Single-axis inertial space orientation attitude control method using magnetic control

Abstract

The application relates to a single-axis inertia space orientation attitude control method using magnetic control, which comprises the following steps: performing optimization control according to the current attitude deviation of a satellite; calculating an optimal magnetic control torque under the current attitude; and calculating an output magnetic torque of a magnetic torque device. The single-axis inertia space orientation attitude control method using magnetic control can be used for single-axis inertia space orientation attitude control of a satellite using magnetic control, can calculate an optimal output torque vector of the magnetic control according to the current expected control torque of the satellite and the direction of the current magnetic field vector, and can guarantee that the actual output torque of two control axes is the same as the expected control torque when the satellite is controlled by the magnetic torque device, so that high-precision attitude control can be realized. Meanwhile, an angular velocity damping strategy is designed for a non-control axis, so that the angular velocity of the axis will not be excessively large.

Description

Technical Field

[0001] This invention relates to the field of spacecraft attitude control technology, and in particular to a single-axis inertial space orientation attitude control method using magnetic control. Background Technology

[0002] Compared to reaction flywheels, control moment gyroscopes, or propulsion systems, magnetic torquers are lightweight, consume less power, and have a simpler structure, making them suitable for attitude control of certain microsatellites where high attitude control precision is not required. Alternatively, in cases where existing control actuators such as flywheels fail during satellite operation, magnetic torquers can be used as backup control to ensure the satellite's solar panels remain oriented towards the sun.

[0003] During their orbital operation, satellites are subject to space disturbance torques such as gravity gradient torque, atmospheric drag torque, and remanent magnetic torque. Compared to actuators like flywheels, magnetic torque converters have a weaker torque output capability. The proportion of space disturbance torques to the magnetic torque converter's output torque is relatively large, which can significantly impact satellite attitude control. Furthermore, the magnetic torque converter's output torque is limited to the direction perpendicular to the Earth's magnetic field, essentially allowing only two-axis control at a time. This prevents it from outputting the optimal desired control torque based on current attitude control deviations. It can even lead to reverse polarity in the actual output torque along certain axes, resulting in low attitude control accuracy. Summary of the Invention

[0004] The present invention aims to solve the technical problems in the prior art by providing a single-axis inertial space orientation attitude control method using magnetic control.

[0005] To solve the above-mentioned technical problems, the technical solution of the present invention is as follows:

[0006] A single-axis inertial space orientation attitude control method using magnetic control includes the following steps:

[0007] Step 1: Based on the current attitude deviation of the satellite, choose to perform angular velocity damping or magnetic output torque optimization control;

[0008] Step 2: Calculate the optimal magnetic control torque under the current attitude;

[0009] Step 3: Calculate the output magnetic moment of the magnetic torquer based on the current magnetic field vector.

[0010] In the above technical solution, step 1 specifically includes:

[0011] Determine the satellite bias quaternion and bias angular velocity, and then determine whether the bias angular velocity exceeds the threshold:

[0012] If the value exceeds the limit, angular velocity damping optimization control is performed; otherwise, magnetic output torque optimization control is performed.

[0013] In the above technical solution, step 2 specifically includes:

[0014] If the deviation angular velocity exceeds the threshold, angular velocity damping optimization control is performed to reduce the angular velocity to a certain range; then, the desired magnetic control torque for attitude control is calculated, and magnetic control output torque optimization control is performed.

[0015] If the deviation angular velocity does not exceed the threshold, the desired magnetic control torque is calculated, and the magnetic control output torque is optimized.

[0016] In the above technical solution, the calculation of the desired magnetic control torque in step 2 is specifically as follows:

[0017] Calculate the desired control torque T based on the angular velocities of each axis of the satellite:

[0018]

[0019]

[0020]

[0021] Where, ω max =max(|ω x |,|ω y |,|ω z |), ω x ω y ω z The angular velocities of the satellite along the X, Y, and Z axes are respectively, T max T is the maximum magnetic torque that the magnetic torquer can output under the current track. x T y T z These are the components of the actual output torque on the X, Y, and Z axes, respectively.

[0022] In the above technical solutions,

[0023] The optimized control of the magnetic output torque in step 2 is as follows:

[0024] Based on the satellite's current attitude deviation, the current desired torque, T, is calculated using the control law. e First, ensure that the projection T of the desired torque in the XY plane is... e_xy The control torque T on the satellite's Z-axis remains constant. z Adjustments were made to ensure that the satellite's desired output torque T was perpendicular to the Earth's magnetic field vector;

[0025] Assumption:

[0026] B x T x +B y T y +B z T z =0

[0027] Among them, T x T y T z These are the components of the actual output torque on the X, Y, and Z axes, respectively, B. x B y B z The components of the geomagnetic field vector in the body coordinate system along the X, Y, and Z axes;

[0028] Therefore, it can be concluded that:

[0029]

[0030] At the same time, if vector B is close to the XY plane of the body coordinate system, it will cause the actual output torque component T on the Z-axis to be affected. z The torque is too large, therefore the output torque on the X, Y, and Z axes is limited proportionally:

[0031]

[0032]

[0033]

[0034] Among them, T max T is the maximum magnetic torque that the magnetic torquer can output under the current track. x_out T y_out T z_out These are the output torques for the X, Y, and Z axes, respectively.

