CN115817856B - Method and device for controlling stable attitude of satellite to solar spin based on pure magnetic control mode - Google Patents
Method and device for controlling stable attitude of satellite to solar spin based on pure magnetic control mode Download PDFInfo
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Abstract
The invention provides a method and a device for controlling the stable attitude of a satellite to a solar spin based on a pure magnetic control mode, and belongs to the technical field of spacecraft attitude control. Wherein the method comprises the following steps: calculating an estimated value of the satellite three-axis angular rate, and respectively calculating target output torque of the satellite in each control stage of the solar spin according to the estimated value of the satellite three-axis angular rate, wherein the solar vector measured by the solar sensor is considered in the target output torque in the solar capture stage, the spin-up stage and the solar spin stabilization stage when the sun is positioned in the measurement view field of the solar sensor; and calculating the target output magnetic moment of the triaxial magnetic torquer according to the target output moment and the geomagnetic vector measurement value, and converting the target output magnetic moment into a control instruction so as to realize satellite attitude control. According to the solar spin stabilizing device, the influence of factors such as a ground shadow area, offset installation of a sun sensor and a solar sailboard, a field of view range of the sun sensor, environmental interference moment and the like is fully considered, and the satellite can be controlled to finally realize the spin stabilizing from any initial state.
Description
Technical Field
The invention belongs to the technical field of spacecraft attitude control, and particularly relates to a method and a device for controlling a stable attitude of a satellite on a solar spin based on a pure magnetic control mode.
Background
Typically, to ensure on-board system power supply, the satellites are in a sun-to-sun orientation for the majority of the time in orbit. Under the sun-facing orientation state, the satellite continuously adjusts the satellite posture by means of the posture control system, so that the solar sailboard is always aligned to the solar vector. The sun-oriented state is critical to the supply of on-board energy, and whether a stable and reliable sun-oriented posture can be established directly determines the success or failure of satellite flight tasks. Therefore, the research on how to adopt the minimum system to realize the control of the sun orientation gesture is significant, especially for satellites with limited configuration on the satellite and working faults of devices.
In 2007, researchers such as Luo Jungong design a spin stabilization attitude control method for a solar sail spacecraft, discuss a spin speed obtaining mode, and perform simulation analysis on attitude control characteristics of the solar sail spacecraft, and simulation results show that the solar sail spacecraft can realize uniaxial pointing through spin. In 2019, researchers such as Xia Xiwang select a sun sensor and a magnetometer as gesture sensors and a magnetic torquer as gesture execution mechanisms, and put forward a pure magnetic control spin-to-sun orientation method and a correction method thereof, so that stable gesture control of satellites on the spin-to-sun is realized. In 2020, researchers such as Liu Shanwu put forward a solar capturing attitude control method based on extremely simple system configuration aiming at a satellite running on a morning and evening solar synchronous orbit, and the method fully utilizes the satellite orbit characteristics, selects a magnetometer and a magnetic torquer as an attitude sensor and an attitude actuator respectively, and realizes the stable attitude control of the satellite on the solar spin by adopting a pitching axis spinning method.
However, in the existing sun-facing orientation attitude control method, the influences of factors such as a ground shadow area, offset installation of a sun sensor and a solar sailboard, a field of view range of the sun sensor, an environmental interference moment and the like in the actual in-orbit operation process of a satellite are not fully considered, so that the application range is limited, and the method is difficult to apply to engineering practice. For most satellites, the satellites inevitably enter an earth shadow area in the process of orbit running, and the output of a sun sensor is invalid, so that a new attitude control method is required to be designed to meet the requirements of different orbit satellites on the daily directional attitude control.
Disclosure of Invention
The invention aims to overcome the defects of the prior art and provides a method and a device for controlling the stable attitude of a satellite to a solar spin based on a pure magnetic control mode. Based on a pure magnetic control mode, under the condition of fully considering the influence of factors such as a ground shadow area, offset installation of a sun sensor and a solar sailboard, a field of view range of the sun sensor, environmental interference moment and the like, the solar spin stabilizing device can control satellites to finally realize the solar spin stabilizing from any initial state, can effectively ensure smooth implementation of flight tasks such as satellite equipment debugging, energy acquisition and the like, and has the advantages of low cost, low power consumption, simple flow, high reliability, good stability and the like, and is suitable for most earth orbit satellites.
An embodiment of a first aspect of the present invention provides a method for controlling a stable attitude of a satellite to a solar spin based on a purely magnetic control mode, including:
calculating an estimated value of the satellite three-axis angular rate;
according to the estimated value of the satellite three-axis angular rate, respectively calculating the target output moment of the satellite in each control stage of the solar spin, wherein each control stage of the solar spin comprises the following steps: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
calculating the target output magnetic moment of the triaxial magnetometers of the satellites according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellites;
and converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite.
