CN113830332B - Ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer - Google Patents

Abstract

The invention relates to an ignition attitude establishment and dynamic tracking method for electric push orbit transfer. The steps include: (1) establishing a two-stage orbit control coordinate system for electric push orbit transfer; (2) establishing electric push ignition based on an arbitrary reference coordinate system Target attitude; (3) Attitude tracking to the sun during the electropush orbit transfer. The present invention is based on the orbit control coordinate system at different stages of electric propulsion orbit transfer, through the attitude offset of the axis where the thrust direction is controlled around the orbit, it can effectively keep the attitude tracking of the sun plane during the ignition process, and meet the energy demand of the attitude tracking the direction of the sun. The problem of multi-target attitude control during electric push ignition is solved permanently.

Description

An ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer

technical field

The invention relates to an ignition attitude establishment and dynamic tracking method for electric push orbit transfer, belonging to the field of spacecraft attitude orbit control.

Background technique

The combined orbit change of chemical propulsion and electric propulsion is adopted, that is, after the separation of the satellite and rocket, the chemical apogee engine is used to realize multiple orbit changes, and after the satellite is sent into a handover orbit at a certain height and a certain inclination angle, electric propulsion is used to realize the orbit during the subsequent orbit transfer. Lifting and orbital rounding. The electric propulsion orbit lifting mode is used to transfer the orbit using electric propulsion to raise the semi-major axis, reduce the eccentricity and reduce the orbit inclination.

According to the electric propulsion orbit transfer strategy, it is mainly divided into the stage of raising the semi-major axis of the orbit and the stage of rounding the orbit. In the stage of raising the semi-major axis of the orbit, the thrust is along the plane formed by the velocity direction (or vertical radial direction) and the orbital angular momentum, and the orbital plane. A certain included angle is used to adjust the orbital inclination; the thrust direction is perpendicular to the semi-major axis during the orbital rounding stage, and a certain included angle with the orbital plane is also used to adjust the orbital inclination. Therefore, the attitude of the electric push orbit change target is complicated.

Due to the small thrust of the electric thruster, it takes a long time to change the orbit of the electric thruster every turn, and the visibility of the arc section is complicated, which requires a complete and reliable independent processing process on the star. At the same time, when electric propulsion is used for orbit transfer, it needs to work in a high-power and high-thrust mode, which has high requirements for the energy of the whole star. It needs to meet the joint adjustment of attitude and sailboard, and automatically align with the sun to ensure energy. The energy angle is defined as the angle between the normal direction of the solar wing and the direction of the sun, and there is a clear precision requirement for the energy angle during the electric propulsion orbit transfer.

Contents of the invention

The technical problem solved by the present invention is: to overcome the deficiencies of the prior art, to propose an ignition attitude establishment and dynamic tracking method for electric push track transfer, so that the ignition attitude combines the requirements of improving the semi-major axis or reducing the eccentricity of the orbit, Adjusting the ignition direction by lowering the orbital inclination, and dynamically tracking the multi-target requirements of the energy demand in the direction of the sun can accurately establish the attitude of the electric push ignition target, and realize long-term and effective dynamic tracking of the attitude towards the sun.

The technical solution of the present invention is:

An ignition attitude establishment and dynamic tracking method for electric push orbit transfer, the steps include:

(1) Establish the orbit control coordinate system for the two phases of electric propulsion orbit transfer

The electric propulsion orbit transfer is divided into two stages, the reference attitude of each stage is set as the orbit control coordinate system of this stage, the electric thruster is installed on the -Z plane of the satellite body system, and the thrust direction is along the Z axis of the satellite body system;

(2) Establish the attitude of the electric push ignition target based on any reference coordinate system

θ_br，ψ_br，将其设置为三轴姿态控制的目标姿态角，则可基于任意参考坐标系将卫星控制到电推点火的目标姿态；Under any reference coordinate system datum, by solving the Euler angle

θ _br , ψ _br , set it as the target attitude angle of the three-axis attitude control, then the satellite can be controlled to the target attitude of electric push ignition based on any reference coordinate system;

(3) Attitude tracking to the sun during the orbit transfer period

After step (2) establishes the attitude of the electric push ignition target from any reference coordinate system, it is converted to the orbit control coordinate system obtained in step (1) as the reference coordinate system to perform attitude control during the electric push orbit transfer; at the same time, the electric push During the orbit transfer, in order to meet the energy requirements, it is necessary to realize the normal direction of the sun wing to the sun by rotating around the axis of the ignition direction and around the sun wing;

Through attitude tracking to the sun and solar panel solar rotation angle control, the angle between the normal direction of the solar panel and the direction of the sun, that is, the energy angle, is controlled within the range of accuracy requirements, and the normal direction of the solar panel can be independently aligned with the sun to ensure energy. .

Further, in step (1), the electric propulsion orbit transfer is implemented, that is, the transfer orbit uses electric propulsion to raise the semi-major axis, reduce the eccentricity and lower the orbit inclination.

