Satellite attitude pointing adjustment method based on multi-body momentum moment transmission

Abstract

The invention discloses a satellite attitude pointing adjustment method based on multi-body momentum moment transmission, and the method comprises the steps: calculating an attitude error value according to a target attitude angle and an initial attitude angle of satellite rotation; calculating a rotation angle acceleration based on the attitude error value; solving the rotation angular acceleration of the satellite body based on the attitude kinetic equation of the moment of momentum; updating the rotation angular velocity of the satellite body and the rotation angular velocity and the rotation angle of the tail end rod of the mechanical arm; judging whether the rotation angle meets the motion range constraint condition of the joint angle, if so, entering the next step, and otherwise, terminating the algorithm; updating a real part and an imaginary part of the quaternion, then calculating three components of an attitude angle of satellite rotation, then calculating the attitude angle of satellite rotation, and then calculating an attitude error value between the attitude angle of satellite rotation and a target attitude angle; and finally, judging whether the attitude error value is smaller than a preset threshold value, if so, performing attitude pointing adjustment according to the three components, and otherwise, returning to the step of calculating the rotational angular acceleration.

Description

Technical Field

[0001] The invention relates to a satellite attitude adjustment technology, and in particular to a satellite attitude pointing adjustment method based on multi-body momentum moment transfer. Background Art

[0002] Satellite attitude control has always been one of the core technologies for satellites to perform on-orbit service missions. Existing attitude control actuators include reaction flywheels, chemical thrusters, control torque gyroscopes (CMGs), and magnetic torquers. Figure 1 )'s attitude control complies with the momentum theorem, that is, the derivative of the momentum of a particle about a fixed point in space with respect to time is equal to the torque of the force on the same point. The flywheel system and the small satellite body achieve the stability and control of the star body's attitude through the exchange of momentum.

[0003] When the external disturbance torque Md acts on the satellite, the system will produce a certain attitude angle deviation relative to the reference axis in the space coordinate system. The attitude sensor on the satellite measures the change of the attitude angle and transmits the deviation to the control system. The flywheel control system changes the angular velocity Ω of the flywheel according to the pre-set control instructions, and then generates the corresponding control torque Mc of the disturbance torque, which can fully absorb the influence of the disturbance torque Md on the attitude of the satellite, thereby realizing the correction of the attitude deviation of the satellite and the stability of the attitude angle, and eliminating the influence of the disturbance torque on the satellite.

[0004] Disadvantages of reaction flywheels: Flywheel control systems have the problem of speed saturation. When the flywheel accelerates or deflects in a certain direction to overcome certain non-periodic disturbances, it will eventually reach the maximum allowable speed. In this extreme state, the flywheel will no longer be able to absorb the excess momentum of the spacecraft and lose control. This state is called saturation, and saturation is a shortcoming that the flywheel system itself cannot overcome. Therefore, in order to avoid this situation, the flywheel system needs to be equipped with another system to unload saturation, which means that the flywheel system cannot operate independently and must be coordinated with a second control method. After the speed reaches saturation, the commonly used unloading method requires the use of fuel.

[0005] The chemical propulsion system currently used in spacecraft (the schematic diagram of the chemical propulsion system can be found in Figure 2 ) are mostly active control, including cold gas propulsion, chemical propulsion, electric propulsion and new concept propulsion. Chemical propulsion systems obtain thrust by converting high pressure into high-speed airflow through the combustion of fuel in the combustion chamber. Similar to rockets, some use solid fuel, some use a single fuel such as hydrazine, and some use a combination of multiple fuels.

[0006] The chemical thrusters on the satellite generate thrust by burning propellants. According to Newton's third law, the ejected gas changes the satellite's attitude under the action of the reaction force. The thrusters on the satellite are usually arranged at different positions or angles. The attitude control system controls the injection time and direction of different thrusters based on the feedback from the sensors, generates different torques, and adjusts the satellite's attitude.

