CN116946392B - Geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias - Google Patents
Geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias Download PDFInfo
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Abstract
The invention relates to a geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias, which comprises the following steps: calculating ignition time and orbital inclination control quantity according to the initial orbit and the target orbit of the satellite; determining the ignition position of the electric thruster, and determining the attitude offset angle of each ignition moment according to the calculated track inclination angle control quantity; calculating the ignition time length of the electric thruster; considering track dip perturbation, determining a control batch according to the track dip control quantity and the ignition duration; calculating perturbation terms and radial component force to horizontal longitude change amounts to obtain nominal semi-long axis offset; control is performed. Aiming at the coupling problem of the electric propulsion dip angle to the flat longitude and the semi-long axis, the invention designs an electric propulsion dip angle control scheme based on multi-dimensional attitude bias, provides a flat longitude drift compensation method of the semi-long axis bias, and solves the influence of the coupling thrust of electric propulsion tangential and radial on the flat longitude and the semi-long axis during electric propulsion control.
Description
Technical Field
The invention relates to the technical field of data processing, in particular to a geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias.
Background
Because the electric propulsion technology has the characteristics of high specific impulse, long service life and small thrust, the effective load mass ratio of the spacecraft can be obviously improved, the working life of the spacecraft is prolonged, and the emission cost is reduced, so that the electric propulsion technology can be widely applied to the spacecraft. Electric propulsion in geosynchronous orbit (GEO) satellites is commonly used for orbital transfer, position maintenance, and other maneuvering tasks of the Geosynchronous Transfer Orbit (GTO). In addition, GEO satellites have maneuvering tasks such as tilt control during orbit.
Where tilt angle control is different from north-south position maintenance, typically, the maintenance range of communications GEO satellite north-south position maintenance is at most ±0.1°, with a control amount of 0.2 °. The control amount of the inclination angle control is larger, and is generally more than 0.3 degrees. The inclination angle control needs a relatively large speed increment, chemical propulsion is easy to realize, but the fuel consumption is large. The electric propulsion is used for realizing, the fuel consumption is small, but the time required by the small thrust is long, so that some precision is sacrificed, and the electric propulsion control device has the advantages of keeping the control quantity uniform and stable and being beneficial to the long-term control of the small thrust.
In addition, to avoid the effect of plume and electromagnetic radiation on the payload and solar cell windsurfing board, the electric thruster is mounted on the satellite at a position close to the back floor side, away from the payload as far as possible, and the thrust is inclined to the geocentric direction (Z direction) through the centroid. Meanwhile, in order to reduce pollution of the plume to the solar sailboard on the north-south side of the satellite, a certain included angle is formed between the thruster and the X, Y, Z direction. The installation position of the electric thruster not only generates normal thrust required by the inclination angle control, but also has component force in tangential and radial directions, and can influence the semi-long axis, the eccentricity and the flat longitude of the track. Consideration is given to how to effectively eliminate the effects of semi-long axis, flat longitude variations during tilt angle control, which puts new demands on the control strategy of the electrically propelled satellite.
Accordingly, there is a need to improve one or more problems in the related art as described above.
It is noted that this section is intended to provide a background or context for the technical solutions of the invention set forth in the claims. The description herein is not admitted to be prior art by inclusion in this section.
Disclosure of Invention
The present invention is directed to a geosynchronous satellite electric propulsion tilt control method based on multi-dimensional attitude bias that, at least in part, addresses one or more of the problems due to the limitations and disadvantages of the related art.
The invention provides a geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias, which comprises the following steps:
calculating ignition time and orbital inclination control quantity according to the initial orbit and the target orbit of the satellite;
determining the ignition position of the electric thruster, and determining the attitude offset angle of each ignition moment according to the calculated track inclination angle control quantity;
calculating the ignition time length of the electric thruster according to the ignition time, the track inclination angle control quantity and the attitude offset angle;
considering track dip perturbation, determining a control batch according to the track dip control quantity and the ignition duration;
calculating perturbation items and the radial component force of the electric thruster to the horizontal longitude change amount to obtain a nominal semi-long axis offset;
and performing control according to the obtained nominal semi-long axis offset and the control batch.
Preferably, the step of performing control according to the obtained nominal semi-major axis offset and the control batch comprises the steps of:
performing chemical propulsion control based on the nominal semi-major axis offset;
and performing electric propulsion control according to the track inclination angle control amount and the control batch.
