Magnetic control method for single-axis sun-oriented stable attitude of optical microsatellite

Abstract

The invention relates to a magnetic control method for a single-axis sun-oriented stable attitude of an optical microsatellite, and relates to the technical field of aerospace. The magnetic control method comprises the following steps: step 1, modeling optical microsatellite attitude dynamics adopting magnetic control; 2, calculating the posture of the single axis to the sun; 3, calculating an attitude control expected moment; and 4, calculating the expected magnetic moment of the magnetic torquer. According to the magnetic control method for the single-axis sun-oriented stable attitude of the optical microsatellite, the single-axis sun-oriented stable attitude control of the satellite can be realized under the condition of extremely low power consumption, and the energy stability of a non-task long-term mode and the reliable operation of the satellite can be ensured.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and in particular to a magnetic control method for a single-axis solar-stabilized attitude of an optical micro-satellite. Background Art

[0002] With the development of microelectronics technology, the payload and electronic components of optical remote sensing satellites are developing in the direction of miniaturization, lightness and high performance, and the overall design of satellites is also designed around the purpose of low mass, low power consumption and high performance. The actuators that provide control torque for optical microsatellites include magnetic torquers and reaction flywheels. In the attitude stabilization control mode, the reaction flywheel is generally used to provide three-axis torque for attitude stabilization control, and the magnetic torquer is used to provide external control torque to unload the excess angular momentum of the reaction flywheel. However, in the long-term operation mode of optical microsatellites such as the Sun-stabilization attitude, this control method has certain disadvantages. The power consumption of the reaction flywheel can be more than ten times that of the magnetic torquer, and once the reaction flywheel fails, it will cause the satellite attitude to become unstable, resulting in energy loss problems and affecting satellite safety. Summary of the invention

[0003] The present invention aims to solve the technical problems in the prior art and provides a magnetic control method for a single-axis solar-stabilized attitude of an optical micro-satellite.

[0004] In order to solve the above technical problems, the technical solutions of the present invention are as follows:

[0005] A magnetic control method for a single-axis solar-stabilized attitude of an optical micro-satellite comprises the following steps:

[0006] Step 1, modeling the attitude dynamics of the optical microsatellite using magnetic control;

[0007] Establish the attitude dynamics model of optical micro-satellite using magnetic control for the design of satellite attitude controller;

[0008] Step 2, calculate the single-axis attitude toward the sun;

[0009] Calculate the expected single-axis attitude toward the sun based on the UTC time obtained by the onboard GPS;

[0010] Step 3, calculation of desired torque for attitude control;

[0011] Combined with the satellite's attitude dynamics model, the desired attitude quaternion is calculated, and the error attitude dynamics model using magnetic control is established; the desired torque of magnetic control is calculated using a proportional differential controller;

[0012] Step 4, calculating the expected magnetic moment of the magnetic torquer;

[0013] According to the desired magnetic control torque calculated by the attitude controller, the desired magnetic moment of the magnetic torquer is calculated as the control instruction of the magnetic torquer.

[0014] In the above technical solution, in step 4, in the process of calculating the expected magnetic moment of the magnetic torquer, the switching control is designed according to the angular velocity of the satellite relative to the sun axis, so as to further enhance the stability of the single-axis sun control.

[0015] In the above technical solution, step 1 is specifically as follows:

[0016] The reference coordinate system for the attitude control of the optical microsatellite adopts the J2000 flat equatorial geocentric inertial coordinate system. As a reference coordinate system, the satellite's body coordinate system Each axis of rotation coincides with the principal axis of inertia; the body coordinate system Relative to the inertial coordinate system The attitude of is expressed using global non-singular quaternion as And satisfy the constraints The attitude angular velocity is expressed as The control torque of the magnetic torquer is expressed as

[0017] The satellite attitude dynamics controlled by magnetic torquer is expressed as:

[0018]

[0019] Where: is a positive definite matrix, representing the moment of inertia of the satellite; Denote Q, ω, q respectively 0 , the derivative of q; I 3 represents the 3×3 identity matrix; is an antisymmetric matrix. For any vector x, Satisfies S(x)y=x×y, where × represents vector cross product;

