CN111422379B - Formation satellite cooperative orbit control method - Google Patents

Formation satellite cooperative orbit control method Download PDF

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CN111422379B
CN111422379B CN202010202067.9A CN202010202067A CN111422379B CN 111422379 B CN111422379 B CN 111422379B CN 202010202067 A CN202010202067 A CN 202010202067A CN 111422379 B CN111422379 B CN 111422379B
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satellite
control
time
orbit
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CN111422379A (en
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杜耀珂
完备
王文妍
王嘉轶
陈桦
龚腾上
刘美师
崔佳
何煜斌
岳杨
贾艳胜
王禹
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Shanghai Aerospace Control Technology Institute
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/242—Orbits and trajectories
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets

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Abstract

A formation satellite cooperative orbit control method is characterized in that a latitude argument corresponding to a main satellite orbit control point is traversed, a formation configuration eccentricity vector phase angle obtained after orbit control through calculation of the latitude argument is compared with current actual formation satellite formation configuration state parameters, the latitude argument which can meet the conditions most is finally selected as the latitude argument corresponding to the control point, and the starting control air injection time of a main satellite and other auxiliary satellites in a formation satellite is calculated through the selected latitude argument.

Description

Formation satellite cooperative orbit control method
Technical Field
The invention relates to a formation satellite cooperative orbit control method, and belongs to the field of high-precision formation and orbit control of satellites.
Background
The satellite is influenced by atmospheric resistance when in operation, the height of the satellite is continuously reduced, the operation period is continuously shortened, and the actual state is continuously deviated from the nominal state. Therefore, for a conventional single satellite, the orbit is always kept controlled periodically based on constraints such as ground track drift or local time drift. Track retention control is achieved by raising the semi-major axis of the track. Likewise, there are also orbit preservation requirements for the formation satellites. When a single star is changed into a multi-star formation, the orbit maintenance of the formation satellites has the requirement of multi-star cooperative control based on formation configuration and inter-star safety constraint. At present, the research aiming at the multi-satellite cooperative orbit control of the formation satellite has the defects of too simple engineering constraint or too many assumed preconditions and the like, and the effective design close to the practical application is lacked.
Disclosure of Invention
The technical problem solved by the invention is as follows: aiming at the problem that the restraint and control method for the multi-satellite cooperative orbit control of the formation satellites in the prior art are too simple to cause insufficient control precision, the formation satellite cooperative orbit control method is provided.
The technical scheme for solving the technical problems is as follows:
a formation satellite cooperative orbit control method comprises the following steps:
(1) according to the current orbit semimajor axis a of the formation main satellite in the formation satellite, the orbit attenuation rate
Figure BDA0002419732370000011
Calculating the range L of the ground track and the angular velocity n of the orbit of the main star to obtain the velocity increment delta v required by the main star in formation;
(2) according to the current formation satellite configuration state, determining the position coordinates of a formation satellite control point, gradually increasing the latitude argument of the control point from 0 degrees to 360 degrees, and calculating the formation configuration eccentricity vector phase angle theta after the orbit control corresponding to the selected latitude argument in real timeFFc；
(3) Corresponding all latitude argument to the formation configuration eccentricity vector phase angle theta after track controlFFcThe state parameter theta of the formation configuration of the current actual formation satelliteFF0Comparing, selecting the state parameter theta of the formation configuration of the satellite relative to the current actual formationFF0After the control of the orbit closest to 90 degrees or 270 degreesFormation configuration eccentricity vector phase angle thetaFFcThe eccentricity vector phase angle theta is formed by the formation configuration after the track controlFFcThe corresponding latitude argument is taken as the latitude argument u of the control point corresponding to the formation satellite control point for implementing the speed increment controlc；
(4) According to the current on-board time T of the formation main star0Actual latitude argument u0And latitude argument u of control point selected by formation main starcCalculating the air injection intermediate time T of the formation main satellitepGuarantee formation of satellites at TpOrbit control is carried out at all times, and the start-control air injection time of other satellites is calculated at the same time;
(5) according to the thrust F of each satellite in the formation satellitesiAnd satellite mass MiCalculating the air injection time T of each satelliteLiAnd jet start control time TpiAnd performing orbit control according to the start control air injection time and the air injection duration of each satellite.
In the step (1), the method for calculating the speed increment Δ v required by the formation masters comprises the following steps:
Figure BDA0002419732370000021
in the step (2), the current latitude argument corresponds to the formation configuration eccentricity vector phase angle theta after the track controlFFcThe calculation method of (2) is as follows:
θFFc＝atan2([p0·sin(θFF0)+2·Δv/n·sin(uc)],[p0·cos(θFF0)+2·Δv/n·cos(uc)])
in the formula, p0Modulus of eccentricity vector, theta, for formation configurationFF0And configuring the state parameters for the current actual formation satellite formation.
In the step (4), the method for calculating the jet intermediate time of the formation satellites comprises the following steps:
Figure BDA0002419732370000022
Tp＝T0+du/n
in the formula, du is the track phase interval from the start control time to the current time.
In the step (5), the calculation method of the start control air injection time and the air injection duration of each satellite comprises the following steps:
Figure BDA0002419732370000023
in the formula, TLiDuration of injection for each satellite, TpiThe starting and controlling time of air injection.
Compared with the prior art, the invention has the advantages that:
(1) according to the formation satellite cooperative orbit control method provided by the invention, the control constraint of a formation satellite orbit is realized by continuously increasing the latitude amplitude value of a control point corresponding to a formation satellite, the latitude amplitude value which can best meet the control requirement by selecting a constraint condition is taken as the optimal latitude amplitude angle of the control point, the calculation of the air injection start control time of a main satellite and an auxiliary satellite in the formation satellite is realized by the latitude amplitude angle, the control quantity is determined by the formation main satellite, the control is started simultaneously, and the control point is planned, so that the method is oriented to engineering requirements, has good applicability, and can solve the problem of insufficient control precision caused by the fact that the constraint and control method for the formation satellite multi-satellite cooperative orbit control in the prior art is too simple;
