Low-orbit remote sensing satellite three-axis fast maneuvering attitude trajectory planning and tracking method

Abstract

The invention discloses a three-axis fast maneuvering attitude trajectory planning and tracking method for a low-orbit remote sensing satellite. Relates to the technical field of aerospace, in particular to the technical field of low-orbit remote sensing satellite three-axis fast maneuvering attitude trajectory planning and tracking. The balance problem of maneuverability and stability in the attitude transformation process is effectively solved, and the superiority of the anti-interference capability is improved. The method comprises the following steps: establishing a satellite kinematics model and a satellite dynamics model; calculating the quaternion of the expected coordinate system relative to the initial coordinate system of the satellite at the starting moment of movement; attitude trajectory planning is completed by improving a tracking differentiator; calculating a total control rate; a real control rate is obtained through an execution mechanism; calculating the quaternion of the satellite body coordinate system relative to the inertial seat system and the angular velocity of the satellite body coordinate system relative to the inertial seat system; a real body quaternion and a real body angular velocity are obtained through a measuring mechanism; and circularly planning the trajectory and tracking.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to the field of three-axis rapid maneuvering attitude trajectory planning and tracking technology for low-orbit remote sensing satellites. Background Art

[0002] With the continuous development of aerospace technology, low-orbit optical remote sensing satellites have been widely used in resource management, earth observation, environmental monitoring and other fields. In multi-point imaging tasks, satellites can obtain images in different directions through multiple maneuvers. The higher the maneuvering efficiency in the same time, the wider the imaging range. In the current attitude maneuvering research, there is a difficult problem of balancing maneuverability and stability. The pursuit of fast response by the controller will sacrifice stability to a certain extent, resulting in reduced image quality; while emphasizing stability may lead to slower maneuvering convergence speed, affecting the range and efficiency of imaging. At the planning level, the existing optimization methods have the problem of high computing power, and the analytical method does not consider the stability of the actuator. Some scholars use the tracking differentiator method for path planning, which considers acceleration smoothing while improving maneuverability, but only considers single-axis maneuvering and does not extend to three axes; at the control level, the industry mostly uses PD control. Although this method is simple and feasible, it has poor anti-interference ability and does not consider actuator saturation. Although other methods consider actuator and environmental interference, they have problems such as difficulty in adjusting parameters or high computing power, which is difficult to implement in engineering. Summary of the invention

[0003] In view of the above problems, the present invention discloses a three-axis rapid maneuvering attitude trajectory planning and tracking method for a low-orbit remote sensing satellite, which relates to the field of aerospace technology, effectively solves the balance problem between maneuverability and stability during the three-axis attitude transformation of the satellite, and improves the superiority of anti-interference capability.

[0004] The method comprises the following steps:

[0005] S1. Establish satellite kinematic model and satellite dynamic model;

[0006] S2, calculate the quaternion of the desired coordinate system relative to the satellite initial coordinate system at the start of the movement;

[0007] S3, completing attitude trajectory planning by improving the tracking differentiator;

[0008] S4. Calculate the total control rate;

[0009] S5. Get the real control rate through the actuator;

[0010] S6, calculating the quaternion of the satellite body coordinate system relative to the inertial coordinate system and the angular velocity of the satellite body coordinate system relative to the inertial coordinate system;

[0011] S7, obtaining the real body quaternion and the real body angular velocity through the measuring mechanism;

[0012] S8, loop planning trajectory and tracking.

[0013] Furthermore, it is characterized in that the kinematic equation of the kinematic model is specifically: Said represents the derivative of the quaternion of the satellite body coordinate system relative to the inertial coordinate system, q0 represents the real part of Q, Q represents the quaternion of the satellite body coordinate system relative to the inertial coordinate system, q represents the imaginary part of Q, represents the derivative of q0, S(q) is the antisymmetric matrix of q, I3 is the third-order unit matrix, ω represents the angular velocity of the satellite body coordinate system relative to the inertial coordinate system, and T represents the controller step size; the dynamic equation of the dynamic model is specifically: Said represents the derivative of ω, represents the positive definite matrix of the satellite's moment of inertia, 3 is a three-dimensional space, S(ω) represents the antisymmetric matrix of ω, and h w ∈ 3 is the total angular momentum of the reaction flywheel, and u′ represents the true control rate.

