CN114735239B - Spacecraft orbit maneuvering control method and device and electronic equipment - Google Patents

Spacecraft orbit maneuvering control method and device and electronic equipment Download PDF

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CN114735239B
CN114735239B CN202210453387.0A CN202210453387A CN114735239B CN 114735239 B CN114735239 B CN 114735239B CN 202210453387 A CN202210453387 A CN 202210453387A CN 114735239 B CN114735239 B CN 114735239B
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spacecraft
orbit
track
nominal
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CN114735239A (en
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费立刚
吕丽红
李基�
耿淑敏
韩宝玲
王兴龙
杨纪伟
张超
崔庆丰
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    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
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    • B64G1/00—Cosmonautic vehicles
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Abstract

The invention provides a method, a device and electronic equipment for controlling spacecraft orbit maneuvering, and relates to the technical field of spacecraft orbit control, wherein an orbit control mode comprises at least one of the following modes: the method comprises a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity maintaining control mode and a satellite point track maintaining control mode, wherein each maintaining control mode has a corresponding trigger condition, and in the orbit process of the spacecraft, whether the spacecraft needs to be subjected to orbit control or not and an orbit control strategy when the spacecraft needs to be controlled can be judged through the number of orbits and an actually selected orbit control mode, so that the running speed of the spacecraft can be adjusted according to a corresponding orbit control speed increment at a target control moment. Therefore, the method belongs to a self-adaptive adjusting method of the spacecraft orbit, and compared with a method for controlling the orbit of the spacecraft by using the control parameters which are injected in advance, the method can effectively solve the technical problem that the spacecraft orbit maintaining accuracy is poor.

