CN113830332B - Ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer - Google Patents
Ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer Download PDFInfo
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Abstract
The invention relates to an ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer, which comprises the following steps: (1) Determining an orbit control coordinate system of two stages of electric propulsion orbit transfer; (2) Establishing an electric-pushing ignition target posture based on any reference coordinate system; and (3) tracking the attitude of the electric propulsion orbit during the transfer to the sun. On the basis of an orbit control coordinate system of an electric propulsion orbit at different stages of transfer, the attitude offset of an axis where the orbit control thrust direction is located is adopted, so that the attitude tracking sun-facing plane in the ignition process is effectively kept, the energy requirement of attitude tracking sun direction is met, and the multi-target attitude control problem during electric propulsion ignition is innovatively solved.
Description
Technical Field
The invention relates to an ignition attitude establishing and dynamic tracking method for electric propulsion orbit transfer, and belongs to the field of spacecraft attitude orbit control.
Background
The method adopts the combination of chemical propulsion and electric propulsion to change the orbit, namely, after separating the satellite and the arrow, a chemical remote engine is firstly adopted to realize multiple times of orbit change, after the satellite is sent into a handover orbit with a certain height and a certain inclination angle, the electric propulsion is utilized to realize the orbit lifting and the orbit rounding during the subsequent orbit transfer. And the electric propulsion track lifting mode is used for lifting the semimajor axis, reducing the eccentricity and reducing the track inclination angle of the transfer track by using electric propulsion.
According to an electric-pushing track transfer strategy, the method mainly comprises a track semimajor axis lifting stage and a track rounding stage, wherein in a plane formed by thrust in a speed direction (or vertical radial direction) and track angular momentum in the track semimajor axis lifting stage, a certain included angle is formed between the thrust and a track plane and is used for adjusting the track inclination angle; in the track rounding stage, the thrust direction is vertical to the semi-long shaft, and a certain included angle is formed between the thrust direction and the plane of the track for adjusting the inclination angle of the track. Therefore, the electric-thrust orbital transfer target attitude is complicated.
Because the thrust of the electric thruster is small, the electric thruster orbit change consumes long time in each circle, the visibility of an arc section is complex, and the autonomous processing flow on the satellite is required to be complete and reliable. Meanwhile, when the electric propulsion is used for rail transfer, the electric propulsion needs to work in a high-power high-thrust mode, the requirement on the whole satellite energy is high, and the requirements on posture and sailboard combined adjustment and automatic sun-facing to guarantee the energy are met. The energy angle is defined as the included angle between the normal direction of the solar wing and the direction of the sun, and the energy angle has definite precision requirement during the transfer of the electric propulsion orbit.
Disclosure of Invention
The technical problem solved by the invention is as follows: the method overcomes the defects of the prior art, and provides the ignition attitude establishment and dynamic tracking method for the electric propulsion track transfer, so that the ignition attitude integrates the multi-target requirements of the orbital transfer requirement for improving the semimajor axis or reducing the eccentricity, the ignition direction adjustment for depressing the track inclination angle and the energy requirement for dynamically tracking the sun direction, the electric propulsion ignition target attitude can be accurately established, and the long-term effective attitude-to-day dynamic tracking is realized.
The technical scheme of the invention is as follows:
an ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer comprises the following steps:
(1) Rail-controlled coordinate system for determining two stages of electric push rail transfer
The electric propulsion orbit transfer is divided into two stages, the reference attitude of each stage is set as an orbit control coordinate system of the stage, an electric thruster is arranged on a Z surface of a satellite body system, and the thrust direction is along the Z axis of the satellite body system;
(2) Establishing electric-pushing ignition target posture based on any reference coordinate system
Under the reference of any reference coordinate system, by resolving Euler angle
θ br ，ψ br If the target attitude angle is set as the target attitude angle of the three-axis attitude control, the satellite can be controlled to the target attitude of the electric propulsion ignition based on any reference coordinate system;
(3) Attitude-to-sun tracking during electric push rail transfer
After establishing the attitude of the electric propulsion ignition target by the arbitrary reference coordinate system in the step (2), converting the attitude control during the electric propulsion track transfer into the reference coordinate system by using the track control coordinate system obtained in the step (1); meanwhile, in order to meet the energy requirement during the transfer of the electric propulsion track, the normal sun alignment of the solar wing is realized by rotating around the axis of the ignition direction and rotating around the solar wing;
the included angle between the normal direction of the solar sailboard and the sun direction, namely the energy angle, is controlled to be within the precision requirement range through attitude sun tracking and sun-facing rotation angle control of the solar sailboard, and the requirement of autonomous sun-facing of the normal direction of the solar sailboard is met to guarantee energy.
