Method for establishing a multi-constraint relative pointing coordinate system of a space vehicle

Abstract

The application provides a method for establishing a multi-constraint relative pointing coordinate system of a space vehicle, comprising the following steps: calculating two rotation angles α and β to make the X axis of the vehicle point to the target according to the relative position vector of the vehicle and the target in the orbit system, to obtain a first pointing coordinate system; calculating the attitude quaternion q from the orbit system to the first pointing coordinate system according to the two rotation angles α and β o→m ; calculating the attitude quaternion q o→m from the orbit system to the first pointing coordinate system and the quaternion q o→b from the orbit system to the vehicle body system, to calculate the attitude quaternion q m→b and the attitude conversion matrix A m→b of the vehicle body system relative to the first pointing coordinate system; calculating a third rotation angle γ according to different constraints; rotating the angles α, β and γ in the order of 231 to obtain a second pointing coordinate system; and calculating the quaternion, angular velocity and feed-forward angular acceleration of the vehicle body system relative to the second pointing coordinate system, to provide necessary conditions for feed-forward control quantity calculation of large dynamic pointing tracking control of the vehicle.

Description

Technical Field

[0001] This invention relates to a method for establishing a multi-constraint relative pointing coordinate system for a spacecraft, belonging to the field of satellite attitude control. Background Technology

[0002] Spacecraft can perform target imaging and situational awareness, which has high civilian and military value. Continuous target tracking, especially when the relative speed between the spacecraft and the target is high, requires establishing a relative pointing coordinate system and calculating the relative angular velocity and angular acceleration of the two satellites for attitude tracking and control. In addition to ensuring the optical payload points to the target, other constraints on the spacecraft must also be met. These include adjusting the main maneuvering axis when pointing to the target; minimizing sunlight exposure to the heat dissipation surface to avoid affecting heat dissipation; ensuring the data transmission antenna is grounded as much as possible while pointing to the target; and minimizing the angular momentum loss of the actuators to reduce energy consumption while pointing to the target.

[0003] Currently, the main method for establishing a relative pointing coordinate system is to use the relative position vector as the pointing axis, with the other two axes defined through a cross product. This method can only satisfy the function of pointing to the target or has a single constraint function, and cannot be applied to different constraints under different tasks. At the same time, the angular velocity and angular acceleration of the body relative to the pointing system can only be obtained through differential methods, which have a lot of noise in practical engineering applications, resulting in poor pointing control accuracy or even unusable conditions. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a method for establishing a multi-constraint relative pointing coordinate system for a spacecraft. This provides a general method for establishing the relative pointing coordinate system of a spacecraft and provides the necessary conditions for solving the feedforward control quantity for the large dynamic pointing and tracking control of the spacecraft.

[0005] The technical solution of this invention is: a method for establishing a multi-constraint relative pointing coordinate system for a spacecraft, comprising the following steps:

[0006] S1. Based on the relative position vector between the aircraft and the target under the orbital system OX0Y0Z0, calculate the angle α of rotation around the OY0 axis of the orbital system to obtain the coordinate system OX1Y1Z1, and then the angle β of rotation around the OZ1 axis of the coordinate system OX1Y1Z1 to obtain the first pointing coordinate system OX2Y2Z2, so that the X-axis of the aircraft's own system points to the target at this time.

[0007] S2. Calculate the attitude quaternion q from the orbital system to the first pointing coordinate system based on the two rotation angles α and β. o→m ;

[0008] S3. Based on the attitude quaternion q from the orbital system to the first pointing coordinate system.o→m The quaternion q from the orbital system to the spacecraft's own system o→b Calculate the attitude quaternion q of the aircraft's self-frame relative to the first pointing coordinate system. m→b and attitude transformation matrix A m→b ;

[0009] S4. Calculate the rotation angle γ of the first pointing coordinate system OX2Y2Z2 around the OX2 axis according to different constraints. The constraints include ordinary target pointing constraints, angular momentum saving constraints, lighting constraints, data transmission pointing constraints, and rolling axis control constraints.

