Attitude control parameter calculation method for satellite with large-flexibility solar wing sailboard

Abstract

The invention relates to an attitude control parameter calculation method for a satellite with a large-flexibility solar panel. The attitude control parameter calculation method comprises the following steps: determining a target angle of rotation of a B axis of a two-dimensional solar panel driving mechanism relative to a zero position during sun tracking rotation of the solar panel; calculating a target rotational inertia by using the target angle and the satellite rotational inertia of the B axis of the two-dimensional solar panel driving mechanism at the zero position; and calculating PID control parameters by using the target rotational inertia. According to the attitude control parameter calculation method, for a satellite with a large-flexibility solar panel, self-adaptive calculation is carried out on the inertia change of the satellite when the B axis of the two-dimensional solar panel driving mechanism rotates to obtain the target rotational inertia, then the PID control parameters are calculated according to the target rotational inertia, and the robustness requirement of attitude control can be met.

Description

Technical Field

[0001] This invention relates to the field of satellite attitude control technology, and in particular to a method for calculating attitude control parameters of a satellite with a large flexible solar panel. Background Technology

[0002] Increased satellite energy demands necessitate larger solar array areas. Satellites with large, flexible solar arrays have significant moments of inertia, accounting for a large proportion of the overall satellite inertia. To maintain solar tracking and ensure energy supply during three-axis Earth alignment, the large, flexible solar arrays must rotate in tandem with the Solar Actuation and Control (SADA) mechanism. This rotation causes substantial changes in the overall satellite's moment of inertia, impacting attitude control; fixed control parameters are insufficient to guarantee robust attitude control. Summary of the Invention

[0003] To address at least some of the aforementioned problems in the prior art, this invention provides a method for calculating the attitude control parameters of a satellite with a large flexible solar panel, comprising the following steps:

[0004] Determine the target angle of the B-axis of the two-dimensional solar panel drive mechanism relative to the zero position when the solar panel rotates to track the sun. The B-axis of the two-dimensional solar panel drive mechanism is fixed to the satellite body and can rotate around the Zb axis of the satellite body coordinate system.

[0005] The target's moment of inertia was calculated using the target angle and the satellite's moment of inertia when the B-axis of the two-dimensional solar panel drive mechanism was at zero position; and

[0006] Calculate PID control parameters using the target moment of inertia.

[0007] Furthermore, when the B-axis is at zero, the satellite's moment of inertia matrix I is:

[0008]

[0009] I xx0 It is the moment of inertia of the satellite about the x-axis of the satellite's body coordinate system, I yy0 It is the moment of inertia of the satellite about the y-axis of the satellite's body coordinate system, I zz0 It is the moment of inertia of the satellite about the z-axis of the satellite's body coordinate system, I xy0 I represents the product of inertia of the satellite about the OXY plane of the satellite's body coordinate system. xz0 I represents the product of inertia of the satellite about the OXZ plane of the satellite's body coordinate system. yz0 This represents the inertial product of the satellite about the OYZ plane of the satellite's body coordinate system;

[0010] Calculate the target's moment of inertia, including the satellite's moment of inertia around the x-axis, y-axis, and z-axis of the satellite's coordinate system.

[0011] I xx =(I xx0 -I yy0 ) / 90*||β0|-90|+I yy0

[0012] I yy =(I yy0 -I xx0 ) / 90*||β0|-90|+I xx0

[0013] I zz =I zz0 ,

[0014] Where β0 is the target angle of rotation of the B-axis relative to the zero position.

[0015] Furthermore, the PID control parameters are calculated using the target moment of inertia, including the proportional control parameter Kp, the integral control parameter Ki, and the derivative control parameter Kd:

[0016] Kp=Iλ 2

[0017] Ki = c pi ×Iλ 2

[0018] Kd=2ζ×Iλ

[0019] λ=2πB w ,

[0020] Among them B w λ is the design frequency, λ is the design bandwidth, and c is the design frequency. pi ζ is the integral coefficient, and ζ is the damping ratio.