Further, in step (1), the orbit transfer goal of the first stage is to increase the semi-major axis and at the same time press the inclination angle. The orbit control thrust in this stage is within the XoOoYo plane of the orbit coordinate system, and maintains a certain angle with +Xo, and its absolute The value does not change to Ψ1, and the range of Ψ1 is 0°-90°, but at every half orbital period, the exact value of the argument depends on the orbit control strategy, and the corresponding argument changes when it is 90° and 270°.

Further, the reference attitude of the attitude control in the first stage, that is, the orbit control coordinate system is set as: first rotate around the Z axis of the orbit coordinate system by an angle Ψ1, which is called the orbit control yaw angle, and then rotate around the Y axis of the orbit coordinate system 90°, so that the -X axis of the orbit control coordinate system points to the ground. At this time, the angle between the +Z axis of the orbit control coordinate system and the forward direction is the required orbit control yaw angle. Change the sign of the angle at 90° and 270°.

Further, in step (1), the transformation matrix C _ORMi1 of the orbit control coordinate system relative to the inertial system in the first stage of the electric push orbit transfer is obtained by rotating the orbit coordinate system C _oi , namely

Among them, Ψ1 is the orbit control yaw angle.

Further, in step (1), the goal of the orbit transfer in the second stage is to reduce the eccentricity and at the same time press the inclination angle. In this stage, the orbit control thrust maintains a certain orbit control angle on the basis of inertial orientation, so as to further adjust the orbit inclination angle , its absolute value does not change to Ψ2, but changes its sign when the argument is 90° and 270°.

Further, in step (1), the reference attitude of the attitude control in the first stage, that is, the orbit control coordinate system is obtained based on the near-focus coordinate system PQW rotation.

Further, in the near focus coordinate system PQW, the P axis points to the perigee direction of the orbit, the W axis points to the positive normal direction of the orbit, and the Q axis, the P axis and the W axis form a right-handed orthogonal coordinate system.

Further, the corresponding relationship between the orbit control coordinate system and the axes of the PQW coordinate system is that the X axis of the orbit control coordinate system corresponds to the PQW coordinate system -P axis, the orbit control coordinate system Y axis corresponds to the PQW coordinate system -W axis, and the orbit control coordinate system Z The axis corresponds to the PQW coordinate system -Q. After converting the corresponding relationship of the coordinate axes, and then rotating Ψ2 around the X axis of the current coordinate system, the orbit control coordinate system C _ORMi2 of the second stage of the electric push orbit transfer is obtained.

Further, first calculate the transformation matrix C _PQWi from the inertial system to the PQW coordinate system according to the orbital elements

Where ω is the argument of perigee, Ω is the right ascension of the ascending node of the orbit, and i is the inclination of the orbit;

Further, the orbit control coordinate system C _ORMi2 of the second stage of electric propulsion orbit transfer can be obtained,

in:

sψ ₂ ＝sin(ψ ₂ ),cψ ₂ ＝cos(ψ ₂ )

sΩ＝sin(Ω),cΩ＝cos(Ω)

si=sin(i), ci=cos(i)

sω=sin(ω),cω=(cosω)

G _k1 =cΩ·sω+sΩ·cω·ci

G _k2 =sΩ·sω−cΩ·cω·ci.

Further, in step (2), first, after the ignition attitude of the electric propulsion orbit transfer is established, the attitude transformation matrix C _br of the satellite system relative to the reference coordinate system should satisfy

C _br ·C _ri =C _ORMi1 or C _br ·C _ri =C _ORMi2

θ_br，ψ_br，则有where C _ri is the attitude conversion matrix from the inertial system to the reference coordinate system, and the three-axis attitude Euler angle of the satellite in the reference coordinate system is

θ _br , ψ _br , then we have

或

or

When establishing the target attitude of electric propulsion ignition for the first stage of electric propulsion orbit transfer, order

When establishing the electric push ignition target attitude for the second stage of electric push orbit transfer, order

θ_br，ψ_br。Given the sequence of Euler angles, it can be obtained by C _br

θ _br , ψ _br .

Furthermore, when the Euler rotation sequence is 3-2-1,

ψ _br =tan2 ^-1 (c ₁₂ ,c ₁₁ )∈[-π,π]

where the function tan2 ^-1 (y,x) is defined as

When the Euler rotation sequence is 3-1-2,

θ _br =tan2 ^-1 (-c ₁₃ ,c ₃₃ )∈[-π,π]

ψ _br =tan2 ^-1 (-c ₂₁ ,c ₂₂ )∈[-π,π]

The rest of the Euler rotation sequence is the same.