Further, in the step (1), electric propulsion track transfer is implemented, namely the transfer track uses electric propulsion to lift the semimajor axis, reduce the eccentricity and reduce the track inclination angle.
Further, in step (1), the orbit transfer target of the first stage is to increase the semimajor axis while the tilt angle is pressed, the orbital control thrust of the stage is in the XoOoYo plane of the orbit coordinate system, an included angle is kept between + Xo, the absolute value of the orbital control thrust does not become Ψ 1, Ψ 1 ranges from 0 to 90 °, but at every half orbit period, the accurate value of the argument is determined according to the orbital control strategy, and the corresponding argument changes positive and negative at 90 ° and 270 °.
Further, the reference attitude, i.e., the orbit control coordinate system, of the first-stage attitude control is set to: firstly, rotating a phi 1 angle around the Z axis of the orbit coordinate system, which is called an orbit control yaw angle, and then rotating the phi 90 angle around the Y axis of the orbit coordinate system, so that the-X axis of the orbit control coordinate system points to the ground, and at the moment, the included angle between the + Z axis of the orbit control coordinate system and the advancing direction is the required orbit control yaw angle, the amplitude of the orbit control yaw angle is unchanged, but the positive and negative values are changed when the amplitude is 90 degrees and 270 degrees.
Further, in the step (1), the transformation matrix C of the orbit control coordinate system of the first stage of the electric propulsion orbit transfer relative to the inertia system is adopted ORMi1 From an orbital coordinate system C oi Is rotated to obtain
Wherein Ψ 1 is a trajectory-controlled yaw angle.
Further, in step (1), the goal of orbit transfer in the second stage is to reduce eccentricity while suppressing inclination, and the orbit control thrust in this stage maintains a certain orbit control angle based on inertial orientation, so as to further adjust the inclination of the orbit, whose absolute value does not change to Ψ 2, but changes positive and negative when the argument is 90 ° and 270 °.
Further, in step (1), the reference attitude of the first-stage attitude control, i.e., the orbit control coordinate system, is obtained by rotation based on the near-focus coordinate system PQW.
Furthermore, in the near focus coordinate system PQW, the P axis points in the near point direction of the track, the W axis points in the normal direction of the track, and the Q axis, the P axis and the W axis form a right-hand orthogonal coordinate system.
Further, the corresponding relation of each axis of the orbit control coordinate system and the PQW coordinate system is that the X axis of the orbit control coordinate system corresponds to PQW coordinate system-P axis, the Y axis of the orbit control coordinate system corresponds to PQW coordinate system-W axis, the Z axis of the orbit control coordinate system corresponds to PQW coordinate system-Q axis, and after the corresponding relation of the coordinate axes is converted, psi 2 is rotated around the X axis of the current coordinate system, namely the orbit control coordinate system C of the second stage of the electric propulsion orbit transfer is obtained ORMi2 。
Further, a transformation matrix C from the inertial system to the PQW coordinate system is first calculated from the orbit components PQWi
Wherein omega is the amplitude angle of the near place of the track, omega is the right ascension of the ascending intersection point of the track, and i is the inclination angle of the track;
further obtaining an orbit control coordinate system C of the second stage of the electric propulsion orbit transfer ORMi2 ，
Wherein:
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), firstly, after the ignition attitude of the electric propulsion orbit transfer is established, the attitude transformation matrix C of the satellite body system relative to the reference coordinate system br Should satisfy
C br ·C ri ＝C ORMi1 Or C br ·C ri ＝C ORMi2
Wherein C ri Setting the three-axis attitude Euler angle of the satellite in the reference coordinate system as
θ br ，ψ br Then there is
When the electric propulsion track is transferred to the first stage to establish the electric propulsion ignition target attitude, the order is given
When the electric propulsion track is transferred to the second stage to establish the electric propulsion ignition target attitude, the order is given
Given the Euler angle, the sequence can be changed by C br Obtaining
θ br ，ψ br 。
Further, when the Euler rotation order is 3-2-1,
ψ br ＝tan2 -1 (c 12 ,c 11 )∈[-π,π]
wherein the function tan2 -1 (y, x) is defined as
When the euler rotation order 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 remaining euler rotation sequences are the same.