[0020] The satellite's expected attitude is defined as the attitude orientation of the body coordinate system relative to the inertial coordinate system, and is expressed by the expected attitude quaternion Representation; the attitude tracking error is defined as the error quaternion:

[0021]

[0022] Where: Represents quaternion multiplication;

[0023] The angular velocity tracking error is:

[0024] ω e =ω-R(Q e )ω d

[0025] Where: d is the desired angular velocity of the spacecraft; the rotation matrix R(Q e ) has the following relationship:

[0026]

[0027] Where: S(q e ) indicates q e The antisymmetric matrix of ; Represents R(Q e ) and satisfy the constraint || R(Q e )||=1.

[0028] In the above technical solution, step 2 is specifically as follows:

[0029] Set UTC time t utc Transformed into Julian time t JD for:

[0030] t JD =t utc / 86400+2451545

[0031] According to the Julian time t JD , calculate the Julian century number T JD for:

[0032] T JD =(t JD -2451545) / 36525

[0033] Obliquity of the ecliptic s By Julian century number T JD The calculation is:

[0034]

[0035] Sun geometric mean ecliptic longitude L 0 The mean anomaly angle M with the sun can be obtained from the Julian century number T JD The calculation is:

[0036]

[0037] According to the solar geometric mean ecliptic longitude L 0 The mean anomaly angle M with the sun gives the sun's ecliptic longitude l s for:

[0038]

[0039] According to the obliquity of the ecliptic s and the sun's ecliptic longitude s, the attitude quaternion of the earth pointing to the sun in the J2000 flat equatorial geocentric inertial coordinate system is calculated as:

[0040]

[0041] Since the attitude toward the sun is equivalent to the stability in inertial space, the expected angular velocity ω of the inertial coordinate system is d for:

[0042] ω d =

[000] T .

[0043] In the above technical solution, step 3 is specifically as follows:

[0044] The satellite's error attitude dynamics are:

[0045]

[0046] Where: is a positive definite matrix, representing the moment of inertia of the satellite; is an antisymmetric matrix. For any vector x, Satisfies S(x)y=x×y, where × represents vector cross product;

[0047] is the attitude angular velocity; is the control torque of the magnetic torquer; Respectively represent Q e ,ω,q e0 ,q e The derivative of .

[0048] The single-axis sun attitude controller is:

[0049] T c =u b +u f

[0050] Where:

[0051]

[0052] Where: K p is the proportional control gain; K d is the differential control gain; K f is the feedforward control gain; q e1 ,q e2 are the error quaternion q e The vector component of 1 ,ω 2 ,ω 3 are the vector components of the attitude angular velocity ω; δ k (ω 1) is the set self-stabilizing axis control gain and meets the following conditions:

[0053]

[0054] Where: T c is the desired control torque, is the angular velocity limit required for the self-stabilizing axis. b is a basic proportional-derivative controller, u f It is a nonlinear compensation controller.

[0055] In the above technical solution, step 4 is specifically as follows:

[0056] The desired control torque T for maintaining single-axis sun stability control is calculated. c Then, the magnetic field strength M of the Earth's magnetic field measured by the magnetometer in the satellite's coordinate system is b , the expected magnetic moment m of the three-axis magnetic torquer is obtained as:

[0057]

[0058] Where: M b1 ,M b2 ,M b3 M b The vector components of M b =(M b1 ,M b2 ,M b3 ) T ; T c1 ,T c2 ,T c3 T c The vector components of T c =(T c1 ,T c2 ,T c3 ) T ;k mag Assign gain to the magnetron.