Further, if the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbital system reference, and the sun direction changes with time. In order to track the energy source, it is necessary to realize the attitude dynamic offset, that is, during the electric propulsion orbit change of the transfer orbit, the +Z axis of the satellite points to Orbit control thrust direction, when the sailboard is required to point to the sun, the system rotates a certain angle around the Z axis of the orbit control coordinate system, so that the XOZ plane of the satellite coincides with the plane composed of the sun direction and thrust direction.

Further, the expression S _i of the sun vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris, and the transformation matrix C _ORMi1 of the orbit control coordinate system relative to the inertial system in the first stage of the electric orbit transfer is converted to the sun vector Expressed in the orbit control coordinate system of the first stage, namely

S _ORM1 ＝[S _ORM1X S _ORM1Y S _ORM1Z ] ^T ＝C _ORMi1 S _i

Then the attitude control target is based on the orbit control coordinate system of the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is ψ _d =tan2 ^-1 (S _ORM1Y , S _ORM1X )∈[-π,π].

Further, if the electric propulsion orbit transfer is in the second stage, the ignition attitude is based on the inertial system reference, and the sun vector direction is fixed in the short term, so the attitude offset angle to ensure energy is also fixed, that is, during the electric propulsion orbit change of the transfer orbit, the satellite The +Z axis points to the thrust direction of the orbit control. When the sail is required to point to the sun, the system rotates around the Z axis of the orbit control coordinate system at a certain angle, so that the XOZ plane of the satellite coincides with the plane composed of the sun direction and the thrust direction.

Further, the expression S _i of the sun vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris, and the conversion matrix C _ORMi2 of the orbit control coordinate system relative to the inertial system in the first stage of the electric push orbit transfer is used to convert the sun vector Expressed in the orbit control coordinate system of the first stage, namely

S _ORM2 ＝[S _ORM2X S _ORM2Y S _ORM2Z ] ^T ＝C _ORMi2 S _i

Then the attitude control target is based on the orbit control coordinate system of the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is ψ _d =tan2 ^-1 (S _ORM2Y , S _ORM2X )∈[-π,π].

Furthermore, the satellite system is defined as: taking a certain feature point of the satellite as the origin, the X-axis and Z-axis are along the direction of the spacecraft’s characteristic axis, and the Y-axis is perpendicular to the Z-axis and X-axis, forming a right-handed orthogonal coordinate system. When the orbit system is at zero attitude, the X, Y, and Z axes of the system coincide with the Xo, Yo, and Zo axes of the orbit coordinate system;

The orbital coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the center of the earth, the Xo axis is in the satellite orbit plane, perpendicular to the Zo axis, and points to the flight direction of the satellite, and the Yo axis is perpendicular to the Zo axis and the Xo axis, and constitutes a right-hand positive Intersecting coordinate system.

The beneficial effect of the present invention compared with prior art is:

(1) The existing technology is mainly aimed at the optimization of the orbit change strategy of the electric push track transfer, but in the practical application of the electric push track transfer control, there are still many engineering issues that need to be considered; according to the electric push track transfer strategy, the main points are: In order to improve the semi-major axis stage of the orbit and the rounding stage of the orbit, the thrust of the semi-major axis stage of the orbit is increased along the plane formed by the velocity direction (or vertical radial direction) and the orbital angular momentum, and there is a certain angle with the orbital plane to adjust the orbital inclination ; In the stage of orbit rounding, the thrust direction is perpendicular to the semi-major axis, and also has a certain angle with the orbit plane to adjust the orbit inclination. Therefore, the attitude of the target of electric propulsion orbit change is complex. The present invention aims at the complex attitude control problem during the electric propulsion orbit transfer, and proposes an ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer, and divides different orbit control stages to establish electric propulsion. The orbit control coordinate system of orbit transfer can accurately establish the target attitude of electric propulsion ignition based on any reference coordinate system, and at the same time meet the multi-target attitude control requirements during electric propulsion orbit transfer, including the orbit change requirements of increasing the semi-major axis or reducing eccentricity , The need to adjust the ignition direction to reduce the orbital inclination angle;

(2) On the basis of the orbit control coordinate system at different stages of electric propulsion orbit transfer, the present invention effectively maintains the attitude tracking of the sun-facing plane during the ignition process through the attitude offset around the axis where the orbit control thrust direction is located, and satisfies the energy required for attitude tracking of the sun direction It innovatively solves the multi-target attitude control problem during the electric push ignition, that is, simultaneously meets the orbit change requirements of increasing the semi-major axis or reducing the eccentricity, the ignition direction adjustment requirements of reducing the orbital inclination angle, and the dynamic tracking of the sun direction. Energy requirements provide an effective technical means for realizing precise attitude control during electric propulsion orbit transfer;

(3) The method of the present invention has been successfully applied in engineering. Based on the on-orbit verification, the attitude of the electric propulsion ignition target can be accurately established, and the long-term effective attitude tracking to the sun can be realized, which effectively promotes the electric propulsion orbit transfer technology in various fields. Engineering applications on space vehicles. Taking advantage of the significant advantages of high specific impulse of electric propulsion can greatly reduce the amount of propellant carried by spacecraft. With the continuous improvement of the requirements for the carrying capacity of spacecraft and the continuous improvement of the level of electric propulsion technology, more and more spacecraft will be equipped with electric propulsion systems. The ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer proposed by the present invention provides A method for establishing the attitude of the ignition target that integrates multi-target requirements during the electric push orbit transfer is established, and ensures that the energy demand for Japan is met during the electric push orbit transfer, which has strong application efficiency and market competitiveness.