[0010] S5. Starting from the orbital system, rotate by angles α, β, and γ according to the 231 rotation sequence to obtain the second pointing coordinate system OX3Y3Z3. Calculate the quaternion q from the orbital system to the second pointing coordinate system. o→q and attitude transformation matrix A o→q ;

[0011] S6. Calculate the quaternion q of the aircraft's self-frame relative to the second pointing coordinate system. q→b The representation of the angular velocity and feedforward angular acceleration of the aircraft's intrinsic coordinate system relative to the second pointing coordinate system in the aircraft's intrinsic coordinate system provides the necessary conditions for the calculation of feedforward control quantities for the large dynamic pointing and tracking control of the aircraft.

[0012] Preferably, the method for calculating angles α and β in step S1 is as follows:

[0013] According to the description of the relative position vector between the spacecraft and the target in the orbital frame, r = [x o y o z o ] T ,

[0014]

[0015]

[0016] Where, x o The y-coordinate represents the relative position vector in the orbital system along the OX0 axis. o The z-axis represents the relative position vector in the orbital frame, along the OY0 axis. o This represents the OZ0 axis coordinate of the relative position vector in the orbital system.

[0017] Preferably, in step S2, the attitude quaternion q from the orbital system to the first pointing coordinate system is calculated based on the two rotation angles α and β. o→m :

[0018]

[0019] in, q represents quaternion multiplication. y Let q be the quaternion representation of the angle α of rotation about the OY0 axis. z The quaternion representation of the angle β of rotation about the OZ1 axis is as follows:

[0020]

[0021] Preferably, in step S3, the attitude quaternion q of the aircraft's own system relative to the first pointing coordinate system is calculated. m→b Specifically:

[0022]

[0023] Where, q m→o Let q be the attitude quaternion from the orbital system to the first pointing coordinate system. o→m The inverse of q m→o =q o→m -1 .

[0024] Preferably, in step S4, when considering the target pointing constraint, setting γ = 0 completes the real-time pointing function of the target;

[0025] When considering saving angular momentum constraints: Let γ be a constant value that can be modified. By adjusting the value of the γ angle, change the main maneuvering axis of the aircraft during the target pointing process, adjust the angular velocity required by the three axes during the target pointing process, and select the maneuvering axis with a smaller moment of inertia.

[0026] Preferably, in step S4, when considering illumination constraints, if the ±Y plane of the spacecraft's main system is a heat dissipation surface, the method for calculating the optimal value of γ using the solar vector is as follows:

[0027] The solar vector R under the spacecraft's own system sun_b =[x sun_b y sun_b z sun_b ] T Transform to the first pointing coordinate system to obtain the solar vector R in the first pointing coordinate system. sun_m :

[0028] R sun_m =A b→m R sun_b =[x sun_m y sun_m z sun_m ] T

[0029] in, x is the attitude transformation matrix from the aircraft's intrinsic coordinate system to the first pointing coordinate system; sun_b ,y sun_b ,z sun_bLet x be the three-axis coordinates of the solar vector in the spacecraft's intrinsic frame. sun_m ,y sun_m ,z sun_m Let the solar vector be the three-axis coordinates in the first pointing coordinate system; then we have:

[0030]

[0031] Preferably, in step S4, when considering the data transmission pointing constraint, the data transmission antenna vector is made coplanar with the vector pointing the aircraft towards the Earth during the target pointing process. The calculation method for the γ angle is as follows:

[0032] Let the data transmission antenna vector in the aircraft's own system be R. antenna_b =[x antenna_b y antenna_b z antenna_b ] T x antenna_b ,y antenna_b ,z antenna_b The data transmission antenna vector is represented by its three-axis coordinates within the aircraft's own system; the vector pointing from the aircraft to the Earth within the aircraft's own system is R. earth_b =[x earth_b y earth_b z earth_b ] T x earth_b ,y earth_b ,z earth_b The three-axis coordinates of the vector pointing from the center of mass of the spacecraft to the center of the Earth in the spacecraft's own system.

