CN111268177B - Distributed closed-loop autonomous position maintaining control method for geostationary orbit satellite - Google Patents
Distributed closed-loop autonomous position maintaining control method for geostationary orbit satellite Download PDFInfo
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Abstract
The invention designs a distributed closed-loop autonomous position keeping control method for a stationary orbit satellite, which disperses the traditional centralized position keeping control quantity to be executed every day, eliminates the perturbation variation quantity every day by utilizing pulse thrust, further combines the feedback control action of a deviation closed loop, and can always maintain the position of the satellite in an accurate range through the combined control of the mean longitude, the eccentricity and the orbit inclination angle, wherein the east-west direction control precision is better than +/-0.05 degrees, and the south-north direction control precision is better than +/-0.01 degrees.
Description
Technical Field
The invention belongs to the field of autonomous orbit control of spacecrafts, and relates to a distributed closed-loop autonomous position maintaining control method for a stationary orbit satellite, which is used for realizing autonomous high-precision orbit control on a satellite.
Background
With the continuous expansion of the application field of satellites in China, the number of satellites in on-orbit operation is increased sharply, and the conventional mode of daily management of the satellites by depending on a ground measurement and control center cannot adapt to the development requirement increasingly, so that intelligent and autonomous operation of the satellites must be realized. The autonomous operation means that the satellite completely utilizes software and hardware equipment of the satellite to autonomously complete various functions and operations required by a flight task without depending on the support of a ground station, wherein the autonomous control of the attitude and the orbit is the key point for realizing the autonomous operation of the satellite. Currently, satellite attitude control is automatically completed on the satellite, but most satellites still depend on a ground station for orbit measurement and control, and only few satellites have the function of on-satellite autonomous orbit control.
The existing satellite autonomous orbit control method refers to a ground centralized type position keeping management mode, when the position of a satellite reaches the boundary of a keeping ring, position keeping control quantity is calculated autonomously on the satellite, proper control time is selected, position keeping operation is implemented autonomously, wherein the longitude flatness control adopts a single-side limit ring mode, the eccentricity control adopts a double-pulse strategy, and an orbit inclination angle control target is zero. According to the perturbation acceleration of the satellite fixed point position and the size of the position retaining ring, the east-west position retaining control period is 1 to 3 weeks generally, and the north-south position retaining control period is about 1 month. In consideration of navigation positioning and control errors, the accuracy index of centralized autonomous position protection control is generally +/-0.2 degrees, and the accuracy is poorer than the management capability of ground +/-0.05 degrees, so that a satellite autonomous high-accuracy orbit control method needs to be further researched.
Disclosure of Invention
The technical problem solved by the invention is as follows: in order to realize the function of autonomous high-precision orbit control on the satellite, a distributed closed-loop position holding control method for a stationary orbit satellite is provided, and the satellite position can be always maintained in an accurate range through the joint control of the mean longitude, the eccentricity and the orbit inclination.
The technical solution of the invention is as follows: a distributed closed-loop autonomous position maintaining control method for a stationary orbit satellite is provided, which comprises three steps of controlling the mean longitude, the eccentricity and the orbit inclination angle, and comprises the following steps:
(1) flat longitude control strategy: firstly, adjusting a semi-major axis of a satellite to a nominal stationary orbit height, secondly, adopting a distributed closed-loop independent east-west position keeping control method, eliminating a perturbation variation of a longitude and latitude degree and a closed-loop control quantity of a longitude deviation fixed point position every day through pulse thrust, selecting the control moment near a near place or a far place, and adopting a combined control mode of longitude and eccentricity.
(2) Eccentricity control strategy: when the variation of eccentricity perturbation exceeds the variation caused by mean longitude control every day, impulse thrust is needed to eliminate the excess part, and meanwhile, eccentricity closed-loop control quantity is introduced; the method adopts a double-pulse mode, is carried out in two batches every day, the control quantity is half of the total speed increment each time, the first time is executed by combining the average longitude control, and the second time is separated by 180 degrees from the first time and has the opposite direction.
(3) Track inclination angle control strategy: a distributed closed-loop autonomous north-south position maintaining control method is adopted, and long-term and long-period perturbation variation of the track inclination angle and closed-loop control quantity of the inclination angle deviating from a target position are eliminated through pulse thrust every day; if the southbound protection is selected, the control time every day is near the horizontal right ascension corresponding to the inclination perturbation direction, and if the northbound protection is selected, the control time every day is near the horizontal right ascension corresponding to the inclination perturbation opposite direction.