Description of drawings

Fig. 1 is a flow chart of the method of the present invention.

Detailed ways

The present invention will be further elaborated below in conjunction with embodiment.

Exemplary embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present invention are shown in the drawings, it should be understood that the invention may be embodied in various forms and should not be limited to the embodiments set forth herein. On the contrary, the embodiment is provided to enable a more thorough understanding of the present invention and to fully convey the method of the present invention to those skilled in the art. The present invention will be described in detail below with reference to the accompanying drawings and examples.

The present invention provides an ignition attitude establishment and dynamic tracking method for electric push track transfer, as shown in Figure 1, the steps are as follows:

First enter step (1) to establish the orbit control coordinate system for the two stages of electric push orbit transfer.

Implement electric push orbit transfer, that is, the transfer orbit uses electric propulsion to elevate the semi-major axis, reduce eccentricity, and reduce orbital inclination. The electric push orbit transfer can be divided into two stages, and the reference attitude of each stage is set as the orbit control coordinate system of the new stage. Generally speaking, the electric thruster is installed on the -Z plane of the satellite body system, and the thrust direction is along the Z axis of the satellite body system.

The satellite system is defined as: taking a certain feature point of the satellite as the origin, the X-axis and Z-axis are both along the direction of the spacecraft’s characteristic axis, and the Y-axis is perpendicular to the Z-axis and X-axis, forming a right-handed orthogonal coordinate system. In this paper, when the satellite is in the zero attitude of the orbit system, the X, Y, and Z axes of the system coincide with the Xo, Yo, and Zo axes of the orbit coordinate system.

The orbit coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the center of the earth, the Xo axis is in the satellite orbit plane, perpendicular to the Zo axis, and points to the flight direction of the satellite, and the Yo axis is perpendicular to the Zo axis and the Xo axis, and constitutes a right hand Orthogonal coordinate system.

The goal of the orbital transfer in the first stage is to increase the semi-major axis and at the same time press the inclination angle. At this stage, the orbit control thrust maintains a certain angle with +Xo in the XoOoYo plane of the orbital coordinate system, and its absolute value remains unchanged at Ψ1, but the positive and negative changes before and after the argument angle is 90° and 270°. The reference attitude of the attitude control in the first stage, that is, the orbit control coordinate system is set as follows: first rotate the angle Ψ1 around the Z axis of the orbit coordinate system, which is called the orbit control yaw angle, and then rotate 90° around the Y axis of the orbit coordinate system, Make the -X axis of the orbit control coordinate system point to the ground. At this time, the angle between the +Z axis of the orbit control coordinate system and the forward direction (towards the east) is the required orbit control yaw angle, and its amplitude remains unchanged, but in The argument angle is 90° and 270° before and after changing the positive and negative. Therefore, the transformation matrix C _ORMi1 of the orbit control coordinate system relative to the inertial system in the first stage of electric push orbit transfer is obtained by rotating the orbit coordinate system C _oi , namely

The goal of the orbital transfer in the second stage is to reduce the eccentricity and at the same time press the inclination angle. At this stage, the orbit control thrust maintains a certain orbit control angle on the basis of the inertial orientation, so as to further adjust the orbit inclination angle. The reference attitude of the attitude control in the first stage, that is, the orbit control coordinate system is obtained by rotating the near-focus coordinate system PQW. In the perifocus coordinate system PQW, the P axis points to the perigee direction of the orbit, the W axis points to the positive normal direction of the orbit, and the Q axis, P axis, and W axis form a right-handed orthogonal coordinate system. The corresponding relationship between the orbit control coordinate system and each axis of the PQW coordinate system is as follows: the X axis of the orbit control coordinate system corresponds to the PQW coordinate system -P axis, the orbit control coordinate system Y axis corresponds to the PQW coordinate system -W axis, and the orbit control coordinate system Z axis corresponds to the PQW Coordinate System - Q. After transforming the corresponding relationship of the coordinate axes, rotate Ψ2 around the X axis of the current coordinate system to obtain the orbit control coordinate system C _ORMi2 of the second stage of the electric push orbit transfer. That is, first calculate the transformation matrix C _PQWi from the inertial system to the PQW coordinate system according to the orbital elements

Where ω is the argument of perigee, Ω is the right ascension of the ascending node of the orbit, and i is the inclination of the orbit.