[0007] Disadvantages of chemical thrusters: First, chemical thrusters consume a lot of fuel to generate thrust and torque, which may cause satellites to face fuel shortages during long-term missions, especially in deep space exploration or long-term missions. Fuel consumption will limit the effective working time and operational capabilities of satellites. Usually, spacecraft need to adjust their attitude frequently, so the propellant consumption is large, and the propellant carried by spacecraft is limited, which requires regular rocket launches for resupply, so the cost of using thrusters is high.

[0008] Secondly, the control accuracy of chemical thrusters is relatively low. Since their thrust is usually large and difficult to adjust accurately, it is difficult to achieve very fine attitude adjustments, so they are not suitable for tasks that require high-precision control. Compared with other more precise attitude control methods, chemical thrusters have certain disadvantages in accuracy and efficiency. These factors limit the application of chemical thrusters in some high-demand tasks, especially for satellite systems that require long-term stable operation and high-precision attitude adjustment. Summary of the invention

[0009] In view of the above-mentioned deficiencies in the prior art, the satellite attitude pointing adjustment method based on multi-body momentum moment transfer provided by the present invention solves the problem that the existing attitude control method relies on fuel or flywheels and other equipment for adjustment, which consumes fuel and thus shortens the satellite's on-orbit life.

[0010] In order to achieve the above-mentioned object of the invention, the technical solution adopted by the present invention is:

[0011] A method for adjusting satellite attitude and pointing based on multi-body momentum moment transfer is provided, which comprises the following steps:

[0012] S1. Calculate the attitude error value according to the target attitude angle and the initial attitude angle of the satellite rotation;

[0013] S2, based on the attitude error value, PD feedback control is used to obtain the rotational angular acceleration of the end rod of the manipulator;

[0014] S3, bringing the rotational angular acceleration and angular velocity of the end rod of the manipulator arm and the rotational angular velocity of the satellite body into the attitude dynamics equation of the satellite based on the momentum moment, and solving to obtain the rotational angular acceleration of the satellite body;

[0015] S4, according to the rotational angular acceleration of the satellite body and the end rod of the robotic arm, respectively update the rotational angular velocity of the satellite body and the rotational angular velocity and rotation angle of the end rod of the robotic arm;

[0016] S5, judging whether the rotation angle satisfies the activity range constraint condition of the joint angle, if so, proceeding to step S6, otherwise terminating the satellite attitude pointing adjustment method;

[0017] S6. Update the real part and the imaginary part of the quaternion according to the angular velocity component of the satellite body's rotation angular velocity on the xyz axis; calculate the three components of the satellite's rotation attitude angle based on the updated real part and the imaginary part;

[0018] S7. Calculate the attitude angle of the satellite rotation according to the three components, and calculate the attitude error value between the attitude angle of the satellite rotation and the target attitude angle;

[0019] S8. Determine whether the attitude error value is less than a preset threshold value. If so, adjust the attitude pointing according to the three components. Otherwise, return to step S2.

[0020] Furthermore, the satellite's attitude dynamics equation based on momentum is expressed as:

[0021]

[0022] Where, J is the moment of inertia of the satellite body; is the angular acceleration of the satellite body; ω bi is the angular velocity of the satellite body; J w is the moment of inertia of the end rod of the robot arm; ω w is the angular velocity of the end rod of the robot arm; T env is the environmental torque; is the angular acceleration of the end rod of the robot arm.

[0023] Furthermore, the expressions for updating the angular velocity of the satellite body and the angular velocity and rotation angle of the end rod of the robotic arm are:

[0024]

[0025]

[0026] φ w (t) = φ w (t―1)+ω w Δt

[0027] Among them, ω bi (t) and ω bi (t―1) is the angular velocity of the satellite at time t and time t-1 respectively; ω w (t) and ωw (t―1) are the rotational angular velocities of the end rod of the manipulator at time t and time t-1 respectively; φ w (t) and φ w (t―1) are the rotation angles of the end rod of the robot arm at time t and time t-1 respectively; Δt is the discrete time step.