The size of the semi-major axis of the satellite corresponding to the nominal stationary orbit height in the step (1) is 42165.795 km.
The formula for calculating the perturbation variation of the longitude flatness in the step (1) is as follows:
in the formula
Acceleration is a perturbation of the longitude degree, and is related to the fixed point position of the satellite;
indicating that the change amount within one day is calculated.
The calculation formula of the closed-loop control quantity of the longitude leveling degree in the step (1) is as follows:
where λ is the current location geographic longitude, λsFor the target fixed-point position geographic longitude, KpλThe longitude control method is a longitude closed-loop feedback control coefficient and has the function of automatically adjusting the drift trend of the longitude and maintaining the longitude within the range required by the control precision index for a long time, and the value is usually 0.01-0.02.
The speed increment calculation formula required by the longitude adjustment control in the step (1) is as follows:
in the formula of omegae＝360.9856(°/day)＝7.2921×10-5(rad/s) is the rotational angular velocity of the earth, Vs3074.7(m/s) is the nominal stationary track speed.
The length of the control arc section required by the longitude and latitude control in the step (1) is as follows:
in the formula Msat(kg) is the current mass of the satellite, FEW(N) effective thrust in the east-west direction, Wpulse(s) is the control of the pulse width length, Ppulse(s) is a pulse control period.
In the step (1), a combined control mode of longitude and eccentricity is adopted, and the longitude and eccentricity control time is selected as follows:
(a) when Δ VT<When 0, performing western-direction protection near the near point, wherein the control starting time corresponds to the true near point angle:
(b) when Δ VT>When 0, the east position is protected near the far place, and the corresponding true near point angle at the control starting time is as follows:
the calculation formula of the variation of the beat rate variation in the step (2) is as follows:
in the formula CrIs reflection of light pressureCoefficient, depending on the satellite surface material, with a value of 1.0 for total absorption, usually 1.5; ssat(m2) The cross section of the satellite perpendicular to the sunlight; p0＝4.65×10-6(N/m2) Is the solar radiation pressure per unit area.
The formula for calculating the variable quantity of the eccentricity caused by the flatness control in the step (2) is as follows:
the judgment condition for applying the eccentricity control in the step (2) is as follows:
Δep>Δeλ
the calculation formula of the eccentric rate perturbation control quantity in the step (2) is as follows:
Δed＝Δep-Δeλ(Δed>0)
the heart rate deviation closed-loop control quantity calculation formula in the step (2) is as follows:
Δec＝Kpe(e-ef)
wherein the current eccentricity is e; e.g. of the typefThe target eccentricity is usually 0; kpeThe control method is an eccentricity closed-loop feedback control coefficient, has the function of automatically adjusting the increasing trend of the eccentricity, and maintains the eccentricity within a circle keeping range for a long time, and the value is usually 0.01.
The calculation formula of the total speed increment required by the heart rate deviation control in the step (2) is as follows:
in the step (2), the eccentricity control is executed in combination with the meridian flatness control in the first batch of every day, and the eccentricity control quantity is superposed on the meridian flatness control quantity
And recording the first control end time as t0The starting time of the second batch is selected to be t0Is executed after +0.5 days, and the control amount is
And in the opposite direction.
The calculation formula of the track inclination angle long-term perturbation variation in the step (3) is as follows:
in the formula of omegamThe ascending crossing point of the white tract, right ascension, imIs the angle between the white road surface and the equatorial plane.
The calculation formula of the long-period perturbation variation of the track inclination angle in the step (3) is as follows:
in the formula ns0.0172(rad/day) 0.9856 (degree/day) mean angular velocity of solar apparent motion, ΩsThe ascending crossing point of the ecliptic tract, right ascension, isIs the angle between the equatorial plane and the equatorial plane.
