Switching control method for rapid maneuvering of attitude of optical remote sensing satellite

Abstract

The invention relates to a switching control method for optical remote sensing satellite attitude rapid maneuver. The switching control method comprises the following steps: step 1, establishing a satellite reference coordinate system and an attitude kinetic equation; step 2, path planning of an attitude maneuvering section; step 3, path planning of an attitude convergence section; step 4, establishing an error kinetic equation for controller design; 5, attitude switching control is carried out; and step 6, rate compensation control. According to the switching control method for optical remote sensing satellite attitude rapid maneuver, the satellite attitude maneuver control efficiency can be further improved, the attitude adjustment time is effectively shortened, the number of imaging targets of a remote sensing satellite in finite time is increased, and the imaging area is expanded; high-efficiency and high-quality multi-target image acquisition is realized; parameters are easy to debug, attitude control efficiency is high, and the method is suitable for attitude control system design of small and medium-sized low-cost optical remote sensing satellites in engineering practice.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and in particular to a switching control method for rapid attitude maneuvering of an optical remote sensing satellite. Background Art

[0002] With the continuous development of the aerospace field, the technology of optical remote sensing satellites is also constantly iterating and updating. In addition to traditional advanced large-scale optical remote sensing satellites such as WorldView and Spot, multiple remote sensing satellite constellations have been built, such as the Jilin-1 constellation and the Pigeon constellation. These constellations have high image timeliness and powerful image acquisition capabilities through a large number of satellite networks. At the same time, the imaging functions of satellites are becoming increasingly rich. They are no longer limited to traditional ground push-scanning imaging, but can also achieve multi-target point imaging, co-track stereo imaging, multi-strip imaging, and inertial space imaging. In order to comprehensively improve the image acquisition capabilities of remote sensing satellite constellations, satellites need to have the ability to quickly adjust and control attitudes to ensure that image data of multiple targets can be captured in a short period of time in an efficient and high-quality manner.

[0003] Traditional high-performance satellite attitude maneuver control relies on high-torque actuators to control torque gyroscopes, but they are expensive and not suitable for the design of small optical remote sensing satellites. For small optical remote sensing satellites, most use low-cost, light-weight reaction flywheels as actuators, but their control torque is relatively small. Summary of the invention

[0004] The present invention aims to solve the technical problems in the prior art and provides a switching control method for rapid attitude maneuvering of an optical remote sensing satellite.

[0005] In order to solve the above technical problems, the technical solutions of the present invention are as follows:

[0006] A switching control method for rapid attitude maneuvering of an optical remote sensing satellite comprises the following steps:

[0007] Step 1, establish the satellite reference coordinate system and attitude dynamics equations;

[0008] Step 2, path planning for the attitude maneuvering phase;

[0009] Step 3: path planning for the attitude convergence segment;

[0010] Step 4, establishing the error dynamics equation for controller design;

[0011] According to the attitude dynamics equation of step 1, the expected attitude quaternion Q of the attitude maneuver process planned in steps 2 and 3 r , desired angular velocity ω r and the expected angular acceleration a r, we get the error dynamics equation of attitude control;

[0012] Step 5, posture switching control;

[0013] The controller is designed using sliding mode control theory;

[0014] Step 6, rate compensation control;

[0015] When the pointing error of the attitude control system meets the preset conditions, rate compensation control is performed.

[0016] In the above technical solution, step 1 is specifically:

[0017] Inertial coordinate system is the J2000 flat equatorial geocentric coordinate system;

[0018] Orbital coordinate system In, O O X O is the flight direction of the satellite, O O Z O From the satellite pointing to the center of the earth, O O Y O With O O X O and O O Z O Follow the right-hand rule;

[0019] Satellite body coordinate system In, O B Z B Coincident with the optical axis of the optical remote sensing satellite payload, O B X B Coincident with the direction of push-scan imaging of optical payload, O B Y B With O B X B and O B Z B Follow the right-hand rule;

[0020] The satellite attitude dynamics equation is:

[0021]

[0022] in, Represents the attitude quaternion; Indicates q v The antisymmetric matrix of ; represents the attitude angular velocity, The satellite's moment of inertia matrix; is the control angular momentum of the reaction flywheel; is the control torque of the reaction flywheel.