Preferably, the control method further comprises the steps of:
and evaluating the control result by using three parameters of the track inclination angle control quantity, the nominal semi-long axis offset quantity and the eccentricity ratio.
Preferably, the calculation process of the step of calculating the ignition timing and the orbital tilt control amount according to the initial orbit and the target orbit of the satellite is as follows:
component i of initial orbital tilt in polar x-axis 0x And a component i on the y-axis 0y The method comprises the following steps of:
component i of the target orbit inclination in the polar x-axis 1x And a component i on the y-axis 1y The method comprises the following steps of:
the control amount of the track dip angle isThe units are degrees:
the control direction of the vector of the track inclination angle control quantity is as follows:
the moment of the ignition middle point is t of each day 1 Time sum t 2 When (1):
wherein i is 0 For initial track pitch, i 1 For a target track pitch angle, a is the nominal semi-major axis offset,for ascending intersection point, right-left>Is the amplitude angle of the near place +.>For the straight-up point angle +.>Is the track epoch time, n is the track angular velocity.
Preferably, ignition is performed using an electric thruster of the ne+sw or se+nw combination.
Preferably, after ignition is performed by adopting an electric thruster combined by NE+SW or SE+NW, the electric thruster generates normal thrust to the satellite by adjusting the X-axis gesture of the rolling shaft; by adjusting the Y-axis attitude of the pitching axis, the electric thruster generates tangential thrust to the satellite.
Preferably, the nominal semi-major axis offset comprises: j (J) 22 Term perturbation versus latitude change amount and radial component versus latitude change amount.
Preferably, J 22 The term perturbation vs. flat longitude change amount calculation process is as follows:
according to drift velocity D of flat longitude and drift initial velocity D 0 Drift acceleration of plane longitudeThe relation is:
integrating the above method to obtain:
at the position ofIn the plane, the solution of the above equation is a parabolic curve with the opening right or left, when the flat longitude drifts the acceleration +.>When the opening is right; drift acceleration +.>When the opening is left, lambda 0 For the initial flat longitude, t is the single-batch continuous ignition duration, and lambda is the drift flat longitude;
flat longitude drift accelerationThe method meets the following conditions:
wherein,=-14.545°；
calculatingThereby deriving for counteracting J 22 Semi-major axis offset a due to term perturbation 1 。
Preferably, the radial component force versus plane longitude change amount calculation process is as follows:
wherein ΔD is the amount of change in the flat longitude drift rate, F 0 For the initial electric thrust, alpha is the attitude offset angle of the rolling axis, beta is the attitude offset angle of the pitching axis, a 2 To offset the semi-long axis offset brought about by the radial force component.
The technical scheme provided by the invention can comprise the following beneficial effects:
aiming at the coupling problem of the electric propulsion dip angle to the flat longitude and the semi-long axis, the invention designs an electric propulsion dip angle control scheme based on multidimensional gesture bias, provides a flat longitude drift compensation method of the semi-long axis bias, and solves the influence of the coupling thrust of the electric propulsion tangential and radial directions on the flat longitude and the semi-long axis during the electric propulsion dip angle control.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the disclosure and together with the description, serve to explain the principles of the disclosure. It will be apparent to those of ordinary skill in the art that the drawings in the following description are merely examples of the disclosure and that other drawings may be derived from them without undue effort.
FIG. 1 illustrates a flow chart of a geosynchronous satellite electric propulsion tilt control method based on multi-dimensional attitude biasing in an exemplary embodiment of the invention;
FIG. 2 illustrates a schematic layout of an electric thruster in an exemplary embodiment of the present invention;
FIG. 3 illustrates a schematic diagram of a tilt-control ignition strategy in an exemplary embodiment of the present invention;
FIG. 4 illustrates a schematic diagram of attitude adjustment in an exemplary embodiment of the invention;
FIG. 5 shows a polar plot of tilt angle change in an exemplary embodiment of the invention;
fig. 6 illustrates a graph of variation in flat longitude at the time of tilt angle control in an exemplary embodiment of the present invention;
FIG. 7 is a polar plot of eccentricity change during tilt angle control in an exemplary embodiment of the present invention;
fig. 8 shows a graph of eccentricity change at the time of tilt angle control in an exemplary embodiment of the present invention.
Detailed Description
Example embodiments will now be described more fully with reference to the accompanying drawings. However, the exemplary embodiments may be embodied in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concept of the example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.