[0059] The present invention has the following beneficial effects:

[0060] The magnetic control method of the optical micro-satellite single-axis solar-stabilized attitude of the present invention has two main advantages over the prior art:

[0061] First, since the magnetic torquer can only provide control torque in a direction perpendicular to the geomagnetic field, it cannot provide stable three-axis control torque. This problem makes it impossible for this magnetic control method to guarantee stable three-axis control of the Sun-stabilized attitude for a long time. The magnetic control method for the single-axis Sun-stabilized attitude of the optical microsatellite of the present invention applies the limited control capability of the magnetic torquer to the accuracy of the sailboard normal vector to the Sun, and efficiently realizes the single-axis Sun-stabilized attitude control. Compared with the three-axis Sun-stabilization of magnetic control. The magnetic control method for the single-axis Sun-stabilized attitude of the optical microsatellite of the present invention has higher accuracy towards the Sun, stronger charging efficiency, and higher system operation stability.

[0062] Second, the single-axis solar stabilization control mode does not require the use of a high-power reaction flywheel, but only requires the use of low-power actuators and sensors such as magnetic torquers, magnetometers, gyroscopes and star sensors, which greatly reduces the power consumption of the satellite in long-term non-mission mode, implements the low-power satellite design criteria, and reduces the development and production costs of satellites. The magnetic control method for the single-axis solar stabilization attitude of the optical microsatellite of the present invention has a simple structure, is easy to implement, and has high reliability. It can greatly improve the robustness of the optical satellite attitude control system in engineering practice, and has important engineering application value for low-cost microsatellite design.

[0063] The magnetic control method for the single-axis stable attitude of an optical micro-satellite towards the sun of the present invention can realize the single-axis stable attitude control of the satellite towards the sun under the condition of extremely low power consumption, and can ensure the energy stability of the non-mission long-term mode and the reliable operation of the satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] The present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments.

[0065] Figure 1 It is the control method principle and flow chart.

[0066] Figure 2 Schematic diagram of attitude quaternion curve.

[0067] Figure 3 Schematic diagram of the deviation quaternion curve.

[0068] Figure 4 It is a schematic diagram of attitude angular velocity curve.

[0069] Figure 5 Schematic diagram of the output magnetic torque curve of the magnetic torquer.

[0070] Figure 6 Schematic diagram of the output torque curve of the magnetic torquer. DETAILED DESCRIPTION

[0071] The inventive concept of the present invention is:

[0072] In order to achieve the low-power satellite design goal, further reduce the satellite's operating power consumption in long-term mode, and improve the reliability of the attitude control system, the present invention proposes a magnetic control method for the single-axis solar-stabilized attitude of an optical microsatellite, which can achieve single-axis solar-stabilized control under extremely low power consumption conditions and has important engineering application value.

[0073] The present invention is described in detail below with reference to the accompanying drawings.

[0074] The magnetic control method of the optical micro-satellite single-axis sun-stabilized attitude of the present invention comprises the following steps:

[0075] Step 1, modeling the attitude dynamics of the optical microsatellite using magnetic control;

[0076] The reference coordinate system for the attitude control of the optical microsatellite adopts the J2000 flat equatorial geocentric inertial coordinate system. As a reference coordinate system, the satellite's body coordinate system Each axis of rotation coincides with the principal axis of inertia. Relative to the inertial coordinate system The attitude is generally expressed using global non-singular quaternion as And satisfy the constraints The attitude angular velocity is expressed as The control torque of the magnetic torquer is expressed as Therefore, the satellite attitude dynamics controlled by the magnetic torquer can be expressed as:

[0077]

[0078] Where: is a positive definite matrix, representing the moment of inertia of the satellite; Denotes Q,ω,q 0 , the derivative of q; I 3 represents the 3×3 identity matrix; is an antisymmetric matrix. For any vector x, It satisfies S(x)y=x×y, where × represents vector cross product.

[0079] The satellite's expected attitude is defined as the attitude orientation of the body coordinate system relative to the inertial coordinate system, and is expressed by the expected attitude quaternion Represented by. The posture tracking error is defined as the error quaternion:

[0080]

[0081] Where: Represents quaternion multiplication.

[0082] The angular velocity tracking error is:

[0083] ωe =ω-R(Q e )ω d (3)

[0084] Where: d is the desired angular velocity of the spacecraft; the rotation matrix R(Q e ) has the following relationship:

[0085]

[0086] Where: S(q e ) indicates q e The antisymmetric matrix of ; Represents R(Q e ) and satisfy the constraint || R(Q e )||=1.

