CN115626304A - Triaxial angular momentum unloading method for geosynchronous orbit satellite - Google Patents
Triaxial angular momentum unloading method for geosynchronous orbit satellite Download PDFInfo
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Abstract
The invention relates to a triaxial angular momentum unloading method for a geosynchronous orbit satellite, which comprises the following steps of: s1, determining angular momentum of a thruster to be unloaded; s2, making an unloading scheme according to the determined angular momentum to be unloaded; and S3, executing angular momentum unloading, and controlling the mechanical arm to adjust the pointing direction of a thruster at the tail end of the mechanical arm. The geosynchronous orbit satellite is utilized, a thruster is required to be periodically adopted for orbit control and angular momentum unloading, and the traditional satellites are respectively used, so that more propellants are required to be consumed. The invention utilizes the thrust controlled by the periodic orbit, adopts the mechanical arm to actively deflect the position and the direction of the thrust, and can utilize the orbit-controlled thrust to unload the angular momentum of the satellite. The fuel is saved, the control times of the thruster can be reduced, and the interference of the starting of the thruster on the satellite attitude is reduced.
Description
Technical Field
The invention relates to a triaxial angular momentum unloading method for a geosynchronous orbit satellite.
Background
Geosynchronous orbit satellites are affected by disturbance forces and disturbance moments in orbit, and the orbits and attitudes of the geosynchronous orbit satellites deviate from nominal positions, so that the orbits and attitudes of the satellites need to be maintained and controlled. Particularly, for a single-wing sailboard geosynchronous orbit satellite, under the sunlight pressure, the angular momentum of the satellite accumulates about 40-100 Nms every day, and the satellite angular momentum actuating mechanism is generally not saturated until the satellite is unloaded every day, so that the satellite can work normally.
For track maintenance control, i.e. by performing engine ignition in the appropriate direction at the appropriate timing to generate the appropriate speed increment, the track control is achieved; the method is characterized in that an expected attitude control moment arm is generated by actively deviating the mass center of an engine thrust vector while the orbit is maintained, the expected attitude control moment is realized under the action of thrust, and the attitude maintenance and control are realized under the condition of almost no extra fuel consumption.
The thruster is arranged at the tail end of the mechanical arm, so that the direction and the position of the thruster can be adjusted, and the attitude control and the rail control can be simultaneously realized theoretically.
Advantages and necessities of the mechanical arm in cooperation with its end-mounted thruster:
1. the geostationary satellite is generally a large-platform satellite, the size of a sailboard and a load is large, and due to the influence of plume of a thruster, the thruster is not convenient to mount on a satellite body, and an effective mode is that the thruster is extended out of the satellite body through a mechanical arm;
2. in engineering, by combining whole star layout, the tracks and postures in all required directions can be kept and controlled by one set of mechanical arm and the thrusters, and the number of the thrusters can be greatly saved;
3. because the mechanical arm and the thruster can provide the optimal thrust direction and thrust position, the method is equivalent to the mode of generating resultant force by combining a plurality of thrusters in different directions, and the fuel utilization efficiency can reach the optimal.
4. For some satellites with attitude scanning or attitude unsteady value offset requirements, the influence of the attitude can be counteracted by controlling the mechanical arm to follow the attitude, so that the thruster is maintained at the optimal position and direction, and the expected orbit and attitude maintenance is completed.
5. In the traditional method, two sets of mechanical arms are required to be configured for attitude maintaining control, or ignition is required to be carried out at two different positions on a track, so that complete three-axis attitude maintaining control can be completed, otherwise only two-axis attitude control can be realized, and three-axis attitude unloading cannot be completed.
Disclosure of Invention
Aiming at the problems, the invention provides a triaxial angular momentum unloading method for a geosynchronous orbit satellite, which adopts a single mechanical arm and a thruster to realize triaxial angular momentum unloading and comprises the following steps:
s1, determining angular momentum of a thruster to be unloaded;
s2, making an unloading scheme according to the determined angular momentum to be unloaded, and determining the direction of the thruster;
and S3, executing angular momentum unloading, and controlling the mechanical arm to adjust the pointing direction of a thruster at the tail end of the mechanical arm.
