A magnetic control method for single-axis sun-pointing stable attitude of optical micro-satellite

Abstract

The present application relates to a kind of optical micro-satellite single-axis sun-stabilized attitude magnetic control method, it is related to aerospace technical field.The magnetic control method includes the following steps: step 1, the attitude dynamics of optical micro-satellite using magnetic control is modeled;Step 2, the attitude of single-axis sun is calculated;Step 3, attitude control expected moment calculation;Step 4, the expected magnetic moment calculation of magnetic moment device.The optical micro-satellite single-axis sun-stabilized attitude magnetic control method of the present application can realize the single-axis sun-stabilized attitude control of satellite under the condition of extremely low power consumption, can guarantee the energy stability of non-mission long-term mode and the reliable operation of satellite.

Description

Technical Field

[0001] The present 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 microsatellite. Background Art

[0002] With the advancement of microelectronics technology, the payloads and electronic components of optical remote sensing satellites are moving towards miniaturization, lightweighting, and high performance. The overall satellite design is also centered around the principles of low mass, low power consumption, and high performance. The actuators that provide control torque for optical microsatellites include a magnetic torquer and a reaction flywheel. In attitude stabilization control mode, the reaction flywheel is generally used to provide the three-axis torque for attitude stabilization control, while the magnetic torquer provides external control torque to unload excess angular momentum from the reaction flywheel. However, in the long-term operation mode of optical microsatellites, such as stabilizing attitude toward the Sun, this control method has certain drawbacks. The power consumption of the reaction flywheel can be more than ten times that of the magnetic torquer. Furthermore, a reaction flywheel failure can lead to satellite attitude instability, resulting in energy loss and compromising 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 microsatellite.

[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 microsatellite comprises the following steps:

[0006] Step 1: Model the attitude dynamics of the optical microsatellite using magnetic control;

[0007] Establish an attitude dynamics model of an optical microsatellite using magnetic control for the design of the satellite's 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 an error attitude dynamics model using magnetic control is established; the desired magnetic control torque is calculated using a proportional differential controller;

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

[0013] According to the magnetic control expected moment calculated by the attitude controller, an expected magnetic moment of the magnetic moment device is calculated as a control instruction of the magnetic moment device.

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

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

[0016] The reference coordinate system of the attitude control of the optical micro-satellite adopts the J2000 equatorial and terrestrial inertial coordinate system As a reference coordinate system, the body coordinate system of the satellite Each rotation axis coincides with the inertia principal axis; the body coordinate system Relative to the inertial coordinate system The attitude of the satellite is expressed by a global non-singular quaternion as And satisfies the constraint condition The attitude angular velocity is expressed as The control moment of the magnetic moment device is expressed as

[0017] The satellite attitude dynamics controlled by the magnetic moment device is expressed as:

[0018]

[0019] In the formula: is a positive definite matrix, representing the rotational inertia of the satellite; respectively represent the derivatives of Q, ω, q0 and q; I3 represents a 3*3 unit matrix; is an anti-symmetric matrix, for any vector x, Satisfies S(x)y=x×y, and × represents vector cross product;

[0020] The desired attitude of the satellite is defined as the attitude of the body coordinate system relative to the inertial coordinate system, which is expressed by the expected attitude quaternion The attitude tracking error is defined as the error quaternion:

[0021]

[0022] In the formula: represents quaternion multiplication;

[0023] The angular velocity tracking error is:

[0024] ω e =ω-R(Q e )ω d

[0025] In the formula: ωd Desired angular velocity of the spacecraft; rotation matrix R(Q e ) has the following relationship:

[0026]

[0027] In the formula, S(q e ) represents the skew-symmetric matrix of q e ; The derivative of R(Q e ) is represented; and the constraint condition ||R(Q e )||=1 is satisfied.

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

[0029] The UTC time t utc is converted into Julian time t JD as follows:

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

[0031] According to the Julian time t JD , the Julian century number T JD is calculated as follows:

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

[0033] The ecliptic-obiquity angle i s is calculated from the Julian century number T JD as follows:

[0034]

[0035] The solar geometric mean longitude L0 and the solar mean anomaly M can be calculated from the Julian century number T JD as follows:

[0036]

[0037] According to the solar geometric mean longitude L0 and the solar mean anomaly M, the solar longitude l s is obtained as follows:

[0038]

[0039] According to the ecliptic-obiquity angle i s and the solar longitude l s , the attitude quaternion of the Earth pointing to the Sun in the J2000 equatorial terrestrial inertial coordinate system is calculated as follows:

[0040]

[0041] Since the sun-pointing attitude is equivalent to the inertial space stabilization, the desired angular velocity ω d is:

[0042] ω d =

[000] T .