Further, the orbit control coordinate system C _ORMi2 of the second stage of electric propulsion orbit transfer can be obtained,

in

sψ ₂ ＝sin(ψ ₂ ),cψ ₂ ＝cos(ψ ₂ )

sΩ＝sin(Ω),cΩ＝cos(Ω)

si=sin(i), ci=cos(i)

sω=sin(ω),cω=(cosω)

G _k1 =cΩ·sω+sΩ·cω·ci

G _k2 = sΩ·sω-cΩ·cω·ci

It may be assumed that this embodiment is in the first stage of electric push orbit transfer, and the orbit control coordinate system is taken as C _ORMi1 .

(2) Establish the attitude of the electric push ignition target based on any reference coordinate system

First of all, after the ignition attitude of the electric propulsion orbit transfer is established, the attitude transformation matrix C _br of the satellite system relative to the reference coordinate system should satisfy

C _br ·C _ri =C _ORMi1 or C _br ·C _ri =C _ORMi2

θ_br，ψ_br，则有where C _ri is the attitude conversion matrix from the inertial system to the reference coordinate system, assuming that the three-axis attitude Euler angle of the satellite in the reference coordinate system is

θ _br , ψ _br , then we have

或

or

When establishing the target attitude of electric propulsion ignition for the first stage of electric propulsion orbit transfer, order

When establishing the electric push ignition target attitude for the second stage of electric push orbit transfer, order

θ_br，ψ_br。当欧拉转动顺序为3-2-1时，Given the sequence of Euler angles, it can be obtained by C _br

θ _br , ψ _br . When the Euler rotation sequence is 3-2-1,

ψ _br =tan2 ^-1 (c ₁₂ ,c ₁₁ )∈[-π,π]

where the function tan2 ^-1 (y,x) is defined as

When the Euler rotation sequence is 3-1-2,

θ _br =tan2 ^-1 (-c ₁₃ ,c ₃₃ )∈[-π,π]

ψ _br =tan2 ^-1 (-c ₂₁ ,c ₂₂ )∈[-π,π]

θ_br，ψ_br，将其设置为三轴姿态控制的目标姿态角，则可基于任意参考坐标系将卫星控制到电推点火的目标姿态。The rest of the Euler rotation sequence is the same. Under any reference coordinate system datum, by solving the Euler angle

θ _br , ψ _br , if they are set as the target attitude angle of the three-axis attitude control, the satellite can be controlled to the target attitude of electric push ignition based on any reference coordinate system.

This embodiment is in the first stage of the electric propulsion orbit transfer, without loss of generality, it is set to establish the target attitude of the electric propulsion ignition based on the inertial system, after the ignition attitude of the electric propulsion orbit transfer is established, the attitude transformation of the satellite system relative to the inertial system The matrix should satisfy

C _bi =C _ORMi1

θ_br，ψ_br，则有Suppose the Euler angle of the three-axis attitude of the satellite in the J2000 inertial system is

θ _br , ψ _br , then we have

At this time, C _ri is a unit matrix, then we have

This implementation is in the first stage of electric push orbit transfer. When establishing the electric push ignition target attitude, let

θ_br，ψ_br。本实施例设定欧拉转动顺序为3-2-1，则Given the sequence of Euler angles, it can be obtained by C _br

θ _br , ψ _br . In this embodiment, the Euler rotation sequence is set as 3-2-1, then

ψ _bi ＝tan2 ^-1 (c ₁₂ ,c ₁₁ )∈[-π,π]

where the function tan2 ^-1 (y,x) is defined as

θ_br，ψ_br，将其设置为三轴姿态控制的目标姿态角，则可基于惯性坐标系将卫星控制到电推点火的目标姿态。Under the reference of the inertial coordinate system, by solving the Euler angle

θ _br , ψ _br , if they are set as the target attitude angle of the three-axis attitude control, the satellite can be controlled to the target attitude of electric push ignition based on the inertial coordinate system.

(3) Attitude tracking to the sun during the orbit transfer period

After step (2) establishes the attitude of the electric push ignition target from any reference coordinate system, the orbit control coordinate system obtained in step (1) is used as the reference coordinate system to perform attitude control during the electric push orbit transfer. At the same time, in order to meet the energy requirements during the electric propulsion orbit transfer, it is necessary to realize the normal direction of the sun wing to the sun by rotating around the axis of the ignition direction and the rotation of the sun wing.

If the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbital system reference, and the direction of the sun changes with time, and the attitude dynamic offset needs to be realized in order to track the energy source. That is, during the electric propulsion orbit change of the transfer orbit, the +Z axis of the satellite points to the direction of the orbit control thrust. When the sail is required to point to the sun, the system rotates around the Z axis of the orbit control coordinate system at a certain angle, so that the XOZ plane of the satellite is in line with the direction of the sun. The planes formed by the thrust directions coincide.