The track inclination perturbation control quantity calculation formula in the step (3) is as follows:
the calculation formula of the closed-loop control quantity of the track inclination angle in the step (3) is as follows:
in the formula ixAnd iyIs a current tilt angle vector and has ix＝sin(i)cos(Ω)，iySin (i) sin (Ω), where i is the satellite orbital inclination and Ω is the ascension at the intersection; i.e. ixfAnd iyfIs a target tilt vector, usually taking the value 0;Kpithe method is an inclination angle closed loop feedback control coefficient and has the functions of automatically adjusting the track inclination angle drift trend and maintaining the inclination angle within the control precision index requirement range for a long time, and the value is usually 0.05.
The velocity increment calculation formula required by the track inclination angle control in the step (3) is as follows:
in the formula,. DELTA.ix＝Δixd+Δixc，Δiy＝Δiyd+Δiyc。
The calculation formula of the length of the control arc segment required by the track inclination angle control in the step (3) is as follows:
in the formula FNSAnd (N) is effective thrust in the north-south direction.
In the step (3), the track inclination angle control time is selected as follows:
(a) if the southward position protection is selected, the inclination angle starting and controlling time of the orbit in each day corresponds to the right ascension:
(b) if the north orientation protection is selected, the inclination angle control starting time of the orbit in each day corresponds to the horizontal right ascension:
compared with the prior art, the invention has the advantages that:
1. the distributed closed-loop autonomous position maintaining control method provided by the invention disperses the traditional centralized position maintaining control quantity to be executed every day, eliminates the perturbation variation quantity every day by utilizing pulse thrust, further combines the action of deviation closed-loop feedback control, can always maintain the satellite orbit position within an accurate range, has the east-west direction control accuracy better than +/-0.05 degrees and the south-north direction control accuracy better than +/-0.01 degrees, provides a high-precision on-satellite autonomous orbit control method, and further improves the autonomous operation capability of the satellite.
2. Compared with the traditional centralized strategy, the distributed bit-keeping control can be completely executed in a normal mode, does not relate to attitude control mode conversion, and is favorable for maintaining the normal working state of the satellite without interruption. Simulation analysis shows that the increment of the track control speed required by the scattered bit security year is about 52m/s, which is equivalent to the level of ground management fuel consumption. In addition, the distributed position protection still adopts a chemical thruster as an actuating mechanism, the precise orbit control function can be realized without additionally configuring electric propulsion and other equipment, the adaptability to the current satellite platform is good, and the popularization and the application are facilitated.
Drawings
FIG. 1 is a schematic diagram of a position maintenance boundary for a geostationary orbit satellite;
FIG. 2 is a simulation curve of the satellite semi-major axis control effect;
FIG. 3 is a simulation curve of the eccentricity control effect of a satellite;
FIG. 4 is a simulation curve of the effect of satellite geographic longitude control;
fig. 5 is a simulation curve of the satellite orbit inclination angle control effect.
Detailed Description
The geostationary orbit satellite has position maintenance control boundary indicators in both the longitudinal and latitudinal directions, where the longitudinal boundary is defined as the maximum allowed longitudinal deviation from the fixed point position and the latitudinal boundary is defined as the maximum allowed orbital inclination, as shown in fig. 1. Therefore, the orbit preservation control of the geostationary orbit satellite is a general term for a series of orbit correction control by a propulsion system during the whole service life of the satellite, and the function of the orbit preservation control is to maintain the free perturbation motion of the satellite in a preservation boundary as much as possible and limit the motion of the satellite to be out of the boundary as much as possible.
The method comprises the following steps that the position holding control of the geostationary orbit satellite is generally divided into an east-west position holding mode and a south-north position holding mode, wherein the south-north position holding control aims to reduce the inclination angle of the satellite by utilizing normal thrust, so that the geographic latitude of the satellite is positioned in a holding boundary; the east-west position keeping comprises a longitude flatness control part and an eccentricity control part, wherein the longitude flatness control aims to maintain the longitude flatness of the satellite near the fixed point longitude by using tangential thrust, so that the longitude flatness of the satellite is positioned in a keeping boundary for removing the longitude daily oscillation; the purpose of eccentricity control is to reduce the eccentricity of the satellite by means of double-pulse tangential thrust, reducing the amplitude of the daily periodic oscillation of the satellite's geographical longitude.