[0033] Transform the data transmission antenna vector and the Earth vector in the aircraft's own system to the first pointing coordinate system to obtain R. antenna_m R earth_m :

[0034] R antenna_m =A b→m R antenna_b =[x antenna_m y antenna_m z antenna_m ] T

[0035] R earth_m =A b→m R earth_b =[x earth_m y earth_m z earth_m ] T

[0036] Where, x antenna_m ,y antenna_m ,z antenna_m The data transmission antenna vector has three-axis coordinates in the first pointing coordinate system; xearth_m ,y earth_m ,z earth_m The vector pointing from the spacecraft to Earth is defined by the three axes of the first pointing coordinate system.

[0037] The method for calculating the angle γ is as follows:

[0038]

[0039] Preferably, in step S4, when considering roll axis control constraints, the attitude transformation matrix from the first pointing coordinate system to the aircraft's own system is used based on the current attitude of the aircraft. The method for calculating the γ angle is as follows:

[0040]

[0041] In the formula, A b→m (3,2) is matrix A b→m The element in the 3rd row and 2nd column.

[0042] Preferably, in step S5, the quaternion q from the orbital system to the second pointing coordinate system is calculated. o→q The specific method is as follows:

[0043]

[0044] in,

[0045] Preferably, in step S6, the quaternion q of the aircraft's self-frame relative to the second pointing coordinate system is calculated. q→b The angular velocity and feedforward angular acceleration of the aircraft's intrinsic system relative to the second pointing coordinate system are expressed in the aircraft's intrinsic system as follows:

[0046] S61. Calculate the quaternion q of the aircraft's self-frame relative to the second pointing coordinate system. q→b :

[0047]

[0048] S62. The angular velocity of the aircraft's intrinsic system relative to the second pointing coordinate system in the aircraft's intrinsic system, expressed as ω. q→b =[w q→bx w q→by w q→bz ] T :

[0049] ω q→b =ω i→b -A q→b ·(ω o→q +A o→q [0 -n 0])

[0050] In the formula, A q→b Let A represent the attitude transformation matrix from the second pointing coordinate system to the aircraft's own coordinate system. o→q ω represents the attitude transformation matrix from the orbital frame to the second pointing coordinate system. i→b ω is the inertial angular velocity of the aircraft, measured by a fiber optic gyroscope; n is the orbital angular velocity of the aircraft; w o→q The calculation method is as follows:

[0051]

[0052] w o→q Let be the angular velocity of the second pointing coordinate system relative to the orbital system in the second pointing coordinate system; where, The derivative of angle α; The derivative of angle β; The relative position vector [x] in the orbital system o y o z o ] T and relative velocity vector The relative velocity vector is calculated using Kalman filtering.

[0053]

[0054] S63. Feedforward angular acceleration, i.e., the angular acceleration of the second pointing coordinate system relative to the orbital system in the spacecraft's own system. Specifically:

[0055]

[0056] In the formula, These are the three-axis coordinates of the feedforward angular acceleration vector in the aircraft's own system. The difference of γ is obtained by first-order inertial filtering. The difference formula and the filtering formula are as follows:

[0057]

[0058]

[0059] The second derivative of angle α; The second derivative of angle β; The derivative of angle γ express The filter value; k represents the quantity in the current control cycle, k-1 represents the quantity in the previous control cycle, T s To control the cycle;

[0060]

[0061]

[0062] Where n is the orbital angular velocity of the spacecraft. These represent the three axes of the relative velocity vector in the orbital frame.

[0063] Compared with the prior art, the present invention has the following advantages:

[0064] (1) This invention can achieve target pointing while changing the third rotation angle γ, thereby changing the main maneuvering axis. That is, it can change the distribution of the relative motion velocity of the two satellites in the two non-pointing axes of the spacecraft's main system, adjusting the axis with the largest maneuvering angular velocity to the axis with the smallest moment of inertia of the spacecraft, or combining the small angular velocity attitude maneuvers of the two axes simultaneously to form a larger angular velocity attitude maneuver. This is beneficial to reducing the angular momentum capability requirement of the momentum wheel, thereby reducing the weight of the entire satellite; it can make the data transmission antenna ground-oriented, making the angle between the data transmission antenna and the earth as small as possible, which is beneficial to increasing the available time for on-board data downlink and ground command uplink, and improving the probability of mission success; it can make the heat dissipation surface shielded from sunlight, minimizing the sun's exposure to the heat dissipation surface, which is beneficial to better heat dissipation, reducing the burden on thermal control, and saving thermal control power consumption; it can make the rolling axis less controlled, reducing the angular momentum output of the rolling axis, which is beneficial to saving energy. If the momentum wheel is configured as CMG, it is also beneficial to avoid CMG outer frame singularity and improve the available angular momentum capability of the CMG's non-pointing axis; it provides a general method for establishing the relative pointing coordinate system of the spacecraft.