According to the current star time and solar ephemeris, the expression S _i of the sun vector in the inertial coordinate system is obtained. After the transformation matrix C _ORMi1 of the orbit control coordinate system relative to the inertial system in the first stage of the electric push orbit transfer, the sun vector is expressed in the second stage The orbit control coordinate system of the first stage, that is,

S _ORM1 ＝[S _ORM1X S _ORM1Y S _ORM1Z ] ^T ＝C _ORMi1 S _i

Then the attitude control target is based on the orbit control coordinate system in the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is

ψ _d ＝tan2 ^-1 (S _ORM1Y ,S _ORM1X )∈[-π,π]

If the electric propulsion orbit transfer is in the second stage, the ignition attitude is based on the inertial system reference, and the sun vector direction is fixed in the short term, so the attitude offset angle to ensure energy is also fixed. That is, during the electric propulsion orbit change of the transfer orbit, the +Z axis of the satellite points to the direction of the orbit control thrust. When the sail is required to point to the sun, the system rotates around the Z axis of the orbit control coordinate system at a certain angle, so that the XOZ plane of the satellite is in line with the direction of the sun. The planes formed by the thrust directions coincide.

According to the current star time and solar ephemeris, the expression S _i of the sun vector in the inertial coordinate system is obtained. After the transformation matrix C _ORMi2 of the orbit control coordinate system relative to the inertial system in the first stage of the electric push orbit transfer, the sun vector is expressed in the second stage The orbit control coordinate system of the first stage, that is,

S _ORM2 ＝[S _ORM2X S _ORM2Y S _ORM2Z ] ^T ＝C _ORMi2 S _i

Then the attitude control target is based on the orbit control coordinate system in the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is

ψ _d ＝tan2 ^-1 (S _ORM2Y ,S _ORM2X )∈[-π,π]

Through the tracking of attitude towards the sun and the control of the sun-facing angle of the solar panel, the angle between the normal direction of the solar panel and the direction of the sun, that is, the energy angle, can be controlled within the range of accuracy requirements, and the normal direction of the solar panel can be independently aligned with the sun to ensure energy.

In this embodiment, the electric propulsion orbit transfer is in the first stage, and the ignition attitude is based on the orbital system reference, and the sun direction changes with time, and the attitude dynamic offset needs to be realized in order to track the energy source. That is, during the electric propulsion orbit change of the transfer orbit, the +Z axis of the satellite points to the direction of the orbit control thrust. When the sail is required to point to the sun, the system rotates around the Z axis of the orbit control coordinate system at a certain angle, so that the XOZ plane of the satellite is in line with the direction of the sun. The planes formed by the thrust directions coincide.

According to the current star time and solar ephemeris, the expression S _i of the sun vector in the inertial coordinate system is obtained. After the transformation matrix C _ORMi1 of the orbit control coordinate system relative to the inertial system in the first stage of the electric push orbit transfer, the sun vector is expressed in the second stage The orbit control coordinate system of the first stage, that is,

S _ORM1 ＝[S _ORM1X S _ORM1Y S _ORM1Z ] ^T ＝C _ORMi1 S _i

Then the attitude control target is based on the orbit control coordinate system in the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is

ψ _d ＝tan2 ^-1 (S _ORM1Y ,S _ORM1X )∈[-π,π]

Through the tracking of attitude towards the sun and the control of the sun-facing angle of the solar panel, the angle between the normal direction of the solar panel and the direction of the sun, that is, the energy angle, can be controlled within the range of accuracy requirements, and the normal direction of the solar panel can be independently aligned with the sun to ensure energy.

To sum up, the ignition attitude establishment and dynamic tracking method of electric push orbit transfer involved in the present invention can accurately establish the ignition target attitude of electric push orbit transfer based on any reference coordinate system according to different target stages of electric push transfer orbit control, And during the period of electric propulsion orbit transfer, it provides the target attitude towards the sun with independent tracking energy. The method of the present invention can simultaneously meet the multi-target attitude control requirements during electric propulsion orbit transfer, including the requirements for orbit change to increase the semi-major axis or reduce eccentricity, the requirements for adjusting the ignition direction to lower the orbital inclination angle, and the energy requirements for dynamically tracking the direction of the sun , providing a technical approach for realizing electric propulsion orbit transfer in orbit.

Although the present invention has been disclosed as above with preferred embodiments, it is not intended to limit the present invention, and any person skilled in the art can use the methods disclosed above and technical content to analyze the present invention without departing from the spirit and scope of the present invention. Possible changes and modifications are made in the technical solution. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present invention, which do not depart from the content of the technical solution of the present invention, all belong to the technical solution of the present invention. protected range.