Further, if the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbit system reference, the sun direction is changed in time, attitude dynamic bias needs to be realized for tracking energy, namely during the electric propulsion orbit transfer of the transfer orbit, the satellite + Z axis points to the orbit control thrust direction, and when the sailboard points to the sun, the body system rotates for 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 formed by the sun direction and the thrust direction.
Further, the expression S of the sun vector in an inertial coordinate system is calculated according to the current satellite time and the solar ephemeris i Transfer the first via the electric push railTransformation matrix C of orbital coordinate system of stage relative to inertial system ORMi1 Expressing the sun vector in the first stage of the orbital coordinate system, i.e.
S ORM1 ＝[S ORM1X S ORM1Y S ORM1Z ] T ＝C ORMi1 ·S i
The attitude control target is biased by an angle psi around the Z axis of the orbit control coordinate system on the basis of the orbit control coordinate system in the first stage d Is psi 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 a short period, so that the attitude offset angle for ensuring energy is also fixed, namely during the electric propulsion orbit transfer of the transfer orbit, the satellite + Z axis points to the orbit control thrust direction, and when the sailboard is required to point to the sun, the body system rotates for a certain angle around the Z axis of the orbit control coordinate system, so that the XOZ plane of the satellite is superposed with the plane formed by the sun direction and the thrust direction.
Further, calculating to obtain the expression S of the solar vector in an inertial coordinate system according to the current star time and the solar ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi2 Expressing the sun vector in the first stage of the orbital coordinate system, i.e.
S ORM2 ＝[S ORM2X S ORM2Y S ORM2Z ] T ＝C ORMi2 ·S i
The attitude control target is offset by an angle psi around the Z-axis of the orbit control coordinate system based on the orbit control coordinate system of the first stage d Is psi d ＝tan2 -1 (S ORM2Y ,S ORM2X )∈[-π,π]。
Further, the satellite system is defined as: taking a certain characteristic point of the satellite as an origin, enabling an X axis and a Z axis to be along the characteristic axis direction of the spacecraft, enabling a Y axis to be vertical to the Z axis and the X axis and forming a right-hand orthogonal coordinate system, and enabling the axis direction of the system X, Y, Z to be coincident with the axis directions of Xo, yo and Zo of the orbital coordinate system when the satellite is in a zero attitude;
the orbital coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the geocentric, 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 forms a right-hand orthogonal coordinate system.
Compared with the prior art, the invention has the beneficial effects that:
(1) The prior art mainly aims at the problem of optimization of a track transfer strategy of electric push track transfer, and has a plurality of engineering problems to be considered in the practical application of implementing electric push track transfer control; according to an electric-pushing track transfer strategy, the method mainly comprises a track semimajor axis lifting stage and a track rounding stage, wherein in a plane formed by thrust and track angular momentum in a speed direction (or vertical radial direction) in the track semimajor axis lifting stage, a certain included angle is formed between the thrust and a track plane and is used for adjusting the track inclination angle; in the track rounding stage, the thrust direction is vertical to the semi-long axis, and a certain included angle is formed between the thrust direction and the track plane for adjusting the track inclination angle. Therefore, the attitude of the electric propulsion orbit transfer target is complex, aiming at the problem of complex attitude control during the electric propulsion orbit transfer, the invention provides an ignition attitude establishing and dynamic tracking method for the electric propulsion orbit transfer, which divides different orbit control stages to establish an orbit control coordinate system for the electric propulsion orbit transfer, can accurately establish the attitude of the electric propulsion ignition target based on any reference coordinate system, and simultaneously meets the multi-target attitude control requirements during the electric propulsion orbit transfer, including the orbit changing requirement of improving a semimajor axis or reducing eccentricity and the ignition direction adjusting requirement of reducing the orbit inclination angle;
(2) On the basis of orbit control coordinate systems of different stages of electric propulsion orbit transfer, the attitude offset of an axis where the orbit control thrust direction is located is adopted, so that the attitude tracking sun-facing plane in the ignition process is effectively kept, the energy requirement of attitude tracking sun direction is met, the multi-target attitude control problem in the electric propulsion ignition period is innovatively solved, namely, the orbit changing requirement of improving the semimajor axis or reducing the eccentricity, the ignition direction adjusting requirement of pressing down the orbit inclination angle and the energy requirement of dynamically tracking the sun direction are met, and an effective technical means is provided for realizing accurate attitude control in the electric propulsion orbit transfer period;
(3) The method of the invention has successfully realized engineering application, on-orbit verification can accurately establish the attitude of the electric propulsion ignition target based on the method, and realize long-term effective attitude-to-sun dynamic tracking, thereby effectively promoting the engineering application of the electric propulsion orbit transfer technology on various spacecrafts. The obvious advantage of high specific impulse of electric propulsion can greatly reduce the carrying amount of the propellant of the spacecraft. With the continuous improvement of the requirement on the bearing capacity of the spacecraft and the continuous progress of the technical level of electric propulsion, more and more spacecrafts are provided with electric propulsion systems, and the ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer provided by the invention provides an ignition target attitude establishment method integrating multi-target requirements during electric propulsion orbit transfer, ensures that the energy-to-day requirements are met during electric propulsion orbit transfer, and has stronger application efficiency and market competitiveness.