Furthermore, the drawings are merely schematic illustrations of embodiments of the invention and are not necessarily drawn to scale. The same reference numerals in the drawings denote the same or similar parts, and thus a repetitive description thereof will be omitted. Some of the block diagrams shown in the figures are functional entities and do not necessarily correspond to physically or logically separate entities.
In this example embodiment, a geosynchronous satellite electric propulsion dip control method based on multidimensional attitude bias is provided, and referring to fig. 1, the detection method includes the following steps:
step S101: and calculating the ignition time and the orbit inclination angle control quantity according to the initial orbit and the target orbit of the satellite.
Step S102: and determining the ignition position of the electric thruster, and determining the attitude offset angle of each ignition moment according to the calculated track inclination angle control quantity. Wherein the attitude offset angle includes a roll axis attitude offset angle and a pitch axis attitude offset angle.
Step S103: and calculating the ignition time length of the electric thruster according to the ignition time, the track inclination angle control quantity and the attitude offset angle. The ignition duration here is the ignition duration of a single batch.
Step S104: and considering track inclination perturbation, determining a control batch according to the track inclination control quantity and the ignition duration.
Step S105: calculating perturbation term and radial component force of the electric thruster to change the plane longitude to obtain nominal semi-long axis offset.
Step S106: and performing control according to the obtained nominal semi-long axis offset and the control batch.
Aiming at the coupling problem of the electric propulsion dip angle to the flat longitude and the semi-long axis, the invention designs an electric propulsion dip angle control scheme based on multidimensional gesture bias, provides a flat longitude drift compensation method of the semi-long axis bias, and solves the influence of the coupling thrust of the electric propulsion tangential and radial directions on the flat longitude and the semi-long axis during the electric propulsion dip angle control.
The geosynchronous satellite electric propulsion dip control method based on multidimensional attitude bias of the invention is described in detail below.
Taking BSS-702 series satellites of Boeing company as an example, a multi-reference satellite BSS-702 series satellite is distributed by a national in-orbit GEO electric thruster, 4 electric thrusters are adopted, two electric thrusters are symmetrically arranged on each of the two sides of the north and the south and are arranged on the outer surface of a back floor at the position close to a north-south partition plate, and according to the installation quadrants of each thruster in the whole satellite system, the thrusters are respectively defined as NE (North east), NW (North west), SE (south east) and SW (south west). Each electric thruster is provided with a vector adjustment mechanism for adjusting the thrust output direction, for example parallel to the Z-axis during transfer of the electric thrusting trajectory, or directed towards the centroid during maintenance control, or directed towards a given orientation according to the angular momentum unloading need.
The layout of the electric thrusters is shown in fig. 2, in which the tangential direction (T) is positive in the X-axis direction, the normal direction (N) is positive in the Y-axis negative direction, and the radial direction (R) is positive in the Z-axis positive direction in the satellite body coordinate system, and the signs of each thruster in each direction are shown in table 1.
TABLE 1 sign of the thruster in each direction
1. According to the initial orbit and the target orbit of the satellite, calculating the ignition time and the orbit inclination angle control quantity
Component i of initial orbital tilt in polar x-axis 0x And a component i on the y-axis 0y The method comprises the following steps of:
component i of the target orbit inclination in the polar x-axis 1x And a component i on the y-axis 1y The method comprises the following steps of:
the control amount of the track dip angle isThe units are degrees:
the control direction of the track inclination angle control quantity vector is as follows:
the moment of the ignition middle point is t of each day 1 Time sum t 2 When (1):
wherein i is 0 For initial track pitch, i 1 For a target track pitch angle, a is the nominal semi-major axis offset,for ascending intersection point, right-left>Is the amplitude angle of the near place +.>For the straight-up point angle +.>Is the track epoch time, n is the track average angular velocity.
The track average angular velocity n is as follows:
wherein,is the gravitational constant.
2. Determining a thermal power thruster and determining an attitude offset angle of each ignition moment according to the calculated track inclination angle control quantity
(1) With reference to fig. 3, the tangential thrust and the normal thrust of the electric thruster of the ne+sw or se+nw combination are offset, and only the speed increment is generated in the radial direction, and the generated moment can be offset, so that the speed increment of the satellite in the normal direction can be realized by a satellite rolling gesture offset mode. When the inclination angle is controlled, the X-axis (rolling) gesture is adjusted, and the two thrusters are ignited to generate normal thrust; when controlling the semi-long axis, the Y-axis (pitch) attitude is adjusted, generating tangential thrust.