Further, the S1 includes the steps of:
s11, calculating the angular momentum to be unloaded of the three shafts under the inertial system;
s12, calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system;
s13, converting the angular momentum to be unloaded under the inertial system into the angular momentum to be unloaded under the satellite system through the attitude transfer matrix;
s14, calculating a transfer matrix from the satellite system to a mechanical arm coordinate system;
and S15, converting the angular momentum to be unloaded in the satellite system into the unloading angular momentum of the mechanical arm coordinate system through the transfer matrix in the S14.
Further, the S2 includes the steps of:
s21, determining an angular momentum unloading intermediate parameter;
s22, determining the tail end position of the angular momentum unloading mechanical arm;
and S23, determining the direction of the thruster at the tail end of the angular momentum unloading mechanical arm.
Further, the S3 includes the steps of:
s31, determining the reciprocating motion times of the mechanical arm;
and S32, judging an ignition interval and controlling the mechanical arm to adjust the direction of the thruster according to the ignition interval.
Further, the angular momentum to be unloaded of the three axes in the inertial system calculated in S11 is:
delta_ATT_H_fire_x_i＝-ATT_H_fire_x_i-Dx×T_day/2；
delta_ATT_H_fire_y_i＝-ATT_H_fire_y_i-Dy×T_day/2；
delta_ATT_H_fire_z_i＝-ATT_H_fire_z_i-Dz×T_day/2；
wherein, ATT _ H _ fire _ x _ i, ATT _ H _ fire _ y _ i and ATT _ H _ fire _ z _ i are the triaxial angular momentum of the satellite inertial system at the next theoretical moment respectively; dx, dy and Dz are solar light pressure interference moments of three satellite shafts under an inertial system; t _ day =86400s is seconds a day;
s12, calculating quaternion from the theoretical ignition moment inertial system to the orbital system:
qio＝quatmultiply([cos(0.5×u_fire)0 0sin(0.5×u_fire)],[0.5 -0.5 -0.5 0.5])
wherein u _ fire is a track argument corresponding to the theoretical ignition moment; quatmultiply (A, B) represents the multiplication of quaternion A and quaternion B;
and (3) calculating quaternion from the theoretical ignition moment inertial system to the satellite body system:
qib＝quatmultiply(qio,qob)；
wherein, qb is a quaternion from a satellite orbital system to a satellite body system, and the value is changed according to the different earth pointing regions of the satellite;
calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system:
Abi＝quat2dcm(qib)；
wherein, quat2dcm (: indicates that quaternion in the parenthesis is converted into corresponding attitude transition matrix;
s13, calculating the three-axis angular momentum to be unloaded ATT _ H _ fire _ b in the satellite body coordinate system, and calculating the angular momentum to be unloaded in the satellite body system:
ATT_H_fire_b＝Abi×[delta_ATT_H_fire_x_i；delta_ATT_H_fire_y_i；delta_ATT_H_fire_z_i]；
s14, calculating a transfer matrix qbs from the satellite system to a mechanical arm coordinate system, wherein three direction speed increments of a radius cutting method are respectively Vr, vt and Vn;
if the mechanical arm is arranged on the south side (+ Ys direction) of the satellite, the speed increment is along the normal positive direction, namely Vn >0;
qbs＝quatmultiply(quatinv(qob),[cos(theta/2)-Vr/sqrt(Vr 2 +Vt 2 )×sin(theta/2)0-Vt/sqrt(Vr 2 +Vt 2 )×sin(theta/2)])；
wherein, the quatinv (A) is quaternion inversion operation, and the calculation process is as follows: a = (a 0; a1; a2; a 3), then quatinv (A) = (a 0; -a1; -a2; -a 3);
theta＝arccos(|Vn/sqrt(Vr 2 +Vt 2 +Vn 2 )|)；
s15, calculating satellite triaxial unloading angular momentum ATT _ H _ fire _ S under a mechanical arm coordinate system
ATT_H_fire_s＝inv(Abs)×ATT_H_fire_b；
Wherein inv (a) is a matrix inversion operation, abs is a posture transfer matrix from the robot arm coordinate system to the satellite body system, abs = quat2dcm (qsb), and qsb = quatnv (qbs).