Drawings
FIG. 1 is a flow chart of the method of the present invention.
Detailed Description
The invention is further illustrated by the following examples.
Exemplary embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. While exemplary embodiments of the 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. Rather, this example is provided so that this disclosure will be thorough and complete, and will fully convey the method of the invention to those skilled in the art. The present invention will be described in detail below with reference to the embodiments with reference to the attached drawings.
The invention provides an ignition attitude establishment and dynamic tracking method for electric propulsion track transfer, which comprises the following steps as shown in figure 1:
firstly, the method enters the step (1) to establish an orbit control coordinate system of two stages of electric propulsion orbit transfer.
And (3) implementing electric propulsion track transfer, namely using electric propulsion to lift the semi-long shaft, reduce the eccentricity and reduce the track inclination angle of the transfer track. The electric-propulsion 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 stage. Generally, the electric thruster is installed on the Z plane of the satellite system, and the thrust direction is along the Z axis of the satellite system.
The satellite body is defined as: a certain characteristic point of the satellite is used as an origin, the X axis and the Z axis are along the characteristic axis direction of the spacecraft, and the Y axis is vertical to the Z axis and the X axis and forms a right-hand orthogonal coordinate system. The axis direction of the system X, Y, Z coincides with the axis direction of the orbital coordinate system Xo, yo and Zo in the zero attitude of the orbital system.
Wherein the orbital coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the geocentric, 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 forms a right-hand orthogonal coordinate system.
The goal of the first stage of orbit transfer is to increase the semi-major axis while suppressing the tilt angle. The orbital control thrust at the stage is in an XoOoYo plane of an orbital coordinate system, a certain included angle is kept between the orbital control thrust and + Xo, the absolute value of the orbital control thrust is not changed to psi 1, and the positive and negative of the orbital control thrust are changed before and after the argument is 90 degrees and 270 degrees. The reference attitude, i.e., the tracking coordinate system, of the first-stage attitude control is set to: firstly, rotating a phi 1 angle around the Z axis of the orbit coordinate system, which is called an orbit control yaw angle, and then rotating the phi 90 angle around the Y axis of the orbit coordinate system, so that the-X axis of the orbit control coordinate system points to the ground, and at the moment, the included angle between the + Z axis of the orbit control coordinate system and the advancing direction (towards the east) is the required orbit control yaw angle, the amplitude of the orbit control yaw angle is unchanged, but the positive and negative values are changed before and after the amplitude is 90 degrees and 270 degrees. Therefore, the transformation matrix C of the orbit control coordinate system relative to the inertia system in the first stage of the electric propulsion orbit transfer ORMi1 From an orbital coordinate system C oi Is rotated to obtain
The goal of the second stage of orbital transfer is to reduce eccentricity while compressing the angle of inclination. The orbit control thrust of the stage keeps a certain orbit control included angle on the basis of inertial orientation, so that the inclination angle of the orbit is further adjusted, the absolute value of the inclination angle is not changed to psi 2, but the positive and negative values are changed before and after the argument is 90 degrees and 270 degrees. The reference attitude of the first-stage attitude control, i.e., the orbit control coordinate system, is obtained by rotation based on the near-focus coordinate system PQW. Near focus coordinate system PQW, P-axisPointing to the near-point direction of the track, pointing to the positive normal direction of the track by the W axis, and forming a right-hand orthogonal coordinate system by the Q axis, the P axis and the W axis. The corresponding relation of the orbit control coordinate system and each axis of the PQW coordinate system is that the X axis of the orbit control coordinate system corresponds to PQW coordinate system-P axis, the Y axis of the orbit control coordinate system corresponds to PQW coordinate system-W axis, and the Z axis of the orbit control coordinate system corresponds to PQW coordinate system-Q. After the corresponding relation of the coordinate axes is converted, the psi 2 is rotated around the X axis of the current coordinate system, and the orbit control coordinate system C of the second stage of the electric propulsion orbit transfer is obtained ORMi2 . Firstly, a transformation matrix C from an inertial system to a PQW coordinate system is calculated according to the track elements PQWi
Wherein omega is the amplitude angle of the near place of the track, omega is the right ascension of the ascending intersection point of the track, and i is the inclination angle of the track.