The invention designs a distributed closed-loop autonomous position maintaining control method for a stationary orbit satellite, which disperses the traditional centralized position maintaining control quantity to be executed every day, eliminates the perturbation variation quantity every day by utilizing pulse thrust, further combines the feedback control action of a deviation closed loop and realizes the control function of a high-precision autonomous orbit on the satellite by the joint control of the horizontal longitude, the eccentricity and the orbit inclination angle. The specific method comprises the following steps:
(1) a flat longitude control policy.
Firstly, adjusting the semi-major axis of the satellite to the height of a nominal stationary orbit to offset the long-term drift of the flatness caused by the non-spherical gravity of the earth and the perturbation of the sun and moon gravity; secondly, a distributed closed-loop independent object position keeping control method is adopted, and longitude perturbation variation caused by earth non-spherical gravity perturbation and longitude deviation fixed point position closed-loop control quantity are eliminated through pulse thrust every day; and then, carrying out east-west position keeping control once every day near a near place or a far place, and adopting a combined control mode of flatness longitude and eccentricity, namely always selecting a position with reduced eccentricity to implement when controlling the flatness longitude drift rate, so that the method is favorable for reducing the longitude daily period oscillation caused by the eccentricity. The method comprises the following specific steps:
(a) the satellite is adjusted to a nominal geostationary orbit altitude. Before carrying out distributed autonomous position keeping control, the size of the semi-major axis of the satellite is firstly adjusted to the semi-major axis of the nominal stationary orbit from the ground, and the average value of the semi-major axis of the nominal stationary orbit considering the influence of the earth aspheric gravity and the daytime gravity perturbation is 42165.795 km. After the adjustment is in place, the east-west position is kept without considering the influence of the longitude drift rate, and only the longitude drift acceleration caused by the non-spherical gravity perturbation of the earth needs to be eliminated.
(b) And calculating a horizontal longitude perturbation control quantity. The longitude drift of an geostationary orbit satellite in the east-west direction consists of two parts: one part is the change of the longitude drift rate caused by the perturbation of the earth's non-spherical gravity, the other part is the eccentricity perturbation generated by the sunlight pressure to cause the satellite longitude daily period oscillation, and the two effects are independent to each other and cause the satellite east-west longitude deviation. According to perturbation analysis, the non-spherical gravity perturbation field harmonic term of the earth causes the flatness drift acceleration, and the analysis shows that the second-order field harmonic term is the most dominant and accounts for about 85 percent. Usually, the table lookup is used for obtaining the horizontal longitude perturbation acceleration of a fixed point position, and the acceleration is injected to a satellite as a constant without autonomous calculation. The daily flatness drift rate variation caused by the perturbation of the field harmonic is:
in the formula
Acceleration is a perturbation of the longitude degree, and is related to the fixed point position of the satellite; delta Tday1(day) represents the amount of change calculated over the day.
(c) And calculating a horizontal longitude closed-loop control quantity. Due to the influence of perturbation analysis errors and control errors, the satellite longitude after control still has a slow long-term drift trend, and if the satellite longitude is not processed, the control precision index requirement range can be drifted out finally. Therefore, by using the formation flying relative position control method for reference, the satellite performs closed-loop position feedback control relative to the fixed point position, geographic longitudes are adopted as control targets in the east-west direction, and if the current position longitude is and the target fixed point position longitude is, the longitude closed-loop control quantity is as follows:
where λ is the current location geographic longitude, λsPointing to a targetLocation geographic longitude, KpλThe longitude control method is a longitude closed-loop feedback control coefficient and has the function of automatically adjusting the drift trend of the longitude and maintaining the longitude within the range required by the control precision index for a long time, and the value is usually 0.01-0.02. Under closed-loop relative position control adjustment, the daily control speed increment is slightly increased or decreased, and the total speed increment is not greatly changed, so that the fuel consumption is hardly influenced.