In a specific embodiment of the present invention, the method for calculating the estimated value of the satellite three-axis angular rate is as follows:
(1) Constructing a state vector
wherein ,
the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega x 、ω y 、ω z Respectively->
In the satellite body coordinate system O b X b Y b Z b O of (2) b X b 、O b Y b 、O b Z b An axial component;
(2) Constructing a state equation and a state transition matrix;
the satellite attitude dynamics equation is constructed as follows:
wherein ,
is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction; />
The control moment is output by the triaxial magnetic torquer;
the attitude dynamics equation is a state equation, and is linearized to obtain a state transition matrix
Wherein, deltaT is the operation period;
(3) Constructing a measurement equation and a measurement matrix;
the triaxial magnetometer is measured twice in successionThe measured geomagnetic vector values are respectively marked as B b And
wherein ,
for the geomagnetic vector measurement value at the current moment, +.>
Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component; />
For the geomagnetic vector measurement value at the previous moment, +.>
Respectively, the geomagnetic vector measured value at the previous moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component;
construction of an observation vector
The following are provided:
order the
To represent the vector of the rotation angle of the satellite body coordinate system in the time period delta T, theta Δ Corresponding direction cosine matrix->
The approximation is expressed as:
wherein ,
for vector->
Is expressed as follows:
then a measurement equation is constructed as follows:
wherein ,
is an error term;
obtaining a measurement matrix according to the measurement equation
The method comprises the following steps:
(4) According to the state transition matrix obtained in the step (2) and the measurement matrix obtained in the step (3), calculating to obtain an estimated value of the satellite triaxial angular rate by using a Kalman filtering method
In a specific embodiment of the present invention, the method further comprises:
the target output torque calculation expression is as follows, either during the initial racemization phase, or during the pair-day acquisition phase, before the sun enters the sun sensor measurement field of view of the satellite, or during the pair-day acquisition phase, when the satellite is located in the ground shadow region:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 1 A positive first control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 1 Is vector B b and Tc1 An included angle between the two.
In a specific embodiment of the present invention, the method further comprises:
in the initial racemization stage, if the angular rate value of the satellite triaxial angular rate in the time greater than or equal to the set first time threshold value is always smaller than or equal to the set first angular rate threshold value, ending the initial racemization stage, and enabling the satellite to enter a daily capturing stage.
In a specific embodiment of the present invention, the method further comprises:
during the pair-day capture phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 2 、K 3 、K 4 Positive second, third and fourth control coefficients respectively;
the unit vector is the optical axis direction of the sun sensor; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 2 Is vector B b and Tc2 An included angle between the two.
In a specific embodiment of the present invention, the method further comprises:
in the sun-to-sun capturing stage, if the unit vector of the optical axis direction of the sun sensor
Solar vector S measured by sun sensor at current moment b Is always in a time greater than or equal to the second time thresholdAnd if the satellite is smaller than or equal to the set first angle threshold value, ending the day capturing stage, and enabling the satellite to enter a spinning stage.
In a specific embodiment of the present invention, the method further comprises:
during the spinning phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 3 Is vector B b and Tc3 An included angle between the two.
In a specific embodiment of the present invention, the method further comprises:
in the spinning stage, if the satellite flies into an earth shadow area, returning the satellite to the sun-to-sun capturing stage;
if the normal vector of the satellite star surface which is required to be aligned with the sun vector is set
Solar vector S measured by sun sensor at current moment b The included angle is always smaller than or equal to the set second angle threshold value within the time larger than or equal to the set third time threshold value, and the satellite triaxial angular rate estimated value is +.>
And the set satellite spin angular rate vector +.>
And if the deviation is always smaller than or equal to the set second angular rate threshold value within the same time, ending the spinning stage, and enabling the satellite to enter a phase of stabilizing the spinning.
In a specific embodiment of the present invention, the method further comprises:
calculating a target output torque in the para-Japanese spin stabilization stage;
Wherein, when the satellite is in the sun-shine region, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
when the satellite is in the ground shadow region, the target output torque calculation expression is as follows:
wherein ,K9 A positive ninth control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 4 Is vector B b and Tc4 An included angle between the two.
An embodiment of a second aspect of the present invention provides a satellite-to-solar spin stabilization attitude control device based on a purely magnetic control method, including:
the triaxial angular rate estimation module is used for calculating an estimated value of the triaxial angular rate of the satellite;
The target output torque calculation module is used for calculating target output torque of the satellite in each control stage of the solar spin according to the estimated value of the satellite three-axis angular rate, and the each control stage of the solar spin comprises: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
the target output magnetic moment calculation module calculates the target output magnetic moment of the triaxial magnetometers of the satellite according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellite;
and the attitude control instruction calculation module is used for converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite.
An embodiment of a third aspect of the present invention provides an electronic device, including:
at least one processor; and a memory communicatively coupled to the at least one processor;
The memory stores instructions executable by the at least one processor, the instructions configured to perform a satellite-to-solar spin stabilization attitude control method based on purely magnetic control.
An embodiment of the fourth aspect of the present invention provides a computer-readable storage medium storing computer instructions for causing a computer to execute the above-described method for controlling a steady attitude of a satellite to a solar spin based on a purely magnetic control method
The invention has the characteristics and beneficial effects that:
based on a pure magnetic control mode, under the condition of fully considering the influence of factors such as a ground shadow area, offset installation of a sun sensor and a solar sailboard, a field of view range of the sun sensor, environmental interference moment and the like, the three-axis magnetometer and the sun sensor are adopted as gesture sensitive devices, and the three-axis magneto torquer is adopted as gesture executing devices, so that the satellite can be controlled to finally realize the stability to the daily spin from any initial state.