Further obtaining an orbit control coordinate system C of the second stage of electric propulsion orbit transfer ORMi2 ，
Wherein
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 is not assumed that the first stage of the track transfer is performed by the electric propulsion in this embodiment, and the track-controlled coordinate system is C ORMi1 。
(2) Establishing electric-pushing ignition target posture based on any reference coordinate system
Firstly, after the ignition attitude of the electric-push orbit transfer is established, the attitude conversion matrix C of the satellite body system relative to the reference coordinate system br Should satisfy
C br ·C ri ＝C ORMi1 Or C br ·C ri ＝C ORMi2
Wherein C is ri For an attitude transformation matrix from an inertial system to a reference coordinate system, assuming that the Euler angles of three-axis attitudes of the satellite in the reference coordinate system are
θ br ，ψ br Then there is
When the electric propulsion track is transferred to the first stage to establish the electric propulsion ignition target attitude, the order is given
When the electric propulsion track is transferred to the second stage to establish the electric propulsion ignition target attitude, the order is given
Given the Euler angle, the sequence can be changed by C br Obtaining
θ br ，ψ br . When the euler rotation order is 3-2-1,
ψ br ＝tan2 -1 (c 12 ,c 11 )∈[-π,π]
wherein the function tan2 -1 (y, x) is defined as
When the euler rotation order 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 remaining Euler rotation sequences are the same. Under the reference of any reference coordinate system, by resolving Euler angle
θ br ，ψ br And set as the target attitude angle of the three-axis attitude control, the satellite can be controlled to the target attitude of the electric propulsion ignition based on an arbitrary reference coordinate system.
In the first stage of the electric propulsion orbit transfer, the attitude of the electric propulsion ignition target established based on the inertial system is set without loss of generality, and after the ignition attitude of the electric propulsion orbit transfer is established, the attitude conversion matrix of the satellite system relative to the inertial system is required to meet the requirement
C bi ＝C ORMi1
Assuming that the three-axis attitude Euler angle of the satellite under the J2000 inertial system is
θ br ，ψ br Then there is
At this time C ri Is a unit array, then
The implementation is in the first stage of electric-push orbit transfer, and when establishing the electric-push ignition target posture, the implementation orders
Given the Euler angle, the sequence can be changed by C br Obtaining
θ br ，ψ br . In this embodiment, the Euler rotation sequence is set to 3-2-1
ψ bi ＝tan2 -1 (c 12 ,c 11 )∈[-π,π]
Wherein the function tan2 -1 (y, x) is defined as
Under the reference of an inertial coordinate system, by calculating the Euler angle
θ br ，ψ br Set it as the target attitude angle for three-axis attitude control, then it can be based on inertiaThe coordinate system controls the satellite to a target attitude for electric propulsion ignition.
(3) Attitude-to-sun tracking during electric push rail transfer
And (3) after establishing the attitude of the electrically-propelled ignition target by the arbitrary reference coordinate system in the step (2), converting the orbit control coordinate system obtained in the step (1) into the reference coordinate system, and controlling the attitude during the transfer of the electrically-propelled orbit. Meanwhile, in order to meet the energy requirement during the transfer of the electric propulsion track, the normal sun alignment of the solar wing needs to be realized through rotation around the axis of the ignition direction and rotation of the solar wing.
If the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbit system reference, the sun direction is time-varying, and attitude dynamic bias needs to be realized for tracking energy. During the electric propulsion orbit change of the transfer orbit, the + Z axis of the satellite points to the orbit control thrust direction, and when the sailboard points to the sun, the body system rotates for a certain angle around the Z axis of the orbit control coordinate system, so that the XOZ plane of the satellite is superposed with the plane formed by the sun direction and the thrust direction.
Calculating to obtain the expression S of the solar vector in an inertial coordinate system according to the current satellite time and the solar ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi1 Expressing the sun vector in an orbital coordinate system of the first stage, i.e.