(d) And selecting the mean longitude control time. To sum up, the total control quantity of the longitude drift rate includes a perturbation control quantity and a closed-loop control quantity, and a velocity increment calculation formula required by the longitude-leveling control is as follows:
in the formula of omegae＝360.9856(°/day)＝7.2921×10-5(rad/s) is the rotational angular velocity of the earth, Vs3074.7(m/s) is the nominal stationary track speed. The length of the control arc segment required by the longitude flatness control is as follows:
in the formula Msat(kg) is the current mass of the satellite, FEW(N) effective thrust in the east-west direction, Wpulse(s) is the control of the pulse width length, Ppulse(s) is a pulse control period. The distributed independent east-west position keeping strategy is used for keeping and controlling east-west positions near a near place or a far place every day, a combined control mode of longitude and eccentricity is adopted, namely, the position with reduced eccentricity is always selected to be implemented when the drift rate of longitude is controlled, and therefore longitude day-cycle oscillation of eccentricity gravitation is favorably reduced. If the speed increment is controlled by Δ VT<And 0, performing western-position protection near the near point, wherein the satellite true near point angle corresponding to the control time is as follows:
if the speed increment is controlled by Δ VT>0, performing east keeping near the far place, and controlling the true near point angle of the satellite corresponding to the moment as follows:
(2) and (4) eccentricity control strategy.
In general, the eccentricity is always reduced in the flatness control, so that the value of the eccentricity is always kept in an allowable small range, and the eccentricity control is not separately performed. However, when the area value of the solar pressure of the satellite is relatively large, or the flatness drift rate control amount required when the satellite is fixed at a point near the equilibrium point is relatively small, active eccentricity control is required. The purpose of the eccentricity control of the geostationary orbit satellite is to reduce the eccentricity of the satellite by using double-pulse tangential thrust, reduce the day-cycle oscillation of longitude and ensure the fixed point precision requirement in the east-west direction. And the distributed eccentricity control strategy is used for executing the centralized control quantity in a distributed mode every day and eliminating the perturbation variation quantity of the eccentricity and the deviation from the target value control quantity every day through pulse thrust control. The method comprises the following specific steps:
(a) and (5) judging the starting and controlling conditions. The perturbation of the sunlight pressure can cause the eccentricity vector to move along the perturbation circle year-round period, the perturbation change rate can be considered as a constant in a short time, and the calculation formula of the daily eccentricity perturbation change quantity is as follows:
in the formula CrThe total absorption value is 1.0, usually 1.5, which is the light pressure reflection coefficient and is related to the surface material of the satellite; ssat(m2) The cross section of the satellite perpendicular to the sunlight; p0＝4.65×10-6(N/m2) Is the solar radiation pressure per unit area. The formula for calculating the variable quantity of the eccentricity caused by the control of the flatness warp is as follows:
if Δ eλ>ΔepAnd if the average longitude degree control quantity is less than the average longitude degree, the eccentricity perturbation variation can be completely offset by the average longitude degree control quantity every day, and the average longitude degree and the eccentricity degree can be maintained within the circle keeping range all the time by adopting a combined control strategy, so that the control index requirement is met.
If Δ eλ<ΔepIf the average longitude and the eccentricity combined control strategy is adopted, the eccentricity is controlled by the average longitude and the eccentricity, and the eccentricity is controlled by the eccentricity control strategy.
(b) And calculating the eccentricity control quantity. The eccentricity control strategy also adopts a distributed position protection control mode, the part of the perturbation variable quantity which is more than the normal warp control quantity is eliminated by using pulse thrust every day, and the calculation formula of the eccentricity perturbation control quantity is as follows:
Δed＝Δep-Δeλ(Δed>0)
due to perturbation analysis errors and control errors, there is a possibility that the controlled offset heart rate still has a slowly increasing trend, and if not processed, the controlled offset heart rate will finally exceed the eccentricity keeping circle range. Therefore, closed-loop control of eccentricity is introduced, and the current eccentricity of the satellite is assumed to be e, and the target eccentricity is assumed to be efFor distributed autonomous bit-keeper efNormally set to 0, the calculation formula of the closed-loop control amount of eccentricity is as follows:
Δec＝Kpe(e-ef)
in the formula KpeThe control method is an eccentricity closed-loop feedback control coefficient, has the function of automatically adjusting the increasing trend of the eccentricity, and maintains the eccentricity within a circle keeping range for a long time, and the value is usually 0.01. The total control quantity of the eccentricity ratio comprises a perturbation control quantity and a closed-loop control quantity, and the total speed increment calculation formula required by the eccentricity ratio control is as follows:
(c) and selecting eccentricity control time. The eccentricity control adopts a double-pulse mode, is divided into two batches every day, decelerates once at a near place and accelerates once at a far place, and the increment of the speed is controlled to be delta V every timeTeAnd 2, controlling the interval of two times by 180 degrees and the directions are opposite, so that the generated drift rate changes are mutually counteracted, and the eccentricity increments are superposed. In order to reduce the starting times of the track control, the eccentricity control is executed in combination with the longitude control in the first batch every day, and when the near place is protected towards the west position or the far place is protected towards the east position, the eccentricity control speed increment delta V is superposed on the longitude control quantityTe/2. Because the eccentricity is always kept in a small range, the true near point angle change range is large, and if the selection is executed at a near point or a far point opposite to the first time, the interval of two speed increment cannot be guaranteed to be 180 degrees. Therefore, the second eccentricity control is executed 0.5 days after the first control is finished by taking the accumulated time as a starting condition, the two times of control are ensured to be separated by 180 degrees and opposite in direction, and the effect of only changing the eccentricity does not influence the track drift rate.