[0028] Furthermore, step S6 further includes:

[0029] S61. Calculate the rate of change of the real and imaginary parts of the quaternion according to the angular velocity component of the satellite body's rotational angular velocity on the xyz axis:

[0030]

[0031] Among them, q0(t-1) is the real part of the quaternion at time t-1; q1(t-1), q2(t-1) and q3(t-1) are all the imaginary parts of the quaternion at time t-1; and are the rates of change of the real and imaginary parts of the quaternion at time t; ω x ,ω y and ω z They are ω bi Angular velocity components on the xyz axis;

[0032] S62. Update the real number of the quaternion according to the rate of change of the real part and the imaginary part of the quaternion:

[0033]

[0034] Among them, the value of k is 0, 1, 2 and 3; when k = 0, q0(t) and q0(t-1) are the real parts at time t and t-1 respectively. is the rate of change of the real part at time t; when k = 1, 2 and 3, q k (t) and q k (t―1) is the kth imaginary part at time t and t-1 respectively, is the rate of change of the imaginary part at time t;

[0035] S63. Calculate the three components of the satellite's rotation attitude angle according to the real and imaginary parts of the quaternion at time t:

[0036]

[0037] θ(t)=arcsin(―2(q1(t)q3(t)―q0(t)q2(t)))

[0038]

[0039] in, θ(t) and ψ(t) are the three components of the attitude angle of the satellite rotation at time t; q1(t), q2(t) and q3(t) are the imaginary parts of the quaternion corresponding to k = 1, 2 and 3 at time t, respectively.

[0040] Furthermore, the method for obtaining the motion range constraint condition of the joint angle includes:

[0041] A1. The satellite's robotic arm is divided into a primary robotic arm and a secondary robotic arm. The primary robotic arm is the robotic arm between the satellite body and the end rod of the robotic arm; the secondary robotic arm is the end rod of the robotic arm;

[0042] A2. Calculate the coordinates of the satellite's center of mass based on the satellite's geometric model:

[0043]

[0044]

[0045]

[0046] Where α is the rotation angle of the first-stage manipulator relative to the satellite; r is the length of the first-stage manipulator; L is the length of the satellite body; η is the angle between the first-stage manipulator pointing vector and the rotation axis to the center of mass position vector; δ x , δ y is the uncertainty of the center of mass position of the equivalent satellite body caused by the other two robotic arms that are not located in the same plane; (X0, Y0) is the coordinate of the center of mass of the satellite body;

[0047] A3. Draw a circle with the joint rotation axis of the secondary manipulator as the center and the distance from the joint rotation axis of the secondary manipulator to the mass center of the secondary manipulator as the radius. Calculate the coordinates of the points where the two rays passing through the mass center of the satellite body are tangent to the circle:

[0048]

[0049]

[0050]

[0051]

[0052] Where d is the distance from the joint rotation axis of the secondary robot to the center of mass of the secondary robot; (X1, Y1) and (X2, Y2) are the coordinates of the two points tangent to the circle respectively;

[0053] A4. Calculate the lower and upper limits of the rotation angle of the end rod of the robot arm based on the coordinates of the two points tangent to the circle:

[0054]

[0055] Among them, θ1 and θ2 are the lower and upper limits of the rotation angle of the end rod of the robot arm respectively;

[0056] A5. Use θ1<φ w <θ2 is used as the range constraint of the joint angle.

[0057] Furthermore, the expression of the rotational angular acceleration of the end rod of the manipulator is obtained by using PD feedback control:

[0058]

[0059] in, K is the angular acceleration of the end rod of the robot arm; p is the proportional gain; K d is the differential gain; e is the attitude error value; is the rate of change of the attitude error value.

[0060] Furthermore, the satellite body is a cubic structure, and three space robotic arms are evenly and orthogonally distributed on the satellite body. The rotation axis of the end rod of each robotic arm corresponds one-to-one to the three rotational degrees of freedom of the satellite body, and the end axis of each robotic arm is parallel to the X, Y, and Z axes of the satellite.

[0061] Compared with the prior art, the present invention has the following beneficial effects:

[0062] Achieve zero-fuel high-precision attitude and pointing adjustment: This solution uses the principle of mechanical arm and momentum transfer, and the coordination of information such as the angular acceleration and angular velocity of the mechanical arm and the satellite body. The satellite attitude and pointing can be adjusted through the movement of the mechanical arm, achieving high-precision attitude control without relying on fuel, avoiding the problems of flywheel unloading and thruster fuel consumption in traditional attitude adjustment methods. Achieving attitude adjustment through the movement of the mechanical arm greatly extends the on-orbit service life of the robotic satellite, and is particularly suitable for satellite application scenarios that perform long-term missions.