Claims (11)

1. An ignition attitude establishment and a dynamic tracking method of electric push orbit transfer, it is characterized in that, the step comprises: (1) Establish the orbit control coordinate system for the two phases of electric propulsion orbit transfer The electric propulsion orbit transfer is divided into two stages, the reference attitude of each stage is set as the orbit control coordinate system of this stage, the electric thruster is installed on the -Z plane of the satellite body system, and the thrust direction is along the Z axis of the satellite body system; Among them, the orbit transfer goal of the first stage is to increase the semi-major axis and at the same time press the inclination angle, and the orbit transfer goal of the second stage is to reduce the eccentricity and at the same time press the inclination angle; The reference attitude of the attitude control in the first stage, that is, the orbit control coordinate system is set as follows: first rotate the angle Ψ1 around the Z axis of the orbit coordinate system, which is called the orbit control yaw angle, and then rotate 90° around the Y axis of the orbit coordinate system, Make the -X axis of the orbit control coordinate system point to the ground. At this time, the angle between the +Z axis of the orbit control coordinate system and the forward direction is the required orbit control yaw angle. Its amplitude remains unchanged, but when the argument angle is 90 ° and 270° change the positive and negative; The reference attitude of the attitude control in the second stage, that is, the orbit control coordinate system is obtained by rotating the near-focus coordinate system PQW; The W axis constitutes a right-handed orthogonal coordinate system, and the corresponding relationship between the orbit control coordinate system and each axis of the PQW coordinate system is: the X axis of the orbit control coordinate system corresponds to the PQW coordinate system -P axis, and the Y axis of the orbit control coordinate system corresponds to the PQW coordinate system -W axis, the Z-axis of the orbit control coordinate system corresponds to the PQW coordinate system-Q axis, after converting the corresponding relationship of the coordinate axes, and then rotating Ψ2 around the X-axis of the current coordinate system, the orbit control coordinate system of the second stage of the electric push orbit transfer is obtained; among them, Ψ2 is the orbital inclination of the second stage; (2) Establish the attitude of the electric push ignition target based on any reference coordinate system Under any reference coordinate system datum, by solving the Euler angle

θ _br , ψ _br , set it as the target attitude angle of the three-axis attitude control, then the satellite can be controlled to the target attitude of electric push ignition based on any reference coordinate system; First of all, after the ignition attitude of the electric propulsion orbit transfer is established, the attitude transformation matrix C _br of the satellite system relative to the reference coordinate system should satisfy C _br ·C _ri =C _ORMi1 or C _br ·C _ri =C _ORMi2 where C _ri is the attitude conversion matrix from the inertial system to the reference coordinate system, and the three-axis attitude Euler angle of the satellite in the reference coordinate system is

θ _br , ψ _br , then we have

or

When establishing the target attitude of electric propulsion ignition for the first stage of electric propulsion orbit transfer, order

When establishing the electric push ignition target attitude for the second stage of electric push orbit transfer, order

Given the sequence of Euler angles, it can be obtained by C _br

θ _br , ψ _br ; When the Euler rotation sequence is 3-2-1,

ψ _br =tan2 ^-1 (c ₁₂ ,c ₁₁ )∈[-π,π] where the function tan2 ^-1 (y,x) is defined as

When the Euler rotation sequence is 3-1-2,

θ _br =tan2 ^-1 (-c ₁₃ ,c ₃₃ )∈[-π,π] ψ _br =tan2 ^-1 (-c ₂₁ ,c ₂₂ )∈[-π,π] The rest of the Euler rotation sequence is the same; (3) Attitude tracking to the sun during the orbit transfer period After step (2) establishes the attitude of the electric push ignition target from any reference coordinate system, it is converted to the orbit control coordinate system obtained in step (1) as the reference coordinate system to perform attitude control during the electric push orbit transfer; at the same time, the electric push During the orbit transfer, in order to meet the energy requirements, it is necessary to realize the normal direction of the sun wing to the sun by rotating around the axis of the ignition direction and around the sun wing; Through attitude tracking to the sun and solar panel solar rotation angle control, the angle between the normal direction of the solar panel and the direction of the sun, that is, the energy angle, is controlled within the range of accuracy requirements, and the normal direction of the solar panel can be independently aligned with the sun to ensure energy. . 2. The ignition attitude establishment and dynamic tracking method of a kind of electric push orbit transfer according to claim 1, characterized in that, in step (1), the electric push orbit transfer is implemented, that is, the transfer orbit uses electric propulsion to lift the semi-major axis , Reduce eccentricity and reduce orbital inclination. 3. The ignition attitude establishment and dynamic tracking method of a kind of electric push track transfer according to claim 1, is characterized in that, in step (1), the first-stage track control thrust is in the XoOoYo plane of the track coordinate system, and +Xo maintains a certain included angle, and its absolute value remains unchanged as Ψ1, and the range of Ψ1 is 0°-90°, but at every interval of half an orbital period, the exact value of the argument depends on the orbit control strategy, and the corresponding argument is 90 ° and 270° to change the positive and negative. 4. The ignition attitude establishment and dynamic tracking method of a kind of electric push orbit transfer according to claim 1, characterized in that, in step (1), the orbit control coordinate system of the first stage of electric push orbit transfer is relative to the inertial system The transformation matrix C _ORMi1 is obtained by rotating the orbital coordinate system C _oi , namely