S ORM1 ＝[S ORM1X S ORM1Y S ORM1Z ] T ＝C ORMi1 ·S i
The attitude control target is offset by an angle psi around the Z-axis of the orbit control coordinate system based on the orbit control coordinate system of the first stage d Is composed of
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 a short period, so that the attitude offset angle for guaranteeing energy is also fixed. During the electric propulsion orbit transfer of the transfer orbit, the + Z axis of the satellite points to the orbit control thrust direction, and when the sailboard is required to point to the sun, the body system rotates for a certain angle around the Z axis of the orbit control coordinate system, so that the XOZ plane of the satellite is superposed with the plane formed by the sun direction and the thrust direction.
Calculating to obtain the expression S of the solar vector in an inertial coordinate system according to the current satellite time and the solar ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi2 Expressing the sun vector in the first stage of the orbital coordinate system, i.e.
S ORM2 ＝[S ORM2X S ORM2Y S ORM2Z ] T ＝C ORMi2 ·S i
The attitude control target is biased by an angle psi around the Z axis of the orbit control coordinate system on the basis of the orbit control coordinate system in the first stage d Is composed of
ψ d ＝tan2 -1 (S ORM2Y ,S ORM2X )∈[-π,π]
Through posture sun-facing tracking and sun-facing corner control of the solar sailboards, the included angle between the normal direction of the solar sailboards and the sun direction, namely the energy angle, can be controlled within the precision requirement range, and the requirement of autonomous sun-facing of the normal direction of the solar sailboards is met so as to guarantee energy.
In the embodiment, the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbit system reference, the sun direction is time-varying, and attitude dynamic bias needs to be realized for tracking energy. During the electric propulsion orbit change of the transfer orbit, the + Z axis of the satellite points to the orbit control thrust direction, and when the sailboard points to the sun, the body system rotates for a certain angle around the Z axis of the orbit control coordinate system, so that the XOZ plane of the satellite is superposed with the plane formed by the sun direction and the thrust direction.
Calculating to obtain the expression S of the sun vector in an inertial coordinate system according to the current satellite time and the sun ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi1 Expressing the sun vector in an orbital coordinate system of the first stage, i.e.
S ORM1 ＝[S ORM1X S ORM1Y S ORM1Z ] T ＝C ORMi1 ·S i
The attitude control target is wound around the Z axis of the orbit control coordinate system on the basis of the orbit control coordinate system in the first stageOffset angle psi d Is composed of
ψ d ＝tan2 -1 (S ORM1Y ,S ORM1X )∈[-π,π]
Through posture sun-facing tracking and sun-facing corner control of the solar sailboards, the included angle between the normal direction of the solar sailboards and the sun direction, namely the energy angle, can be controlled within the precision requirement range, and the requirement of autonomous sun-facing of the normal direction of the solar sailboards is met so as to guarantee energy.
In summary, the ignition attitude establishing and dynamic tracking method for the electric propulsion orbit transfer can accurately establish the ignition target attitude of the electric propulsion orbit transfer based on any reference coordinate system according to different target stages controlled by the electric propulsion transfer orbit, and provide the counterglow target attitude of the autonomous tracking energy during the electric propulsion orbit transfer. The method can simultaneously meet the multi-target attitude control requirements during the electric propulsion track transfer period, including the track changing requirement for improving the semimajor axis or reducing the eccentricity, the ignition direction adjusting requirement for reducing the track inclination angle and the energy requirement for dynamically tracking the sun direction, and provides a technical approach for realizing the electric propulsion track transfer on the track.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make possible variations and modifications of the present invention using the method and the technical contents disclosed above without departing from the spirit and scope of the present invention, and therefore, any simple modifications, equivalent changes and modifications made to the above embodiments according to the technical essence of the present invention are all within the scope of the present invention.