[0087] Step 2, calculate the single-axis attitude toward the sun;

[0088] The vector direction of the solar attitude is the direction of the Earth pointing to the Sun. To calculate the single-axis solar attitude, the first thing to do is to calculate the solar vector. In the J2000 flat equatorial geocentric inertial coordinate system, the solar direction vector r sun According to the UTC time of satellite GPS utc The specific calculation process is as follows. utc Transformed into Julian time t JD for:

[0089] t JD =t utc / 86400+2451545 (5)

[0090] According to the Julian time t JD , we can calculate the Julian century number T JD for:

[0091] T JD =(t JD -2451545) / 36525 (6)

[0092] Obliquity of the ecliptic s The Julian century number T JD The calculation is:

[0093]

[0094] Sun geometric mean ecliptic longitude L 0 The mean anomaly angle M with the sun can be obtained from the Julian century number T JD The calculation is:

[0095]

[0096] According to the solar geometric mean ecliptic longitude L 0 The angle of anomaly with the sun is M, and the sun's ecliptic longitude is l s for:

[0097]

[0098] According to the obliquity of the ecliptic s and the sun's ecliptic longitude s , we can calculate that the attitude quaternion of the earth pointing to the sun in the J2000 flat equatorial geocentric inertial coordinate system is:

[0099]

[0100] Since the attitude toward the sun is equivalent to inertial space stability, the expected angular velocity ω of the inertial coordinate system is d for:

[0101] ω d =[0 0 0] T (11)

[0102] Step 3, calculation of desired torque for attitude control;

[0103] According to the satellite attitude dynamics model in step 1, combined with the expected attitude information of the sun calculated in the previous step, the satellite error attitude dynamics is further calculated as:

[0104]

[0105] All variables in formula (12) have been introduced and explained in step 1.

[0106] Since the magnetic torquer can only provide limited control torque in the direction perpendicular to the geomagnetic field, and in the single-axis solar attitude mode, the normal operation of the satellite is conditional on the normal vector of the solar sail panel coinciding with the solar vector. In order to stably achieve this control goal, the main control capability of the magnetic torquer is used to ensure that the normal vector of the solar sail panel coincides with the solar vector, and a small amount of control capability is used to ensure the attitude self-stabilization of the normal vector axis system of the solar sail panel. Therefore, the single-axis solar attitude controller is designed as:

[0107] T c =u b +u f (13)

[0108] Where:

[0109]

[0110] Where: K p is the proportional control gain; K d is the differential control gain; K fis the feedforward control gain;

[0111] q e1 ,q e2 are the error quaternion q e The vector component of 1 ,ω 2 ,ω 3 are the vector components of the attitude angular velocity ω; δ k (ω 1 ) is the set self-stabilizing axis control gain and meets the following conditions:

[0112]

[0113] Where: T c is the desired control torque, is the angular velocity limit required by the self-stabilizing axis, u b is a basic proportional-derivative controller, u f It is a nonlinear compensation controller.

[0114] Step 4, calculating the expected magnetic moment of the magnetic torquer;

[0115] The desired control torque T for maintaining single-axis sun stability control is calculated. c Then, the magnetic field strength M of the Earth's magnetic field measured by the magnetometer in the satellite's coordinate system is b , we can get the expected magnetic moment m of the three-axis magnetic torquer as:

[0116]

[0117] Where: M b1 ,M b2 ,M b3 M b The vector components of M b =(M b1 ,M b2 ,M b3 ) T ; T c1 ,T c2 ,T c3 are the vector components of Tc, and T c =(T c1 ,T c2 ,T c3 ) T , k mag Assign gain to the magnetron.

[0118] The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite of the present invention is described in detail below with reference to the accompanying drawings.

[0119] The magnetic control method of the optical micro-satellite single-axis sun-stabilized attitude of the present invention has the following control process: Figure 1 As shown, the specific steps include:

[0120] Firstly, an attitude dynamics model of an optical microsatellite using magnetic control is established for the design of the satellite's attitude controller.