[0063] A decoupled robotic arm layout design is proposed: by evenly combining the space robotic arms on the CubeSat, a maximally decoupled design is formed, which enables high-precision attitude adjustment control by relying solely on the end of the robotic arm. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] Figure 1 This is the schematic diagram of the reaction flywheel scheme.

[0065] Figure 2 This is a schematic diagram of a chemical propulsion system.

[0066] Figure 3The figure is a flow chart of a satellite attitude pointing adjustment method based on multi-body momentum moment transfer.

[0067] Figure 4 Design diagram for satellite layout.

[0068] Figure 5 This is the geometric model diagram of the joint angle range under the controllable posture angle.

[0069] Figure 6 It is the attitude angle error diagram in the simulation embodiment. DETAILED DESCRIPTION

[0070] The specific implementation modes of the present invention are described below so that those skilled in the art can understand the present invention. However, it should be clear that the present invention is not limited to the scope of the specific implementation modes. For those of ordinary skill in the art, as long as various changes are within the spirit and scope of the present invention as defined and determined by the attached claims, these changes are obvious, and all inventions and creations utilizing the concept of the present invention are protected.

[0071] To facilitate the understanding of this solution, the following are definitions of some professional jargons in the technical solution:

[0072] 1. The attitude of a satellite refers to the direction of the satellite's body coordinate system in the spatial reference system. In order to conveniently describe the physical quantities of satellite motion and establish the corresponding satellite motion equations, multiple coordinate systems need to be established. Generally, the orbital coordinate system is used as the spatial reference coordinate system. The current attitude angle of the satellite is determined by the conversion relationship between the satellite's body coordinate system and the orbital coordinate system.

[0073] 2. The origin of the orbital coordinate system is set at the center of mass of the satellite, where OZ o The axis points to the center of the earth, OX o Axis and OZ o The axes are perpendicular to each other and lie in the orbital plane, pointing in the direction of the satellite's flight, and OY is established using the right-hand coordinate system rule. o The axis direction of the axis.

[0074] 3. The coordinate origin O of the body coordinate system is set at the center of mass of the satellite. The three axes of the coordinate system coincide with the principal axes of inertia of the satellite. The rotation angle between the satellite's body coordinate system and the orbital coordinate system is defined as the satellite's attitude angle.

[0075] 4. Quaternion is a four-dimensional number consisting of a real part and three imaginary parts, represented as q = (q0, q1, q2, q3), which is used to represent rotation operations in three-dimensional space because it can avoid the gimbal lock problem and support efficient calculation. Quaternion is often used in robotics and spacecraft attitude control.

[0076] 5. The attitude angle is used to describe the rotation angle of an object in three-dimensional space relative to a fixed coordinate system, usually expressed by Euler angles (Roll, Pitch, Yaw). The attitude angle can intuitively describe the rotation of an object around the x, y, and z axes, and is widely used in attitude control in the fields of aviation, aerospace, and robotics.

[0077] 6. The desired attitude is the ideal attitude angle or direction of a system (such as a satellite or robot) in the control target. It is the reference angle or direction used by the control system to guide or adjust the current attitude so that the system reaches the target position. The desired attitude is usually achieved through a feedback control system.

[0078] 7. Angular momentum is a measure of the rotational motion of an object and is equal to the product of the object's moment of inertia and its angular velocity. Angular momentum is conserved during rotation unless an external torque is applied to the object. The concept of angular momentum is used in physics and engineering to describe the inertial properties of rotating objects.

[0079] 8. A robotic arm is a mechanical device with multiple joints that can mimic the movements of a human arm. A robotic arm is usually composed of multiple rigid rods and rotating or sliding joints, and can be used for operations such as grasping, handling, and assembly. Robotic arms are widely used in industrial automation, medical and aerospace fields.

[0080] 9. Robotic satellite: refers to a space robot system in which the mass of the base is similar to that of the robotic arm. The robotic arm controls the posture and position of the base by adjusting its own movement to achieve specific mission objectives.