(2) the formation satellite cooperative orbit control method adopted by the invention reduces the damage of orbit control to formation configuration as much as possible, adopts a method with simultaneous starting and the same control quantity, ensures that inter-satellite safety can be guaranteed even under abnormal conditions by planning latitude argument of the control point, and can be independently realized on the satellite, the flow is clear, and the calculation precision is high.
Drawings
FIG. 1 is a flow chart of the control of the cooperative orbit of the formation satellites provided by the invention;
FIG. 2 is a schematic diagram of the cooperative orbital control of the formation satellites provided by the invention;
Detailed Description
A formation satellite cooperative orbit control method is characterized in that through continuously increasing latitude argument corresponding to a control point, a formation configuration eccentricity vector phase angle obtained by calculation of the latitude argument after orbit control is compared with current actual formation satellite formation configuration state parameters, the latitude argument which can meet the conditions most is finally selected as the latitude argument corresponding to the control point, and the start control air injection time of a main satellite and other auxiliary satellites in a formation satellite is calculated by using the selected latitude argument, as shown in figure 1, the specific flow is as follows:
(1) according to the current orbit semimajor axis a of the formation main satellite in the formation satellite, the orbit attenuation rate
Figure BDA0002419732370000031
Calculating a speed increment delta v required by the formation masters according to the range L of the ground track and the track angular speed n of the masters, namely six tracks, wherein:
the calculation method of the speed increment delta v required by the formation masters comprises the following steps:
Figure BDA0002419732370000032
(2) according to the current formation satellite configuration state, determining the position coordinates of a formation satellite control point, gradually increasing the latitude argument of the control point from 0 degrees to 360 degrees, and calculating the formation configuration eccentricity vector phase angle theta after the orbit control corresponding to the selected latitude argument in real timeFFcOne formation satellite corresponds to the same control point on the orbit, the latitude argument of the control point is gradually increased, and the parameter corresponding to the latitude argument is calculated for comparison in the subsequent steps, wherein:
formation configuration eccentricity vector phase angle theta after control of corresponding track of current latitude argumentFFcThe calculation method of (2) is as follows:
θFFc＝atan2([p0·sin(θFF0)+2·Δv/n·sin(uc)],[p0·cos(θFF0)+2·Δv/n·cos(uc)])
in the formula, p0Eccentricity vector for formation configurationModulus of magnitude, θFF0Configuring state parameters for the current actual formation satellite formation;
(3) corresponding all latitude argument to the formation configuration eccentricity vector phase angle theta after track controlFFcThe state parameter theta of the formation configuration of the current actual formation satelliteFF0Comparing, selecting the state parameter theta of the formation configuration of the satellite relative to the current actual formationFF0Track controlled formation eccentricity vector phase angle theta closest to 90 DEG or 270 DEGFFcThe eccentricity vector phase angle theta is formed by the formation configuration after the track controlFFcThe corresponding latitude argument is taken as the latitude argument u of the control point corresponding to the formation satellite control point for implementing the speed increment controlc；
In the step, in the process of gradually increasing the latitude argument, each selected latitude argument can calculate a formation configuration eccentricity vector phase angle, each formation configuration eccentricity vector phase angle is compared with a current actual formation satellite formation configuration state parameter area, each angle is closer to 90 degrees or 270 degrees than the current actual formation satellite formation configuration state parameter area, the closest latitude argument is selected, and the latitude argument corresponding to the formation configuration eccentricity vector phase angle is used as the latitude argument corresponding to the control point for implementing the speed increment control on the formation satellite;
(4) according to the current on-board time T of the formation main star0Actual latitude argument u0And latitude argument u of control point of formation main starcCalculating the air injection intermediate time T of the formation main satellitepEnsure formation to be star at TpAnd starting orbit control at the moment, and simultaneously calculating the start control air injection moments of other satellites, wherein:
the control starting air injection time calculation method of each satellite is as follows:
Figure BDA0002419732370000051
Tp＝T0+du/n
in the formula, du is the track phase interval from the control starting moment to the current moment;
(5) calculating the air injection time of each satellite according to the actual thruster thrust and the satellite quality of the formation main satellite and other satellites, and performing orbit control according to the acquired start control air injection time and air injection time of each satellite, wherein:
the calculation method of the start control air injection time and the air injection duration of each satellite comprises the following steps:
Figure BDA0002419732370000052
in the formula, TLiDuration of injection for each satellite, TpiThe starting and controlling time of air injection.
The following is further illustrated with reference to specific examples:
in the embodiment, the situation of the formation satellite is shown in fig. 2, the semi-major axis a of the current orbit of the formation main satellite in the formation satellite, the orbit decay rate
Figure BDA0002419732370000053
The range L of the ground track and the angular velocity n of the orbit of the main satellite are respectively as follows: a is 6900km,
Figure BDA0002419732370000054
l is 3km, n is 0.0011 rad/s; the speed increment delta v needed by the formation masters can be calculated as follows: Δ v ═ 0.025 m/s;
determining the position coordinates of the control points of the formation satellites, gradually increasing the latitude argument of the control points from 0 degrees to 360 degrees, and calculating the formation configuration eccentricity vector phase angle theta after the orbit control corresponding to the selected latitude argument in real timeFFcThe state parameter theta of the formation configuration of the satellite with the current actual formationFF0In the comparison process, the vector phase angle theta of the eccentricity of the formation configuration can be obtained when the latitude argument is 315.2 degreesFFcThe latitude argument u which is closest to 90 degrees or 270 degrees and is used for implementing speed increment control and corresponds to the current formation satellite control point by 315.2 degreesc；
Current on-board time T of formation main star in formation satellite0384000s, current latitude breadthAngle u0Calculating the air injection middle time of the formation main star as 180 degrees: t isp＝386142.194s；
Taking the main satellite with the mass of 1500kg and the thrust of 2.4N; the auxiliary satellite mass is 1480kg, the thrust is 2.2N, and can be calculated as follows: the main satellite air injection time length is as follows: t isL main15.625 s; the initial gas injection time is as follows: t isp main＝386134.815s；
The jet time length of the satellite is as follows: t isL auxiliary16.818 s; the initial gas injection time is as follows: t isp auxiliary＝386133.785s；
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 (5)