Further, the Abs calculation process is:
if qsb = quatinv (qbs) = (q) 0 ；q 1 ；q 2 ；q 3 )，q 0 ，q 1 ，q 2 ，q 3 Are quaternion values respectively;
Abs＝[1-2(q 0 2 +q 3 2 ),2(q 1 q 2 +q 0 q 3 ),2(q 1 q 3 -q 0 q 2 )；2(q 1 q 2 -q 0 q 3 ),1-2(q 1 2 +q 3 2 ),2(q 2 q 3 +q 0 q 1 )；2(q 1 q 3 +q 0 q 2 ),2(q 2 q 3 -q 0 q 1 ),1-2(q 1 2 +q 2 2 )]。
in summary, the geosynchronous orbit satellite needs to regularly adopt thrusters for orbit control and angular momentum unloading, and the traditional satellites need to consume more propellants when being respectively carried out.
The invention utilizes the thrust controlled by the periodic orbit, adopts the mechanical arm to actively deflect the position and the direction of the thrust, and can utilize the orbit-controlled thrust to unload the angular momentum of the satellite; the fuel is saved, the control times of the thruster can be reduced, and the interference of the starting of the thruster on the satellite attitude is reduced;
the invention is not only suitable for the traditional angular momentum unloading scheme of multiple rail ignition of 30847Hr per rail, but also is suitable for the working condition that each rail is ignited for only 1 time, thereby enlarging the application range of the invention; the characteristic that the mechanical arm can move in multiple degrees of freedom is utilized;
the invention is not only suitable for the satellite fixed pointing condition, but also suitable for the attitude maneuver condition.
Drawings
FIG. 1 is a schematic view of an angular momentum unloading strategy of the present invention using a single mechanical arm plus a thruster.
Detailed Description
The following describes in detail a three-axis angular momentum offloading method for geosynchronous orbit satellites according to the present invention with reference to the accompanying drawings and the following detailed description.
S1, determining angular momentum to be unloaded during the working period of the thruster;
s2, making an unloading scheme according to the determined angular momentum to be unloaded;
and S3, executing angular momentum unloading, and adjusting the pointing direction of a thruster at the tail end of the mechanical arm by controlling the mechanical arm.
Wherein S1 comprises the following steps:
s11, calculating the angular momentum to be unloaded of the three shafts under the inertial system according to the following formula, wherein the unit is Nms:
delta_ATT_H_fire_x_i＝-ATT_H_fire_x_i-Dx×T_day/2；
delta_ATT_H_fire_y_i＝-ATT_H_fire_y_i-Dy×T_day/2；
delta_ATT_H_fire_z_i＝-ATT_H_fire_z_i-Dz×T_day/2；
wherein, ATT _ H _ fire _ x _ i, ATT _ H _ fire _ y _ i and ATT _ H _ fire _ z _ i are the three-axis angular momentum of the satellite inertial system at the next theoretical moment in the unit of Nms respectively; dx, dy and Dz are sunlight pressure interference moments of the satellite triaxial in the inertial system (the sunlight pressure interference moments are generally approximate to constant values in the inertial system), and the unit is Nm; t _ day =86400s is seconds of the day.
S12, calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system;
and (3) calculating a quaternion from a theoretical ignition moment inertia system to an orbital system:
qio＝quatmultiply([cos(0.5×u_fire)0 0sin(0.5×u_fire)],[0.5 -0.5 -0.5 0.5])
wherein u _ fire is a track amplitude corresponding to the theoretical ignition moment and is unit rad; quatmultiply (A, B) represents the multiplication of quaternion A and quaternion B;
and (3) calculating quaternion from the theoretical ignition moment inertial system to the satellite body system:
qib＝quatmultiply(qio,qob)；
wherein, qb is a quaternion from a satellite orbital system to a satellite body system, and the value is related to the task characteristics of the satellite and changes according to different satellite pointing earth areas;
calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system:
Abi＝quat2dcm(qib)；
wherein, quat2dcm (a) represents the conversion of quaternion a into a corresponding attitude transfer matrix;
s13, calculating the three-axis angular momentum to be unloaded ATT _ H _ fire _ b in a satellite body coordinate system, wherein the unit is Nms:
calculating angular momentum to be unloaded under the satellite system, wherein the unit is Nms:
ATT_H_fire_b＝Abi×[delta_ATT_H_fire_x_i；delta_ATT_H_fire_y_i；delta_ATT_H_fire_z_i]；
s14, calculating a transfer matrix qbs from the satellite system to the mechanical arm coordinate system;
the three-direction speed increment of the diameter cutting method is respectively Vr, vt and Vn.