[0081] 10. Joints are the connection points between adjacent rigid rods in a robotic arm, allowing the two parts to move relative to each other. Joints can be revolute joints (providing rotational motion) or sliding joints (providing linear motion). Joints are the core components of the flexible movement of the robotic arm. Through the combination of multiple joints, the robotic arm can achieve complex spatial motion.

[0082] 11. The end rod is the farthest part of the robotic arm and is used to interact directly with objects or the environment.

[0083] 12. The satellite body is the main structural part of the satellite, including the main control system, propulsion system, communication system and other payloads. The satellite body is the core component of attitude control and structural support, and usually cooperates with the attitude control system to achieve space missions.

[0084] 13.Euler angles: A set of angular parameters that describe the posture of a rigid body in three-dimensional space, expressed as three angles (Roll, Pitch, and Yaw). These angles describe how an object rotates from an initial reference coordinate system to its current posture and are commonly used in aviation, aerospace, and robotics.

[0085] like Figure 4As shown, in this scheme, the satellite body is a cubic structure, and three space robotic arms are evenly and orthogonally distributed on the satellite body. The rotation axis of the end rod of each robotic arm corresponds to the three rotational degrees of freedom of the satellite body, and the end axis of each robotic arm is parallel to the X, Y, and Z axes of the satellite. After the satellite adopts the above structure, it can ensure the maximum decoupling between the robotic arms, so that each robotic arm can independently control the adjustment of a certain attitude angle, and through symmetry, the center of mass of the satellite is close to the center of mass of the satellite body, which is convenient for subsequent analysis.

[0086] refer to Figure 3 , Figure 3 A flow chart of a satellite attitude pointing adjustment method based on multi-body momentum moment transfer is shown; Figure 3 As shown, the method S includes steps S1 to S8.

[0087] In step S1, the attitude error value e=φ is calculated based on the target attitude angle and the initial attitude angle of the satellite rotation. target ―φ, where φ target is the target attitude angle, φ is the initial attitude angle of satellite rotation;

[0088] In step S2, based on the attitude error value, PD feedback control is used to obtain the rotational angular acceleration of the end rod of the manipulator:

[0089]

[0090] in, K is the angular acceleration of the end rod of the robot arm; p is the proportional gain; K d is the differential gain; e is the attitude error value; is the rate of change of the attitude error value.

[0091] In step S3, the rotational angular acceleration and angular velocity of the end rod of the manipulator and the rotational angular velocity of the satellite body are introduced into the attitude dynamics equation of the satellite based on the momentum moment, and the rotational angular acceleration of the satellite body is obtained by solving.

[0092] During implementation, the expression of the attitude dynamics equation of the satellite based on the angular momentum is as follows:

[0093]

[0094] Where, J is the moment of inertia of the satellite body; is the angular acceleration of the satellite body; ω bi is the angular velocity of the satellite body; J w is the moment of inertia of the end rod of the robot arm; ω w is the angular velocity of the end rod of the robot arm; T envis the environmental torque; is the angular acceleration of the end rod of the robot arm.

[0095] In step S4, the angular velocity of the satellite body and the angular velocity and the rotation angle of the end rod of the robotic arm are updated respectively according to the angular acceleration of the satellite body and the end rod of the robotic arm.

[0096] During implementation, this solution preferably updates the expressions of the angular velocity of the satellite body and the angular velocity and rotation angle of the end rod of the robotic arm as follows:

[0097]

[0098]

[0099] φ w (t) = φ w (t―1)+ω w Δt

[0100] Among them, ω bi (t) and ω bi (t―1) is the angular velocity of the satellite at time t and time t-1 respectively; ω w (t) and ω w (t―1) are the rotational angular velocities of the end rod of the manipulator at time t and time t-1 respectively; φ w (t) and φ w (t―1) are the rotation angles of the end rod of the robot arm at time t and time t-1 respectively; Δt is the discrete time step.

[0101] In step S5, it is determined whether the rotation angle satisfies the activity range constraint condition of the joint angle. If so, the process proceeds to step S6, otherwise, the satellite attitude pointing adjustment method is terminated.