[0065] (2) This invention solves the analytical solutions of the attitude quaternion, relative angular velocity and feedforward angular acceleration of the aircraft relative to the established coordinate system that satisfies the constraints, providing the necessary conditions for the calculation of feedforward control quantities for the large dynamic pointing and tracking control of the aircraft. Compared with the traditional pointing system obtained by cross product, the relative angular velocity and feedforward acceleration obtained by difference have the disadvantage of large noise. The analytical solution has higher accuracy, and thus the attitude control accuracy is higher. Attached Figure Description

[0066] Figure 1 This is a schematic diagram illustrating the steps of the method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to the present invention.

[0067] Figure 2 This represents the transformation relationship from the orbital system to the first pointing coordinate system OX2Y2Z2, and then to the second pointing coordinate system OX3Y3Z3 in this invention. Detailed Implementation

[0068] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings.

[0069] A spacecraft in a GEO orbit has its camera payload mounted on the +X plane of its mainframe. Like most spacecraft, its heat dissipation surface is located on the ±Y plane of its mainframe. The data transmission antenna is mounted on the +Z plane. During mission execution, the +X plane must be pointed towards the target in real time. The angle between the solar vector and the heat dissipation surface must not exceed 10° to avoid affecting heat dissipation.

[0070] like Figure 1 The diagram illustrates the steps involved in establishing a multi-constraint relative pointing coordinate system for a spacecraft. An example of implementing this method is shown below.

[0071] S1. Based on the relative position vectors of the spacecraft and the target in the orbital frame OX0Y0Z0, calculate the two rotation angles that make the X-axis of the spacecraft's own frame point towards the target. First, rotate the spacecraft around the OY0 axis of the orbital frame by an angle α to obtain the coordinate system OX1Y1Z1. Then, rotate the spacecraft around the OZ1 axis of the coordinate system OX1Y1Z1 by an angle β to obtain the coordinate system OX2Y2Z2. Define the coordinate system OX2Y2Z2 as the first pointing coordinate system.

[0072] The rotational relationships between the orbital system OX0Y0Z0, the coordinate system OX1Y1Z1, and the first pointing coordinate system OX2Y2Z2 are as follows: Figure 2 As shown.

[0073] According to the description of the relative position vector between the spacecraft and the target in the orbital frame, r = [x o y o z o ] T Calculate the angle α of rotation about the OY0 axis of the orbital system and the angle β of rotation about the OZ1 axis of the coordinate system OX1Y1Z1.

[0074]

[0075]

[0076] Where, x o This represents the coordinates of the relative position vector along the OX1 axis in the orbital system, y o This represents the relative position vector in the orbital system along the OY1 axis, z. o This represents the OZ1 axis coordinate of the relative position vector in the orbital system.

[0077] S2. Calculate the attitude quaternion q from the spacecraft's orbital system to the first pointing coordinate system. o→m ;

[0078] Calculate the attitude quaternion q from the orbital system to the first pointing coordinate system based on the two rotation angles α and β. o→m ;

[0079]

[0080] in, q represents quaternion multiplication. y Let q be the quaternion of the attitude from the orbital system to the coordinate system OX1Y1Z1. z The attitude quaternion from coordinate system OX1Y1Z1 to the first pointing coordinate system OX2Y2Z2:

[0081]

[0082] S3. Calculate the attitude quaternion q of the aircraft's self-frame relative to the first pointing coordinate system. m→b and attitude transformation matrix A m→b The specific process is as follows:

[0083] S31. Calculate the attitude quaternion q of the aircraft's self-frame relative to the first pointing coordinate system. m→b .