[0023] In the above technical solution, the satellite body coordinate system Relative to the inertial coordinate system The attitude uses quaternion Represents and satisfies the constraints The attitude angular velocity is expressed as The attitude angular acceleration is expressed as

[0024] In the above technical solution, step 2 is specifically as follows:

[0025] The planning of the attitude maneuver phase is based on the orbital coordinate system is the reference coordinate system, given the attitude maneuver Euler angle Planning Euler Angles and angular velocity The differential equation is:

[0026]

[0027] in, is the planned attitude maneuver angular acceleration, which is specifically designed as follows:

[0028]

[0029] Where r is the planned maximum angular acceleration; is the planned maximum angular velocity; h is the smoothing coefficient of the convergence segment; sign(a) represents the sign function operation; a is the intermediate calculation variable, and its specific calculation method is:

[0030] θ e =(θ r -θ d )+hω r

[0031]

[0032] |θ e |≤rh 2 :a=ω r +θ e / h

[0033] Among them, θ e represents the planning error variable;

[0034] The maximum maneuvering angular acceleration r satisfies the following conditions:

[0035]

[0036] in, is the maximum control torque of the flywheel; J is the moment of inertia of the attitude maneuvering axis;

[0037] Maximum maneuvering angular velocity Requirements:

[0038]

[0039] in, is the maximum controlled angular momentum of the flywheel.

[0040] In the above technical solution, in step 2, during the pitch maneuver, the expected attitude in the orbital coordinate system is calculated as:

[0041]

[0042] in, Represents the planning attitude quaternion in the orbital coordinate system; represents the planned angular velocity in the orbital coordinate system; represents the planned angular acceleration in the orbital coordinate system;

[0043] Convert the desired attitude information in the orbital coordinate system into the desired attitude quaternion Q in the inertial coordinate system r , desired angular velocity ω r and the expected angular acceleration a r for:

[0044]

[0045] Among them, Q oI is the attitude quaternion of the orbital coordinate system relative to the inertial coordinate system; ω o represents the orbital angular velocity; Representing quaternions The rotation transformation matrix of represents the desired angular velocity ω r The differential of .

[0046] In the above technical solution, during pitch maneuver control, step 3 is specifically as follows:

[0047] Convert the current attitude quaternion to the orbital coordinate system:

[0048]

[0049] in, Represents the quaternion intermediate variable in the orbital coordinate system, which is used for subsequent planning calculations; (Q oI ) -1 Represents the orbital quaternion Q oI The inverse of

[0050] Calculate the current pitch angle θ based on the quaternion of the orbital coordinate system s for:

[0051] θ s =arcsin(2(q s0 q s2 -q s1 q s3 ))

[0052] At the pitch angle θ s is the expected angle, repeat step 2 to obtain the expected posture information of the plan.

[0053] In the above technical solution, step 4 is specifically as follows:

[0054] The attitude control error is defined as:

[0055]

[0056] in, represents the inverse of the planning quaternion, Represents quaternion multiplication;

[0057] The angular velocity tracking error is defined as:

[0058] ω e =ω-R(Q e )ω r

[0059] Among them, ω r is the planned angular velocity of the satellite;

[0060] The rotation matrix R(Q e ) has the following relationship:

[0061]

[0062] in, The error angular velocity ω e The antisymmetric matrix of ;

[0063] And satisfy the constraint condition || R(Q e )||=1;

[0064] According to the attitude dynamics equation in step 1, combined with the attitude control error and angular velocity tracking error, the error dynamics equation of attitude control is obtained as follows:

[0065]

[0066] in, represents the error quaternion vector part q ev The antisymmetric matrix of .

[0067] In the above technical solution, step 5 is specifically as follows:

[0068] The sliding surface σ is defined as:

[0069] σ=λω e +q ev

[0070] Where, λ is the sliding surface gain;

[0071] The pointing error requirement for a given satellite is:

[0072]

[0073] Among them, ||q ev ||,||ω e || respectively represent q ev ,ω e The vector norm of ; The index requirements for satellite attitude control;

[0074] When the attitude control system pointing error Combined with the path planning of the attitude maneuvering segment in step 2, the attitude controller is:

[0075] T w =-k σ σ+ω × (Jω+h w )+JR(Q e ) r

[0076] Among them, k σ represents the convergence gain of attitude control;

[0077] When the attitude control system pointing error Combined with the path planning of the attitude convergence segment in step 3, the attitude controller is:

[0078]

[0079] Among them, k σ represents the convergence gain of attitude control; k s is the gain of switching control; ε is the dead zone of switching control; |σ| represents the modulus value of the sliding surface σ.