If the mechanical arm is installed on the south side (+ Ys direction) of the satellite, the speed increment is along the normal positive direction, namely Vn >0.
qbs＝quatmultiply(quatinv(qob),[cos(theta/2)-Vr/sqrt(Vr 2 +Vt 2 )×sin(theta/2)0-Vt/sqrt(Vr 2 +Vt 2 )×sin(theta/2)])；
Wherein, the quatinv (A) is quaternion inversion operation, and the calculation process is as follows: a = (a 0; a1; a2; a 3), then quatinv (A) = (a 0; -a1; -a2; -a 3);
theta＝arccos(|Vn/sqrt(Vr 2 +Vt 2 +Vn 2 )|)；
s15, calculating satellite three-axis unloading angular momentum ATT _ H _ fire _ S under the mechanical arm coordinate system
ATT_H_fire_s＝inv(Abs)×ATT_H_fire_b；
Wherein inv (A) is matrix inversion operation, abs is an attitude transfer matrix from a mechanical arm coordinate system to a satellite body system, the attitude transfer matrix Abs is obtained by conversion of quaternion qsb, and quaternion qsb and quaternion qbs are in inverse operation relationship with each other; qsb = quatinv (qbs) = (q) 0 ；q 1 ；q 2 ；q 3 ) Abs = quat2dcm (qsb); in particular, the amount of the solvent to be used,
Abs＝[1-2(q 0 2 +q 3 2 ),2(q 1 q 2 +q 0 q 3 ),2(q 1 q 3 -q 0 q 2 )；2(q 1 q 2 -q 0 q 3 ),1-2(q 1 2 +q 3 2 ),2(q 2 q 3 +q 0 q 1 )；2(q 1 q 3 +q 0 q 2 ),2(q 2 q 3 -q 0 q 1 ),1-2(q 1 2 +q 2 2 )]。
ATT _ H _ fire _ s is a three-dimensional vector that can be expressed as (ATT _ H _ fire _ s (1); ATT _ H _ fire _ s (2); ATT _ H _ fire _ s (3)); wherein
ATT_H_fire_x_s＝ATT_H_fire_s(1)；
ATT_H_fire_y_s＝ATT_H_fire_s(2)；
ATT_H_fire_z_s＝ATT_H_fire_s(3)。
S21, determining angular momentum unloading intermediate parameters
If sqrt (ATT _ H _ fire _ x _ s) 2 +ATT_H_fire_y_s 2 +ATT_H_fire_z_s 2 )<yuzhi _1, yuzhi _1represents a first judgment threshold value of unloading, and when the formula is met, the three-axis angular momentum unloading is not required to be executed at this time; then the
R＝0
delta_L＝0
Alpha＝0
Beta＝0
Variable definitions are shown in fig. 1:
r is an XsOZs plane angular momentum unloading equivalent moment arm in the unit of m; delta _ L is the angular momentum unloading arm of the Ys shaft, and the unit is m; the Alpha needs the deflection angle of the mechanical arm, unit rad, for generating the angular momentum unloading in the direction Ys; beta is the included angle between the projection of the tail end position of the mechanical arm in the XsOZs plane and Xs, and unit rad;
otherwise, if sqrt (ATT _ H _ fire _ x _ s) 2 +ATT_H_fire_z_s 2 )<yuzhi _2 and yuzhi _2are second judgment thresholds for unloading, and when the formula is met, the XsOZs face angular momentum does not need to be unloaded at this time, but the Ys axis angular momentum needs to be unloaded;
R＝0
delta_L＝2×R_limit
Beta＝pi/4
Alpha＝arcsin(-ATT_H_fire_y_s/F/(0.5×mod((ufire_t-ufire_0),2×pi)/we)/delta_L)；
wherein, R _ limit is the maximum radius, unit and m allowed by the end position of the mechanical arm for unloading angular momentum in the XsOZs plane;
pi represents a circumference ratio;
f is the thrust magnitude in N;
ufire _ t is the track amplitude corresponding to the ignition end point, and unit rad;
ufire _0 is the track amplitude corresponding to the ignition starting point, in units of rad;
we is the satellite orbital angular velocity (and also the earth's autonomous angular velocity when in geosynchronous orbit), in units rad.
If | Alpha | > Alpha _ yuzhi, then Alpha = sign (Alpha) × Alpha _ yuzhi
Wherein, alpha _ yuzhi is the maximum value allowed by the deviation of the tail end of the mechanical arm from the original pointing angle when the Ys needs to be unloaded, and the number is injected from the ground according to the accumulation condition of the angular momentum of the Ys and does not exceed pi/6 (30 degrees).