Claims (11)
1. An ignition attitude establishing and dynamic tracking method for electric propulsion orbit transfer is characterized by comprising the following steps:
(1) Rail-controlled coordinate system for determining two stages of electric push rail transfer
The electric propulsion orbit transfer is divided into two stages, the reference attitude of each stage is set as an orbit control coordinate system of the stage, an electric thruster is arranged on a Z surface of a satellite body system, and the thrust direction is along the Z axis of the satellite body system; the track transfer target of the first stage is to improve the semimajor axis and simultaneously press the inclination angle, and the track transfer target of the second stage is to reduce the eccentricity and simultaneously press the inclination angle;
the reference attitude, i.e., the orbit control coordinate system, of the first-stage attitude control is set to: firstly, rotating a phi 1 around the Z axis of an orbit coordinate system, namely an orbit control yaw angle, and then rotating the phi 90 around the Y axis of the orbit coordinate system to enable the-X axis of the orbit control coordinate system to point to the ground, wherein the included angle between the + Z axis of the orbit control coordinate system and the advancing direction is the required orbit control yaw angle, the amplitude of the orbit control yaw angle is unchanged, and the positive and negative are changed when the amplitude is 90 degrees and 270 degrees;
the reference attitude of the attitude control of the second stage, namely the rotation of the orbit control coordinate system based on the near focus coordinate system PQW; in a near focus coordinate system PQW, a P axis points to the near point direction of a track, a W axis points to the normal direction of the track, a Q axis, the P axis and the W axis form a right-hand orthogonal coordinate system, and the corresponding relationship of each axis of the orbit control coordinate system and a PQW coordinate system is as follows: the X axis of the orbit control coordinate system corresponds to a PQW coordinate system-P axis, the Y axis of the orbit control coordinate system corresponds to a PQW coordinate system-W axis, the Z axis of the orbit control coordinate system corresponds to a PQW coordinate system-Q axis, and after the corresponding relation of coordinate axes is converted, psi 2 is rotated around the X axis of the current coordinate system, so that the orbit control coordinate system of the second stage of the electric propulsion orbit transfer is obtained; wherein Ψ 2 is an orbital inclination angle of the second stage;
(2) Establishing electric-pushing ignition target attitude based on arbitrary reference coordinate system
Under the reference of any reference coordinate system, by resolving Euler angle
θ br ，ψ br If the target attitude angle is set as the target attitude angle of the three-axis attitude control, the satellite can be controlled to the target attitude of the electric propulsion ignition based on any reference coordinate system;
firstly, after the ignition attitude of the electric propulsion orbit transfer is established, an attitude conversion matrix C of a satellite body system relative to a reference coordinate system br Should satisfy
C br ·C ri ＝C ORMi1 Or C br ·C ri ＝C ORMi2
Wherein C is ri Setting the three-axis attitude Euler angle of the satellite in the reference coordinate system as
θ br ，ψ br Then there is
When the electric propulsion track is transferred to the first stage to establish the electric propulsion ignition target attitude, the order is given
When the electric propulsion track is transferred to the second stage to establish the electric propulsion ignition target attitude, the order is given
Given Euler angle rotation order, can be changed from C br Obtaining
θ br ，ψ br ；
When the euler rotation order is 3-2-1,
ψ br ＝tan2 -1 (c 12 ,c 11 )∈[-π,π]
wherein the function tan2 -1 (y, x) is defined as
When the euler rotation order 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 other Euler rotation sequences are the same;
(3) Attitude-to-sun tracking during electric push rail transfer
After establishing the attitude of the electric propulsion ignition target by the arbitrary reference coordinate system in the step (2), converting the attitude control during the electric propulsion track transfer into the reference coordinate system by using the track control coordinate system obtained in the step (1); meanwhile, in order to meet the energy requirement during the transfer of the electric propulsion track, the normal sun alignment of the solar wing is realized by rotating around the axis of the ignition direction and rotating around the solar wing;
the included angle between the normal direction of the solar sailboard and the sun direction, namely the energy angle, is controlled to be within the precision requirement range through attitude sun tracking and sun-facing rotation angle control of the solar sailboard, and the requirement of autonomous sun-facing of the normal direction of the solar sailboard is met to guarantee energy.
2. The ignition attitude establishing and dynamic tracking method for electric propulsion orbit transfer as claimed in claim 1, wherein in step (1), the electric propulsion orbit transfer is implemented, i.e. the transfer orbit uses electric propulsion to lift the semi-major axis, reduce eccentricity and reduce the orbit inclination angle.
3. The method for establishing and dynamically tracking the ignition attitude of the electric-propulsion orbit transfer as claimed in claim 1, wherein in the step (1), the orbit control thrust of the first stage is in an XoOoYo plane of the orbit coordinate system, and keeps a certain included angle with + Xo, the absolute value of the orbit control thrust does not become Ψ 1, Ψ 1 ranges from 0 ° to 90 °, but at every half orbit period, the accurate value of the argument depends on the orbit control strategy, and the corresponding argument changes positive and negative at 90 ° and 270 °.