[0021] Furthermore, the axis fixed to axis B and rotating about an axis perpendicular to axis B is axis A of the two-dimensional solar panel drive mechanism;

[0022] When the A-axis of the solar panel is parallel to the Yb-axis of the satellite's coordinate system, the B-axis is at zero.

[0023] The present invention also provides a satellite attitude control method, comprising:

[0024] The PID control parameters were calculated using the attitude control parameter calculation method for satellites with large flexible solar panels.

[0025] The PID control parameters are input into the PID controller, which calculates the control torque and sends it to the actuator. The actuator generates the corresponding torque to drive the satellite to adjust its attitude.

[0026] Furthermore, the control torque is calculated using the following formula:

[0027]

[0028] In the formula, T x It is the X-axis control torque, T y It is the Y-axis control torque; T z It is the Z-axis control torque. θ, ψ are the roll, pitch, and yaw attitude deviation angles used for satellite control, and ω x ,ω y ,ω z To control the angular velocity of the orbital system deviation, K p K i K d These are proportional control parameters, integral control parameters, derivative control parameters, and integral term control parameters. Limit the amplitude using the absolute value of K. p1 It is the proportional control parameter for the scroll axis, K. i1 K is the integral control parameter for the rolling axis. d1 It is the differential control parameter of the rolling shaft, K p2 It is the pitch axis proportional control parameter, K i2 It is the pitch axis integral control parameter, K d2 It is the pitch axis differential control parameter, K p3 It is the yaw axis proportional control parameter, K i3 It is the yaw axis integral control parameter, K d3 It is the differential control parameter of the yaw axis.

[0029] The present invention has at least the following beneficial effects:

[0030] The attitude control parameter calculation method of the present invention is designed for satellites with large flexible solar panels. It adaptively calculates the change in satellite inertia when the B-axis of the two-dimensional solar panel drive mechanism rotates to obtain the target rotational inertia, and then calculates PID control parameters based on the target rotational inertia, which can meet the robustness requirements of attitude control. Attached Figure Description

[0031] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the various embodiments of the present invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope.

[0032] Figure 1 The diagram shows the relationship between the two-dimensional SADA zero position and rotation direction and the satellite body coordinate system.

[0033] Figure 2 The flowchart illustrates a method for calculating attitude control parameters of a satellite with a large flexible solar panel according to an embodiment of the present invention. Detailed Implementation

[0034] It should be noted that the components in the accompanying drawings may be shown exaggerated for illustrative purposes and may not be to scale.

[0035] In this invention, the various embodiments are merely intended to illustrate the solutions of the invention and should not be construed as limiting.

[0036] In this invention, unless otherwise specified, the quantifiers “a” and “one” do not exclude scenarios involving multiple elements.

[0037] It should also be noted that, in the embodiments of the present invention, only a portion of the parts or components may be shown for clarity and simplicity. However, those skilled in the art will understand that, under the teachings of the present invention, the required parts or components can be added as needed for specific scenarios.

[0038] It should also be noted that within the scope of this invention, the terms "same", "equal", and "equal to" do not mean that the two values ​​are absolutely equal, but allow for a certain reasonable error. In other words, the terms also cover "substantially the same", "substantially equal", and "substantially equal to".

[0039] It should also be noted that in the description of this invention, the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not explicitly or implicitly suggest that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0040] Furthermore, the embodiments of the present invention describe the process steps in a specific order; however, this is only for the convenience of distinguishing each step, and is not intended to limit the order of the steps. In different embodiments of the present invention, the order of each step can be adjusted according to the process.

[0041] As satellite energy demands increase, the area of ​​the solar panels increases, leading to a corresponding increase in rotational inertia. The rotation of the solar panels causes significant changes in the overall satellite's rotational inertia, impacting attitude control. This invention employs adaptive calculation based on changes in satellite inertia and designs a method for calculating satellite attitude control parameters based on the adaptive inertia calculation results. By identifying the inertia and designing an adaptive parameter algorithm, the overall satellite attitude control parameters can reduce the impact of inertia changes during solar panel rotation on satellite attitude, ensuring that the satellite attitude control accuracy meets the required specifications.