The invention uses the measured value of the triaxial magnetometer as input to calculate and obtain the triaxial angular rate of the satellite, combines the measured value of the solar sensor to design a gesture control scheme comprising control stages of initial racemization, solar capture, spin-up, solar spin stabilization and the like, and realizes control moment output by using the triaxial magnetometers. The invention can effectively ensure smooth implementation of flight tasks such as satellite equipment debugging, energy acquisition and the like, has the advantages of low cost, low power consumption, simple flow, high reliability, good stability and the like, and can meet the requirements of different orbit satellites on daily directional attitude control.
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Fig. 1 is an overall flowchart of a satellite-to-solar spin stable attitude control method based on a purely magnetic control mode according to an embodiment of the invention.
Detailed Description
The invention provides a method and a device for controlling the stable attitude of a satellite to a solar spin based on a pure magnetic control mode, which are further described in detail below with reference to the accompanying drawings and specific embodiments.
An embodiment of a first aspect of the present invention provides a method for controlling a stable attitude of a satellite to a solar spin based on a purely magnetic control mode, including:
calculating an estimated value of the satellite three-axis angular rate;
according to the estimated value of the satellite three-axis angular rate, respectively calculating the target output moment of the satellite in each control stage of the solar spin, wherein each control stage of the solar spin comprises the following steps: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
calculating a target output magnetic moment of the triaxial magnetometers according to the target output moment and geomagnetic vector measurement values obtained by triaxial magnetometers of the satellites;
And converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite.
In one embodiment of the invention, the satellite is set to run in 10:30 sun-synchronous orbit at an orbit height of 545km at the point of intersection. Satellite triaxial moment of inertia [ 1.37.1.69.2.05 ]]kg·m 2 The satellite residual magnetism is 0.012 A.m 2 The satellite surface reflection coefficient is 0.6. Setting the measurement error of the triaxial magnetometer to be 3×10 -4 μT; setting an optical axis of a sun sensor and a satellite star O b X b 、O b Y b 、O b Z b The included angles of the axes are 69.2952 degrees, 150 degrees and 110.7048 degrees respectively, the view field range of the solar sensor is 90 degrees multiplied by 90 degrees, and the two-axis precision is 0.1 degrees; the maximum output magnetic moment of the triaxial magnetic torquer is set to be [ 2.52.3.21.2.52 ]]A·m 2 . Setting the initial angular rate of satellite triaxial at [ -1.5-1.5-1.5]The initial posture triaxial Euler angle is [160 20 60 ]]°。
In this embodiment, the method for controlling the stable attitude of the satellite to the solar spin based on the pure magnetic control mode has an overall flow shown in fig. 1, and includes the following steps:
(1) Calculating an estimated value of the satellite three-axis angular rate;
in this embodiment, an extended Kalman filtering method (ExtendedKalmanFilter, EKF) is used to estimate the satellite triaxial angular rate using the triaxial magnetometer measurements as input.
(1-1) constructing a state vector;
in this embodiment, the state vector
The definition is as follows:
wherein ,
the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega x 、ω y 、ω z Respectively->
At O b X b Y b Z b O of (2) b X b 、O b Y b 、O b Z b An axial component.
(1-2) constructing a state equation and a state transition matrix;
in this embodiment, because the requirement on the accuracy of the angular velocity estimation is not high, factors such as the space disturbance moment in the satellite attitude dynamics equation are ignored, and the satellite attitude dynamics equation is constructed as follows:
wherein ,
is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix with a centroid inertial principal axis coordinate system, the diagonal element J is x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction; />
The control moment is output by the triaxial magnetic torquer.
The attitude dynamics equation is a state equation, and the state transition matrix is obtained by linearizing the equation
Wherein, deltaT is the set operation period.
(1-3) constructing a measurement equation and a measurement matrix;
geomagnetic vector measured values obtained by two continuous measurements of the triaxial magnetometer are respectively recorded as B b And
wherein ,
for the geomagnetic vector measurement value at the current moment, +.>
Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component; />
For the geomagnetic vector measurement value at the previous moment, +.>
Respectively, the geomagnetic vector measured value at the previous moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component.
Construction of an observation vector
The difference between two adjacent measurements for a triaxial magnetometer:
in the case where Δt is relatively small, the change in the geomagnetic field vector in a short time can be ignored, and it is approximately considered that the difference between the two measurements before and after the magnetometer is caused only by the satellite attitude change. Let the triaxial angular rate of the satellite be
For the vector representing the rotation angle of the satellite body coordinate system in the time period delta T, under the assumption of small angle, theta Δ Corresponding direction cosine matrix->
Can be approximated as:
wherein ,
for vector->
Is expressed as follows:
further, a measurement equation was constructed as follows:
wherein ,
is an error term.
According to the measurement equation, a measurement matrix can be obtained
The method comprises the following steps:
(1-4) kalman filtering;
according to the state transition matrix obtained in the step (1-2) and the measurement matrix obtained in the step (1-3), obtaining the estimated value of the satellite three-axis angular rate through solution according to the general algorithm flow of the extended Kalman filtering
The method is used for controlling each stage of the gesture controller.