[0084]

[0085] Where, q o→b Let q be the quaternion from the orbital system to the spacecraft's intrinsic system. m→o Let q be the attitude quaternion from the orbital system to the first pointing coordinate system. o→m The inverse of q m→o =q o→m -1 .

[0086] S32. Based on the attitude quaternion q m→b Calculate the attitude transformation matrix A from the first pointing coordinate system to the aircraft's own coordinate system. m→b .

[0087] Remember q m→b =[q0 q1 q2 q3] T Then we have:

[0088]

[0089] S4. Calculate the rotation angle γ of the OX2 axis around the first pointing coordinate system OX2Y2Z2 according to different constraints. The constraints include ordinary target pointing constraints, angular momentum saving constraints, lighting constraints, data transmission pointing constraints, and rolling axis (X-axis) control constraints.

[0090] Specifically as follows:

[0091] S41. Let γ = 0 or a constant value that can be modified.

[0092] When considering the real-time pointing constraint of the target: setting γ=0 can complete the real-time pointing function of the target.

[0093] When considering saving angular momentum: Let γ be a constant that can be modified. When the orbits of the spacecraft and the target satellite remain unchanged, the relative motion angle, angular velocity, and angular acceleration of the two satellites can be determined. By adjusting the value of the γ angle, the main maneuvering axis during the spacecraft's pointing towards the target can be changed, i.e., the angular velocity required for the three axes during the pointing process can be adjusted. By selecting the maneuvering axis with a smaller moment of inertia, the purpose of saving angular momentum can be achieved. Since the determination method of the γ angle varies depending on the relative orbital conditions, it needs to be determined by ground simulation in advance. For example, under the condition of a rendezvous orbit, the Y-axis is the main maneuvering axis when γ = 0, while under the condition of a rendezvous orbit, the Y-axis is the main maneuvering axis when γ = 90°.

[0094] S42. Using real-time measured solar vectors and considering illumination constraints—that is, aiming to minimize or eliminate solar radiation on a certain surface of the spacecraft—calculate the optimal value of γ. Taking the ±Y surface of the spacecraft as the heat dissipation surface as an example:

[0095] The solar vector R under the spacecraft's own system sun_b =[x sun_b y sun_b z sun_b ] T Transform to the first pointing coordinate system to obtain the solar vector R in the first pointing coordinate system. sun_m :

[0096] R sun_m =A b→m R sun_b =[x sun_m y sun_m z sun_m ] T (6)

[0097] Among them, A b→m =A m T →b , is the attitude transformation matrix from the aircraft's own coordinate system to the first pointing coordinate system. sun_b ,y sun_b ,z sun_b x represents the three-axis coordinates of the solar vector within the spacecraft's intrinsic frame. sun_m ,y sun_m ,z sun_m Let be the three-axis coordinates of the solar vector in the first pointing coordinate system. Then we have:

[0098]

[0099] It can achieve the function of avoiding sunlight in real time on the ±Y plane during the target pointing process.

[0100] S43. Considering the data transmission pointing constraint, calculate the γ angle in real time: Based on the plane where the aircraft's data transmission antenna is installed, ensure that during the pointing process towards the target, the data transmission antenna vector and the vector pointing the aircraft towards the Earth are coplanar, thus creating conditions for real-time data transmission to the ground.

[0101] Let the data transmission antenna vector in the aircraft's own system be R. antenna_b =[x antenna_b y antenna_b z antenna_b ] T x antenna_b ,y antenna_b ,z antenna_b Here are the three-axis coordinates of the data transmission antenna vector within the aircraft's own system. The vector pointing from the aircraft to the Earth within the aircraft's own system is R. earth_b =[x earth_b y earth_b z earth_b ] T x earth_b ,y earth_b ,z earth_b Let γ be the three-axis coordinates of the vector pointing from the spacecraft's center of mass to the Earth's center in the spacecraft's intrinsic frame. To align the data transmission antenna as close to the ground as possible, it is necessary to calculate the angle γ so that the projection vector of the spacecraft's Earth-pointing vector in the spacecraft's intrinsic YOZ plane coincides with the projection vector of the data transmission antenna vector in the spacecraft's intrinsic YOZ plane.