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

[0044] The error attitude dynamics of the satellite is:

[0045]

[0046] In the formula: is a positive definite matrix, representing the moment of inertia of the satellite; is an anti-symmetric matrix, for any vector x, satisfies S(x)y=x×y, and × represents vector cross product;

[0047] is the attitude angular velocity; is the control torque of the magnetic torque device; respectively represent the derivatives of Q e ,ω,q e0 and q e .

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

[0049] T c =u b +u f

[0050] In the formula:

[0051]

[0052] In the formula: K p is the proportional control gain; K d is the derivative control gain; K f is the feedforward control gain; q e1 , q e2 are the vector components of the error quaternion q e ; 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:

[0053]

[0054] In the formula: 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 is calculated to maintain the single-axis sun stability control. 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 for the single-axis solar-stabilized attitude of an optical microsatellite of the present invention has two main advantages over the prior art:

[0061] First, since the magnetic torquer can only provide a control torque in a direction perpendicular to the Earth's magnetic field, it cannot provide a stable three-axis control torque. This problem makes this magnetic control method unable to guarantee stable three-axis solar stabilization attitude control for a long time. The magnetic control method for the single-axis solar stabilization 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 solar stabilization attitude control. Compared with the three-axis solar stabilization of magnetic control. The magnetic control method for the single-axis solar stabilization attitude of the optical microsatellite of the present invention has higher solar accuracy, strong charging efficiency, and high system operation stability.

[0062] Second, the single-axis solar stabilization control mode does not require a high-power reaction flywheel. Instead, it uses low-power actuators and sensors such as a magnetic torquer, magnetometer, gyroscope, and star sensor. This significantly reduces the satellite's power consumption during long-term non-mission mode, implements low-power satellite design principles, and reduces satellite development and production costs. The magnetic control method for single-axis solar stabilization of an optical microsatellite of the present invention has a simple structure, is easy to implement, and has high reliability. It can significantly improve the robustness of optical satellite attitude control systems in actual engineering applications and has important engineering application value for low-cost microsatellite design.

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

[0064] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0065] Figure 1 Schematic diagram of 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 the 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 stabilization of an optical microsatellite. This method can achieve single-axis solar stabilization control under extremely low power consumption conditions and has important engineering application value.

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

[0074] The magnetic control method for the single-axis solar-stabilized attitude of an optical microsatellite of the present invention comprises the following steps:

[0075] Step 1: Model 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 posture 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; represents the derivative of Q,ω,q0,q; I3 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 direction of the body coordinate system relative to the inertial coordinate system, which is expressed by the expected attitude quaternion The attitude tracking error is defined as the error quaternion:

[0080]

[0081] Where: Represents quaternion multiplication.

[0082] The angular velocity tracking error is:

[0083] ω e =ω-R(Q e )ω d (3)

[0084] Where: ω 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 ) represents q e The antisymmetric matrix of ; Represents R(Qe ) and satisfy the constraint || R(Q e )||=1.

[0087] Step 2: Calculate the single-axis attitude toward the sun;

[0088] The direction of the solar attitude vector is the direction of the Earth pointing to the Sun. To calculate the single-axis solar attitude, the first step is to calculate the solar vector. In the J2000 flat equatorial geocentric inertial coordinate system, the solar direction vector r sun UTC time t according to 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 Julian time t JD , the Julian century number T can be calculated JD for:

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

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

[0093]

[0094] The solar geometrical mean ecliptic longitude L0 and the solar mean anomaly M can be obtained from the Julian century number T JD The calculation is:

[0095]

[0096] According to the solar geometric mean ecliptic longitude L0 and the solar mean anomaly angle M, the solar ecliptic longitude l s for:

[0097]

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

[0099]

[0100] Since the attitude toward the sun is equivalent to the stability of the inertial space, 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 and the expected attitude information relative to the sun calculated in the previous step, the satellite error attitude dynamics are 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 Earth's magnetic field, and in the single-axis solar attitude mode, the normal operation of the satellite requires that the normal vector of the solar panel coincide 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 panel coincides with the solar vector, and a small amount of control capability is used to ensure the attitude self-stabilization of the normal axis system of the solar panel. Therefore, the single-axis solar attitude controller is designed as follows:

[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 f is the feedforward control gain;

[0111] q e1 ,q e2 are the error quaternion q e The vector components of 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 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, ub is a basic proportional-derivative controller, u f It is a nonlinear compensation controller.