[0035] The present invention has the following beneficial effects:

[0036] The present invention provides a single-axis inertial space orientation attitude control method using magnetic control, which addresses the problem of single-axis inertial space orientation attitude control of satellites using magnetic control. Based on the satellite's current desired control torque and the current magnetic field vector direction, the optimal magnetic control output torque vector can be calculated.

[0037] The method of this invention ensures that when the satellite is attitude controlled via a magnetic torque converter, the actual output torque of the two control axes is the same as the desired control torque, thereby achieving high-precision attitude control. Simultaneously, an angular velocity damping strategy is designed for the non-control axes to prevent excessive angular velocities on these axes.

[0038] The single-axis inertial space orientation attitude control method of the present invention, which uses magnetic control, is described with the Z-axis as the non-control axis. In fact, any vector of the celestial body can be taken as the non-control axis, and the effect of the present invention can still be achieved. Attached Figure Description

[0039] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments.

[0040] Figure 1 A schematic diagram for optimizing the magnetic control output torque.

[0041] Figure 2 A schematic diagram of the calculation process for satellite magnetic control optimization.

[0042] Figure 3 This is a simulation diagram of the space interference torque of a satellite.

[0043] Figure 4 This is a schematic diagram of the quaternion for satellite attitude deviation.

[0044] Figure 5 This is a schematic diagram of the angular velocity of satellite attitude deviation. Detailed Implementation

[0045] The inventive concept of this invention is as follows:

[0046] This invention presents a single-axis inertial space orientation attitude control method using magnetic control. By eliminating attitude control of unnecessary axes, it optimizes the control output torque by utilizing the characteristic that the output torque is perpendicular to the Earth's magnetic field. While ensuring that the overall output torque is perpendicular to the magnetic field direction, it also ensures that the actual output torque of the two control axes matches the desired torque, thereby achieving high-precision single-axis inertial space orientation attitude control for satellites.

[0047] The present invention provides a single-axis inertial space orientation attitude control method using magnetic control, comprising the following steps:

[0048] Step 1: Based on the current attitude deviation of the satellite, choose to perform angular velocity damping or magnetic output torque optimization control;

[0049] Determine the satellite deviation quaternion and deviation angular velocity, and judge whether the deviation angular velocity exceeds the threshold: if it does, perform angular velocity damping optimization control; otherwise, perform magnetic control output torque optimization control.

[0050] Step 2: Calculate the optimal magnetic control torque under the current attitude;

[0051] If angular velocity damping optimization control is performed, the desired magnetic control torque for damping is calculated; if magnetic control output torque optimization control is performed, the desired magnetic control torque for attitude control is calculated.

[0052] If the deviation angular velocity exceeds the threshold, angular velocity damping optimization control is performed. After the angular velocity is reduced to a certain range, magnetic control output torque optimization control is performed to calculate the desired magnetic control torque for attitude control.

[0053] Step 3: Calculate the output magnetic moment of the magnetic torquer based on the current magnetic field vector;

[0054] Calculate the output magnetic moment of the magnetic torquer based on the current magnetic field vector.

[0055] The present invention will now be described in detail with reference to the accompanying drawings.

[0056] The single-axis inertial space orientation attitude control method using magnetic control of the present invention is described below:

[0057] 1. Basic Principles of Satellite Magnetically Controlled Torque Output

[0058] Assuming the desired output torque of the satellite is T, the magnetic moment is generally generated using the cross product law, as shown in the following formula:

[0059]

[0060] Where K > 0 is the control gain coefficient.

[0061] Therefore, the torque acting on the satellite is:

[0062]

[0063] From the above formula, we can see that:

[0064] (1) When the magnetic field vector B is perpendicular to the desired control torque T, the above equation simplifies to: At this point, the magnetic torque is always the same as the desired control torque.

[0065] (2) When the magnetic field vector B is not perpendicular to the desired control torque T, the magnetic field vector B can be decomposed into a component B along the direction of T. T The component B perpendicular to the T direction N The above equation simplifies to:

[0066]

[0067] The actual output torque along the direction of the desired control torque is:

[0068]

[0069] From the above formula, we can see that T MT A positive value indicates that the angle between the actual output torque and the desired output torque is acute. Furthermore, the component of the magnetic field strength in the direction of the desired torque, B, is known. T The smaller the size, the higher the control efficiency.