[0102] Transform the data transmission antenna vector and the Earth vector in the aircraft's own system to the first pointing coordinate system to obtain R. antenna_m R earth_m

[0103] R antenna_m =A b→m R antenna_b =[x antenna_m y antenna_m z antenna_m ] T (8)

[0104] R earth_m =A b→m R earth_b =[x earth_m y earth_m z earth_m ] T (9)

[0105] Then the angle γ can be calculated as:

[0106]

[0107] Where, x antenna_m ,y antenna_m ,zantenna_m Let x be the three-axis coordinates of the data transmission antenna vector in the first pointing coordinate system. earth_m ,y earth_m ,z earth_m The vector pointing from the spacecraft to Earth is represented by the three axes of the first pointing coordinate system.

[0108] If there are constraints on the direction of the relay satellite in the data transmission, then the Earth vector can be replaced with the vector pointing from the spacecraft to the relay satellite.

[0109] S44. Considering the roll axis control constraints, calculate the attitude transformation matrix from the first pointing coordinate system to the aircraft's own system based on the current attitude of the aircraft. Then, the 231 rotation order Euler angle from the first pointing coordinate system to the aircraft's main system is calculated, and the rotation order angle of 1 is taken as γ. At this point, angle γ minimizes the attitude control torque output of the rolling shaft, achieving the goal of saving angular momentum and energy; its calculation method is as follows:

[0110]

[0111] In the formula, A b→m (3,2) is matrix A b→m The element in the 3rd row and 2nd column.

[0112] S5. Based on the three rotation angles α, β, and γ, following the 231 rotation sequence, calculate the following: first, rotate the orbital system around the Y0 axis of the orbital system OX0Y0Z0 by an angle α to obtain the coordinate system OX1Y1Z1; then, rotate the coordinate system around the Z1 axis of the coordinate system OX1Y1Z1 by an angle β to obtain the first pointing coordinate system OX2Y2Z2; finally, rotate the coordinate system around the X2 axis of the coordinate system OX2Y2Z2 by an angle γ to obtain the second pointing coordinate system, denoted as coordinate system OX3Y3Z3. Calculate the attitude transformation matrix A from the orbital system to the second pointing coordinate system. o→q and quaternion q o→q ;

[0113] S51. Based on the angles α, β, and γ of the three rotations, calculate the attitude quaternion q from the orbital system to the second pointing coordinate system in the 231 rotation sequence. o→q . Figure 2 The coordinate system OX3Y3Z3 is the second pointing coordinate system.

[0114]

[0115] in, q o→m Let be the attitude quaternion from the orbital system to the first pointing coordinate system.

[0116] S52. Based on the quaternion q o→q Calculate the attitude transformation matrix A from the orbital frame to the second pointing coordinate system. o→qThe quaternion-to-attitude conversion matrix will not be elaborated upon further.

[0117] S6. Calculate the quaternion q of the aircraft's self-frame relative to the second pointing coordinate system. q→b The angular velocity and angular acceleration of the aircraft's main system relative to the second pointing coordinate system are represented in the aircraft's main system.

[0118] S61. Calculate the quaternion q of the aircraft's self-frame relative to the second pointing coordinate system. q→b .

[0119]

[0120] S62. The angular velocity of the aircraft's intrinsic system relative to the second pointing coordinate system in the aircraft's intrinsic system, expressed as ω. q→b =[w q→bx w q→by w q→bz ] T .

[0121] ω q→b =ω i→b -A q→b ·(ω o→q +A o→q [0-n 0]) (14)

[0122] In the formula, A q→b Let A represent the attitude transformation matrix from the second pointing coordinate system to the aircraft's own coordinate system. o→q ω represents the attitude transformation matrix from the orbital frame to the second pointing coordinate system. i→b The inertial angular velocity of the aircraft can be measured by a fiber optic gyroscope.

[0123]

[0124] w o→q The angular velocity of the second pointing coordinate system relative to the orbital system is expressed in the second pointing coordinate system.