By adjusting the roll axis attitude, as shown in fig. 4a, a normal thrust is generated. Thrust force F in the normal direction Y Is that
（1）
Thrust force F in radial direction Z Is that
（2）
By adjusting the pitch axis attitude, tangential thrust is generated as shown in fig. 4 b. Tangential direction thrust F X Is that
（3）
Thrust force F in radial direction Z Is that
（4）
Wherein F is 0 For the initial electric thrust, F 'is the electric thrust after attitude bias, F' and F 0 The sizes are the same. Alpha is the roll axis attitude offset angle, and beta is the pitch axis attitude offset angle.
(2) Attitude offset angle magnitude
According to formula (1), when the rolling attitude angle is offset by 90 degrees, the thrust only has components in the normal direction, so that the highest efficiency is realized, but the satellite continuous ignition time is longer, and the attitude offset cannot be continuously maintained. Therefore, the attitude control capability of the satellite platform needs to be considered when the attitude offset angle is selected.
3. Calculating the ignition time length of the electric thruster according to the ignition time, the track inclination angle control quantity and the attitude offset angle
For Gaussian perturbation equation (orbit dip control amount is）
Wherein r is the ground center distance, u is the phase, n is the average angular velocity of the track, a is the nominal semi-major axis offset, e is the eccentricity,for speed increment, phase->Speed increment per second>The method comprises the following steps:
wherein u is 0 For the initial phase position,for the total speed increment, T is the single-batch continuous ignition duration, and T is the total time, and the following steps are obtained:
integrating along the ignition arc segment can obtain the track inclination angle size i:
wherein the total speed is increasedThe inclination angle change amount can be regarded as the inclination angle change amount when the phase u is 0 during pulse ignition; u (u) 1 Is the phase after the small thrust action time t. Public factor->Representing small thrustThe arc section work efficiency under the effect.
The change efficiency of the inclination angle is different in different time lengths before and after the control moment, the calculation is performed according to a control formula of the normal thrust to the track inclination angle, and the calculation result is shown in the following table 2:
TABLE 2 efficiency of varying track pitch for different firing durations
And calculating the change efficiency of different ignition time lengths before and after the ignition time to the inclination angle adjustment, comprehensively considering the change inclination angle, the efficiency and the satellite electric thrust index, and selecting the optimal ignition time length of each batch. For example, a comparatively medium total control period and an amount of track inclination change may be selected.
4. Taking track dip perturbation into consideration, determining a control batch according to the track dip control quantity and the ignition duration
Because the electric thrust value is small, the large tilt angle control is often completed in a plurality of days, and therefore, the tilt angle perturbation is required to be added to the consideration of the target tilt angle. The orbit inclination angle is affected by the perturbation of the earth non-spherical shape, the sun-moon attraction, the solar pressure and the like, the annual perturbation amplitude of the orbit inclination angle is about 0.75-0.95 degrees, and the perturbation direction is along the 90 degrees of the right ascension. According to the long term perturbation equation of the geosynchronous orbit dip angle, the daily perturbation speed of the orbit dip angle is as follows:
（5）
in the method, in the process of the invention,the yellow meridian at the intersection point of the white-channel rise is known from an astronomical calendar:
（6）
wherein T is relative julian day, relative 1 month and 1 day of 1950.
Corresponding to 2023, 3 and 15, the relative julian day is 26737, and the white-channel rising intersection point is yellow channelThe tilt perturbation speed was 0.0025 DEG/day.
Taking the shot quantity into consideration, determining a control batch according to the track inclination angle control quantity and the single ignition duration.
5. Calculating perturbation term and radial component force of electric thruster to change amount of plane longitude to obtain nominal semi-long axis offset
According to the attitude rotation angle, the component force of each direction of the thrusters is calculated, and as can be seen from table 1, when the two thrusters at opposite angles are used for generating normal thrust, the component force is always generated in the radial direction, and normal control is performed 2 times a day, and the component force in the radial direction is the ground pointing direction, so that the plane longitude of the satellite can be changed.