[0102] Combine the following Figure 5 The geometric model of the joint angle range under the controllable posture angle is shown, and the method for obtaining the activity range constraint conditions of the joint angle in this scheme is described:

[0103] A1. The satellite's robotic arm is divided into a primary robotic arm and a secondary robotic arm. The primary robotic arm is the robotic arm between the satellite body and the end rod of the robotic arm; the secondary robotic arm is the end rod of the robotic arm. Assuming that the robotic arm includes 5 connecting rods, then the primary robotic arm is the first connecting rod to the fourth connecting rod; if the robotic arm includes 6 connecting rods, then the primary robotic arm is the first connecting rod to the fifth connecting rod.

[0104] A2. Calculate the coordinates of the satellite's center of mass based on the satellite's geometric model:

[0105]

[0106]

[0107]

[0108] Where α is the rotation angle of the first-stage manipulator relative to the satellite; r is the length of the first-stage manipulator; L is the length of the satellite body; η is the angle between the first-stage manipulator pointing vector and the rotation axis to the center of mass position vector; δ x , δ y is the uncertainty of the center of mass position of the equivalent satellite body caused by the other two robotic arms that are not located in the same plane; (X0, Y0) is the coordinate of the center of mass of the satellite body;

[0109] A3. Draw a circle with the joint rotation axis of the secondary manipulator as the center and the distance from the joint rotation axis of the secondary manipulator to the mass center of the secondary manipulator as the radius. Calculate the coordinates of the points where the two rays passing through the mass center of the satellite body are tangent to the circle:

[0110]

[0111]

[0112]

[0113]

[0114] Where d is the distance from the joint rotation axis of the secondary robot to the center of mass of the secondary robot; (X1, Y1) and (X2, Y2) are the coordinates of the two points tangent to the circle respectively;

[0115] A4. Calculate the lower and upper limits of the rotation angle of the end rod of the robot arm based on the coordinates of the two points tangent to the circle:

[0116]

[0117] Among them, θ1 and θ2 are the lower and upper limits of the rotation angle of the end rod of the robot arm respectively;

[0118] A5. Use θ1<φ w <θ2 is used as the range constraint of the joint angle.

[0119] In step S6, the real part and the imaginary part of the quaternion are updated according to the angular velocity component of the angular velocity of the satellite body on the xyz axis; based on the updated real part and the imaginary part, the three components of the attitude angle of the satellite rotation are calculated; step S6 further includes:

[0120] S61. Calculate the rate of change of the real and imaginary parts of the quaternion according to the angular velocity component of the satellite body's rotational angular velocity on the xyz axis:

[0121]

[0122] Among them, q0(t-1) is the real part of the quaternion at time t-1; q1(t-1), q2(t-1) and q3(t-1) are all the imaginary parts of the quaternion at time t-1; and are the rates of change of the real and imaginary parts of the quaternion at time t; ω x ,ω y and ω z They are ω bi Angular velocity components on the xyz axis;

[0123] S62. Update the real number of the quaternion according to the rate of change of the real part and the imaginary part of the quaternion:

[0124]

[0125] Among them, the value of k is 0, 1, 2 and 3; when k = 0, q0(t) and q0(t-1) are the real parts at time t and t-1 respectively. is the rate of change of the real part at time t; when k = 1, 2 and 3, q k (t) and q k (t―1) is the kth imaginary part at time t and t-1 respectively, is the rate of change of the imaginary part at time t;

[0126] S63. Calculate the three components of the satellite's rotation attitude angle according to the real and imaginary parts of the quaternion at time t:

[0127]

[0128] θ(t)=arcsin(―2(q1(t)q3(t)―q0(t)q2(t)))

[0129]

[0130] in, θ(t) and ψ(t) are the three components of the attitude angle of the satellite rotation at time t; q1(t), q2(t) and q3(t) are the imaginary parts of the quaternion corresponding to k = 1, 2 and 3 at time t, respectively.