[0125] n is the orbital angular velocity of the spacecraft; It can be determined by the relative position vector [x] in the orbital system. o y o z o ] T and relative velocity vector The relative velocity vector can be calculated using Kalman filtering.

[0126]

[0127] In the formula, The derivative of angle α; The derivative of angle β;

[0128] S63. Feedforward angular acceleration, i.e., the angular acceleration of the second pointing coordinate system relative to the orbital system in the spacecraft's own system, is expressed as follows:

[0129]

[0130] In the formula, These are the three-axis coordinates of the feedforward angular acceleration vector in the aircraft's intrinsic system. The difference of γ is obtained by first-order inertial filtering. The difference formula and the filtering formula are as follows:

[0131]

[0132]

[0133] The second derivative of angle α; The second derivative of angle β; The derivative of angle γ express The filter value. k represents the quantity in the current control cycle, k-1 represents the quantity in the previous control cycle, and T... s To control the cycle.

[0134]

[0135]

[0136] Where n is the orbital angular velocity of the spacecraft. These represent the three axes of the relative velocity vector in the orbital frame.

[0137] The contents not described in detail in this specification are existing technologies known to those skilled in the art.

Claims

1. A method for establishing a multi-constrained relative pointing coordinate system for a space vehicle, characterized by Includes the following steps: S1. Based on the relative position vectors of the spacecraft and the target in the orbital system OX0Y0Z0, calculate the angle of rotation around the OY0 axis of the orbital system. α Obtain the coordinate system OX1Y1Z1, and then rotate it around the OZ1 axis of the coordinate system OX1Y1Z1 by an angle. β We obtain the first pointing coordinate system OX2Y2Z2, so that the X-axis of the aircraft's own system points to the target at this time; S2, based on the two corners α、β Calculate the attitude quaternion from the orbital frame to the first pointing coordinate system. ; S3. Based on the attitude quaternions from the orbital system to the first pointing coordinate system. Quaternions from orbital system to spacecraft system Calculate the attitude quaternions of the aircraft's self-frame relative to the first pointing coordinate system. and attitude transformation matrix ; S4. Calculate the rotation angle of the first pointing coordinate system OX2Y2Z2 around the OX2 axis according to different constraints. γ The constraints include general target pointing constraints, angular momentum saving constraints, lighting constraints, data transmission pointing constraints, and fewer control constraints for rolling axes; S5. Starting from the orbital system, rotate the angle according to the 231 revolution sequence. α, β, γ Obtain the second pointing coordinate system OX3Y3Z3, and calculate the quaternion from the orbital system to the second pointing coordinate system. and attitude transformation matrix ; S6. Calculate the quaternion of the aircraft's self-frame relative to the second pointing coordinate system. The representation of the angular velocity and feedforward angular acceleration of the aircraft's intrinsic system relative to the second pointing coordinate system in the aircraft's intrinsic system provides the necessary conditions for the calculation of feedforward control quantities for the large dynamic pointing and tracking control of the aircraft. In step S4, when considering the constraint on the target orientation, let At that time, it completes the real-time pointing function of the target; When considering the conservation of angular momentum constraint: Let This is a constant value that can be modified by adjusting the number of entries. The angle value changes the main maneuver axis during the aircraft's target pointing process, adjusts the angular velocity required for the three axes during the target pointing process, and selects the maneuver axis with a smaller moment of inertia; In step S4, when considering illumination constraints, if the ±Y planes of the spacecraft's main system are heat dissipation surfaces, the calculation is combined with the solar vector. The method for finding the optimal value is as follows: The solar vector under the spacecraft's own system Transform to the first pointing coordinate system to obtain the solar vector in the first pointing coordinate system. R sun_m : in, , is the attitude transformation matrix from the aircraft's own coordinate system to the first pointing coordinate system; The three-axis coordinates of the solar vector in the spacecraft's intrinsic frame. Let the solar vector be the three-axis coordinates in the first pointing coordinate system; then we have: In step S4, considering the data transmission pointing constraint, the data transmission antenna vector is made coplanar with the vector pointing the aircraft towards the Earth during the target pointing process. The method for calculating angles is as follows: Let the data transmission antenna vector in the aircraft's main system be... , The data transmission antenna vector is defined by its three-axis coordinates within the aircraft's own system; the vector pointing from the aircraft to the Earth within the aircraft's own system is... , The three-axis coordinates of the vector pointing from the center of mass of the spacecraft to the center of the Earth in the spacecraft's own system. Transform the data transmission antenna vector and the Earth vector in the aircraft's own system to the first pointing coordinate system, and obtain... R antenna_m , R earth_m : in, The data transmission antenna vector is represented by its three-axis coordinates in the first pointing coordinate system. The vector pointing from the spacecraft to Earth is defined by the three axes of the first pointing coordinate system. but The method for calculating angles is as follows: In step S4, when considering roll axis control constraints, based on the current attitude of the aircraft, the attitude transformation matrix from the first pointing coordinate system to the aircraft's own system is used. ,calculate The method for angles is as follows: In the formula, For matrix The element in the 3rd row and 2nd column.