(2) Carrying out attitude control on each control stage of the solar spin by using the satellite triaxial angular rate estimated value obtained in the step 1);
The satellite comprises the following control stages of spin: an initial racemization stage, a daily capture stage, a spinning-up stage and a daily spin stabilization stage; the method comprises the following specific steps:
(2-1) an initial racemization stage;
when the satellite is separated from the carrier or attitude control is not performed for a long time, the satellite may have initial angular momentum, and initial racemization is required first.
(2-1-1) using the satellite triaxial angular rate estimation value obtained in the step (1)
The control law is designed as follows to obtain the target output force of the triaxial magnetic torquer in the initial racemization stageMoment T c1 ：
wherein ,K1 A positive first control coefficient (reference value of 0.01 in this embodiment).
The control moment T c1 The first term on the right of the equation is used to damp the triaxial angular rate and the second term is used to cancel the satellite triaxial angular rate coupling effect.
(2-1-2) geomagnetic vector measurement value B at the present time measured by means of a triaxial magnetometer b Combining the target output torque T obtained in the step (2-1-1) c1 Calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,α1 Is vector B b and Tc1 An included angle between the two.
And (2-1-3) taking the triaxial magnetic torquer as a control component, converting the target output magnetic moment calculated in the step (2-1-2) into a control instruction of the triaxial magnetic torquer, and driving the triaxial magnetic torquer to control the satellite attitude.
(2-1-4) in the initial racemization stage, if the angular rate value of the satellite in the time period of the triaxial angular rate being greater than or equal to the set first time threshold (the reference value of the embodiment is 100 s) is always smaller than or equal to the set first angular rate threshold (the reference value of the embodiment is 0.2 DEG/s), the initial racemization stage is ended, and the satellite enters the step (2-2) for the daily acquisition stage.
(2-2) a daily capture phase;
after the initial racemization of the satellite is completed, the satellite enters a sun-to-sun capturing stage, a sun vector is captured by a sun sensor, and the unit vector of the optical axis direction of the sun sensor is overlapped with the sun vector through control.
(2-2-1) considering that only a triaxial magnetic torquer is adopted as a satellite of a posture execution device before the sun enters a measurement view field of a sun sensor, only a control moment perpendicular to a local geomagnetic vector of the satellite can be output, and the posture movement capability of the satellite is poor, so that the sun is slowly searched by adopting a control law of an initial racemization stage until the sun enters the measurement view field of the sun sensor.
(2-2-2) when the sun enters the measuring field of view of the sun sensor, the sun sensor can obtain the sun vector S by measurement b . The control law is designed to obtain the target output torque T of the triaxial magnetic torquer during the sun capturing stage when the sun is positioned in the measurement view field of the sun sensor of the satellite c2 Aligning the optical axis of the sun sensor with the sun vector direction:
wherein ,K2 、K 3 、K 4 Positive second, third and fourth control coefficients (reference values of the embodiment are 0.0005, 0.001 and 0.02 respectively);
the unit vector is the unit vector of the optical axis direction of the sun sensor, and is determined by the installation angle of the sun sensor on the star; />
The sun vector measured by the sun sensor at the previous moment. The control moment T c2 The first term on the right of the equation is used to control +.>
And S is equal to b Coinciding, the second term and the third term are used for damping the satellite triaxial angular velocity and S b The vertical component, the fourth term, is used to cancel the satellite triaxial angular rate coupling effect. The field of view of the sun sensor is assumed to be less than or equal to 90 degrees in the invention.
It should be noted that if the field of view of the sun sensor is greater than 90 °, the field of view may be determined according to the vector
Sum vector S b The first term in the above equation is adaptively modified based on a custom positive correlation function, step function, clipping function, etc.
In the present embodiment, the torque T is controlled c2 Is characterized by designing a satellite triaxial angular velocity medium and S b The vertical component carries out damping control quantity and eliminates the control quantity of the triaxial angular rate coupling effect of the satellite, so that the satellite can stably and accurately align the optical axis of the sun sensor with the direction of the sun vector under the underactuated condition.
(2-2-3) geomagnetic vector measurement value B at the present time measured by means of a triaxial magnetometer b Combining the target output torque T obtained in the step (2-2-2) c2 Calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,α2 Is vector B b and Tc2 An included angle between the two.
And (2-2-4) taking the triaxial magnetic torquer as a control component, converting the target output magnetic moment calculated in the step (2-2-3) into a control instruction of the triaxial magnetic torquer, and driving the triaxial magnetic torquer to control the satellite attitude.
(2-2-5) in the phase of capturing sun, if the sun sensor has a unit vector of the optical axis direction
With its measured solar vector S b The included angle of the satellite is always smaller than or equal to a set first angle threshold (the reference value of the embodiment is 10 degrees) within the time of being larger than or equal to a second time threshold (the reference value of the embodiment is 25 s), the day capturing stage is ended, and the satellite enters the spinning stage in the step (2-3); otherwise, continuing to control the attitude of the day capturing stage according to the steps (2-2-2) - (2-2-4).
And (2) returning to the step (2-2-1) again if the satellite flies into the ground shadow area in the attitude control process of the daily capturing stage.
(2-3) a spinning stage;
the satellite enters a spinning stage after the sun is captured, and on the basis that the optical axis of the sun sensor is roughly aligned with the sun vector, the normal vector of the satellite body surface of the designated satellite is further overlapped with the sun vector, and the satellite is controlled to spin around the normal vector of the satellite body surface.