Therefore, the amount of change in flat longitude has mainly two parts: first is J at the equator 22 Term perturbation will change the semi-long axis, resulting in daily changes in longitude; secondly, the radial component will change the flat longitude of the satellite. Both parts bring about a change in the flat longitude by biasing the semi-long axis to offset the change in the flat longitude. The calculation method of each part is as follows:
（1）J 22 item perturbation vs. latitude change
The drift velocity and drift acceleration according to the flat longitude are related as follows:
（7）
integrate the above-mentioned method
（8）
At the position ofIn the plane, the solution of the above equation is a parabolic curve with the opening right or left, when the flat longitude drifts the acceleration +.>When the opening is right; drift acceleration +.>When the opening is left; lambda (lambda) 0 For an initial flat longitude, t is the single batch continuous firing duration.
Flat longitude drift accelerationThe method meets the following conditions:
wherein,=-14.545°。
calculatingThereby deriving for counteracting J 22 Semi-major axis offset a due to term perturbation 1 。
(2) Radial component to plane longitude change
The inclination angle adjustment is carried out 2 times every day, the radial force in the same direction counteracts the change of the eccentricity, the change of the flat longitude is a superposition effect, and the calculation formula of the change of the flat longitude is as follows:
（9）
wherein m is satellite quality, deltaD is the change amount of the flat longitude drift rate, F 0 For the initial electric thrust, alpha is the attitude offset angle of the rolling axis, beta is the attitude offset angle of the pitching axis, a 2 To offset the semi-long axis offset brought about by the radial force component.
And summing the two offset amounts to obtain the total offset amount of the semi-long axis.
6. Implementing control based on the obtained nominal semi-major axis offset and control batch
The control comprises two parts: firstly, realizing nominal semi-long axis bias by batch pushing; secondly, the dip angle control is realized by implementing the multi-batch electric pushing control.
7. Evaluating whether the control result meets the expected value
Analyzing the effect of the control strategy, and changing the track inclination angle control quantity, the nominal semi-major axis offset and the eccentricity as follows:
(1) Track inclination angle control amount
The track inclination angle control quantity is mainly determined by normal speed increment, and theoretically, the pitching biasDegree, scroll bias +.>The normal velocity increment brought by the degree is:
（10）
wherein F is the resultant force of the electric thruster,to control the duration. For example, the normal velocity increment is 1m/s, changing the track pitch angle by 0.0186. Normal velocity increment->The inclination increment brought is:
（11）
(2) Track semi-long shaft
The orbit semi-long axis is mainly determined by the tangential velocity increment, and the pitch bias is theoretically realizedThe tangential velocity increment brought by the degree is:
（12）
wherein F is the resultant force of the electric thruster,to control the duration. For example, the tangential velocity is increased by 1m/s, changing the orbit semi-major axis by 27.43 km.
The change amount of the semi-long axis of the track is as follows:
（13）
(3) Eccentricity of orbit
Since the ignition trails of the thrusters NW, NE and the thrusters SW, SE differ by 180 °, the control actions of the two pairs of thrusters on the eccentricity are diametrically opposite, and if the sum of the control amounts of the thrusters NW, NE in the radial direction is exactly equal to the thrusters SW, SE, the radial control forces of the thrusters NW, NE on the eccentricity are exactly offset in the control period of one day. Meanwhile, in order to eliminate the influence of tangential thrust on the eccentricity, the control amounts of the thrusters NW and SW are required to be kept equal, and the control amounts of the thrusters NE and SE are equal, so that the control on the dip angle direction is ensured not to influence the eccentricity change of the track.
The implementation of the above method will be described below using specific examples.
Taking the BSS-702 series satellites of Boeing company as an example, the in-country orbital GEO electric propulsion lays out multi-reference satellites BSS-702 series satellites. Satellite fixed point 87.5 deg. E, retaining ring + -0.1 deg.. The initial orbit inclination angle is 0.96 degrees, the ascending intersection point red-warp is 93.5 degrees, the target orbit inclination angle is 0.61 degrees, and the ascending intersection point red-warp is 93.5 degrees. The initial trajectory is shown in table 3 below:
TABLE 3 initial satellite orbit
1. Calculating ignition time and control quantity
According to the initial orbit and the target orbit of the satellite, the ignition time and the control quantity are calculated, and the components of the initial inclination angle on the polar coordinates are as follows:
the components of the target dip angle on the polar coordinates are:
track inclination angle control amountThe method comprises the following steps:
the control direction of the track inclination angle control quantity vector is as follows:
=93.5°。
the moment of the ignition middle point is t of each day 1 Time sum t 2 When (1):
wherein,for ascending intersection point, right-left>Is the amplitude angle of the near place +.>For the straight-up point angle +.>Is the track epoch time, n isAverage angular velocity of the track.
