during the maneuvering process is determined based on the three-axis attitude angle α0 and angular velocity at the initial moment of the attitude maneuver and angular acceleration Target attitude angle α at the end of attitude maneuver m , angular velocity and angular acceleration The expected three-axis attitude angle α and angular velocity at each moment are calculated by formula (2) as well as the time deviation relative to the start of the attitude maneuver. and angular acceleration

[0121] 4) Calculation of feedforward torque compensation: feedforward control quantity ω r_last is the expected value of the angular velocity of the previous beat, Δt is the control period, m k To adjust the coefficient, it can be adjusted according to the dynamic characteristics.

[0122] 5) Design of model predictive control controller: To meet the high control quality in the process of spacecraft attitude control, the objective function is established:

[0123]

[0124]

[0125] Where N p represents the prediction interval, S and Q are n×n semi-positive definite diagonal matrices, which are the terminal cost matrix and the running cost matrix respectively, x [k+j|k] Represents the state predicted at the kth moment in time k+j.

[0126] Furthermore, the discretization model is converted into a multi-step form:

[0127] X [k] =Φx [k|k] +ΓU [k] ;

[0128]

[0129] Control quantity can be split into expressions:

[0130] In the formula, is the feedforward control quantity at the kth moment; Δu [k|k] is the model prediction control quantity at the kth moment. Since the control quantity is a small correction to the feedforward control quantity, the correction of Δu [k|k] The scope is constrained to simplify the calculation and facilitate the actual engineering calculation.

[0131] The objective function is organized as:

[0132]

[0133] In the formula,

[0134]

[0135] Except for the last two items in the above formula which are affected by the model prediction control quantity, the rest are not affected by the control quantity, so the performance index is simplified to:

[0136]

[0137] Convert the objective function into the standard quadratic programming form:

[0138]

[0139] In the formula, H=Γ T ΩΓ; in the formula, the superscript T represents the transpose of the matrix, the superscript ref represents the reference state trajectory, the subscript [k] represents the multi-step state quantity or control quantity at time k, the subscript [k|k] represents the state at time k, and the subscript [k+i|k] represents the state at time k+i predicted at time k.

[0140] The optimal control sequence ΔU can be obtained using a general method [k]

[0141] Based on the above content, the feedforward model predictive control law is:

[0142]

[0143] M [k] It is the command torque of the stereo imaging satellite, used to output to the control torque gyro group; is the feedforward control quantity at the kth moment; Δu [k|k] is the model prediction control quantity at the kth moment; J is the inertia tensor of the whole star; ω=[ω x ω y ω z ] T is the three-axis angular velocity of the spacecraft;

[0144] Step 3: Calculate the frame angle command angular velocity of the control moment gyro group according to the three-axis command torque of the stereoscopic imaging satellite body obtained in the previous step. In this embodiment, the control moment gyro group includes five control moment gyros (CMGs) installed in a pyramid configuration.

[0145] The control moment gyro group consists of five control moment gyros (CMGs) installed in a pyramid configuration. Each CMG has a nominal angular momentum of h, and the combined angular momentum of the CMGs is h CMGs :

[0146] hCMGs =h·(A cmg ·sinδ+B cmg ·cosδ)·I n×1 ;

[0147] in, is the CMGs frame angle vector matrix, δ i is the frame angle of the i-th control moment gyro; A cmg and B cmg is the matrix related to the installation position of CMGs, A cmg Each column is the angular momentum direction of each CMG when the frame angle is at 90°, B cmg Each column of is the direction of the angular momentum of the CMG when the frame angle of each CMG is at zero position. n×1 is an n×1 unit vector.