[0080] In the above technical solution, step 6 is specifically as follows:

[0081] When the attitude control system pointing error Additional compensating controls are:

[0082] T a =k a ω e ;

[0083] Among them, k a represents the compensation control gain, T a Indicates the additional compensation control torque.

[0084] The present invention has the following beneficial effects:

[0085] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention can further improve the efficiency of satellite attitude maneuvering control, effectively shorten the time for attitude adjustment, increase the number of imaging targets of the remote sensing satellite within a limited time, and expand the imaging area.

[0086] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention takes into account accurate attitude stability and realizes high-efficiency and high-quality multi-target image acquisition.

[0087] The switching control method for rapid maneuvering of the optical remote sensing satellite attitude of the present invention has a simple structure, easy parameter adjustment, high attitude control efficiency, and is suitable for the attitude control system design of small and medium-sized low-cost optical remote sensing satellites in engineering practice. BRIEF DESCRIPTION OF THE DRAWINGS

[0088] The present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments.

[0089] Figure 1 Schematic diagram of the reference coordinate system for satellite control.

[0090] Figure 2 The present invention is a structural block diagram of a control system applicable to the switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to the present invention.

[0091] Figure 3 Schematic diagram of attitude quaternion.

[0092] Figure 4 Schematic diagram of error quaternion.

[0093] Figure 5 Schematic diagram of error angular velocity.

[0094] Figure 6 Schematic diagram of three-axis Euler angles. DETAILED DESCRIPTION

[0095] The inventive concept of the present invention is:

[0096] In order to realize the rapid maneuvering control of the attitude of the optical remote sensing satellite under a small control torque, the present invention designs a switching control method for the rapid maneuvering of the attitude of the optical remote sensing satellite through a switching control method. By switching the attitude planning path and the attitude controller in different control stages, not only the rapid convergence of the attitude maneuvering is realized, but also the high stability control of the satellite imaging is realized, thereby ensuring the high-quality and high-efficiency image acquisition.

[0097] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention further improves the efficiency of satellite attitude maneuvering control and realizes high-efficiency and high-quality remote sensing satellite multi-target image acquisition.

[0098] The switching control method for rapid maneuvering of an optical remote sensing satellite attitude of the present invention comprises the following steps:

[0099] Step 1: Establish the satellite reference coordinate system and attitude dynamics equations:

[0100] like Figure 1 As shown, the reference coordinate system for satellite attitude control mainly includes: Inertial coordinate system Orbital coordinate system and the satellite body coordinate system

[0101] Inertial coordinate system It is the reference datum for the satellite to perform attitude control. All attitude information will eventually be unified into this coordinate system for attitude determination and control design. For optical remote sensing satellites, the inertial coordinate system often uses the J2000 flat equatorial geocentric coordinate system.

[0102] Orbital coordinate system is the reference coordinate system for satellite earth imaging mission, where O O X O is the flight direction of the satellite, O O Z O From the satellite pointing to the center of the earth, O O Y O With O O X O and O O Z O Follow the right-hand rule.

[0103] Satellite body coordinate system The definition of is also closely related to satellite imaging control. B Z B Coincident with the optical axis of the optical remote sensing satellite payload, O B X B Coincident with the direction of push-scan imaging of optical payload, O B Y B With O B X B and O B Z B Follow the right-hand rule.

[0104] Satellite body coordinate system Relative to the inertial coordinate system The attitude uses quaternion Represents and satisfies the constraints The attitude angular velocity is expressed as The attitude angular acceleration is expressed as

[0105] The satellite attitude dynamics equation is:

[0106]

[0107] in, Represents the attitude quaternion; Indicates q v The antisymmetric matrix of ; represents the attitude angular velocity, The satellite's moment of inertia matrix; is the control angular momentum of the reaction flywheel; is the control torque of the reaction flywheel.

[0108] Step 2, path planning for the attitude maneuvering phase:

[0109] The path planning of the attitude maneuvering stage adopts the angular velocity control logic of acceleration-holding-deceleration. The acceleration and deceleration stages are designed according to the maximum torque of the reaction flywheel, and the angular velocity of the holding stage is designed according to the maximum angular momentum of the reaction flywheel, ensuring that the attitude maneuvering stage performs attitude control with the maximum control capability of the flywheel and realizes rapid attitude adjustment. The planning of the attitude maneuvering stage is based on the orbital coordinate system. is the reference coordinate system, given the attitude maneuver Euler angle Planning Euler Angles and angular velocity The differential equation is:

[0110]

[0111] in, is the planned attitude maneuver angular acceleration, which is specifically designed as follows:

[0112]

[0113] Where r is the planned maximum angular acceleration; is the planned maximum angular velocity; h is the smoothing coefficient of the convergence segment; sign(a) represents the sign function operation; a is the intermediate calculation variable. The above formula shows that if the absolute value of the angular velocity exceeds the designed maximum angular velocity, the acceleration is designed to be 0; h is the smoothing coefficient of the convergence segment; a is an intermediate calculation variable, and its specific calculation method is:

[0114] θ e =(θ r -θ d )+hω r

[0115]

[0116] |θ e |≤rh 2 :a=ω r +θ e / h (4)

[0117] Among them, θ e represents the planning error variable

[0118] Maximum maneuvering angular acceleration r Requirements:

[0119]

[0120] in, is the maximum control torque of the flywheel; J is the moment of inertia of the attitude maneuvering axis.

[0121] Maximum maneuvering angular velocity Requirements:

[0122]

[0123] in, is the maximum controlled angular momentum of the flywheel.

[0124] Taking the pitch maneuver as an example, the expected attitude in the orbital coordinate system is calculated as:

[0125]

[0126] in, Represents the planning attitude quaternion in the orbital coordinate system; represents the planned angular velocity in the orbital coordinate system; represents the planned angular acceleration in the orbital coordinate system.

[0127] Convert the desired attitude information in the orbital coordinate system into the desired attitude quaternion Q in the inertial coordinate system r , desired angular velocity ω r and the expected angular acceleration a r for:

[0128]

[0129] Among them, Q oI is the attitude quaternion of the orbital coordinate system relative to the inertial coordinate system; ω o represents the orbital angular velocity; Representing quaternions The rotation transformation matrix of represents the desired angular velocity ω r The differential of .

[0130] Step 3: Path planning for the attitude convergence segment:

[0131] In the attitude convergence stage, the attitude pointing accuracy is relatively easy to meet the expected index requirements, while the convergence of angular velocity is the main factor limiting the attitude maneuvering time. In order to achieve high-precision and fast angular velocity control, the following still takes the pitch maneuvering control as an example to explain the attitude planning in the attitude convergence stage in detail. Convert the current attitude quaternion to the orbital coordinate system:

[0132]

[0133] in, Represents the quaternion intermediate variable in the orbital coordinate system, which is used for subsequent planning calculations; (Q oI ) -1 Represents the orbital quaternion Q oI The inverse of.

[0134] Calculate the current pitch angle θ based on the quaternion of the orbital coordinate system s for:

[0135] θ s =arcsin(2(q s0 q s2 -q s1 q s3 )) (10)

[0136] At the pitch angle θ s is the expected angle, and formulas (2)-(8) are repeated to obtain the expected posture information of the plan.

[0137] Step 4, establish the error dynamics equation for controller design:

[0138] The expected attitude quaternion Q of the attitude maneuver process planned according to step 2 and step 3 r , desired angular velocity ω r and the expected angular acceleration a r , define the attitude control error as:

[0139]

[0140] in, represents the inverse of the planning quaternion, Represents quaternion multiplication.

[0141] The angular velocity tracking error is defined as:

[0142] ω e =ω-R(Q e )ω r (12)

[0143] Among them, ωr is the planned angular velocity of the satellite; the rotation matrix R(Q e ) has the following relationship:

[0144]

[0145] in, The error angular velocity ω e The antisymmetric matrix of .

[0146] And satisfy the constraint condition || R(Q e )||=1.

[0147] According to the attitude dynamics equation of formula (1) in step 1, combined with the control error defined in formulas (11)-(12), the error dynamics equation of attitude control is:

[0148]

[0149] in, represents the error quaternion vector part q ev The antisymmetric matrix of .

[0150] Step 5, attitude switching control:

[0151] The satellite ensures fast attitude maneuver control. The controller is designed using sliding mode control theory. It is also robust to model parameter uncertainty and external disturbances. The sliding surface is defined as:

[0152] σ=λω e +q ev (15)

[0153] Where λ is the sliding surface gain.

[0154] The pointing error requirement for a given satellite is:

[0155]

[0156] Among them, ||q ev ||,||ω e || respectively represent q ev ,ω e The vector norm of ; The index requirements for satellite attitude control.

[0157] When the attitude control system pointing error Combined with the path planning of the attitude maneuvering segment in step 2, the attitude controller is designed as follows:

[0158] T w =-kσσ+ω × (Jω+h w )+JR(Qe ) r (17)

[0159] Among them, k σ Represents the convergence gain of attitude control.