[0014] Furthermore, the calculation formula of the quaternion of the desired coordinate system relative to the satellite initial coordinate system at the start of the maneuver is: q OG is the quaternion of the desired coordinate system relative to the orbital coordinate system, q iO0 is the quaternion of the initial coordinate system relative to the satellite orbit coordinate system at the start of the maneuver, Represents a tensor product operation.

[0015] Furthermore, the posture trajectory planning is specifically as follows: using the rotation axis e iG0 Get the quaternion q of the satellite's desired coordinate system relative to the initial coordinate system G0 , angular velocity ω G0 and acceleration a G0 ; The iG0 Indicates q iG0 The rotation axis, Indicates the expected planning angle; Said represents the planned expected angular velocity; Said represents the planned expected angular acceleration.

[0016] Further, the method of completing the posture trajectory planning by improving the tracking differentiator is specifically as follows: the planner is represented as a second-order discrete system; the second-order discrete system is specifically as follows:

[0017] The t represents an index in the discrete time series.

[0018] Furthermore, the total control rate is the sum of the adaptive control rate and the compensation control rate, and its calculation formula is as follows: u = u a +u b , where u represents the total control rate, and u a represents the adaptive control rate, the u a =[u a1 ,u a2 ,u a3 ], and its calculation process is as follows: j = [1, 2, 3], δ k is a positive constant, k is an adaptive gain, and u max To control the torque limit, the s aj For a The elements in s a is the sliding vector, the s a =ω e +k 2 q e13 , the q e13 for q e The imaginary part of q e is the deviation quaternion, The q n Represents the real body quaternion, the q G represents the real-time expected quaternion in the inertial system; the ω e is the deviation angular velocity, the ω e =ω n -r(q e )ω G , the ω n represents the real body angular velocity, the ω G represents the real-time expected angular velocity in the inertial system, and R(q e ) represents the transformation matrix from the target coordinate system to the body coordinate system; the u b Represents the compensation control rate; Said The C e =R(q e ).

[0019] Furthermore, the actuator is a flywheel model, and the total control rate u is input into the flywheel model to obtain the real control rate u′.

[0020] Furthermore, the angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite dynamics model; and the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite kinematics model.

[0021] Furthermore, the measuring mechanism includes a gyro model and a star sensor model; the real body quaternion and the real body angular velocity are obtained by the measuring mechanism as follows: the angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is input into the gyro model to obtain the real body angular velocity ω n ; Input the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system into the star sensor model to obtain the real body quaternion q n .

[0022] Furthermore, the cyclic planning trajectory and tracking are specifically as follows: the real body angular velocity ω n and the real body quaternion q n Enter step S4 to complete the planning and tracking of the desired trajectory.

[0023] The beneficial effects of the present invention are:

[0024] The present invention can realize trajectory planning from any satellite attitude to imaging attitude, plan the shortest three-axis path and perform path tracking while considering external interference and flywheel constraints. A tracking controller based on adaptive parameters and compensation estimation is designed for attitude tracking, and adaptive parameters are used to improve anti-interference capability, and actuator saturation is considered to control the satellite to smoothly track the planned trajectory to reach the target attitude. Compared with the current advanced control methods FAMF and PD methods, this method comprehensively considers the maneuverability and stability of the satellite, and ensures attitude stability during imaging while improving maneuverability. The attitude angular velocity and angular acceleration obtained by attitude planning are used to design compensation estimates to improve tracking capabilities. This method is simple to adjust parameters and easy to implement in engineering. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 This is a flow chart of a posture trajectory planning and tracking method according to an embodiment of the present invention;

[0026] Figure 2 A schematic diagram of a shaft angle curve planned for an embodiment of the present invention;

[0027] Figure 3 A schematic diagram of a planning axis angular velocity curve according to an embodiment of the present invention;

[0028] Figure 4 A schematic diagram of a planned axis angular acceleration curve according to an embodiment of the present invention;