[0148] The synthetic torque M generated by the control torque gyro group is:

[0149]

[0150] Among them, A cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro when the frame angle of the control moment gyro is at 90°; B cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro rotating at high speed when the frame angle of the control moment gyro is at zero position; C is the Jacobi matrix of the torque output of the control moment gyro group;

[0151] The operating rate of the control moment gyro group is:

[0152]

[0153] in, is the frame angle command angular velocity; It is a singular robust inverse instruction operation law; It is the zero motion singularity avoidance operation law; is the nominal frame angle regression operation law;

[0154] It should be noted that there are multiple implementation methods for the zero motion singular avoidance operation law and the frame angle regression operation law, which are well-known contents. This embodiment mainly describes the singular robust inverse instruction operation law, which is as follows:

[0155] Singular robust inverse instruction operation law:

[0156]

[0157] λ is the adjustment coefficient; according to the extreme value condition, we know that:

[0158] According to the maneuvering time and the attitude, angular velocity and angular acceleration information of each imaging strip at the initial and terminal imaging moments according to the mission planning provided by this embodiment, and the stereoscopic imaging satellite attitude control method based on feedforward model prediction provided by this embodiment, the MPC control effect diagram in the embodiment of the present invention is obtained, and the traditional PID control effect diagram is given, such as Figure 3-Figure 12 As shown in the figure, by comparing the two control effects, we can see that:

[0159] 1) Figure 3 is a curve showing the change of the three-axis attitude angle during satellite stereoscopic imaging in an embodiment of the present invention, Figure 4 The three-axis attitude angular velocity change curve of the satellite during stereoscopic imaging in the embodiment of the present invention. The working mode equals 6 for the ground operation mode, the satellite flies at zero attitude; the working mode equals 5 for the attitude maneuvering mode, the satellite maneuvers its attitude to a specified attitude; the working mode equals 13 for the stereoscopic imaging mode, the satellite performs passive push-broom imaging at a specified attitude. Both MPC control and PID control can maneuver the attitude angle and angular velocity to a specified attitude.

[0160] 2) Figure 5 is the MPC control attitude angle error curve in the embodiment of the present invention, Figure 6 It is the PID control attitude angle error curve in the embodiment of the present invention. It can be seen that when the working mode is the stereo imaging mode, the attitude angle error of the MPC control is smaller than the attitude angle error of the PID control. Figure 7 It is a partial enlarged graph curve of the MPC control attitude angle error in the embodiment of the present invention, Figure 8 This is a partial enlarged curve of the PID control attitude angle error in the embodiment of the present invention. It can be seen that the attitude angle error is better than 0.0002° (3σ) during stereo imaging under MPC control, and the attitude angle error is better than 0.02° (3σ) during stereo imaging under PID control. Note that generally, imaging is performed 3s after entering the stereo imaging mode, so the statistical attitude angle error uses the data 3s after entering the stereo imaging mode.

[0161] 3) Fig. 9 is the MPC control attitude angular velocity error curve in the embodiment of the present invention, Fig.10 It is the PID control attitude angular velocity error curve in the embodiment of the present invention. It can be seen that when the working mode is the stereo imaging mode, the attitude angular velocity error of the MPC control is smaller than the attitude angular velocity error of the PID control. Fig.11 It is a partial enlarged graph curve of the MPC control attitude angular velocity error in an embodiment of the present invention, Fig.12This is a partial enlarged curve of the PID control attitude angular velocity error in the embodiment of the present invention. It can be seen that the attitude angular velocity error during stereo imaging under MPC control is better than 0.0003° / s (3σ), and the attitude angular velocity error during stereo imaging under PID control is better than 0.002° / s (3σ).

[0162] In summary, by using the stereoscopic imaging satellite attitude control method based on feedforward model prediction of the present invention, the whole satellite can maneuver quickly and stabilize quickly, and the pointing accuracy and attitude stability are better than traditional PID control, which greatly improves the attitude maneuvering process of stereoscopic imaging and gives full play to the satellite performance.

[0163] The embodiments of the present invention are described in detail above with reference to the accompanying drawings, but the present invention is not limited to the described embodiments. For those skilled in the art, various changes, modifications, substitutions and variations of these embodiments including components are made without departing from the principles and spirit of the present invention, and still fall within the scope of protection of the present invention.