(3) And (4) a track inclination angle control strategy.
The inclination angle of the orbit of the static orbit satellite influenced by the perturbation of the sun-moon gravitation has long-term drift, the control is kept through the north-south position, and the inclination angle of the orbit is reduced by utilizing the normal thrust of the orbit, so that the satellite is always positioned in a keeping boundary in the north-south latitude direction. The latitudinal direction is usually defined as the equatorial plane with respect to the epoch time and with high accuracy requirements as the true equatorial plane with respect to the epoch time, so the orbital inclination control should be with respect to the instantaneous flat or instantaneous true inertial frame, rather than the J2000 inertial frame. And the distributed autonomous north-south position protection control is implemented by distributing the centralized position protection control quantity to each day and eliminating the inclination perturbation variation quantity and the deviation target position control quantity by pulse thrust control. The method comprises the following specific steps:
(a) and calculating the tilt perturbation control quantity. The perturbation analysis shows that the orbit inclination angle is changed for a long time and a long period due to the sun-moon gravitational perturbation, the drift rate can be regarded as a constant in a short time, and the calculation formula of the change quantity of the orbit inclination angle in the long-term perturbation is as follows:
in the formula of omegamThe ascending crossing point of the white tract, right ascension, imIs the angle between the white road surface and the equatorial plane. The calculation formula of the long-period perturbation variation quantity of the combined track inclination angle is as follows:
in the formula ns0.0172(rad/day) 0.9856 (degree/day) mean angular velocity of solar apparent motion, ΩsThe ascending crossing point of the ecliptic tract, right ascension, isIs the angle between the equatorial plane and the equatorial plane. The calculation formula of the perturbation control quantity of the track inclination angle obtained every day is as follows:
(b) and calculating the closed-loop control quantity of the inclination angle. Due to the influence of perturbation analysis errors and control errors, the inclination angle of the controlled orbit still has a slow long-term drift trend, and if the inclination angle of the controlled orbit is not processed, a retaining ring range can be drifted out finally. Therefore, by using the formation flying relative position control method for reference, the satellite carries out closed-loop position feedback control relative to the fixed point position, the orbit inclination angle is adopted as a control target in the north-south direction, and the current inclination angle vector i is assumed to be (i is equal tox,iy)TTarget dip angle vector if＝(ixf,iyf)TIf the target for the distributed autonomous attitude keeping control is usually selected to be 0, the calculation formula of the closed-loop control amount of the track inclination angle is as follows:
in the formula ix＝sin(i)cos(Ω)，iySin (i) sin (Ω), where i is the satellite orbital inclination and Ω is the ascension at the intersection; kpiThe method is an inclination angle closed loop feedback control coefficient and has the functions of automatically adjusting the track inclination angle drift trend and maintaining the inclination angle within the control precision index requirement range for a long time, and the value is usually 0.05.