4. The ignition attitude establishing and dynamic tracking method for electric propulsion orbit transfer as claimed in claim 1, wherein in step (1), the orbit control coordinate system of the first stage of electric propulsion orbit transfer is relative to the transformation matrix C of the inertial system ORMi1 From an orbital coordinate system C oi Is rotated to obtain
Wherein Ψ 1 is a trajectory-controlled yaw angle.
5. The method for establishing and dynamically tracking ignition attitude of electric propulsion orbit transfer as claimed in claim 1, wherein in step (1), the orbit control thrust of the second stage maintains a certain orbit control included angle on the basis of inertial orientation, so as to further adjust the orbit inclination angle, the absolute value of which is not changed to Ψ 2, but the positive and negative values are changed when the argument is 90 ° and 270 °.
6. The method as claimed in claim 1, wherein the transformation matrix C from the inertial system to the PQW coordinate system is first calculated according to the orbit elements PQWi
Wherein omega is the amplitude angle of the near place of the track, omega is the right ascension of the ascending intersection point of the track, and i is the inclination angle of the track;
further obtaining an orbit control coordinate system C of the second stage of electric propulsion orbit transfer ORMi2 ，
Wherein:
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 establishing and dynamic tracking method for electric propulsion orbit transfer according to claim 1, characterized in that if the electric propulsion orbit transfer is in the first stage, the ignition attitude is based on the orbit system reference, the sun direction is time-varying, attitude dynamic bias is required to be realized for tracking energy, that is, during the electric propulsion orbit transfer of the transfer orbit, the satellite + Z axis points to the orbit control thrust direction, and when the sailboard points to the sun, the body system rotates around the Z axis of the orbit control coordinate system by a certain angle, so that the XOZ plane of the satellite coincides with the plane formed by the sun direction and the thrust direction.
8. The method for establishing ignition attitude and dynamically tracking electric-propulsion orbit transfer as claimed in claim 7, wherein the expression S of the solar vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi1 Expressing the sun vector in the first stage of the orbital coordinate system, i.e.
S ORM1 ＝[S ORM1X S ORM1Y S ORM1Z ] T ＝C ORMi1 ·S i
The attitude control target is offset by an angle psi around the Z-axis of the orbit control coordinate system based on the orbit control coordinate system of the first stage d Is psi d ＝tan2 -1 (S ORM1Y ,S ORM1X )∈[-π,π]。
9. The method as claimed in claim 1, wherein if the electric propulsion orbit transfer is in the second stage, the ignition attitude is based on the inertial system reference, the sun vector direction is fixed in a short period, so that the attitude offset angle for energy source guarantee is also fixed, that is, during the electric propulsion orbit transfer of the electric propulsion orbit, the + Z axis of the satellite points to the direction of the orbit control thrust, and when the sailboard points to the sun, the body system rotates around the Z axis of the orbit control coordinate system by a certain angle, so that the XOZ plane of the satellite coincides with the plane formed by the sun direction and the thrust direction.
10. The method for establishing ignition attitude and dynamically tracking electric-propulsion orbit transfer according to claim 9, wherein the expression S of the solar vector in the inertial coordinate system is calculated according to the current star time and the solar ephemeris i Conversion matrix C of the orbit control coordinate system relative to the inertial system in the first stage of orbit transfer by electric propulsion ORMi2 Expressing the sun vector in the first stage of the orbital coordinate system, i.e.
S ORM2 ＝[S ORM2X S ORM2Y S ORM2Z ] T ＝C ORMi2 ·S i
The attitude control target is offset by an angle psi around the Z-axis of the orbit control coordinate system based on the orbit control coordinate system of the first stage d Is psi d ＝tan2 -1 (S ORM2Y ,S ORM2X )∈[-π,π]。
11. The method for building and dynamically tracking ignition attitude of electric-push orbit transfer according to claim 1, wherein the system of the satellite is defined as: taking a certain characteristic point of the satellite as an origin, enabling an X axis and a Z axis to be along the characteristic axis direction of the spacecraft, enabling a Y axis to be vertical to the Z axis and the X axis and forming a right-hand orthogonal coordinate system, and enabling the axis direction of the system X, Y, Z to be coincident with the axis directions of Xo, yo and Zo of the orbital coordinate system when the satellite is in a zero attitude;
the orbital coordinate system is defined as: the origin Oo is the center of mass of the satellite, the Zo axis points to the geocentric, 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 forms a right-hand orthogonal coordinate system.
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