(2-3-1) in the spinning stage and when the sun is located in the measurement field of view of the sun sensor of the satellite, designing the following control law to obtain the target output torque T of the triaxial magnetic torquer in the spinning stage c3 The normal vector of the satellite plane of the appointed satellite is overlapped with the sun vector, and the satellite plane is rotated around the normal vector of the satellite plane:
wherein ,K5 、K 6 、K 7 、K 8 Fifth, sixth, seventh, and eighth control coefficients respectively positive (the reference values of this embodiment are respectively 0.0004, 0.0008, 0.01, and 0.004);
a normal unit vector of a satellite body plane which is required to be aligned with a sun vector is specified; omega spin The spin angle rate value is set (1 DEG/s is a reference value in this example). The control moment T c3 The first term on the right of the equation is used to control +.>
And S is equal to b Coinciding, the second term and the third term are used for damping the satellite triaxial angular velocity and S b The vertical component, the fourth term, is used for spin-up and the fifth term is used to eliminate the satellite triaxial angular rate coupling effect.
It should be noted that, the advantage of this embodiment is that the control scheme of making the satellite sun sensor stably align with the sun vector first and then making the normal vector of the designated satellite star surface align with the sun vector is designed, so that the reliability and stability of the satellite for realizing the stable control of the solar spin by measuring the sun vector by the sun sensor can be significantly improved under the condition of limited field of view and under-actuated control of the sun sensor.
(2-3-2) geomagnetic vector measurement value B at the present time measured by means of a triaxial magnetometer b Combining the target output torque T obtained in the step (2-3-1) c3 Calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,α3 Is vector B b and Tc3 An included angle between the two.
And (2-3-3) taking the triaxial magnetic torquer as a control component, converting the target output magnetic moment calculated in the step (2-3-2) into a control instruction of the triaxial magnetic torquer, and driving the triaxial magnetic torquer to control the satellite attitude.
(2-3-4) during the spinning phase, when the normal vector of the satellite star surface is specified
Solar vector S measured by sun sensor b The included angle is always smaller than or equal to the set second angle threshold (10 ° in the present embodiment) within the time greater than or equal to the set third time threshold (25 s in the present embodiment), and the estimated value of the satellite triaxial angular rate>
And the set satellite spin angular rate vector +.>
The deviation between the two phases is always smaller than or equal to a set second angular rate threshold value (the reference value of the embodiment is 0.1 degrees/s) in the same time, the spinning stage is ended, and the satellite enters a phase of stabilizing the daily spin in the step (2-4); otherwise, according to the steps of(2-3-1) - (2-3-3) continuing the attitude control of the spinning stage.
And (2) returning to the step (2-2-1) again if the satellite flies into the ground shadow area in the gesture control process of the spinning stage.
(2-4) a phase of stabilization to the daily spin;
after the solar spin stabilization stage is entered, the satellite has certain axiality along the normal vector direction of the designated satellite body plane, so that the designated satellite body plane still points to the sun vector when the satellite enters the earth shadow area, and the controlled attitude is maintained.
(2-4-1) when the satellite is in the sunshine zone, designing the following control law to obtain the target output torque T of the triaxial magnetic torquer at the stable phase of the solar spin c4 The normal vector of the satellite star surface is overlapped with the sun vector, and the satellite star surface spins around the normal vector of the star surface:
wherein ,K5 、K 6 、K 7 、K 8 Fifth, sixth, seventh, and eighth control coefficients respectively positive (the reference values of this embodiment are respectively 0.0004, 0.0008, 0.01, and 0.004);
a normal unit vector of a satellite body plane which is required to be aligned with a sun vector is specified; omega spin The spin angle rate value is set (1 DEG/s is a reference value in this example). The control moment T c4 The first term on the right of the equation is used to control +.>
And S is equal to b Coinciding, the second term and the third term are used for damping the satellite triaxial angular velocity and S b The vertical component, the fourth term, is used for spin-up and the fifth term is used to eliminate the satellite triaxial angular rate coupling effect.
When the satellite is in the earth shadow region, the following control law is designed to obtain the phase three of the stable phase of the solar spinTarget output torque T of shaft magnetic torquer c4 To control the satellite to maintain a spinning state:
wherein ,K9 A positive ninth control coefficient (reference value of 0.01 in this embodiment). At the control moment T c4 The first term on the right of the equation is used to control the satellite to maintain the spin angular rate and the second term is used to eliminate the satellite triaxial angular rate coupling effect.
(2-4-2) geomagnetic vector measurement value B at the present time measured by means of a triaxial magnetometer b Combining the target output torque T obtained in the step (2-4-1) c4 Calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,α4 Is vector B b and Tc4 An included angle between the two.
And (2-4-3) taking the triaxial magnetic torquer as a control component, converting the target output magnetic moment calculated in the step (2-4-2) into a control instruction of the triaxial magnetic torquer, and driving the triaxial magnetic torquer to control the satellite attitude.
In order to achieve the above embodiments, a second aspect of the present invention provides a satellite-to-solar spin stabilization attitude control device based on a purely magnetic control method, including:
the triaxial angular rate estimation module is used for calculating an estimated value of the triaxial angular rate of the satellite;
The target output torque calculation module is used for calculating target output torque of the satellite in each control stage of the solar spin according to the estimated value of the satellite three-axis angular rate, and the each control stage of the solar spin comprises: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
the target output magnetic moment calculation module calculates the target output magnetic moment of the triaxial magnetometers of the satellite according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellite;
and the attitude control instruction calculation module is used for converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite.