Among them, Ψ1 is the orbit control yaw angle. 5. The ignition attitude establishment and dynamic tracking method of a kind of electric propulsion orbit transfer according to claim 1, characterized in that, in step (1), the orbit control thrust in the second stage maintains a certain orbit on the basis of inertial orientation Control the included angle, so as to further adjust the orbital inclination, its absolute value does not change to Ψ2, but the positive and negative changes when the argument angle is 90° and 270°. 6. The ignition attitude establishment and dynamic tracking method of a kind of electric push orbit transfer according to claim 1, is characterized in that, first calculate the conversion matrix C _PQWi of the inertial system to the PQW coordinate system according to the orbit elements

Where ω is the argument of perigee, Ω is the right ascension of the ascending node of the orbit, and i is the inclination of the orbit; Further, the orbit control coordinate system C _ORMi2 of the second stage of electric propulsion orbit transfer can be obtained,

in: sψ ₂ ＝sin(ψ ₂ ),cψ ₂ ＝cos(ψ ₂ ) sΩ＝sin(Ω),cΩ＝cos(Ω) si=sin(i), ci=cos(i) sω=sin(ω),cω=(cosω) G _k1 =cΩ·sω+sΩ·cω·ci G _k2 =sΩ·sω−cΩ·cω·ci. 7. The ignition attitude establishment and dynamic tracking method for electric orbit transfer according to claim 1, characterized in that, if the electric orbit transfer is in the first stage, the ignition attitude is based on the orbit system reference, and the direction of the sun is time-varying , in order to track the energy source, it is necessary to realize the attitude dynamic offset, that is, during the transfer orbit electric propulsion orbit change, the +Z axis of the satellite points to the direction of the orbit control thrust, and when the sail is required to point to the sun, the system rotates by a certain amount around the Z axis of the orbit control coordinate system Angle so that the XOZ plane of the satellite coincides with the plane formed by the sun direction and the thrust direction. 8. The ignition attitude establishment and dynamic tracking method of a kind of electric push orbit transfer according to claim 7, characterized in that, the expression S _i of the sun vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris, and is obtained by The transformation matrix C _ORMi1 of the orbit control coordinate system in the first stage of electric propulsion orbit transfer relative to the inertial system expresses the sun vector in the orbit control coordinate system of the first stage, namely S _ORM1 ＝[S _ORM1X S _ORM1Y S _ORM1Z ] ^T ＝C _ORMi1 S _i Then the attitude control target is based on the orbit control coordinate system of the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is ψ _d =tan2 ^-1 (S _ORM1Y , S _ORM1X )∈[-π,π]. 9. The ignition attitude establishment and dynamic tracking method for electric orbit transfer according to claim 1, characterized in that, if the electric orbit transfer is in the second stage, the ignition attitude is based on the inertial frame reference, and the sun vector direction is short-term Internally fixed, so the attitude offset angle is also fixed to ensure energy. That is, during the electric propulsion orbit change of the transfer orbit, the +Z axis of the satellite points to the direction of the orbit control thrust. When the sail is required to point to the sun, the system revolves around the orbit control coordinate system Z The axis is rotated by a certain angle so that the XOZ plane of the satellite coincides with the plane formed by the direction of the sun and the direction of thrust. 10. The ignition attitude establishment and dynamic tracking method of a kind of electric push orbit transfer according to claim 9, characterized in that, the expression S _i of the sun vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris, and is obtained by The conversion matrix C _ORMi2 of the orbit control coordinate system in the first stage of electric propulsion orbit transfer relative to the inertial system expresses the sun vector in the orbit control coordinate system of the first stage, namely S _ORM2 ＝[S _ORM2X S _ORM2Y S _ORM2Z ] ^T ＝C _ORMi2 S _i Then the attitude control target is based on the orbit control coordinate system of the first stage, and the Z-axis offset angle ψ _d around the orbit control coordinate system is ψ _d =tan2 ^-1 (S _ORM2Y , S _ORM2X )∈[-π,π]. 11. The ignition attitude establishment and dynamic tracking method for electric push orbit transfer according to claim 1, characterized in that the satellite system is defined as: taking a certain feature point of the satellite as the origin, and the X-axis and the Z-axis are both along the spaceflight The Y axis is perpendicular to the Z axis and the X axis, and constitutes a right-handed orthogonal coordinate system. When the satellite is in the zero attitude of the orbit system, the X, Y, Z axis directions of the system and the Xo, Yo, Zo axis directions of the orbit coordinate system coincide; The orbital coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the center of the earth, the Xo axis is in the satellite orbit plane, perpendicular to the Zo axis, and points to the flight direction of the satellite, and the Yo axis is perpendicular to the Zo axis and the Xo axis, and constitutes a right-hand positive Intersecting coordinate system.

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