[0042] Satellites with large, flexible solar panels use reaction wheels as control actuators for attitude control. Their solar panel control mechanism is a two-dimensional solar panel drive mechanism (SADA), with key characteristics including:

[0043] (1) The relationship between the zero position and rotation direction of the two-dimensional SADA and the coordinate system of the satellite body is shown in [reference]. Figure 1 .

[0044] (2) The solar panels are basically symmetrical about axis A and axis B respectively.

[0045] (3) The center of mass of the entire star is close to the B-axis of the two-dimensional SADA.

[0046] like Figure 1 As shown, after the solar array deploys, the axis fixed to the satellite body and capable of rotating around the Zb axis of the satellite body coordinate system is defined as the B-axis of SADA. The axis fixed to the B-axis and rotating around an axis perpendicular to the B-axis is defined as the A-axis of SADA. When the A-axis of the solar array is parallel to the Yb axis of the satellite body coordinate system, it is defined as the zero position of the B-axis; when the normal of the solar array surface is in the negative direction of the Zb axis, it is defined as the zero position of the A-axis. In the zero position state, according to the right-hand rule, the right thumb points in the negative direction of the Zb axis, and the rotation direction of the other four fingers is the direction of increasing positive rotation angle of the B-axis; the right thumb points in the negative direction of the Yb axis, and the rotation direction of the other four fingers is the direction of increasing positive rotation angle of the A-axis.

[0047] The satellite's body coordinate system is OcXbYbZb, with the origin Oc defined as the satellite's center of mass. OcXb, OcYb, and OcZb are parallel to the three axes of the satellite's mechanical coordinate system and point in the same direction.

[0048] The theoretical basis of satellite attitude control is introduced below:

[0049] In the field of satellite attitude control, the celestial dynamics equations (also known as attitude dynamics equations) are the core equations describing the relationship between a satellite's rotational motion around its center of mass and the torques acting on it. They form the theoretical foundation for attitude control system design (such as PID control). Essentially, they are based on the angular momentum theorem, describing the dynamic relationship between the change in a satellite's angular momentum and external torques.

[0050] In the satellite's body coordinate system (OcXbYbZb), the equations of motion for the celestial body are expressed as follows:

[0051]

[0052] Where J is the satellite's moment of inertia matrix, a 3×3 symmetric matrix that describes the satellite's mass distribution's inertial resistance to rotation, in the form of:

[0053]

[0054] J xx J represents the moment of inertia of a satellite about the X-axis of its body coordinate system. yy J represents the moment of inertia of the satellite about the Y-axis of its body coordinate system. zz J represents the moment of inertia of the satellite about the Z-axis of the satellite's body coordinate system. xy J represents the product of inertia of the satellite about the OXY plane of the satellite's body coordinate system. xz J represents the inertial product of the satellite about the OXZ plane of the satellite's body coordinate system. yz The product of inertia of the satellite about the OYZ plane of the satellite's body coordinate system; ω = [ω x ω y ω z ] T It is the angular velocity of the satellite relative to the inertial frame (the component in the body coordinate system), which is a vector and is measured in rad / s;

[0055] ω is the derivative of angular velocity with respect to time (angular acceleration vector); ω = (Jω) is the Coriolis torque (virtual torque introduced by coordinate system rotation);

[0056] u represents the control torque, which is the torque actively applied by the control system to counteract interference and track the target attitude; h w h represents the angular momentum of the flywheel (if a reaction flywheel is configured). w It is the projection of the flywheel's spin angular momentum onto the body coordinate system; T d This refers to disturbance torque (external environmental disturbances, such as solar radiation pressure, Earth's non-spherical gravity, geomagnetic torque, etc.). Undesirable torques from the external environment can disrupt a satellite's attitude stability.