Otherwise, if the absolute value ATT _ H _ fire _ y _ s is smaller than yuzhi _1, it represents that the angular momentum of the surface of the XsOZs needs to be unloaded at this time, but the angular momentum of the axis Ys does not need to be unloaded;
R＝sqrt(ATT_H_fire_x_s 2 +ATT_H_fire_z_s 2 )/F/(mod((ufire_t-ufire_0),2×pi)/we)；
wherein F is the thrust magnitude in N; ufire _ t is the track amplitude corresponding to the ignition end point, and is unit rad; ufire _0 is the track amplitude corresponding to the ignition starting point, in units of rad; we is the satellite orbital angular velocity (also the earth autonomous angular velocity when in geosynchronous orbit), in units rad; sqrt represents the root at the second degree, mod (x, y) represents the remainder of x divided by y.
If R > R _ limit, then R = R _ limit
delta_L＝0；
Alpha＝0；
Beta＝atan2(-ATT_H_fire_x_s,-ATT_H_fire_z_s)
Wherein the atan2 (y, x) value range is-pi to pi, specifically:
when x >0, atan2 (y, x) = arctan (y/x), the range of values is (-pi/2, pi/2);
when x <0, atan2 (y, x) = pi + arctan (y/x), the range of values is (pi/2, 1.5 × pi);
when x =0 and y >0, atan2 (y, x) = pi/2;
when x =0 and y <0, atan2 (y, x) = -pi/2;
when x =0 and y =0, atan2 (y, x) =0;
otherwise, the triaxial angular momentum all needs to be unloaded
R＝sqrt(ATT_H_fire_x_s 2 +ATT_H_fire_z_s 2 )/F/(mod((ufire_t-ufire_0),2×pi)/we)
If R > R _ limit, then R = R _ limit
delta_L＝0.2
Alpha＝arcsin(-ATT_H_fire_y_s/F_lowthrust/(0.5×mod((ufire_t-ufire_0),2×pi)/we)/delta_L)
If | Alpha | > Alpha _ yuzhi, then Alpha = sign (Alpha) × Alpha _ yuzhi
Beta＝atan2(-Xs,-Zs)
In order to ensure that the logic described above is true, it is necessary to satisfy: yuzhi _1>, yuzhi_2;
s22, determining the tail end position of the angular momentum unloading mechanical arm;
projection Pos _ crash _ x _1 of the unloading position 1 of the outer ring of the mechanical arm on an Xs axis: unit m
Pos_thrust_x_1＝(R+0.5×delta_L)×cos(Beta)
Projection of mechanical arm outer ring unloading position 1 on Zs axis Pos _ threst _ z _1: unit m
Pos_thrust_z_1＝-(R+0.5×delta_L)×sin(Beta)
Projection of mechanical arm inner ring unloading position 2 on Xs axis Pos _ threst _ x _2: unit m
Pos_thrust_x_2＝(R-0.5×delta_L)×cos(Beta)
Projection Pos _ crash _ z _2 of the unloading position 2 of the inner ring of the mechanical arm on the Zs axis: unit m
Pos_thrust_z_2＝-(R-0.5×delta_L)×sin(Beta)
S23, determining the direction of a thruster at the tail end of the angular momentum unloading mechanical arm;
n_x＝ATT_H_fire_z_s/sqrt(ATT_H_fire_x_s 2 +ATT_H_fire_z_s 2 )
n_y＝0
n_z＝-ATT_H_fire_x_s/sqrt(ATT_H_fire_x_s 2 +ATT_H_fire_z_s 2 )
n1＝[n_x n_y n_z]
n2＝[-n_x-n_y-n_z]
the unloading position 1 of the outer ring of the mechanical arm corresponds to the direction vec1 of the thruster
vec1＝cos(Alpha)×[0 1 0]+(1-cos(Alpha))×(dot(n1,[0 1 0]))×n1+sin(Alpha)×(cross(n1,[0 1 0]))；
The unloading position 2 of the inner ring of the mechanical arm corresponds to the direction vec2 of the thruster
vec2＝cos(Alpha)×[0 1 0]+(1-cos(Alpha))×(dot(n2,[0 1 0]))×n2+sin(Alpha)×(cross(n2,[0 1 0]))；
Wherein: dot (A, B) is a dot product of two 3 × 1 vectors A and B; cross (a, B) is a cross product of two 3 × 1 vectors a and B;
s31, determining the reciprocating motion frequency N _ cycle of the mechanical arm, wherein the N _ cycle is an integer which is larger than or equal to 1, and the smaller the N _ cycle is, the larger the interference on other two shafts is when the Ys shaft is unloaded, and the advantage is that the mechanical arm action frequency is smaller; and vice versa. In actual engineering, it is generally recommended that the N _ cycle is between 1 and 6, as the N _ cycle is smaller, within an allowable range, as well as in actual engineering.