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

[0115] The desired control torque T is calculated to maintain the single-axis sun stability control. 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 microsatellite according to the present invention will be described in detail below with reference to the accompanying drawings.

[0119] The magnetic control method of the optical microsatellite single-axis solar stabilization 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, the desired magnetic torque of the magnetic torquer is calculated based on the desired magnetic control torque calculated by the attitude controller, which serves as the control instruction for the magnetic torquer. Similarly, during the calculation of the desired magnetic torque, the switching control design is based on the angular velocity of the satellite's solar axis, further enhancing the stability of single-axis solar control.

[0124] Table 1 Parameters related to the embodiment

[0125]

[0126] The simulation parameters for this embodiment are shown in Table 1, which provides the satellite's moment of inertia parameters, magnetoractuator parameters, and controller parameters. To fully demonstrate the effectiveness and practicality of the method proposed in this invention, the design of this embodiment specifically includes three parts, with the entire process lasting 5000 seconds. From 0 to 2300 seconds, the satellite is in angular velocity damping mode, the first mode after satellite separation. From 2300 to 5000 seconds, the satellite is in single-axis solar stabilization mode, aimed at pointing the normal of the satellite's solar panels toward the sun to ensure satellite energy stability.

[0127] The simulation results of this embodiment are as follows Figure 2-Figure 6 shown. Figure 2 The figure is the attitude quaternion curve 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 for the entire process shows that the satellite's X-axis attitude is in a continuous rotational change. However, after the Y-axis and Z-axis enter the single-axis solar control mode at 2300, the deviation quaternion basically converges to 0, ensuring stable single-axis solar attitude control. Figure 4 The following is a graph of the three-axis attitude angular velocity during the entire process. It can be observed from the graph that after 2300s, the angular velocity of the X-axis is less than 1° / s, indicating a low-speed spinning state. This also meets the preset angular velocity switching threshold, and 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 solar-stabilized attitude of an optical microsatellite of the present invention can realize the single-axis solar-stabilized attitude control of the satellite under the condition of extremely low power consumption, and can ensure the energy stability and reliable operation of the satellite in the non-mission long-term mode.

[0129] Obviously, the above embodiments are merely example for clearly illustrating but not limitation to the embodiments. Based on the above description, other different forms of changes or variations can be made by those skilled in the art. Here, all the embodiments need not and can not be enumerated. The obvious changes or variations derived from the above description are still within the protection scope of the present application.

Claims

1. A magnetic control method for a single-axis solar-stabilized attitude of an optical microsatellite, characterized in that: The following steps are involved: Step 1: Model the attitude dynamics of the optical microsatellite using magnetic control; Establish an attitude dynamics model of an optical microsatellite using magnetic control for the design of the satellite's 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; According to UTC time t utc Obliquity of the ecliptic i s and the sun's ecliptic longitude s , according to the obliquity of the ecliptic i 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 towards the sun is equivalent to the stability of the inertial space, the expected angular velocity ω of the inertial coordinate system is d for: oh d =[000] T Step 3, calculation of desired torque for attitude control; 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 magnetic control torque is calculated using a proportional differential controller; Step 4, calculation of 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; 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 can be expressed as: Where: is a positive definite matrix, representing the moment of inertia of the satellite; Represent the derivatives of Q, ω, q0, q respectively; I3 represents the 3×3 identity 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 direction of the body coordinate system relative to the inertial coordinate system, which is expressed by the expected attitude quaternion express; The pose 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: ω d is the desired angular velocity of the spacecraft; the rotation matrix R(Q e ) has the following relationship: Where: S(q e ) represents q e The antisymmetric matrix of ; Represents R(Q e ) and satisfy the constraint || R(Q e )||=1; 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; Represents 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 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 meets the following conditions: 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 is a nonlinear compensation controller; Step 4 is as follows: The desired control torque T is calculated to maintain the single-axis sun stability control. 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.

2. The magnetic control method for the single-axis solar stabilization of an optical microsatellite according to claim 1, characterized in that: In step 4, during the calculation of the desired magnetic moment of the magnetic torquer, the switching control is designed according to the angular velocity of the satellite relative to the solar axis to further enhance the stability of the single-axis solar control.

3. The magnetic control method for the single-axis solar-stabilized attitude of an optical microsatellite 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 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 solar geometric mean ecliptic longitude L0 and the solar mean anomaly M are calculated from the Julian century number T JD The calculation is: According to the solar geometric mean ecliptic longitude L0 and the solar mean anomaly 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 towards the sun is equivalent to the stability of the inertial space, the expected angular velocity ω of the inertial coordinate system is d for: oh d =[000] T 。

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