[0057] The satellite uses a classic PID control law, in the following form:

[0058]

[0059] In the formula, T x It is the X-axis (rolling) control torque, T y It is the Y-axis (pitch) control torque; T z It is the Z-axis (yaw) control torque. θ, ψ are the attitude deviation angles used for satellite control, ω x ,ω y ,ω z To control the angular velocity of the orbital system deviation, K p K i K d These are proportional control parameters, integral control parameters, derivative control parameters, and integral term control parameters. Limit the absolute value. Set a maximum permissible absolute value for the integral term. When the absolute value of the integral term exceeds the maximum permissible absolute value, forcibly limit the absolute value of the integral term to the maximum permissible absolute value.

[0060] Figure 2 The flowchart illustrates a method for calculating attitude control parameters of a satellite with a large flexible solar panel according to an embodiment of the present invention.

[0061] like Figure 2 As shown, a method for calculating attitude control parameters of a satellite with large flexible solar panels includes the following steps:

[0062] Step 1: Determine the target angle of the B-axis of the two-dimensional solar panel drive mechanism relative to the zero position when the solar panel rotates to track the sun.

[0063] The target angle β0 of the B-axis of SADA relative to the zero position rotation changes with the rotation of the solar panel.

[0064] Step 2: Calculate the target's rotational inertia using the target angle and the satellite's rotational inertia when the B-axis of the two-dimensional solar panel drive mechanism is at zero position.

[0065] When the B-axis is at zero, the satellite's moment of inertia matrix I is:

[0066]

[0067] I xx0 It is the moment of inertia of the satellite about the x-axis of the satellite's body coordinate system, I yy0 It is the moment of inertia of the satellite about the y-axis of the satellite's body coordinate system, I zz0 It is the moment of inertia of the satellite about the z-axis of the satellite's body coordinate system, I xy0 I represents the product of inertia of the satellite about the OXY plane of the satellite's body coordinate system. xz0 I represents the product of inertia of the satellite about the OXZ plane of the satellite's body coordinate system. yz0 This represents the satellite's inertial product about the OYZ plane of its body coordinate system.

[0068] Calculate the target's moment of inertia, including the satellite's moment of inertia around the x-axis, y-axis, and z-axis of the satellite's coordinate system.

[0069]

[0070] Step 3: Calculate the PID control parameters using the target moment of inertia, including the proportional control parameter Kp, integral control parameter Ki, and derivative control parameter Kd:

[0071]

[0072] Among them B w λ is the design frequency, λ is the design bandwidth, and c is the design frequency. pi ζ is the integral coefficient, and ζ is the damping ratio, which is typically 0.707.

[0073] The significant change in the overall satellite's rotational inertia caused by the rotation of the solar arrays is mainly due to the rotation along the B-axis. Therefore, to simplify the calculation, this invention only considers the rotation along the B-axis and uses the angle of rotation along the B-axis to calculate the control parameters.

[0074] A satellite attitude control method includes:

[0075] The PID control parameters were calculated using the attitude control parameter calculation method for satellites with large flexible solar panels.

[0076] The PID control parameters are input to the PID controller, which calculates the control torque and sends it to the actuator. The actuator generates the corresponding torque to drive the satellite to adjust its attitude. The actuator can be the satellite's reaction flywheel, thruster, etc.

[0077] The control torque is calculated using the following formula:

[0078]

[0079] In the formula, T x It is the X-axis control torque, T y It is the Y-axis control torque; T z It is the Z-axis control torque. θ, ψ are the roll, pitch, and yaw attitude deviation angles used for satellite control, and ω x ,ω y ,ω z To control the angular velocity of the orbital system deviation, K p K i K d These are proportional control parameters, integral control parameters, derivative control parameters, and integral term control parameters. Limit the amplitude using the absolute value of K. p1It is the proportional control parameter for the scroll axis, K. i1 K is the integral control parameter for the rolling axis. d1 It is the differential control parameter of the rolling shaft, K p2 It is the pitch axis proportional control parameter, K i2 It is the pitch axis integral control parameter, K d2 It is the pitch axis differential control parameter, K p3 It is the yaw axis proportional control parameter, K i3 It is the yaw axis integral control parameter, K d3 It is the differential control parameter of the yaw axis.