[0029] Figure 5 This is a schematic diagram of a deviation quaternion curve according to an embodiment of the present invention;

[0030] Figure 6 This is a schematic diagram of a deviation angular velocity curve according to an embodiment of the present invention;

[0031] Figure 7 Schematic diagram of side swing angle comparison curve according to an embodiment of the present invention;

[0032] Figure 8 Schematic diagram of a comparison curve of lateral swing angular velocity according to an embodiment of the present invention;

[0033] Fig. 9 A schematic diagram of a three-axis deviation angle curve according to an embodiment of the present invention;

[0034] Fig.10 Schematic diagram of three-axis deviation angular velocity curves according to an embodiment of the present invention. DETAILED DESCRIPTION

[0035] The technical solution of the present invention will be described clearly and completely below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0036] Embodiment 1,

[0037] This embodiment provides a low-orbit remote sensing satellite three-axis rapid maneuvering attitude trajectory planning and tracking method, and the relevant parameters of the embodiment are shown in Table 1:

[0038] Table 1

[0039]

[0040]

[0041] The flowchart of the method is as follows Figure 1 , the method comprises the following steps:

[0042] S1. Establish satellite kinematic model and satellite dynamic model.

[0043] The relevant operations of step S1 are introduced with a specific example:

[0044] The kinematic equation of the kinematic model is specifically: Said represents the derivative of the quaternion of the satellite body coordinate system relative to the inertial coordinate system, wherein Q represents the quaternion of the satellite body coordinate system relative to the inertial coordinate system, wherein Q = (q0, q1, q2, q3) T =(q0,q T ) T, q0 represents the real part of Q, q represents the imaginary part of Q, q1, q2 and q3 represent the imaginary unit of Q, represents the derivative of q0, S(q) is the antisymmetric matrix of q, I3 is the third-order unit matrix, ω represents the angular velocity of the satellite body coordinate system relative to the inertial coordinate system, and ω=(ω1,ω2,ω3) T , ω1, ω2 and ω3 represent the angular velocities of the three axes of the body coordinate system, T represents the controller step size, and the symbol “·” represents the operation of derivation; the dynamic equation of the dynamic model is specifically: Said represents the derivative of ω, represents the positive definite matrix of the satellite's moment of inertia, 3 is a three-dimensional space, S(ω) represents the antisymmetric matrix of ω, and h w ∈ 3 is the total angular momentum of the reaction flywheel, and u′ represents the true control rate.

[0045] S2. Coordinate transformation: Calculate the quaternion of the desired coordinate system relative to the satellite's initial coordinate system at the start of the operation.

[0046] The relevant operations of step S2 are introduced with a specific example:

[0047] The calculation formula of the quaternion of the initial coordinate system at the start time of the maneuver relative to the satellite orbit coordinate system is: q OG is the quaternion of the desired coordinate system relative to the orbital coordinate system, q iO0 is the quaternion of the initial coordinate system relative to the satellite orbit coordinate system at the start of the maneuver; iO0 The calculation formula is The q i0 is the initial quaternion of the satellite, and the q O0 is the initial orbit quaternion.

[0048] S3. Complete attitude trajectory planning by improving the tracking differentiator.

[0049] The relevant operations of step S3 are introduced with a specific example:

[0050] When the satellite is in multi-point imaging mission mode, different imaging points correspond to different roll angles. The initial state of the satellite is an arbitrary attitude, and the satellite sets the target attitude to roll angles of 10°, -10°, 20° and -20° relative to the orbital coordinate system at 0s, 200s, 300s and 400s respectively.

[0051] In order to realize the shortest path rotation of the satellite's three-axis attitude, the shortest path is designed by using the three-axis attitude transformation rotation axis invariance constraint, and the satellite's three-axis angular acceleration and angular velocity are constrained. The attitude trajectory planning is completed by improving the tracking differentiator. The attitude trajectory planning is specifically: using the rotation axis e iG0 Get the quaternion q of the satellite's desired coordinate system relative to the initial coordinate system G0 , angular velocity ω G0 and acceleration a G0 , The iG0 Indicates q iG0 The rotation axis, Said Indicates q iG0 The imaginary part, the θ iG0 represents the axis angle, the θ iG0 =2arccos(q iG0 ), represents the expected planning angle (planning axis angle); Said represents the planned expected angular velocity (planned axis angular velocity); Said represents the planned expected angular acceleration (planned axis angular acceleration), such as Figure 2 , 3 As shown in FIG. 4 , it can be seen that the planning trajectory of the present invention comprehensively considers the limitations of the actuator and the maneuverability of the satellite.