[0131] In step S7, the attitude angle of the satellite rotation is calculated according to the three components, and the attitude error value between the attitude angle of the satellite rotation and the target attitude angle is calculated;

[0132] In step S8, it is determined whether the attitude error value is less than a preset threshold value. If so, the attitude pointing adjustment is performed according to the three components, otherwise, the process returns to step S2.

[0133] The following is a description of the effect of the satellite attitude pointing adjustment method of this scheme in combination with simulation:

[0134] Given the target attitude angle of [0.1, 0.08, 0.13] rad, assuming that the rotation angle α of the first-level manipulator of the three space manipulators of the satellite relative to the satellite is 120°, the initial angle of the joint rotation angle (that is, the initial attitude angle of the satellite rotation) is 81°, the damping coefficient of the joint maximum value is 1000N*m / deg, and the environmental torque interference is a sinusoidal interference with an amplitude of 0.1Nm, 0.13Nm, 0.08Nm and a frequency of 2Hz / s. By limiting the joint angle range and controlling it, the attitude angle error diagram is obtained as shown in the figure Figure 6 As shown, Figure 6 The nested graph in the figure is an enlarged graph of the attitude angle error at 14.5s to 15s, which is used to show the specific error accuracy. From the nested graph, we can see that the maximum accuracy error is less than 1x10e5.

[0135] exist Figure 6 The three lines represent the error changes of the three attitude angles, and the curves phi, theta and psi correspond to θ(t) and ψ(t), the horizontal axis is time t(s), and the vertical axis is the attitude angle error value (rad), which reflects that the three attitude angle errors of the satellite gradually tend to 0 with the control, which indicates the effectiveness of the control.

[0136] Combination Figure 6 It can be found that the satellite's attitude quickly reaches the ideal state and is able to overcome the interference of environmental torque. Finally, the stable accuracy can reach 1x10^-5, achieving zero fuel consumption and high-precision rapid alignment.

Claims

1. A satellite attitude pointing adjustment method based on multi-body momentum moment transfer, characterized in that: Includes steps: S1. Calculate the attitude error value according to the target attitude angle and the initial attitude angle of the satellite rotation; S2, based on the attitude error value, PD feedback control is used to obtain the rotational angular acceleration of the end rod of the manipulator; S3, bringing the rotational angular acceleration and angular velocity of the end rod of the manipulator arm and the rotational angular velocity of the satellite body into the attitude dynamics equation of the satellite based on the momentum moment, and solving to obtain the rotational angular acceleration of the satellite body; S4, according to the rotational angular acceleration of the satellite body and the end rod of the robotic arm, respectively update the rotational angular velocity of the satellite body and the rotational angular velocity and rotation angle of the end rod of the robotic arm; S5, judging whether the rotation angle satisfies the activity range constraint condition of the joint angle, if so, proceeding to step S6, otherwise terminating the satellite attitude pointing adjustment method; S6. Update the real part and the imaginary part of the quaternion according to the angular velocity component of the satellite body's rotation angular velocity on the xyz axis; calculate the three components of the satellite's rotation attitude angle based on the updated real part and the imaginary part; S7. Calculate the attitude angle of the satellite rotation according to the three components, and calculate the attitude error value between the attitude angle of the satellite rotation and the target attitude angle; S8. Determine whether the attitude error value is less than a preset threshold value. If so, adjust the attitude pointing according to the three components. Otherwise, return to step S2.

2. The satellite attitude pointing adjustment method according to claim 1, characterized in that: The expression of the attitude dynamics equation of the satellite based on the angular momentum is: Where, J is the moment of inertia of the satellite body; is the angular acceleration of the satellite body; ω bi is the angular velocity of the satellite body; J w is the moment of inertia of the end rod of the robotic arm; ω w is the angular velocity of the end rod of the robot arm; T env is the environmental torque; is the angular acceleration of the end rod of the robot arm.

3. The satellite attitude pointing adjustment method according to claim 2, characterized in that: The expressions for updating the angular velocity of the satellite body and the angular velocity and rotation angle of the end rod of the robotic arm are: f w (t)=φ w (t―1)+ω w Δt Among them, ω bi (t) and ω bi (t―1) is the angular velocity of the satellite at time t and time t-1 respectively; ω w (t) and ω w (t―1) are the rotational angular velocities of the end rod of the manipulator at time t and time t-1 respectively; φ w (t) and φ w (t―1) are the rotation angles of the end rod of the robot arm at time t and time t-1 respectively; Δt is the discrete time step.