2. Thruster combination selection and attitude offset
The tilt control can be implemented according to the above-described combination of ne+sw or se+nw, and the present embodiment selects the ne+sw combination. When the NW and SW thrusters respectively ignite, the tangential direction forces cancel each other, the normal direction forces cancel each other, the radial direction forces are superimposed in the +z direction, and the original force is oriented in the +z direction.
Considering the attitude control capability of a satellite platform, selecting an attitude offset angle + -40 DEG and t 1 Posture offset 40 DEG, t at moment of ignition 2 The attitude at the time of ignition is offset by-40 °.
3. Calculating the ignition duration
The change amount of the inclination angle and the change efficiency are comprehensively considered, and the ignition duration is selected to be 2.5 hours before and after the ignition moment.
Because the long arc section ignites, the inclination change efficiency has a loss, takes the ignition middle moment as the midpoint, calculates the inclination change efficiency of ignition for 5 hours as follows:
the speed increment is:
the control quantity of the track dip angle is as follows:
=0.0109°
rolling bias is 40 degrees on the basis, and the track inclination angle control quantity is as follows:
/>
4. determining control batches
According to the total control amount of the dip angle of 0.35 degrees, the control amount of the dip angle of each batch is 0.00701 degrees, the dip angle of 2 batches of controllable dip angles is 0.01402 degrees every day, the dip angle is increased by 0.0025 degrees every day in consideration of the perturbation effect, and the total control days are as follows:
two batches were controlled a day, and a total of 30.5 days 61 batches were found to be required based on the calculated results for 30.3 days.
5. Calculating nominal semi-major axis offset
（1）J 22 Offset of term perturbation
It is known that:=87.4°，/>when =87.6°, d=0, drift acceleration +.>Is that
= -0.000686 °/day 2
From the formulaObtain->Obtaining D 0 Nominal semi-long axis offset = 0.01656 °/day: />。
(2) Offset of radial velocity change plane longitude
The satellite continuously ignites for 5 hours in a batch, and the radial speed increment is as follows:
the satellite fires 2 batches per day for 10 hours continuously with a radial velocity increment of 0.966 m/s and a flat longitude drift velocity of 0.036 °/day.
The semi-major axis offset is:
in sum, the satellite fixed point is at 87.5 DEG E, the east-west holding range is + -0.1 DEG, and the acceleration is shifted according to the longitudeCan be calculated to counteract J 22 The semi-major axis offset of the term perturbation is-1.3 km; the satellite was continuously fired for 5 hours with a radially generated velocity increment of 0.483 m/s, firing 2 batches per day, with a flat longitude drift velocity of 0.036 °/day, and a semi-major axis offset of 2.82 km to offset the radial force component. The total offset of the semi-major axis is thus 1.52 km.
6. Control implementation
The control comprises two parts:
first, 1 batch push realizes nominal semi-long axis bias. Selecting the remote point moment as the ignition time:
55 min 48 sec =8
The tangential velocity increment is:
and secondly, implementing electric pushing control to realize inclination angle control. The tilt control was performed starting from GEO satellite at 3 month 15 of 2023, setting the orbital tilt control amount to 0.35 °, firing two pairs of thrusters nw+se diagonally, performing two times per day at 55 minutes 48 seconds and 20 minutes 48 seconds of satellite 8, firing a total of 61 batches with a control strategy of roll axis offset of 40 degrees for each batch for 5 hours in view of control efficiency and tolerable loss of accuracy. Due to the difference in control efficiency of the thrusters in the tangential direction, the half-major axis cannot be completely counteracted, resulting in a reduction of about 500 meters per day. In order to avoid the situation of pushing things during the electric push dip angle control, the original control strategy is optimized, and the pitching bias is added on the basis of the rolling bias by 40 degrees so as to offset the effect of reducing the semi-long axis caused by tangential thrust difference.
7. Evaluating whether the control result meets the expected value
Fig. 5 shows that the satellite orbit inclination is reduced from 0.96 ° (TOD coordinate system) to 0.61 ° by 61 batch control, and the total is reduced by 0.35 °, thereby achieving the control objective.
It can be seen from fig. 6 that the first 16 control strategies are 40 degrees offset from the roll axis, and that the flat longitude shifts eastward due to the difference in control efficiency of the two thrusters in the tangential direction, and the radial force component. The latter 45 control strategies consider the longitude holding range, increase the pitch attitude bias on the basis of the roll axis bias, and Ping Jingdu drift west first and east after reaching the darcy.