It should be noted that the foregoing explanation of the embodiment of the method for controlling the steady state of the satellite to the daily spin based on the purely magnetic control method is also applicable to the device for controlling the steady state of the satellite to the daily spin based on the purely magnetic control method in this embodiment, and will not be repeated herein. According to the satellite-to-solar spin stable attitude control device based on the purely magnetic control mode, an estimated value of the three-axis angular rate of a satellite is calculated; according to the estimated value of the satellite three-axis angular rate, respectively calculating the target output moment of the satellite in each control stage of the solar spin, wherein each control stage of the solar spin comprises the following steps: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor; calculating a target output magnetic moment of the triaxial magnetometers according to the target output moment and geomagnetic vector measurement values obtained by triaxial magnetometers of the satellites; and converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite. Based on the pure magnetic control mode, under the condition of fully considering factors such as the ground shadow area, the offset installation of the sun sensor and the solar sailboard, the field of view range of the sun sensor, the environmental interference moment and the like, the satellite can be controlled to finally realize the stability of the solar spin from any initial state, the smooth implementation of flight tasks such as satellite equipment debugging, energy acquisition and the like can be effectively ensured, and the method has the advantages of low cost, low power consumption, simple flow, high reliability, good stability and the like, and is suitable for most earth orbit satellites.
To achieve the above embodiments, an embodiment of a third aspect of the present invention provides an electronic device, including:
at least one processor; and a memory communicatively coupled to the at least one processor;
the memory stores instructions executable by the at least one processor, the instructions configured to perform a satellite-to-solar spin stabilization attitude control method based on purely magnetic control.
To achieve the above embodiments, a fourth aspect of the present invention provides a computer-readable storage medium storing computer instructions for causing the computer to execute the above method for controlling a steady attitude of a satellite to a solar spin based on a purely magnetic control method.
It should be noted that the computer readable medium described in the present disclosure may be a computer readable signal medium or a computer readable storage medium, or any combination of the two. The computer readable storage medium can be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or a combination of any of the foregoing. More specific examples of the computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer diskette, a hard disk, a Random Access Memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing. In the context of this disclosure, a computer-readable storage medium may be any tangible medium that can contain, or store a program for use by or in connection with an instruction execution system, apparatus, or device. In the present disclosure, however, the computer-readable signal medium may include a data signal propagated in baseband or as part of a carrier wave, with the computer-readable program code embodied therein. Such a propagated data signal may take any of a variety of forms, including, but not limited to, electro-magnetic, optical, or any suitable combination of the foregoing. A computer readable signal medium may also be any computer readable medium that is not a computer readable storage medium and that can communicate, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device. Program code embodied on a computer readable medium may be transmitted using any appropriate medium, including but not limited to: electrical wires, fiber optic cables, RF (radio frequency), and the like, or any suitable combination of the foregoing.
The computer readable medium may be contained in the electronic device; or may exist alone without being incorporated into the electronic device. The computer readable medium carries one or more programs which, when executed by the electronic device, cause the electronic device to perform a method for controlling the steady attitude of a satellite relative to a solar spin according to the above embodiment.
Computer program code for carrying out operations of the present disclosure may be written in one or more programming languages, including an object oriented programming language such as Java, smalltalk, C ++ and conventional procedural programming languages, such as the "C" programming language or similar programming languages. The program code may execute entirely on the user's computer, partly on the user's computer, as a stand-alone software package, partly on the user's computer and partly on a remote computer or entirely on the remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any kind of network, including a Local Area Network (LAN) or a Wide Area Network (WAN), or may be connected to an external computer (for example, through the Internet using an Internet service provider).
In the description of the present specification, a description referring to terms "one embodiment," "some embodiments," "examples," "specific examples," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present application. In this specification, schematic representations of the above terms are not necessarily directed to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, the different embodiments or examples described in this specification and the features of the different embodiments or examples may be combined and combined by those skilled in the art without contradiction.
Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include at least one such feature. In the description of the present application, the meaning of "plurality" is at least two, such as two, three, etc., unless explicitly defined otherwise.
Any process or method descriptions in flow charts or otherwise described herein may be understood as representing modules, segments, or portions of code which include one or more executable instructions for implementing specific logical functions or steps of the process, and further implementations are included within the scope of the preferred embodiment of the present application in which functions may be executed out of order from that shown or discussed, including substantially concurrently or in reverse order, depending on the functionality involved, as would be understood by those reasonably skilled in the art of the embodiments of the present application.