1. A formation satellite cooperative orbit control method is characterized by comprising the following steps:
(1) according to the current orbit semimajor axis a of the formation main satellite in the formation satellite, the orbit attenuation rate
Figure FDA0003009694590000011
Calculating the range L of the ground track and the angular velocity n of the orbit of the main star to obtain the velocity increment delta v required by the main star in formation;
(2) according to the current formation satellite configuration state, determining the position coordinates of a formation satellite control point, gradually increasing the latitude argument of the control point from 0 degrees to 360 degrees, and calculating the formation configuration eccentricity vector phase angle theta after the orbit control corresponding to the selected latitude argument in real timeFFc；
(3) Corresponding all latitude argument to the formation configuration eccentricity vector phase angle theta after track controlFFcThe state parameter theta of the formation configuration of the current actual formation satelliteFF0Comparing, selecting the state parameter theta of the formation configuration of the satellite relative to the current actual formationFF0Track controlled formation eccentricity vector phase angle theta closest to 90 DEG or 270 DEGFFcThe eccentricity vector phase angle theta is formed by the formation configuration after the track controlFFcThe corresponding latitude argument is taken as the latitude argument u of the control point corresponding to the formation satellite control point for implementing the speed increment controlc；
(4) According to the current on-board time T of the formation main star0Actual latitude argument u0And latitude argument u of control point selected by formation main starcCalculating the air injection intermediate time T of the formation main satellitepGuarantee formation of satellites at TpOrbit control is carried out at all times, and the start-control air injection time of other satellites is calculated at the same time;
(5) according to the thrust F of each satellite in the formation satellitesiAnd satellite mass MiCalculating the air injection time T of each satelliteLiAnd jet start control time TpiAnd performing orbit control according to the start control air injection time and the air injection duration of each satellite.
2. The method according to claim 1, wherein the method comprises:
in the step (1), the method for calculating the speed increment Δ v required by the formation masters comprises the following steps:
Figure FDA0003009694590000021
wherein, DeltaL is the variation of the ground track range, ReThe radius of the earth.
3. The method according to claim 1, wherein the method comprises:
in the step (2), the current latitude argument corresponds to the formation configuration eccentricity vector phase angle theta after the track controlFFcThe calculation method of (2) is as follows:
θFFc＝a tan 2([p0·sin(θFF0)+2·Δv/n·sin(uc)],[p0·cos(θFF0)+2·Δv/n·cos(uc)])
in the formula, p0Modulus of eccentricity vector, theta, for formation configurationFF0And configuring the state parameters for the current actual formation satellite formation.
4. The method according to claim 1, wherein the method comprises:
in the step (4), the method for calculating the jet intermediate time of the formation satellites comprises the following steps:
Figure FDA0003009694590000022
Tp＝T0+du/n
in the formula, du is the track phase interval from the start control time to the current time.
5. The method according to claim 1, wherein the method comprises:
in the step (5), the calculation method of the start control air injection time and the air injection duration of each satellite comprises the following steps:
Figure FDA0003009694590000031
in the formula, TLiDuration of injection for each satellite, TpiThe starting and controlling time of air injection.
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CN114460952B (en) * 2022-01-17 2023-03-24 上海卫星工程研究所 Double-star cooperative orbit transfer method and system for initializing elliptical orbit flight accompanying configuration
CN116166049B (en) * 2023-04-25 2023-07-07 中国西安卫星测控中心 Inter-star distance maintaining control method for unstable multi-star serial formation system
CN116750210B (en) * 2023-07-12 2024-02-20 银河航天(北京)网络技术有限公司 Method for controlling orbit of satellite formation