[0121] Then, the expected single-axis attitude toward the sun is calculated based on the UTC time obtained by the onboard GPS.

[0122] Then, combined with the satellite's attitude dynamics model, the desired attitude quaternion is calculated, and an error attitude dynamics model using magnetic control is established; the desired torque of magnetic control is calculated using a proportional differential controller, and this controller ensures stable single-axis solar control through switching control.

[0123] Finally, according to the desired magnetic control torque calculated by the attitude controller, the desired magnetic moment of the magnetic torquer is calculated as the control instruction of the magnetic torquer. Similarly, in the process of calculating the desired magnetic moment, the switching control is designed according to the angular velocity of the satellite's solar axis to further enhance the stability of the single-axis solar control.

[0124] Table 1 Parameters of the embodiment

[0125]

[0126] The simulation parameters of this embodiment are shown in Table 1, which gives the satellite's moment of inertia parameters, magnetic torquer parameters and controller parameters. In order to fully illustrate the effectiveness and practicality of the method proposed in the present invention, the design of the embodiment specifically includes three parts, and the whole process takes a total of 5000s. Among them, the satellite is in angular velocity damping mode from 0 to 2300s, which is the first mode after the satellite is separated; from 2300 to 5000s, the satellite is in a single-axis solar stabilization mode, the purpose of which is to point the normal of the satellite solar sail panel to the sun to ensure the stability of the satellite energy.

[0127] The simulation results of this embodiment are as follows Figure 2-Figure 6 shown. Figure 2 This is the attitude quaternion curve diagram of the whole process. It can be observed from the figure that the attitude of the satellite is in a state of continuous change. Figure 3 The deviation quaternion curve of the whole process can be clearly observed from this figure that the satellite's X-axis attitude is in a continuous rotation change, while the Y-axis and Z-axis basically converge to 0 after entering the single-axis solar control mode at 2300, ensuring stable single-axis solar attitude control. Figure 4 The three-axis attitude angular velocity curve diagram of the whole process. It can be observed from the figure that after 2300s, the angular velocity of the X-axis is lower than 1° / s, which is in a low-speed spinning state and meets the pre-set angular velocity switching threshold. The angular velocity processing of the Y-axis and Z-axis is completely stable. Figure 5 This is the output magnetic torque curve of the magnetic torquer. Figure 6 This is the output torque curve of the magnetic torquer.

[0128] The magnetic control method for the single-axis stable attitude of an optical micro-satellite towards the sun of the present invention can realize the single-axis stable attitude control of the satellite towards the sun under the condition of extremely low power consumption, and can ensure the energy stability of the non-mission long-term mode and the reliable operation of the satellite.

[0129] 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 magnetic control method for a single-axis solar-stabilized attitude of an optical micro-satellite, characterized in that: The following steps are involved: Step 1, modeling the attitude dynamics of the optical microsatellite using magnetic control; Establish the attitude dynamics model of optical micro-satellite using magnetic control for the design of satellite attitude controller; Step 2, calculate the single-axis attitude toward the sun; Calculate the expected single-axis attitude toward the sun based on the UTC time obtained by the onboard GPS; Step 3, calculation of desired torque for attitude control; Combined with the satellite's attitude dynamics model, the desired attitude quaternion is calculated, and the error attitude dynamics model using magnetic control is established; the desired torque of magnetic control is calculated using a proportional differential controller; Step 4, calculating the expected magnetic moment of the magnetic torquer; According to the desired magnetic control torque calculated by the attitude controller, the desired magnetic moment of the magnetic torquer is calculated as the control instruction of the magnetic torquer.

2. The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite according to claim 1, characterized in that: In step 4, in the process of calculating the expected magnetic moment of the magnetic torquer, the switching control is designed according to the angular velocity of the satellite relative to the sun axis to further enhance the stability of the single-axis sun control.