Logic and/or steps represented in the flowcharts or otherwise described herein, e.g., a ordered listing of executable instructions for implementing logical functions, can be embodied in any computer-readable medium for use by or in connection with an instruction execution system, apparatus, or device, such as a computer-based system, processor-containing system, or other system that can fetch the instructions from the instruction execution system, apparatus, or device and execute the instructions. For the purposes of this description, a "computer-readable medium" can be any means that can contain, store, communicate, propagate, or transport the program for use by or in connection with the instruction execution system, apparatus, or device. More specific examples (a non-exhaustive list) of the computer-readable medium would include the following: an electrical connection (electronic device) having one or more wires, a portable computer diskette (magnetic device), a Random Access Memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber device, and a portable compact disc read-only memory (CDROM). Additionally, the computer-readable medium may even be paper or other suitable medium upon which the program is printed, as the program may be electronically captured, via, for instance, optical scanning of the paper or other medium, then compiled, interpreted or otherwise processed in a suitable manner, if necessary, and then stored in a computer memory.
It is to be understood that portions of the present application may be implemented in hardware, software, firmware, or a combination thereof. In the above-described embodiments, the various steps or methods may be implemented in software or firmware stored in a memory and executed by a suitable instruction execution system. For example, if implemented in hardware, as in another embodiment, may be implemented using any one or combination of the following techniques, as is well known in the art: discrete logic circuits having logic gates for implementing logic functions on data signals, application specific integrated circuits having suitable combinational logic gates, programmable Gate Arrays (PGAs), field Programmable Gate Arrays (FPGAs), and the like.
Those of ordinary skill in the art will appreciate that all or a portion of the steps carried out in the method of the above-described embodiments may be implemented by a program to instruct related hardware, where the program may be stored in a computer readable storage medium, and where the program, when executed, includes one or a combination of the steps of the method embodiments.
In addition, each functional unit in each embodiment of the present application may be integrated in one processing module, or each unit may exist alone physically, or two or more units may be integrated in one module. The integrated modules may be implemented in hardware or in software functional modules. The integrated modules may also be stored in a computer readable storage medium if implemented as software functional modules and sold or used as a stand-alone product.
The above-mentioned storage medium may be a read-only memory, a magnetic disk or an optical disk, or the like. Although embodiments of the present application have been shown and described above, it will be understood that the above embodiments are illustrative and not to be construed as limiting the application, and that variations, modifications, alternatives, and variations may be made to the above embodiments by one of ordinary skill in the art within the scope of the application.
Claims (5)
1. A satellite-to-solar spin stable attitude control method based on a pure magnetic control mode is characterized by comprising the following steps:
calculating an estimated value of the satellite three-axis angular rate;
according to the estimated value of the satellite three-axis angular rate, respectively calculating the target output moment of the satellite in each control stage of the solar spin, wherein each control stage of the solar spin comprises the following steps: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
calculating the target output magnetic moment of the triaxial magnetometers of the satellites according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellites;
Converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite;
the method for calculating the estimated value of the satellite three-axis angular rate comprises the following steps:
(1) Constructing a state vector
wherein ,
the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega x 、ω y 、ω z Respectively->
In the satellite body coordinate system O b X b Y b Z b O of (2) b X b 、O b Y b 、O b Z b An axial component;
(2) Constructing a state equation and a state transition matrix;
the satellite attitude dynamics equation is constructed as follows:
wherein ,
is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction; />
The control moment is output by the triaxial magnetic torquer;
the attitude dynamics equation is a state equation, and is linearized to obtain a state transition matrix
Wherein, deltaT is the operation period;
(3) Constructing a measurement equation and a measurement matrix;
geomagnetic vector measured values obtained by two continuous measurements of the triaxial magnetometer are respectively recorded as B b And
wherein ,
for the geomagnetic vector measurement value at the current moment, +. >
Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component; />
For the geomagnetic vector measurement value at the previous moment, +.>
Respectively at the previous momentMagnetic vector measurement value is in satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component;
construction of an observation vector
The following are provided:
order the
To represent the vector of the rotation angle of the satellite body coordinate system in the time period delta T, theta Δ Corresponding direction cosine matrix->
The approximation is expressed as:
wherein ,
for vector->
Is expressed as follows:
then a measurement equation is constructed as follows:
wherein ,
is an error term;
obtaining a measurement matrix according to the measurement equation
The method comprises the following steps:
(4) According to the state transition matrix obtained in the step (2) and the measurement matrix obtained in the step (3), calculating to obtain an estimated value of the satellite triaxial angular rate by using a Kalman filtering method
The target output torque calculation expression is as follows, either during the initial racemization phase, or during the pair-day acquisition phase, before the sun enters the sun sensor measurement field of view of the satellite, or during the pair-day acquisition phase, when the satellite is located in the ground shadow region:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 1 A positive first control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 1 Is vector B b and Tc1 An included angle between the two;
during the pair-day capture phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 2 、K 3 、K 4 Positive second, third and fourth control coefficients respectively; />
The unit vector is the optical axis direction of the sun sensor; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 2 Is vector B b and Tc2 An included angle between the two;
during the spinning phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 3 Is vector B b and Tc3 An included angle between the two;
in the phase of stabilizing the solar spin, when the satellite is in the sun-irradiated region, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; / >
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b In the case of a centroid inertial principal axis coordinate system,matrix J is a diagonal matrix, diagonal elements J of matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
in the phase of the para-solar spin stabilization, when the satellite is in the earth shadow region, the target output torque calculation expression is as follows:
wherein ,K9 A positive ninth control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 4 Is vector B b and Tc4 An included angle between the two.