[0070] Meanwhile, the actual output torque perpendicular to the desired control torque direction is:

[0071]

[0072] Where θ is the angle between the magnetic field vector B and the desired control torque T. For B N The unit vector of direction. Because T MN Perpendicular to T, this torque increases the angular momentum perpendicular to the direction of the magnetic field vector, and the torque in this direction is maximum when the angle between B and T is 45°.

[0073] The above analysis shows that when the desired control torque is not perpendicular to the magnetic field vector, the actual control torque cannot be guaranteed to be in the same direction as the desired control torque. This results in the actual control performance being such that while attitude deviations in some axes converge, attitude deviations in other axes gradually diverge, thus affecting control accuracy.

[0074] 2. Optimize magnetic control output torque

[0075] Based on the characteristic that the magnetic control torque is always perpendicular to the magnetic field vector, satellite magnetic control is essentially a two-axis attitude control. Therefore, one axis of the satellite can be designated as a non-control axis, and control of that axis can be abandoned, while attitude control is only applied to the other two axes. That is, by controlling the two axes of the satellite according to the characteristics of magnetic dual-axis attitude control, the output torque of both controls is guaranteed to be the same as the desired torque. Through attitude control of the two control axes, arbitrary pointing of the non-control axis can be achieved.

[0076] For example, assuming that only the Z-axis of the satellite is required to point in a certain direction of inertial space, then the Z-axis is a non-control axis, and the satellite attitude control system only performs attitude control on the X-axis and Y-axis of the satellite. The specific process is as follows.

[0077] Based on the satellite's current attitude deviation, the current desired torque, T, is calculated using the control law. e First, ensure that the projection T of the desired torque in the XY plane is... e_xy The control torque T on the satellite's Z-axis remains constant. z Adjustments were made to ensure that the satellite's desired output torque T was perpendicular to the Earth's magnetic field vector, such as... Figure 1 As shown.

[0078] Depend on Figure 1 It can be seen that the actual output torque vector T is perpendicular to the geomagnetic field vector B in the body coordinate system, and the endpoint of T is at T. e -T e_xy Online. Meanwhile, T e -T e_xy Since the line is parallel to the Z-axis of the body coordinate system, it can be concluded that as long as vector B is not perpendicular to the Z-axis of the body coordinate system and is in line with the desired control torque T...e If they are not parallel, then there must exist a moment vector T that is perpendicular to vector B. Therefore, we can conclude that:

[0079] B x T x +B y T y +B z T z =0

[0080] Therefore, it can be concluded that:

[0081]

[0082] At the same time, if vector B is close to the XY plane of the body coordinate system, it will cause the actual output torque component T on the Z-axis to be affected. z The torque is too large, therefore the output torque of each axis is limited proportionally:

[0083]

[0084]

[0085]

[0086] Among them, T max T is the maximum magnetic torque that the magnetic torquer can output under the current track. x_out T y_out T z_out These are the output torques for the X, Y, and Z axes, respectively.

[0087] By optimizing the magnetic control output torque as described above, the output torque of the two control axes is ensured to be consistent with the desired torque, thereby ensuring that the two control axes can achieve high control accuracy.

[0088] 3. Angular velocity damping

[0089] Because the aforementioned control process did not control the non-control axis of the celestial body, namely the Z-axis, if the calculated control torque along the Z-axis maintains the same polarity over a long period, and considering the long-term influence of space disturbance torques, the angular velocity of this axis will gradually diverge. Therefore, it is necessary to dampen the angular velocity of this axis when it exceeds a certain threshold.

[0090] Based on the angular velocities of each axis of the satellite, the desired control torque T is recalculated.

[0091]

[0092]

[0093]

[0094] Where, ω max =max(|ω x |,|ω y |,|ω z |).

[0095] Once the angular velocity is reduced to a certain range, attitude control is restarted.

[0096] After optimizing the magnetic output torque, the output magnetic torque of the torque converter is calculated according to the formula above, as follows:

[0097]

[0098] Therefore, the overall execution process of this invention is as follows:

[0099] First, based on the current attitude deviation of the satellite, we choose to perform angular velocity damping or magnetic output torque optimization control.

[0100] The optimal magnetic control torque under the current attitude is calculated;

[0101] Furthermore, the output magnetic moment of the magnetic torquer is calculated based on the current magnetic field vector, such as... Figure 2 As shown.

[0102] The following describes an application example of the single-axis inertial space orientation attitude control method using magnetic control according to the present invention.