[0052] The method of completing the posture trajectory planning by improving the tracking differentiator is as follows: the planner is represented as a second-order discrete system; the second-order discrete system is as follows: The t represents an index in a discrete time series; The calculation formula is: Said is a saturation function, s=sign(y), f=sign(f′-fix(f′))+fix(f′), sign() represents a sign function, The r l represents the maximum angular acceleration of the axis angle, The r l represents the maximum angular acceleration of the three axes, Said represents the maximum angular acceleration of the x-axis, represents the maximum angular acceleration of the y-axis, represents the maximum angular acceleration of the z-axis, and y represents an intermediate calculation variable. h is a smoothing factor, and fix() is an upward rounding function. This method can balance the stability and rapidity of the planned path by adjusting the size of the smoothing factor h.

[0053] S4. Calculate the total control rate.

[0054] The relevant operations of step S4 are introduced with a specific example:

[0055] The total control rate is designed as the sum of the adaptive control rate and the compensation control rate. The adaptive control rate uses an adaptive gain to ensure that the attitude and angular velocity errors are close to zero and explicitly considers the control input saturation; the compensation control rate is used to improve the dynamic performance and attitude stability during attitude tracking. This method can prove that the whole system has global stability in the presence of bounded disturbances and parameter uncertainties.

[0056] The total control rate is the sum of the adaptive control rate and the compensation control rate, and its calculation formula is: u = u a +u b , where u represents the total control rate, and u a represents the adaptive control rate, the u a =[u a1 ,u a2 ,u a3 ], and its calculation process is as follows: j = [1, 2, 3], δ k is a positive constant, k is an adaptive gain, and the derivative of k is: The γ is a constant representing the convergence rate; the q e0 Indicates q e The real part of ω ei Represents the vector ω e The i-th element in q ei Represents vector q e The i-th element of s i represents the i-th element of vector s, Represents ω e The transpose of Indicates q e13 The transpose of q e13 Indicates q e The imaginary part, u max To control the torque limit, the s aj is the sliding vector s a The elements in s a =ω e +k 2 q e13 , the q e13 Q e The imaginary part of qe =[q e0 ,q e1 ,q e2 ,q e3 ] is the deviation quaternion, the q e1 ,q e2 and q e3 for q e The imaginary unit, The q n Represents the real body quaternion, the q G Represents the real-time expected quaternion in the inertial system; The q OG0 Represents the quaternion of the desired coordinate system at the time of maneuver relative to the orbital coordinate system. The planned trajectory is based on the trajectory at the time of maneuver and needs to consider the real-time changing orbital quaternion. The q O represents the real-time orbital coordinate system; the ω e =[ω e1 ,ω e2 ,ω e3 ] is the deviation angular velocity, the ω e1 ,ω e2 and ω e3 ω e The first row of elements, the ω e =ω n -R(q e )ω G , the ω n represents the real body angular velocity, the ω G represents the real-time expected angular velocity in the inertial system, and R(q e ) represents the transformation matrix from the target coordinate system to the body coordinate system; the ω G =ω OG +ω G0 , the ω OG represents the angular velocity of the orbital coordinate system relative to the inertial coordinate system in the desired coordinate system, The ω O represents the real-time orbital angular velocity, the R(q OG0 ) is q OG0 The rotation matrix of Figure 5 and Figure 6 As shown, the control accuracy and attitude stability of the satellite during imaging meet the imaging control requirements; b Represents the compensation control rate; Said The C e =R(q e ).

[0057] S5. Obtain the actual control rate through the actuator.

[0058] The relevant operations of step S5 are introduced with a specific example:

[0059] The actuator is a flywheel model. In the simulation, the total control rate u is input into the flywheel model to obtain a real control rate u′ with noise.