4. The satellite attitude pointing adjustment method according to claim 2, characterized in that: Step S6 further comprises: S61. Calculate the rate of change of the real and imaginary parts of the quaternion according to the angular velocity component of the satellite body's rotational angular velocity on the xyz axis: Among them, q0(t-1) is the real part of the quaternion at time t-1; q1(t-1), q2(t-1) and q3(t-1) are all the imaginary parts of the quaternion at time t-1; and are the rates of change of the real and imaginary parts of the quaternion at time t; ω x ,ω y and ω z They are ω bi Angular velocity components on the xyz axis; S62. Update the real number of the quaternion according to the rate of change of the real part and the imaginary part of the quaternion: Among them, the value of k is 0, 1, 2 and 3; when k = 0, q0(t) and q0(t-1) are the real parts at time t and t-1 respectively. is the rate of change of the real part at time t; when k = 1, 2 and 3, q k (t) and q k (t―1) is the kth imaginary part at time t and t-1 respectively, is the rate of change of the imaginary part at time t; S63. Calculate the three components of the satellite's rotation attitude angle according to the real and imaginary parts of the quaternion at time t: θ(t)=arcsin(―2(q1(t)q3(t)―q0(t)q2(t))) in, θ(t) and ψ(t) are the three components of the attitude angle of the satellite rotation at time t; q1(t), q2(t) and q3(t) are the imaginary parts of the quaternion corresponding to k = 1, 2 and 3 at time t, respectively.

5. The satellite attitude pointing adjustment method according to claim 1, characterized in that: Methods for obtaining the range constraints of joint angles include: A1. The satellite's robotic arm is divided into a primary robotic arm and a secondary robotic arm. The primary robotic arm is the robotic arm between the satellite body and the end rod of the robotic arm; the secondary robotic arm is the end rod of the robotic arm; A2. Calculate the coordinates of the satellite's center of mass based on the satellite's geometric model: Where α is the rotation angle of the first-stage manipulator relative to the satellite; r is the length of the first-stage manipulator; L is the length of the satellite body; η is the angle between the first-stage manipulator pointing vector and the rotation axis to the center of mass position vector; δ x , δ y is the uncertainty of the center of mass position of the equivalent satellite body caused by the other two robotic arms that are not located in the same plane; (X0, Y0) is the coordinate of the center of mass of the satellite body; A3. Draw a circle with the joint rotation axis of the secondary manipulator as the center and the distance from the joint rotation axis of the secondary manipulator to the mass center of the secondary manipulator as the radius. Calculate the coordinates of the points where the two rays passing through the mass center of the satellite body are tangent to the circle: Where d is the distance from the joint rotation axis of the secondary robot to the center of mass of the secondary robot; (X1, Y1) and (X2, Y2) are the coordinates of the two points tangent to the circle respectively; A4. Calculate the lower and upper limits of the rotation angle of the end rod of the robot arm based on the coordinates of the two points tangent to the circle: Among them, θ1 and θ2 are the lower and upper limits of the rotation angle of the end rod of the robot arm respectively; A5. Use θ1<φ w <θ2 is used as the range constraint of the joint angle, φ w The rotation angle of the end rod of the robot arm.

6. The satellite attitude pointing adjustment method according to claim 1, characterized in that: The expression of the angular acceleration of the end rod of the manipulator obtained by PD feedback control is: in, K is the angular acceleration of the end rod of the robot arm; p is the proportional gain; K d is the differential gain; e is the attitude error value; is the rate of change of the attitude error value.

7. The satellite attitude pointing adjustment method according to any one of claims 1 to 6, characterized in that: The satellite body is a cubic structure, and three space robotic arms are evenly and orthogonally distributed on the satellite body. The rotation axis of the end rod of each robotic arm corresponds one-to-one to the three rotational degrees of freedom of the satellite body, and the end axis of each robotic arm is parallel to the X, Y, and Z axes of the satellite.

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