[0080] While some embodiments of the present invention have been described in this application, those skilled in the art will understand that these embodiments are merely illustrative. Numerous variations, alternatives, and improvements will arise in those skilled in the art under the teachings of this invention without departing from its scope. The appended claims are intended to define the scope of the invention and thereby cover methods and structures within the scope of the claims themselves and their equivalents.

Claims

1. A method of computing attitude control parameters for a satellite having a large flexible solar wing panel, characterized by, The method comprises the following steps: determining a target angle of a B-axis of a two-dimensional solar sail drive mechanism relative to a zero position when a solar wing sailboard is tracking the sun, the B-axis being fixed to a satellite body and being capable of rotating about a Zb-axis of a satellite body coordinate system; calculating a target moment of inertia using the target angle and a moment of inertia of the B-axis of the two-dimensional solar sail drive mechanism at the zero position; calculating a PID control parameter using the target moment of inertia. When the B-axis is at the zero position, a moment of inertia matrix I of the satellite is:

2. The method of claim 1, wherein, The calculation of the target moment of inertia comprises a moment of inertia of the satellite about an x-axis of the satellite body coordinate system, a moment of inertia of the satellite about a y-axis of the satellite body coordinate system, and a moment of inertia of the satellite about a z-axis of the satellite body coordinate system: I xx0 is the moment of inertia of the satellite about the x-axis of the satellite body coordinate system, I yy0 is the moment of inertia of the satellite about the y-axis of the satellite body coordinate system, I zz0 is the moment of inertia of the satellite about the z-axis of the satellite body coordinate system, I xy0 denotes the product of inertia of the satellite about the OX plane of the satellite body coordinate system, I xz0 denotes the product of inertia of the satellite about the OY plane of the satellite body coordinate system, I yz0 denotes the product of inertia of the satellite about the OZ plane of the satellite body coordinate system; where β0 is the target angle of the B-axis relative to the zero position. I xx = (I xx0 - I yy0 ) / 90*||β0|-90|+I yy0 I yy = (I yy0 - I xx0 ) / 90*||β0|-90|+I xx0 I zz = I zz0 , The calculation of the PID control parameter using the target moment of inertia comprises a proportional control parameter Kp, an integral control parameter Ki, and a differential control parameter Kd:

3. The method of claim 2, wherein, Kd = 2ζ×Iλ Kp = I λ 2 Ki = c pi x I λ 2 An A-axis of the two-dimensional solar sail drive mechanism is fixed to the B-axis and rotates about an axis perpendicular to the B-axis; λ = 2πB w , where B w is the design frequency, λ is the design bandwidth, c pi is the integral coefficient, and ζ is the damping ratio.

4. The method of claim 1, wherein, When the A-axis of the solar wing sailboard is parallel to a Yb-axis of the satellite body coordinate system, the B-axis is at the zero position. The method comprises:

5. A method of satellite attitude control, characterized by, calculating the PID control parameter by the method for calculating an attitude control parameter of a large-flexibility solar wing sailboard satellite according to any one of claims 1 to 4; inputting the PID control parameter into a PID controller, the PID controller calculating a control moment and sending the control moment to an executing mechanism, the executing mechanism generating a corresponding moment to drive the satellite to adjust an attitude. The control moment is calculated by the following formula:

6. The method of claim 5, wherein, ​ wherein T x is the X-axis control torque, T y is the Y-axis control torque; T z is the Z-axis control torque, θ, ψ are the roll, pitch, yaw attitude deviation angles for satellite control, ω x , ω y , ω z are the orbit system deviation angular velocities for control, K p , K i , K d are the proportional control parameters, integral control parameters, and differential term control parameters, respectively, and the absolute value of the integral term is limited, K p1 is the roll-axis proportional control parameter, K i1 is the roll-axis integral control parameter, K d1 is the roll-axis differential term control parameter, K p2 is the pitch-axis proportional control parameter, K i2 is the pitch-axis integral control parameter, K d2 is the pitch-axis differential term control parameter, K p3 is the yaw-axis proportional control parameter, K i3 is the yaw-axis integral control parameter, and K d3 is the yaw-axis differential term control parameter.

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