2. The method according to claim 1, wherein the method further comprises:
in the initial racemization stage, if the angular rate value of the satellite triaxial angular rate in the time greater than or equal to the set first time threshold value is always smaller than or equal to the set first angular rate threshold value, ending the initial racemization stage, and enabling the satellite to enter a daily capturing stage.
3. The method according to claim 1, wherein the method further comprises:
in the sun-to-sun capturing stage, if the unit vector of the optical axis direction of the sun sensor
Solar vector S measured by sun sensor at current moment b And (3) the included angle is always smaller than or equal to a set first angle threshold value within the time of being larger than or equal to a second time threshold value, ending the daily capturing phase, and enabling the satellite to enter the spinning-up phase.
4. The method according to claim 1, wherein the method further comprises:
in the spinning stage, if the satellite flies into an earth shadow area, returning the satellite to the sun-to-sun capturing stage;
if the normal vector of the satellite star surface which is required to be aligned with the sun vector is set
Solar vector S measured by sun sensor at current moment b The included angle is always smaller than or equal to the set second angle threshold value within the time larger than or equal to the set third time threshold value, and the satellite triaxial angular rate estimated value is +.>
And the set satellite spin angular rate vector +.>
And if the deviation is always smaller than or equal to the set second angular rate threshold value within the same time, ending the spinning stage, and enabling the satellite to enter a phase of stabilizing the spinning.
5. The utility model provides a satellite is to day spin stable gesture controlling means based on pure magnetic control mode which characterized in that includes:
the triaxial angular rate estimation module is used for calculating an estimated value of the triaxial angular rate of the satellite;
the target output torque calculation module is used for calculating target output torque of the satellite in each control stage of the solar spin according to the estimated value of the satellite three-axis angular rate, and the each control stage of the solar spin comprises: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;
The target output magnetic moment calculation module calculates the target output magnetic moment of the triaxial magnetometers of the satellite according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellite;
the attitude control instruction calculation module is used for converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite;
the method for calculating the estimated value of the satellite three-axis angular rate comprises the following steps:
(1) Constructing a state vector
wherein ,
the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega x 、ω y 、ω z Respectively->
In the satellite body coordinate system O b X b Y b Z b O of (2) b X b 、O b Y b 、O b Z b An axial component;
(2) Constructing a state equation and a state transition matrix;
the satellite attitude dynamics equation is constructed as follows:
wherein ,
is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction; />
The control moment is output by the triaxial magnetic torquer;
the attitude dynamics equation is a state equation, and is linearized to obtain a state transition matrix 
Wherein, deltaT is the operation period;
(3) Constructing a measurement equation and a measurement matrix;
geomagnetic vector measured values obtained by two continuous measurements of the triaxial magnetometer are respectively recorded as B b And
wherein ,
for the geomagnetic vector measurement value at the current moment, +.>
Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component; />
For the geomagnetic vector measurement value at the previous moment, +.>
Respectively, the geomagnetic vector measured value at the previous moment is in a satellite body coordinate system O b X b 、O b Y b 、O b Z b An axial component;
construction of an observation vector
The following are provided:
order the
To represent the vector of the rotation angle of the satellite body coordinate system in the time period delta T, theta Δ Corresponding direction cosine matrix->
The approximation is expressed as:
wherein ,
for vector->
Is expressed as follows:
then a measurement equation is constructed as follows:
wherein ,
is an error term;
obtaining a measurement matrix according to the measurement equation
The method comprises the following steps:
(4) According to the state transition matrix obtained in the step (2) and the measurement matrix obtained in the step (3), calculating to obtain an estimated value of the satellite triaxial angular rate by using a Kalman filtering method
The target output torque calculation expression is as follows, either during the initial racemization phase, or during the pair-day acquisition phase, before the sun enters the sun sensor measurement field of view of the satellite, or during the pair-day acquisition phase, when the satellite is located in the ground shadow region:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 1 A positive first control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 1 Is vector B b and Tc1 An included angle between the two;
during the pair-day capture phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 2 、K 3 、K 4 Positive second, third and fourth control coefficients respectively; />
The unit vector is the optical axis direction of the sun sensor; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 2 Is vector B b and Tc2 An included angle between the two;
during the spinning phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/->
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 3 Is vector B b and Tc3 An included angle between the two;
in the phase of stabilizing the solar spin, when the satellite is in the sun-irradiated region, the target output torque calculation expression is as follows:
wherein ,
k is the estimated value of the three-axis angular rate of the satellite 5 、K 6 、K 7 、K 8 Positive fifth, sixth, seventh and eighth control coefficients, respectively; />
For setting the satellite star surface to be aligned with the sun vectorA normal unit vector; omega spin For a set spin angular rate value; s is S b Sun vector measured for sun sensor at present moment,/- >
The sun vector measured by the sun sensor at the previous moment; />
Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O b X b Y b Z b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J x 、J y 、J z Respectively represent satellite windings O b X b 、O b Y b 、O b Z b Moment of inertia of the rotation in the axial direction;
in the phase of the para-solar spin stabilization, when the satellite is in the earth shadow region, the target output torque calculation expression is as follows:
wherein ,K9 A positive ninth control coefficient;
calculating a target output magnetic moment M of the triaxial magnetic torquer:
wherein ,Bb The geomagnetic vector measurement value at the current moment; alpha 4 Is vector B b and Tc4 An included angle between the two.
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