S32, in the range of 0-1/(2 XN _ cycle) and 8230in the ignition region, in the range of 2 XN _ cycle-2/(2 XN _ cycle) to 2 XN _ cycle-1/(2 XN _ cycle), the tail end of the mechanical arm is positioned at the unloading position 1 of the outer ring of the mechanical arm, and the mechanical arm is controlled to enable the thruster to point to vec1;
in the section of 1/(2 XN _ cycle) -2/(2 XN _ cycle), \ 8230, 2 XN _ cycle-1/(2 XN _ cycle) -1 of the ignition section, the tail end of the mechanical arm is at the unloading position 2 of the inner ring of the mechanical arm, and the mechanical arm is controlled to lead the thruster to point to the vec2.
The invention has the following beneficial effects:
1. the geosynchronous orbit satellite is utilized, a thruster is required to be periodically adopted for orbit control and angular momentum unloading, and the traditional satellites are respectively used, so that more propellants are required to be consumed. The invention utilizes the thrust controlled by the periodic orbit, adopts the mechanical arm to actively deflect the position and the direction of the thrust, and can utilize the orbit-controlled thrust to unload the angular momentum of the satellite. The fuel is saved, the control times of the thruster can be reduced, and the interference of the starting of the thruster on the satellite attitude is reduced;
2. the invention is not only suitable for the traditional angular momentum unloading scheme of multiple rail ignition of 30847Hr per rail, but also is suitable for the working condition that each rail is ignited for only 1 time, thereby enlarging the application range of the invention;
3. by utilizing the characteristic that the mechanical arm can move in multiple degrees of freedom, the invention is not only suitable for the situation of fixed pointing of the satellite, but also suitable for the situation of attitude maneuver (the attitude maneuver, namely qb is time-varying).
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be limited only by the attached claims.
Claims (10)
1. A triaxial angular momentum unloading method for a geosynchronous orbit satellite is characterized by comprising the following steps of:
s1, determining angular momentum of a thruster to be unloaded;
s2, making an unloading scheme according to the determined angular momentum to be unloaded, and determining the direction of the thruster;
and S3, executing angular momentum unloading, and controlling the mechanical arm to adjust the pointing direction of a thruster at the tail end of the mechanical arm.
2. The method of unloading triaxial angular momentum of a geosynchronous orbiting satellite according to claim 1, wherein said S1 comprises the steps of:
s11, calculating the angular momentum to be unloaded of the three shafts under the inertial system;
s12, calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system;
s13, converting the angular momentum to be unloaded under the inertial system into the angular momentum to be unloaded under the satellite system through the attitude transfer matrix;
s14, calculating a transfer matrix from the satellite system to a mechanical arm coordinate system;
and S15, converting the angular momentum to be unloaded in the satellite system into the unloading angular momentum of the mechanical arm coordinate system through the transfer matrix in the S14.
3. The method for triaxial angular momentum offloading of a geosynchronous orbit satellite according to claim 1, wherein said S2 comprises the steps of:
s21, determining an angular momentum unloading intermediate parameter;
s22, determining the tail end position of the angular momentum unloading mechanical arm;
and S23, determining the direction of a thruster at the tail end of the angular momentum unloading mechanical arm.
4. The method for triaxial angular momentum offloading of a geosynchronous orbit satellite according to claim 1, wherein said S3 comprises the steps of:
s31, determining the reciprocating motion times of the mechanical arm;
and S32, judging an ignition interval and controlling the mechanical arm to adjust the direction of the thruster according to the ignition interval.