FIGS. 7 and 8 show the eccentricity change during tilt angle control, wherein each batch of electric propulsion adjustment has a speed increment in the radial direction due to the installation angle of the electric thruster, and the single 5-hour control speed increment is about 0.52 m/s, which affects the eccentricityThe control effect is counteracted by 2 times of control effect, and the influence on the eccentricity is small.
It is to be understood that the terms "center," "longitudinal," "transverse," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," and the like in the above description are directional or positional relationships as indicated based on the drawings, merely to facilitate description of embodiments of the invention and to simplify the description, and do not indicate or imply that the devices or elements referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus should not be construed as limiting embodiments of the invention.
Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include one or more such feature. In the description of the embodiments of the present invention, the meaning of "plurality" is two or more, unless explicitly defined otherwise.
In the embodiments of the present invention, unless explicitly specified and limited otherwise, the terms "mounted," "connected," "secured" and the like are to be construed broadly and include, for example, either permanently connected, removably connected, or integrally formed; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communicated with the inside of two elements or the interaction relationship of the two elements. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
In embodiments of the invention, unless expressly stated or limited otherwise, a first feature "above" or "below" a second feature may include both the first and second features being in direct contact, or may include both the first and second features not being in direct contact but being in contact with each other through additional features therebetween. Moreover, a first feature being "above," "over" and "on" a second feature includes the first feature being directly above and obliquely above the second feature, or simply indicating that the first feature is higher in level than the second feature. The first feature being "under", "below" and "beneath" the second feature includes the first feature being directly under and obliquely below the second feature, or simply means that the first feature is less level than the second feature.
In the description of the present specification, a description referring to terms "one embodiment," "some embodiments," "examples," "specific examples," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present invention. In this specification, schematic representations of the above terms are not necessarily directed to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, one skilled in the art can combine and combine the different embodiments or examples described in this specification.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.
Claims (5)
1. The geosynchronous satellite electric propulsion dip angle control method based on multidimensional attitude bias is characterized by comprising the following steps of:
calculating ignition time and orbital inclination control quantity according to the initial orbit and the target orbit of the satellite;
determining the ignition position of the electric thruster, and determining the attitude offset angle of each ignition moment according to the calculated track inclination angle control quantity;
calculating the ignition time length of the electric thruster according to the ignition time, the track inclination angle control quantity and the attitude offset angle;
considering track dip perturbation, determining a control batch according to the track dip control quantity and the ignition duration;
calculating perturbation items and the radial component force of the electric thruster to the horizontal longitude change amount to obtain a nominal semi-long axis offset;
performing control according to the obtained nominal semi-major axis offset and the control batch;
the calculation process of the step of calculating the ignition time and the orbit inclination angle control amount according to the initial orbit and the target orbit of the satellite is as follows:
component i of initial orbital tilt in polar x-axis 0x And a component i on the y-axis 0y The method comprises the following steps of:
；
component i of the target orbit inclination in the polar x-axis 1x And a component i on the y-axis 1y The method comprises the following steps of:
；
the control amount of the track dip angle isThe units are degrees:
；
the control direction of the vector of the track inclination angle control quantity is as follows:
；
the moment of the ignition middle point is t of each day 1 Time sum t 2 When (1):
；
wherein i is 0 For initial track pitch, i 1 For a target track pitch angle, a is the nominal semi-major axis offset,for ascending intersection point, right-left>Is the amplitude angle of the near place +.>For the straight-up point angle +.>The track epoch time, n is the track angular velocity;
the nominal semi-major axis offset comprises: j (J) 22 Term perturbation versus latitude change amount and radial component versus latitude change amount;
J 22 the term perturbation vs. flat longitude change amount calculation process is as follows:
according to drift velocity D of flat longitude and drift initial velocity D 0 Drift acceleration of plane longitudeThe relation is:
；
integrating the above method to obtain:
；
at the position ofIn the plane, the solution of the equation is parabolic with the opening right or left, when the plane longitude drifts the accelerationWhen the opening is right; drift acceleration +.>When the opening is left, lambda 0 For the initial flat longitude, t is the single-batch continuous ignition duration, and lambda is the drift flat longitude;
flat longitude drift accelerationThe method meets the following conditions:
；
wherein,=-14.545°；
calculatingThereby deriving for counteracting J 22 Semi-major axis offset a due to term perturbation 1 ；
The radial component force versus plane longitude change amount is calculated as follows:
；
wherein m is satellite quality, deltaD is the change amount of the flat longitude drift rate, F 0 For the initial electric thrust, alpha is the attitude offset angle of the rolling axis, beta is the attitude offset angle of the pitching axis, a 2 To offset the semi-long axis offset brought about by the radial force component.