(c) And selecting the tilt angle control time. In summary, the total control quantity of the inclination angle vector includes two parts, namely perturbation control quantity and closed-loop control quantity, and the calculation formula of the velocity increment required by the track inclination angle control is as follows:
in the formula,. DELTA.ix＝Δixd+Δixc，Δiy＝Δiyd+Δiyc. The calculation formula of the length of the control arc segment required by the track inclination angle control is as follows:
in the formula FNSAnd (N) is effective thrust in the north-south direction. If the southward position protection is selected, the inclination angle starting and controlling time of the orbit in each day corresponds to the right ascension:
if the north orientation protection is selected, the inclination angle control starting time of the orbit in each day corresponds to the horizontal right ascension:
(4) simulation verification
The distributed closed-loop position maintaining control method provided by the invention is subjected to mathematical simulation verification to obtain a fixed-point position 120-degree E static trackFor example, the simulation sets the satellite mass to Msat2500kg, light pressure reflection area Ssat＝50m2East-west and south-north direction rail-controlled effective thrust FEW＝FNSControl pulse width length W as 18Npulse0.256s, pulse control period Ppulse20.032 s. Under the combined control action of the flat longitude, the eccentricity and the orbital inclination, the variation curve of the semi-major axis of the satellite in one year is shown in figure 2, the variation curve of the eccentricity is shown in figure 3, the variation curve of the geographical longitude is shown in figure 4, and the variation curve of the orbital inclination is shown in figure 5. The longitude of the satellite is always kept within the range of 120 degrees E +/-0.05 degrees, the eccentricity is always kept within 0.0002, the inclination angle of the orbit is always kept within 0.01 degrees, the increment of the speed required by the combination control of the longitude and the eccentricity of the satellite all the year is about 2.562m/s, and the increment of the speed required by the inclination control of the orbit all the year is about 49.464 m/s.
According to the distributed closed-loop autonomous position maintaining control method provided by the invention, the traditional centralized position maintaining control quantity is distributed to be executed every day, the perturbation variation quantity of every day is eliminated by using pulse thrust, the deviation closed-loop feedback control action is further combined, the satellite orbit position can be always maintained in an accurate range, the east-west direction control accuracy is better than +/-0.05 degrees, the south-north direction control accuracy is better than +/-0.01 degrees, and the autonomous operation capability of the satellite is further improved.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Claims (11)
1. A distributed closed-loop autonomous position maintaining control method for a stationary orbit satellite is characterized by comprising the following steps:
(1) and performing flatness control, comprising:
(1.1) adjusting the semi-long axis of the satellite to the height of a nominal stationary orbit so as to offset the long-term drift of the flatness caused by the non-spherical gravity of the earth and the perturbation of the gravity of the sun and the moon;
(1.2) adopting a distributed closed-loop independent east-west position maintaining control method, and eliminating the perturbation variation of the longitude through impulse thrust every day and the closed-loop control quantity of the longitude deviating from the fixed point position;
(1.3) selecting a longitude leveling control moment, and determining the speed increment and the control arc segment length required by longitude leveling control: performing east-west position keeping control once in the vicinity of a near place or a far place every day, selecting the position in the vicinity of the near place or the far place at the control moment, and adopting a combined control mode of flat longitude and eccentricity;
(2) performing eccentricity control, comprising:
(2.1) judging the starting control condition for applying the eccentricity control according to the relation between the eccentricity perturbation variation and the eccentricity variation caused by the longitude control;
(2.2) calculating an eccentricity perturbation control quantity, an eccentricity closed-loop control quantity and a total speed increment required by eccentricity control;
(2.3) selecting an eccentricity control moment;
(3) performing orbital inclination control comprising:
(3.1) calculating a tilt perturbation control quantity: a distributed closed-loop autonomous north-south position maintaining control method is adopted, long-term and long-period perturbation variation of the track inclination angle is eliminated through pulse thrust every day, and finally daily track inclination angle perturbation control quantity is obtained;
(3.2) calculating the inclination angle closed-loop control quantity of the inclination angle deviating from the target position;
(3.3) selecting the inclination angle control time: the velocity increment required for track pitch control and the control arc length are determined.
2. The distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as claimed in claim 1 wherein: the semi-major axis of the satellite corresponding to the nominal stationary orbit height in the step (1.1) is 42165.795 km;
the calculation formula of the horizontal longitude perturbation variation is as follows:
in the formula
Acceleration is a perturbation of the longitude degree, and is related to the fixed point position of the satellite; delta Tday1(day) represents calculating the change amount within one day;
the calculation formula of the horizontal longitude closed-loop control quantity is as follows:
where λ is the current location geographic longitude, λsFor the target fixed-point position geographic longitude, KpλClosed loop feedback control coefficients for longitude.