5. The method for offloading angular momentum of geosynchronous orbit satellite of claim 2, wherein the angular momentum to be offloaded for the three axes of the computed inertia system in S11 is:
delta_ATT_H_fire_x_i＝-ATT_H_fire_x_i-Dx×T_day/2；
delta_ATT_H_fire_y_i＝-ATT_H_fire_y_i-Dy×T_day/2；
delta_ATT_H_fire_z_i＝-ATT_H_fire_z_i-Dz×T_day/2；
wherein, ATT _ H _ fire _ x _ i, ATT _ H _ fire _ y _ i and ATT _ H _ fire _ z _ i are the triaxial angular momentum of the satellite inertial system at the next theoretical moment respectively; dx, dy and Dz are sunlight pressure interference moments of the three satellite axes under an inertial system; t _ day =86400s is seconds of the day.
6. The method for unloading triaxial angular momentum of a geosynchronous orbit satellite according to claim 5, wherein S12, the quaternion of the inertial system to the orbital system at the theoretical ignition moment is calculated:
qio＝quatmultiply([cos(0.5×u_fire)0 0sin(0.5×u_fire)],[0.5-0.5-0.5 0.5])；
wherein u _ fire is a track argument corresponding to the theoretical ignition moment; quatmultiply (A, B) represents the multiplication of quaternion A and quaternion B;
and (3) calculating quaternion from a theoretical ignition moment inertial system to a satellite body system:
qib＝quatmultiply(qio,qob)；
wherein, qb is a quaternion from a satellite orbital system to a satellite body system, and the value is changed according to the difference of the satellite pointing to the earth region;
calculating an attitude transfer matrix from a theoretical ignition moment inertial system to a satellite body system:
Abi＝quat2dcm(qib)；
where quat2dcm (: indicates that the quaternion in the parenthesis is converted to the corresponding attitude transition matrix.
7. The method for unloading three-axis angular momentum of a geosynchronous orbit satellite according to claim 6, wherein S13, the angular momentum to be unloaded ATT _ H _ fire _ b in the three-axis system of the satellite is calculated, and the angular momentum to be unloaded in the system of the satellite is calculated:
ATT_H_fire_b＝Abi×[delta_ATT_H_fire_x_i；delta_ATT_H_fire_y_i；
delta_ATT_H_fire_z_i]。
8. the method for unloading triaxial angular momentum of a geosynchronous orbit satellite according to claim 7, wherein S14, calculating a transfer matrix qbs of the satellite system to a robot arm coordinate system, wherein the three-direction velocity increments of the tangent method are Vr, vt and Vn;
if the mechanical arm is arranged on the south side (+ Ys direction) of the satellite, the speed increment is along the normal positive direction, namely Vn >0;
qbs＝quatmultiply(quatinv(qob),[cos(theta/2)-Vr/sqrt(Vr 2 +Vt 2 )×sin(theta/2)
0-Vt/sqrt(Vr 2 +Vt 2 )×sin(theta/2)])；
wherein, the quatinv (A) is quaternion inversion operation, and the calculation process is as follows: a = (a 0; a1; a2; a 3) then quatinv (A) = (a 0; -a1; -a2; -a 3);
theta＝arccos(|Vn/sqrt(Vr 2 +Vt 2 +Vn 2 )|)。
9. the method for unloading three-axis angular momentum of a geosynchronous orbit satellite according to claim 8, wherein S15, the three-axis unloading angular momentum of the satellite ATT _ H _ fire _ S under the robot coordinate system is calculated,
ATT_H_fire_s＝inv(Abs)×ATT_H_fire_b；
wherein inv (a) is a matrix inversion operation, abs is a posture transfer matrix from a mechanical arm coordinate system to a satellite body system, abs = quat2dcm (qsb), and qsb = quatinv (qbs).
10. The method of claim 9, wherein the Abs calculation procedure is:
if qsb = quatinv (qbs) = (q) 0 ；q 1 ；q 2 ；q 3 )，q 0 ，q 1 ，q 2 ，q 3 The quaternion qsb values are respectively;
Abs＝[1-2(q 0 2 +q 3 2 ),2(q 1 q 2 +q 0 q 3 ),2(q 1 q 3 -q 0 q 2 )；2(q 1 q 2 -q 0 q 3 ),1-2(q 1 2 +q 3 2 ),2(q 2 q 3 +
q 0 q 1 )；2(q 1 q 3 +q 0 q 2 ),2(q 2 q 3 -q 0 q 1 ),1-2(q 1 2 +q 2 2 )]。
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