2. The control method of claim 1, wherein the step of performing control based on the derived nominal semi-major axis offset and the control batch comprises:
performing chemical propulsion control based on the nominal semi-major axis offset;
and performing electric propulsion control according to the track inclination angle control amount and the control batch.
3. The control method according to claim 1, characterized in that the control method further comprises the steps of:
and evaluating the control result by using three parameters of the track inclination angle control quantity, the nominal semi-long axis offset quantity and the eccentricity ratio.
4. A control method according to claim 1, characterized in that ignition is performed with an electric thruster of ne+sw or se+nw combination.
5. The control method according to claim 4, wherein after ignition by an electric thruster using a combination of ne+sw or se+nw, the electric thruster generates a normal thrust to the satellite by adjusting the X-axis attitude of the rolling axis; by adjusting the Y-axis attitude of the pitching axis, the electric thruster generates tangential thrust to the satellite.
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Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104015938A (en) * | 2014-05-07 | 2014-09-03 | 北京理工大学 | Position keeping method of electric propulsion stationary orbit satellite |
| FR3047813A1 (en) * | 2016-02-16 | 2017-08-18 | Airbus Defence & Space Sas | METHOD FOR CONTROLLING SATELLITE ATTITUDE GUIDANCE, SATELLITE, SATELLITE PLURALITIES AND ASSOCIATED COMPUTER PROGRAM |
| CN108490963A (en) * | 2018-02-08 | 2018-09-04 | 中国空间技术研究院 | Position keeping method under full electric propulsion satellite electric thruster fault mode and system |
| CN114394260A (en) * | 2021-11-29 | 2022-04-26 | 中国西安卫星测控中心 | Small-thrust south-north keeping control analysis calculation method for geostationary satellite |
| CN115510378A (en) * | 2022-09-22 | 2022-12-23 | 中国西安卫星测控中心 | Method for calculating nominal semi-major axis perturbation of geostationary orbit |
| CN115626303A (en) * | 2022-09-27 | 2023-01-20 | 上海航天控制技术研究所 | Method for maintaining inclination angle of geosynchronous orbit by adopting electric propulsion |
| CN115993777A (en) * | 2022-11-24 | 2023-04-21 | 中国西安卫星测控中心 | Track perturbation model inversion-based diameter-cut joint control decoupling iteration calibration method |
-
2023
- 2023-07-12 CN CN202310847379.9A patent/CN116946392B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104015938A (en) * | 2014-05-07 | 2014-09-03 | 北京理工大学 | Position keeping method of electric propulsion stationary orbit satellite |
| FR3047813A1 (en) * | 2016-02-16 | 2017-08-18 | Airbus Defence & Space Sas | METHOD FOR CONTROLLING SATELLITE ATTITUDE GUIDANCE, SATELLITE, SATELLITE PLURALITIES AND ASSOCIATED COMPUTER PROGRAM |
| CN108490963A (en) * | 2018-02-08 | 2018-09-04 | 中国空间技术研究院 | Position keeping method under full electric propulsion satellite electric thruster fault mode and system |
| CN114394260A (en) * | 2021-11-29 | 2022-04-26 | 中国西安卫星测控中心 | Small-thrust south-north keeping control analysis calculation method for geostationary satellite |
| CN115510378A (en) * | 2022-09-22 | 2022-12-23 | 中国西安卫星测控中心 | Method for calculating nominal semi-major axis perturbation of geostationary orbit |
| CN115626303A (en) * | 2022-09-27 | 2023-01-20 | 上海航天控制技术研究所 | Method for maintaining inclination angle of geosynchronous orbit by adopting electric propulsion |
| CN115993777A (en) * | 2022-11-24 | 2023-04-21 | 中国西安卫星测控中心 | Track perturbation model inversion-based diameter-cut joint control decoupling iteration calibration method |
Non-Patent Citations (1)
| Title |
|---|
| 赵义平 ; 李峰 ; 经姚翔 ; 李大伟 ; .GEO电推进卫星轨道漂移策略研究.航天器工程.2016,(第04期),全文. * |
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