2. The method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to claim 1, characterized in that: In step S1 α、β The angle calculation method is as follows: Based on the description of the relative position vectors of the aircraft and the target in the orbital frame , ； in, This represents the coordinates of the relative position vector along the OX0 axis in the orbital system. This represents the OY0 axis coordinates of the relative position vector in the orbital system. This represents the OZ0 axis coordinate of the relative position vector in the orbital system.

3. The method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to claim 1, characterized in that: In step S2, based on the two corners α、β Calculate the attitude quaternion from the orbital system to the first pointing coordinate system. : in," " represents quaternion multiplication, For rotation about the OY0 axis α Quaternion representation of angles, For rotation about the OZ1 axis β The quaternion representation of angles is as follows: 。 4. The method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to claim 1, characterized in that: In step S3, the attitude quaternions of the aircraft's self-frame relative to the first pointing coordinate system are calculated. Specifically: in, The attitude quaternion from the orbital system to the first pointing coordinate system The reverse, that is .

5. The method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to claim 1, characterized in that: In step S5, the quaternion from the orbital system to the second pointing coordinate system is calculated. The specific method is as follows: in, .

6. The method for establishing a multi-constraint relative pointing coordinate system for a spacecraft according to claim 1, characterized in that: In step S6, the quaternion of the aircraft's self-frame relative to the second pointing coordinate system is calculated. The angular velocity and feedforward angular acceleration of the aircraft's intrinsic system relative to the second pointing coordinate system are expressed in the aircraft's intrinsic system as follows: S61. Calculate the quaternion of the aircraft's self-frame relative to the second pointing coordinate system. : S62. Representation of the angular velocity of the aircraft's intrinsic frame relative to the second pointing coordinate system in the aircraft's intrinsic frame. : In the formula, This represents the attitude transformation matrix from the second pointing coordinate system to the aircraft's own coordinate system. This represents the attitude transformation matrix from the orbital frame to the second pointing coordinate system. The inertial angular velocity of the aircraft is measured by a fiber optic gyroscope. n The orbital angular velocity of the spacecraft; The calculation method is as follows: Let be the angular velocity of the second pointing coordinate system relative to the orbital system in the second pointing coordinate system; where, Indicates angle α The derivative; Indicates angle β The derivative; , The relative position vector in the orbital frame and relative velocity vector The relative velocity vector is calculated using Kalman filtering. ； S63. Feedforward angular acceleration, i.e., the angular acceleration of the second pointing coordinate system relative to the orbital system in the spacecraft's own system. Specifically: In the formula, , , These are the three-axis coordinates of the feedforward angular acceleration vector in the aircraft's own system. for The difference is obtained by first-order inertial filtering. The difference formula and the filtering formula are as follows: Indicates angle α The second derivative; Indicates angle β The second derivative; Indicates angle The derivative, express The filtered value; k This indicates the quantity in the current control cycle. k -1 indicates the amount from the previous control cycle. T s To control the cycle; in, n The orbital angular velocity of the spacecraft. , , These represent the three axes of the relative velocity vector in the orbital frame.

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