[0160] When the attitude control system pointing error Combined with the path planning of the attitude convergence segment in step 3, the attitude controller is designed as follows:

[0161]

[0162] Among them, k σ represents the convergence gain of attitude control; k s is the gain of switching control; ε is the dead zone of switching control; |σ| represents the modulus value of the sliding surface σ.

[0163] Among them, k s is the gain of the switching control.

[0164] Step 6, rate compensation control:

[0165] In order to further speed up the convergence of the attitude maneuver convergence phase, when the attitude control system pointing error Perform rate compensation control, and the additional compensation control design is:

[0166] T a =k a ω e (19)

[0167] Among them, k a represents the compensation control gain, T a Indicates the additional compensation control torque.

[0168] The present invention is described in detail below with reference to the accompanying drawings.

[0169] The structural block diagram of the system to which the switching control method for rapid maneuvering of an optical remote sensing satellite attitude of the present invention is applicable is as follows: Figure 2 As shown. The satellite imaging Euler angle θ in the given orbital coordinate system d , the attitude maneuvering path is divided into the attitude maneuvering section and the convergence section for planning respectively. Then, according to the pointing accuracy index of attitude control, a switching attitude controller with different structures is designed, which ensures the maneuvering efficiency and high stability of the attitude maneuvering section and the convergence section at the same time. In addition, in order to further accelerate the convergence speed of the attitude convergence section, angular velocity compensation control is added in its control process. Finally, the designed method realizes high-efficiency attitude maneuvering adjustment.

[0170] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention illustrates the efficiency of attitude maneuvering control through attitude maneuvering control of a satellite pitch axis. The curve of attitude quaternion is as follows: Figure 3 As shown, the error quaternion and angular velocity of attitude control are as follows Figure 4-Figure 5 As shown, the Euler angle curves of the three axes are as follows Figure 6 As shown in the figure, it can be observed that the satellite maneuvers the pitch axis attitude from 15° to -15°, achieves the predetermined control stability within 24 seconds, meets the index requirements of imaging control, and fully converges the attitude within 26 seconds, achieving fast and stable imaging control.

[0171] The relevant parameters in this embodiment are shown in the following table:

[0172]

[0173] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention can further improve the efficiency of satellite attitude maneuvering control, effectively shorten the time for attitude adjustment, increase the number of imaging targets of the remote sensing satellite within a limited time, and expand the imaging area.

[0174] The switching control method for rapid attitude maneuvering of an optical remote sensing satellite of the present invention takes into account accurate attitude stability and realizes high-efficiency and high-quality multi-target image acquisition.

[0175] The switching control method for rapid maneuvering of the optical remote sensing satellite attitude of the present invention has a simple structure, easy parameter adjustment, high attitude control efficiency, and is suitable for the attitude control system design of small and medium-sized low-cost optical remote sensing satellites in engineering practice.

[0176] Obviously, the above embodiments are merely examples for the purpose of clear explanation, and are not intended to limit the implementation methods. For those skilled in the art, other different forms of changes or modifications can be made based on the above description. It is not necessary and impossible to list all the implementation methods here. The obvious changes or modifications derived therefrom are still within the scope of protection of the invention.

Claims

1. A switching control method for rapid attitude maneuvering of an optical remote sensing satellite, characterized in that: The following steps are involved: Step 1, establish the satellite reference coordinate system and attitude dynamics equations; Step 2, path planning for the attitude maneuvering phase; Step 3: path planning for the attitude convergence segment; Step 4, establishing the error dynamics equation for controller design; According to the attitude dynamics equation of step 1, the expected attitude quaternion Q of the attitude maneuver process planned in steps 2 and 3 r , desired angular velocity ω r and the expected angular acceleration a r , we get the error dynamics equation of attitude control; Step 5, posture switching control; The controller is designed using sliding mode control theory; Step 6, rate compensation control; When the pointing error of the attitude control system meets the preset conditions, rate compensation control is performed.

2. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: Step 1 is as follows: Inertial coordinate system is the J2000 flat equatorial geocentric coordinate system; Orbital coordinate system In, O O X O is the flight direction of the satellite, O O Z O From the satellite pointing to the center of the earth, O O Y O With O O X O and O O Z O Follow the right-hand rule; Satellite body coordinate system In, O B Z B Coincident with the optical axis of the optical remote sensing satellite payload, O B X B Coincident with the direction of push-scan imaging of optical payload, O B Y B With O B X B and O B Z B Follow the right-hand rule; The satellite attitude dynamics equation is: in, Represents the attitude quaternion; Indicates q v The antisymmetric matrix of ; represents the attitude angular velocity, The satellite's moment of inertia matrix; is the control angular momentum of the reaction flywheel; is the control torque of the reaction flywheel.

3. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 2, characterized in that: Satellite body coordinate system Relative to the inertial coordinate system The attitude uses quaternion Represents and satisfies the constraints The attitude angular velocity is expressed as The attitude angular acceleration is expressed as 4. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: Step 2 is as follows: The planning of the attitude maneuver phase is based on the orbital coordinate system is the reference coordinate system, given the attitude maneuver Euler angle Planning Euler Angles and angular velocity The differential equation is: in, is the planned attitude maneuver angular acceleration, which is specifically designed as follows: Where r is the planned maximum angular acceleration; is the planned maximum angular velocity; h is the smoothing coefficient of the convergence segment; sign(a) represents the sign function operation; a is the intermediate calculation variable, and its specific calculation method is: i e =(θ r -θ d )+hω r |θ e |≤rh 2 :a=ω r +θ e / h Among them, θ e represents the planning error variable; The maximum maneuvering angular acceleration r satisfies the following conditions: in, is the maximum control torque of the flywheel; J is the moment of inertia of the attitude maneuvering axis; Maximum maneuvering angular velocity Requirements: in, is the maximum controlled angular momentum of the flywheel.

5. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 4, characterized in that: In step 2, during the pitch maneuver, the expected attitude in the orbital coordinate system is calculated as: in, Represents the planning attitude quaternion in the orbital coordinate system; represents the planned angular velocity in the orbital coordinate system; represents the planned angular acceleration in the orbital coordinate system; Convert the desired attitude information in the orbital coordinate system into the desired attitude quaternion Q in the inertial coordinate system r , desired angular velocity ω r and the expected angular acceleration a r for: Among them, Q oI is the attitude quaternion of the orbital coordinate system relative to the inertial coordinate system; ω o represents the orbital angular velocity; Representing quaternions The rotation transformation matrix of represents the desired angular velocity ω r The differential of .

6. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: During pitch maneuver control, step 3 is as follows: Convert the current attitude quaternion to the orbital coordinate system: in, Represents the quaternion intermediate variable in the orbital coordinate system, which is used for subsequent planning calculations; (Q oI ) -1 Represents the orbital quaternion Q oI The inverse of Calculate the current pitch angle θ based on the quaternion of the orbital coordinate system s for: θ s arcsin(2(q). s0 q s2 -q s1 q s3 )) At the pitch angle θ s is the expected angle, repeat step 2 to obtain the expected posture information of the plan.

7. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: Step 4 is as follows: The attitude control error is defined as: in, represents the inverse of the planning quaternion, Represents quaternion multiplication; The angular velocity tracking error is defined as: oh e =ω-R(Q e )oh r Among them, ω r is the planned angular velocity of the satellite; The rotation matrix R(Q e ) has the following relationship: in, The error angular velocity ω e The antisymmetric matrix of ; And satisfy the constraint condition || R(Q e )||=1; According to the attitude dynamics equation in step 1, combined with the attitude control error and angular velocity tracking error, the error dynamics equation of attitude control is obtained as follows: in, represents the error quaternion vector part q ev The antisymmetric matrix of .

8. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: Step 5 is as follows: The sliding surface σ is defined as: I am e +q ev Where, λ is the sliding surface gain; The pointing error requirement for a given satellite is: Among them, ||q ev ||,||ω e || respectively represent q ev ,ω e The vector norm of ; The index requirements for satellite attitude control; When the attitude control system pointing error Combined with the path planning of the attitude maneuvering segment in step 2, the attitude controller is: T w =-k σ s+w × (Jω+h w )+JR(Q e )a r Among them, k σ represents the convergence gain of attitude control; When the attitude control system pointing error Combined with the path planning of the attitude convergence segment in step 3, the attitude controller is: Among them, k σ represents the convergence gain of attitude control; k s is the gain of switching control; ε is the dead zone of switching control; |σ| represents the modulus value of the sliding surface σ.

9. The switching control method for rapid attitude maneuvering of an optical remote sensing satellite according to claim 1, characterized in that: Step 6 is as follows: When the attitude control system pointing error Additional compensating controls are: T a =k a oh e ; Among them, k a represents the compensation control gain, T a Indicates the additional compensation control torque.

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