[0103] A satellite uses a magnetic torque converter for attitude control, maintaining its inertial spatial orientation along its Z-axis. The satellite's relevant parameters are shown in the table below:

[0104] Table 1 Satellite Parameters

[0105]

[0106] The average magnetic field strength of the Earth at this orbital altitude is approximately 36,000 nT, and the calculated average single-axis magnetostrictive torque of the satellite is approximately 1.8 mNm. The satellite is subjected to disturbances in orbit including gravity gradient torque, atmospheric drag torque, remanent magnetic torque, and optical pressure torque. The overall space disturbance torque simulation is as follows: Figure 3 As shown.

[0107] The maximum single-axis interference torque of the satellite is approximately 0.35 mNm, accounting for 20% of the single-axis magnetron torque.

[0108] The actual control effect of the present invention is as follows: Figure 4 As shown, the attitude quaternion control accuracy of the satellite's X-axis and Y-axis is 0.0075 (3σ), corresponding to an angle of approximately 0.859°.

[0109] Meanwhile, the simulation results of the deviation angular velocity of each axis are as follows: Figure 5 As shown, the angular velocity of the non-control axis Z-axis is within 0.149° / s.

[0110] This invention presents a single-axis inertial space orientation attitude control method using magnetic control. By eliminating attitude control of unnecessary axes, it optimizes the control output torque by utilizing the characteristic that the output torque is perpendicular to the Earth's magnetic field. While ensuring that the overall output torque is perpendicular to the magnetic field direction, it also ensures that the actual output torque of the two control axes matches the desired torque, thereby achieving high-precision single-axis inertial space orientation attitude control for satellites.

[0111] This invention presents a single-axis inertial space orientation attitude control method using magnetic control. Addressing the problem of single-axis inertial space orientation attitude control for satellites using magnetic control, the optimal magnetic control output torque vector can be calculated based on the satellite's current desired control torque and the current magnetic field vector direction. This method ensures that when the satellite is controlled via the magnetic torque converter, the actual output torque of both control axes is the same as the desired control torque, thus achieving high-precision attitude control. Simultaneously, an angular velocity damping strategy is designed for non-control axes to prevent excessive angular velocities on these axes.

[0112] This invention is described using the Z-axis as the non-control axis. In fact, any vector of the celestial body can be used as the non-control axis, and the effect of this invention can still be achieved.

[0113] Obviously, the above embodiments are merely illustrative examples for clear explanation and are not intended to limit the implementation. Those skilled in the art will recognize that other variations or modifications can be made based on the above description. It is neither necessary nor possible to exhaustively list all possible implementations here. However, obvious variations or modifications derived therefrom are still within the scope of protection of this invention.

Claims

1. A method for single-axis inertial space orientation attitude control using magnetic control, characterized by, Includes the following steps: Step 1: Based on the current attitude deviation of the satellite, choose to perform angular velocity damping or magnetic output torque optimization control; Step 2: Calculate the optimal magnetic control torque under the current attitude; Step 3: Calculate the output magnetic moment of the magnetic torquer based on the current magnetic field vector; Step 1 specifically includes: Determine the satellite bias quaternion and bias angular velocity, and then determine whether the bias angular velocity exceeds the threshold: If the value exceeds the limit, angular velocity damping optimization control is performed; otherwise, magnetic output torque optimization control is performed. Step 2 specifically includes: If the deviation angular velocity exceeds the threshold, angular velocity damping optimization control is performed to reduce the angular velocity to a certain range; then, the desired magnetic control torque is calculated, and magnetic control output torque optimization control is performed. If the deviation angular velocity does not exceed the threshold, the desired magnetic control torque is calculated, and the magnetic control output torque is optimized and controlled. The calculation of the desired magnetic control torque in step 2 is as follows: Calculate the desired control torque based on the angular velocities of each axis of the satellite. : in, , , , These represent the angular velocities of the satellite along its X, Y, and Z axes, respectively. This represents the maximum magnetic torque that the magnetic torquer can output under the current track conditions. , , These are the components of the actual output torque on the X, Y, and Z axes, respectively. The optimized control of the magnetic output torque in step 2 is as follows: Based on the satellite's current attitude deviation, the current desired torque is calculated using the control law. First, ensure the projection of the desired torque onto the XY plane. The control torque on the satellite's Z-axis remains unchanged. Adjustments were made to ensure the satellite's desired output torque. It can be perpendicular to the Earth's magnetic field vector; Assumption: in, , , These represent the components of the actual output torque along the X, Y, and Z axes, respectively. , , The components of the geomagnetic field vector in the body coordinate system along the X, Y, and Z axes; Therefore, it can be concluded that: At the same time, if vector If the plane is close to the XY plane of the body coordinate system, it will result in the component of the actual output torque on the Z-axis. The torque is too large, therefore the output torque on the X, Y, and Z axes is limited proportionally: in, This represents the maximum magnetic torque that the magnetic torquer can output under the current track conditions. , , These are the output torques for the X, Y, and Z axes, respectively.

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