3. The distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 2 wherein: the velocity increment calculation formula required for the flat longitude control is as follows:
in the formula of omegae＝360.9856(°/day)＝7.2921×10-5(rad/s) is the rotational angular velocity of the earth, Vs3074.7(m/s) is the nominal stationary track speed;
the control arc segment length required by the flat longitude control is as follows:
in the formula Msat(kg) is the current mass of the satellite, FEW(N) effective thrust in the east-west direction, Wpulse(s) is the control of the pulse width length, Ppulse(s) is a pulse control period.
4. A method of decentralized closed-loop autonomous position maintenance control of geostationary orbiting satellites according to claim 3, characterized in that:
(a) when Δ VT<0, then the vehicle is moving near the near pointAnd (4) performing line-wise west position protection, wherein the start control moment corresponds to the true approach point angle:
(b) when Δ VT>When 0, the east position is protected near the far place, and the corresponding true near point angle at the control starting time is as follows:
5. a method of decentralized closed-loop autonomous position maintenance control of geostationary orbiting satellites according to claim 3, characterized in that: in the step (2.1), the calculation formula of the variation of the beat of the heart rate is as follows:
in the formula CrIs the light pressure reflection coefficient; ssat(m2) The cross section of the satellite perpendicular to the sunlight; p0＝4.65×10-6(N/m2) Is the solar radiation pressure per unit area, Vs3074.7(m/s) is the nominal stationary track speed;
the calculation formula of the eccentricity variation caused by the flat longitude control is as follows:
6. the distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 5 wherein: the judgment conditions for applying the eccentricity control are as follows:
Δep>Δeλ。
7. the distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 5 wherein: the formula for calculating the perturbation control quantity of the eccentricity is as follows:
Δed＝Δep-Δeλ，Δed>0；
the calculation formula of the closed-loop control quantity of the eccentricity ratio is as follows:
Δec＝Kpe(e-ef)；
wherein e is the current eccentricity of efTo target eccentricity size, KpeThe control coefficient is closed-loop feedback control coefficient of eccentricity;
the total velocity increment calculation formula required by eccentricity control is as follows:
8. the distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 7 wherein: selecting the eccentricity control moment, specifically: eccentricity control is executed by combining the first batch of daily flatness control, and eccentricity control quantity is superposed on flatness control quantity
And recording the first control end time as t0The starting time of the second batch is selected to be t0Is executed after +0.5 days, and the control amount is
And in the opposite direction.
9. The distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 7 wherein: in the step (3.1), the calculation formula of the track inclination angle long-term perturbation variation is as follows:
in the formula of omegamThe ascending crossing point of the white tract, right ascension, imIs an included angle between a white road surface and an equatorial plane;
the calculation formula of the long-period perturbation variation of the track inclination angle is as follows:
in the formula ns0.0172(rad/day) 0.9856 (degree/day) mean angular velocity of solar apparent motion, ΩsThe ascending crossing point of the ecliptic tract, right ascension, isIs an included angle between the ecliptic plane and the equatorial plane;
the track inclination perturbation control quantity calculation formula is as follows:
definition of ixAnd iyIs a current tilt angle vector and has ix＝sin(i)cos(Ω)，iySin (i) sin (Ω), where i is the satellite orbital inclination and Ω is the ascension at the intersection; Δ ix1、Δiy1、Δix2、Δiy2、Δixd、ΔiydAre all as follows ixAnd iyA base variation.
10. The distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 9 wherein: the calculation formula of the track inclination angle closed-loop control quantity is as follows:
ixfand iyfIs a target dip angle vector; kpiAnd the control coefficient is a closed loop feedback control coefficient of the inclination angle.
11. The distributed closed-loop autonomous position maintenance control method of geostationary orbiting satellites as set forth in claim 10 wherein: the velocity increment calculation formula required by the track inclination angle control is as follows:
in the formula,. DELTA.ix＝Δixd+Δixc，Δiy＝Δiyd+Δiyc；
The calculation formula of the length of the control arc segment required by the track inclination angle control is as follows:
in the formula FNS(N) is effective thrust in the north-south direction;
the track inclination angle control time is selected as follows:
(a) if the southward position protection is selected, the inclination angle starting and controlling time of the orbit in each day corresponds to the right ascension:
(b) if the north orientation protection is selected, the inclination angle control starting time of the orbit in each day corresponds to the horizontal right ascension:
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