[0060] S6. Calculate the quaternion of the satellite body coordinate system relative to the inertial coordinate system and the angular velocity of the satellite body coordinate system relative to the inertial coordinate system.

[0061] The relevant operations of step S6 are introduced with a specific example:

[0062] The angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite dynamics model; the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite kinematics model.

[0063] S7. Obtain the real body quaternion and the real body angular velocity through the measuring mechanism.

[0064] The relevant operations of step S7 are introduced with a specific example:

[0065] The measuring mechanism includes a gyro model and a star sensor model; in the simulation, the angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is input into the gyro model to obtain the real body angular velocity ω with noise n In the simulation, the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system is input into the star sensor model to obtain the real body quaternion q with noise n .

[0066] S8, loop planning trajectory and tracking.

[0067] The loop planning trajectory and tracking are specifically as follows: the real body angular velocity ω n and the real body quaternion q n Enter step S4 to complete the planning and tracking of the desired trajectory.

[0068] Embodiment 2,

[0069] This embodiment provides a comparative experiment of a three-axis rapid maneuvering attitude trajectory planning and tracking method for a low-orbit remote sensing satellite; the process is as follows:

[0070] First, the maneuvering efficiency of the present invention is compared with that of the FAMF and PD methods for multi-point imaging tasks. Figure 7 and 8As shown, a schematic diagram of a roll angle comparison curve and a schematic diagram of a roll angular velocity comparison curve are shown. From the curves, it can be seen that the present invention can achieve faster convergence in the roll angle and the roll angular velocity compared with the other two methods.

[0071] It is assumed that when the satellite orbital angular velocity is 3 times the standard deviation (99.7% confidence level) less than or equal to 0.001° / s within the 30s time window, the maneuver is considered to be in a stable state. The maneuvering stability time is defined as the start time of the stable state minus the start time of the maneuver. The attitude maneuvering stability time of the three methods is shown in Tables 2 and 3. In order to quantify the maneuvering efficiency, the attitude maneuvering time of the present invention is regarded as a 100% comparison benchmark. It can be seen from Tables 2 and 3 that the maneuvering efficiency of the present invention is higher than that of the other two methods, which proves that the path planned by the present invention is the shortest path of the three axes. The control method of the present invention has superior tracking performance and improves the maneuvering ability of the satellite.

[0072] Table 2

[0073] Target swing angle The present invention FAMF PD 10° 73.875s(100％) 86.625s(117％) 84.750s(115％) -10° 35.375s(100％) 49.125s(139％) 52.625s(149％) 20° 41.375s(100％) 54.875s(133％) 57.250s(138％) -20° 46.250s(100％) 59.500s(129％) 61.875s(134％)

[0074] Table 3

[0075]

[0076] The present invention also conducts comparative experiments on attitude stability and control accuracy for the ground push-scan mode. The satellite requires attitude stability and control accuracy to be small enough in the ground push-scan imaging mode to ensure the imaging quality. The attitude planning and control process sets the initial state of the satellite to a certain arbitrary attitude, and the target state to a 30-degree side swing angle relative to the orbital coordinate system. The deviation angle value and angular velocity value obtained by the experiment are as follows: Fig. 9 , 10 shown.

[0077] The attitude stability and control deviation of the present invention are regarded as 100% comparison benchmarks, and the comparison of attitude stability and attitude control deviation after satellite stabilization is shown in Table 3. Experiments show that the attitude stability and control deviation of the present invention when the attitude is stable are better than those of the other two methods, from which it can be seen that the controller used in the present invention is superior in anti-interference ability, and the method realizes the progressive tracking of the planned trajectory and the interference suppression of external noise.

Claims

1. A method for three-axis rapid maneuvering trajectory planning and tracking of a low-orbit remote sensing satellite, characterized in that: The method comprises the following steps: S1. Establish satellite kinematic model and satellite dynamic model; S2, calculate the quaternion of the desired coordinate system relative to the satellite initial coordinate system at the start of the movement; S3, completing attitude trajectory planning by improving the tracking differentiator; S4. Calculate the total control rate; S5. Get the real control rate through the actuator; S6, calculating the quaternion of the satellite body coordinate system relative to the inertial coordinate system and the angular velocity of the satellite body coordinate system relative to the inertial coordinate system; S7, obtaining the real body quaternion and the real body angular velocity through the measuring mechanism; S8, loop planning trajectory and tracking.