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CN101794154A (en) * 2009-11-25 2010-08-04 哈尔滨工业大学 Decoupling control method for relative orbits and attitudes of formation satellites
CN103257653A (en) * 2013-05-22 2013-08-21 上海新跃仪表厂 Satellite team configuring control method based on fuel consumption optimization
CN110316402A (en) * 2019-06-03 2019-10-11 上海航天控制技术研究所 A kind of satellite attitude control method under formation control mode
CN110450991A (en) * 2019-08-16 2019-11-15 西北工业大学 The method of micro-nano satellite cluster capture space non-cooperative target
CN110849377A (en) * 2019-11-20 2020-02-28 上海航天控制技术研究所 Space target relative navigation system and method based on constellation cooperation

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101794154A (en) * 2009-11-25 2010-08-04 哈尔滨工业大学 Decoupling control method for relative orbits and attitudes of formation satellites
CN103257653A (en) * 2013-05-22 2013-08-21 上海新跃仪表厂 Satellite team configuring control method based on fuel consumption optimization
CN110316402A (en) * 2019-06-03 2019-10-11 上海航天控制技术研究所 A kind of satellite attitude control method under formation control mode
CN110450991A (en) * 2019-08-16 2019-11-15 西北工业大学 The method of micro-nano satellite cluster capture space non-cooperative target
CN110849377A (en) * 2019-11-20 2020-02-28 上海航天控制技术研究所 Space target relative navigation system and method based on constellation cooperation

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