3. The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite according to claim 1 or 2, characterized in that: Step 1 is as follows: The reference coordinate system for the attitude control of the optical microsatellite adopts the J2000 flat equatorial geocentric inertial coordinate system. As a reference coordinate system, the satellite's body coordinate system Each axis of rotation coincides with the principal axis of inertia; the body coordinate system Relative to the inertial coordinate system The attitude of is expressed using global non-singular quaternion as And satisfy the constraints The attitude angular velocity is expressed as The control torque of the magnetic torquer is expressed as The satellite attitude dynamics controlled by magnetic torquer is expressed as: Where: is a positive definite matrix, representing the moment of inertia of the satellite; Respectively represent the derivatives of Q, ω, q0, q; I3 represents the 3×3 unit matrix; is an antisymmetric matrix, for any vector Satisfies S(x)y=x×y, where × represents vector cross product; The satellite's expected attitude is defined as the attitude orientation of the body coordinate system relative to the inertial coordinate system, and is expressed by the expected attitude quaternion express; The attitude tracking error is defined as the error quaternion: Where: Represents quaternion multiplication; The angular velocity tracking error is: oh e =ω-R(Q e )oh d Where: d is the desired angular velocity of the spacecraft; the rotation matrix R(Q e ) has the following relationship: Where: S(q e ) indicates q e The antisymmetric matrix of ; Represents R(Q e ) and satisfy the constraint || R(Q e )||=1.

4. The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite according to claim 1 or 2, characterized in that: Step 2 is as follows: Set UTC time t utc Transformed into Julian time t JD for: t JD =t utc / 86400+2451545 According to the Julian time t JD , calculate the Julian century number T JD for: T JD =(t JD -2451545) / 36525 Obliquity of the ecliptic s By Julian century number T JD The calculation is: The geometrical mean ecliptic longitude L0 and the mean anomaly M of the sun are calculated from the Julian century number T JD The calculation is: According to the solar geometrical mean ecliptic longitude L0 and the solar mean anomaly angle M, the solar ecliptic longitude l is obtained. s for: According to the obliquity of the ecliptic s and the sun's ecliptic longitude s , the attitude quaternion of the earth pointing to the sun in the J2000 flat equatorial geocentric inertial coordinate system is calculated as: Since the attitude toward the sun is equivalent to the stability in inertial space, the expected angular velocity ω of the inertial coordinate system is d for: oh d =[0 0 0] T 。 5. The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite according to claim 1 or 2, characterized in that: Step 3 is as follows: The satellite's error attitude dynamics are: Where: is a positive definite matrix, representing the moment of inertia of the satellite; is an antisymmetric matrix, for any vector Satisfies S(x)y=x×y, where × represents vector cross product; is the attitude angular velocity; is the control torque of the magnetic torquer; Respectively represent Q e ,ω,q e0 ,q e The derivative of . The single-axis sun attitude controller is: T c =u b +u f Where: Where: K p is the proportional control gain; K d is the differential control gain; K f is the feedforward control gain; q e1 ,q e2 are the error quaternion q e The vector components of the attitude angular velocity ω; ω1, ω2, ω3 are the vector components of the attitude angular velocity ω; δ k (ω1) is the set self-stabilizing axis control gain and satisfies the following conditions: Where: T c is the desired control torque, is the angular velocity limit required for the self-stabilizing axis. b is a basic proportional-derivative controller, u f It is a nonlinear compensation controller.

6. The magnetic control method for the single-axis solar-stabilized attitude of an optical micro-satellite according to claim 5, characterized in that: Step 4 is as follows: The desired control torque T for maintaining single-axis sun stability control is calculated. c Then, the magnetic field strength M of the Earth's magnetic field measured by the magnetometer in the satellite's coordinate system is b , the expected magnetic moment m of the three-axis magnetic torquer is obtained as: Where: M b1 ,M b2 ,M b3 M b The vector components of M b =(M b1 ,M b2 ,M b3 ) T ; T c1 ,T c2 ,T c3 T c The vector components of T c =(T c1 ,T c2 ,T c3 ) T ;k mag Assign gain to the magnetron.

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