2. The method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite according to claim 1, characterized in that: The kinematic equation of the kinematic model is specifically: Said represents the derivative of the quaternion of the satellite body coordinate system relative to the inertial coordinate system, q0 represents the real part of Q, Q represents the quaternion of the satellite body coordinate system relative to the inertial coordinate system, q represents the imaginary part of Q, represents the derivative of q0, S(q) represents the antisymmetric matrix of q, I3 represents the third-order unit matrix, ω represents the angular velocity of the satellite body coordinate system relative to the inertial coordinate system, and T represents the controller step size; the dynamic equation of the dynamic model is specifically: Said represents the derivative of ω, represents the positive definite matrix of the satellite's moment of inertia, 3 represents a three-dimensional space, S(ω) represents the antisymmetric matrix of ω, and h w ∈ 3 is the total angular momentum of the reaction flywheel, the u ′ Indicates the true control rate.

3. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite according to claim 1, the calculation formula of the quaternion of the expected coordinate system relative to the satellite initial coordinate system at the start of the maneuver is: q OG is the quaternion of the desired coordinate system relative to the orbital coordinate system, q io0 is the quaternion of the initial coordinate system relative to the satellite orbit coordinate system at the start of the maneuver, Represents a tensor product operation.

4. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite according to claim 3, the attitude trajectory planning is specifically: using the rotation axis e ig0 Get the quaternion q of the satellite's desired coordinate system relative to the initial coordinate system G0 , angular velocity ω G0 and acceleration a G0 ; The iG0 Indicates q iG0 The rotation axis, Indicates the expected planning angle; Said represents the planned expected angular velocity; Said represents the planned expected angular acceleration.

5. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite according to claim 4, the attitude trajectory planning is completed by improving the tracking differentiator: the planner is represented as a second-order discrete system; the second-order discrete system is specifically: The t represents an index in the discrete time series.

6. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite in claim 5, the total control rate is the sum of the adaptive control rate and the compensation control rate, and its calculation formula is specifically: u = u a +u b , where u represents the total control rate, and u a represents the adaptive control rate, u a =[u a1 ,u a2 ,u a3 ], and its calculation process is as follows: j = [1, 2, 3], δ k is a positive constant, k is an adaptive gain, and u max To control the torque limit, the s aj For a The elements in s a is the sliding vector, the s a =ω e +k 2 q e13 , the q e13 Q e The imaginary part of q e is the deviation quaternion, The q n Represents the real body quaternion, the q G represents the real-time expected quaternion in the inertial system; the ω e is the deviation angular velocity, the ω e =ω n -r(q e )ω G , the ω n represents the real body angular velocity, the ω G represents the real-time expected angular velocity in the inertial system, and r(q e ) represents the transformation matrix from the target coordinate system to the body coordinate system; the u b Represents the compensation control rate; Said The C e =R(q e ).

7. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite in claim 6, the actuator is a flywheel model, and the total control rate u is input into the flywheel model to obtain the real control rate u ′ .

8. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite as described in claim 7, the angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite dynamics model; the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system is calculated by a satellite kinematics model.

9. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite in claim 8, the measuring mechanism comprises a gyro model and a star sensor model; the real body quaternion and the real body angular velocity obtained by the measuring mechanism are specifically as follows: the angular velocity ω of the satellite body coordinate system relative to the inertial coordinate system is input into the gyro model to obtain the real body angular velocity ω n ; Input the quaternion Q of the satellite body coordinate system relative to the inertial coordinate system into the star sensor model to obtain the real body quaternion q n .

10. According to the method for three-axis rapid maneuvering attitude trajectory planning and tracking of a low-orbit remote sensing satellite in claim 8, the cyclic planning trajectory and tracking are specifically: the real body angular velocity ω n and the real body quaternion q n Enter step S4 to complete the planning and tracking of the desired trajectory.

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