Claims

1. A stereoscopic imaging satellite attitude control method based on feedforward model prediction, characterized in that: The steps include: Step 1: Perform attitude trajectory planning according to the operation information of the stereo imaging satellite, and judge whether the maneuverability meets the requirements through the angular momentum of the stereo imaging satellite. The operation information of the stereo imaging satellite includes the maneuvering time planned for the imaging mission, the attitude, angular velocity and angular acceleration of each imaging strip at the initial and terminal imaging moments; Step 2: Construct a predictive controller based on a feedforward model. The expression is as follows: M [k] It is the command torque of the stereo imaging satellite, used to output to the control torque gyro group; is the feedforward control quantity at the kth moment; Δu [k|k] is the model prediction control quantity at the kth moment; J is the inertia tensor of the whole star; ω=[ω x ω y ω z ] T is the three-axis angular velocity of the spacecraft; Step 3: Calculate the frame angle command angular velocity of the control torque gyro group according to the three-axis command torque obtained in the previous step. The control moment gyro group comprises five control moment gyros, which are installed in a pyramid configuration; The method for obtaining the three-axis command torque is as follows: The nominal angular momentum of each control moment gyro is h, and the combined angular momentum of the control moment gyro group is h CMGs : h CMGs =h·(A cmg ·sinδ+B cmg ·cosδ)·I n×1 ; in, is the CMGs frame angle vector matrix, δ i A is the frame angle of the i-th control moment gyro, i = 1 ~ n; cmg and B cmg is the matrix related to the installation position of CMGs, A cmg Each column is the angular momentum direction of each CMG when the frame angle is at 90°, B cmg Each column is the angular momentum direction of each CMG when the frame angle is at zero. n×1 is an n×1 unit vector; The synthetic torque M generated by the control torque gyro group is: Among them, A cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro when the frame angle of the control moment gyro is at 90°; B cmg Each column of the matrix is ​​the angular momentum direction of the control moment gyro rotating at high speed when the frame angle of the control moment gyro is at zero position; C is the Jacobi matrix of the torque output of the control moment gyro group; The operating rate of the control moment gyro group is: in, is the frame angle command angular velocity; It is a singular robust inverse instruction operation law; It is the zero motion singularity avoidance operation law; is the nominal frame angle regression operation law; Singular robust inverse instruction operation law: λ is the adjustment coefficient; according to the extreme value condition, we know that:

2. The stereoscopic imaging satellite attitude control method based on feedforward model prediction according to claim 1 is characterized in that: The posture trajectory planning method: Set trajectory planning coefficients a0, a1, a2, a3, a4, a5, a6, and a7. The trajectory planning coefficient expressions are as follows: a7＝α0； Among them, t m is the duration of the attitude maneuver; ai is the polynomial coefficient, where i = 0-7, α j Set as the three-axis attitude angle θ, ψ, the calculation formula is the same, where j = 0, m, the subscript 0 represents the starting attitude angle of the attitude maneuver, and the subscript m represents the terminal attitude angle of the attitude maneuver; α0 is the three-axis attitude angle at the initial moment of the attitude maneuver, α m It is the three-axis attitude angle at the end of attitude maneuver.

3. The stereoscopic imaging satellite attitude control method based on feedforward model prediction according to claim 2 is characterized in that: According to the polynomial coefficients, calculate the target attitude α and angular velocity at time t Angular acceleration α=a0t 7 +a1t 6 +a2t 5 +a3t 4 +a4t 3 +a5t 2 +a6t+a7; Where t is relative to the start time of the attitude maneuver, 0≤t≤tm, and α is the attitude angle and roll angle of the three axes respectively. The calculation formulas for the pitch angle θ and the yaw angle ψ are the same.

4. The stereoscopic imaging satellite attitude control method based on feedforward model prediction according to claim 3 is characterized in that: The method for judging whether the maneuverability of the stereoscopic imaging satellite meets the requirements is as follows: The three-axis attitudes are calculated according to the set iteration time of the maneuvering process, and the Euler angular velocity is calculated based on the three-axis attitudes: Then calculate the angular momentum of the stereo imaging satellite: in is the value of the principal axis of inertia of the stereo imaging satellite; The angular momentum during the maneuvering process is calculated to determine whether it exceeds the threshold set by the control torque gyro angular momentum envelope. If it exceeds the threshold, it reflects that the angular velocity maneuverability cannot meet the maneuvering time requirements and the subsequent attitude maneuvering action is stopped. If it does not exceed the threshold, the subsequent attitude maneuver is performed.

5. The stereoscopic imaging satellite attitude control method based on feedforward model prediction according to claim 1, characterized in that: The feedforward model-based predictive controller includes a maneuvering process target posture determination module, a feedforward torque compensation amount calculation module, and a posture model predictive control module.

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