Description

Spacecraft orbit maneuvering control method and device and electronic equipment
Technical Field
The invention relates to the technical field of spacecraft orbit control, in particular to a spacecraft orbit maneuver control method, a spacecraft orbit maneuver control device and electronic equipment.
Background
In the actual engineering business execution process, a large part of spacecraft orbit maneuvering control is calculated by a ground system according to measurement data and a control target in advance through a ground computing system, and the control parameters are processed to be injected to a spacecraft from a ground station to execute control operation. However, in the actual operation process of the spacecraft, some situations which cannot be expected in advance may be encountered, and then the orbit deviation may occur when the spacecraft is controlled by using the control parameters calculated in advance, so that the orbit of the spacecraft is not maintained accurately enough.
Disclosure of Invention
The invention aims to provide a spacecraft orbit maneuver control method, a spacecraft orbit maneuver control device and electronic equipment, so as to solve the technical problem of poor spacecraft orbit maintaining accuracy in the conventional spacecraft orbit control method.
In a first aspect, the invention provides a spacecraft orbit maneuver control method, which comprises the following steps: acquiring an orbit control mode and an orbit number of a spacecraft; wherein the trajectory control mode comprises at least one of: a semi-long shaft maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition; judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition of the orbit control mode and the number of the orbits; if so, determining an orbit control strategy of the spacecraft based on the orbit control mode and the orbit number; wherein the trajectory control strategy comprises: at least one control time and a track control speed increment corresponding to each control time; and adjusting the running speed of the spacecraft based on the corresponding track control speed increment at the target control time.
In an alternative embodiment, the trigger condition of the semi-major axis maintaining control mode comprises: the orbit flat semi-major axis of the spacecraft reaches a preset semi-major axis limit value; if the track control mode comprises: if the semi-major axis maintains the control mode, determining the orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits comprises: acquiring a nominal half-length axis value, a nominal track average speed and a maximum value of a track half-length axis at the current moment; determining a first adjustment amount of the track semimajor axis based on the maximum value of the track semimajor axis and the preset semimajor axis limit value; determining a first track control speed increment based on a first adjustment to the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed; determining a first control time corresponding to the first orbit control speed increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value; a first trajectory control strategy is determined based on the first control time and the first trajectory control speed increment.
In an alternative embodiment, the determining a first control time corresponding to the first orbit control velocity increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value includes: judging whether the mean anomaly angle of the spacecraft is larger than 180 degrees; if not, using the formula
Figure BDA0003617806840000021
Calculating the first control moment; wherein, t c1 Representing the first control moment, t representing the current moment, M representing the mean anomaly of the spacecraft,
Figure BDA0003617806840000022
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000023
representing the nominal half-major axis value; if it is greater than the above value, the formula is used
Figure BDA0003617806840000024
And calculating the first control moment.
In an alternative embodiment, the triggering condition of the tilt angle maintenance control mode includes: the orbit inclination angle of the spacecraft exceeds a preset inclination angle range; if the track control mode comprises: determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits in a tilt angle maintenance control mode, wherein the determining the orbit control strategy comprises the following steps: acquiring a nominal inclination angle value, a nominal track average speed and a nominal half-major axis value; determining an orbit inclination angle adjustment amount based on the nominal inclination angle value and the current orbit inclination angle value; determining a second track control speed increment based on the track pitch adjustment and the nominal track average speed; determining a second control time corresponding to the second orbit control speed increment based on the nominal half-major axis value, the orbit ascension point right ascension of the spacecraft and the mean-anomaly angle of the spacecraft; a second trajectory control strategy is determined based on the second control time and the second trajectory control speed increment.
In an alternative embodiment, the determining a second control time corresponding to the second orbit control speed increment based on the nominal half-major axis value, the orbit ascension of the spacecraft, and the mean-paraxial angle of the spacecraft includes: if the time when the spacecraft moves to the ascending intersection point is earlier than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure BDA0003617806840000031
Calculating the second control moment; wherein, t c2 Representing the second control moment, t representing the current moment, ω representing the orbital ascent crossing right ascension of the spacecraft, M representing the mean-anomaly angle of the spacecraft,
Figure BDA0003617806840000032
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000033
representing the nominal half-major axis value; if the time when the spacecraft moves to the ascending intersection point is later than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure BDA0003617806840000034
And calculating the second control moment.
In an alternative embodiment, the triggering condition of the eccentricity maintenance control mode includes: the variation of the orbit eccentricity ratio of the spacecraft exceeds a preset eccentricity ratio variation limit value; if the track control mode comprises: an eccentricity maintaining control mode, and then determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, including: acquiring a nominal eccentricity and a nominal track average speed; determining a variation in orbital eccentricity of the spacecraft based on the nominal eccentricity and a current orbital eccentricity; determining a third track control speed increment and a fourth track control speed increment based on the variation in track eccentricity and the nominal track average speed; determining a third control time corresponding to the third orbit control speed increment and a fourth control time corresponding to the fourth orbit control speed increment based on the orbit ascension point right ascension of the spacecraft, the mean-anomaly angle of the spacecraft and the variation of the orbit eccentricity; determining a third trajectory control strategy based on the third control time, the third trajectory control speed increment, the fourth control time, and the fourth trajectory control speed increment.
In an optional embodiment, the triggering condition of the sub-satellite point trajectory maintenance control mode includes:
Figure BDA0003617806840000041
wherein λ represents a current longitude of the spacecraft,
Figure BDA0003617806840000042
nominal longitude representing the target latitude circle, N representing the number of intervals that the regression orbit trajectory net divides a latitude circle equally, Δ λ max Representing a preset longitude boundary value; if the track control mode comprises: if the track of the sub-satellite points maintains the control mode, determining the orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits comprises the following steps: acquiring a nominal track average speed, the nominal longitude, the preset longitude boundary value, a nominal semi-major axis value and a semi-major axis attenuation speed caused by atmospheric resistance; determining a semi-major axis deviation amount based on the nominal longitude, the preset longitude boundary value, the nominal semi-major axis value, and the semi-major axis decay rate; determining a second adjustment quantity of the orbit semi-major axis based on the semi-major axis deviation quantity, the nominal semi-major axis value and the current orbit semi-major axis value of the spacecraft; determining a fifth track control speed increment based on a second adjustment to the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed; determining a fifth control moment corresponding to the fifth orbit control speed increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value; based on the fifth control time and the fifth trackThe control speed increment determines a fourth trajectory control strategy.
In a second aspect, the present invention provides a spacecraft orbit maneuver control device, comprising: the acquisition module is used for acquiring the orbit control mode and the number of orbits of the spacecraft; wherein the trajectory control mode comprises at least one of: a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition; the judging module is used for judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition of the orbit control mode and the number of the orbits; the determining module is used for determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits under the condition that the spacecraft needs to be subjected to orbit control; wherein the trajectory control strategy comprises: at least one control time and a track control speed increment corresponding to each control time; and the adjusting module is used for adjusting the running speed of the spacecraft based on the corresponding track control speed increment at the target control moment.
In a third aspect, the present invention provides an electronic device, comprising a memory and a processor, wherein the memory stores a computer program operable on the processor, and the processor executes the computer program to implement the steps of the method according to any of the foregoing embodiments.
In a fourth aspect, the invention provides a computer readable medium having non-volatile program code executable by a processor, the program code causing the processor to perform the method of any of the preceding embodiments.
The invention provides a spacecraft orbit maneuver control method, which comprises the following steps: acquiring an orbit control mode and the number of orbits of a spacecraft; wherein the track control mode comprises at least one of: a semi-long shaft maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition; judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition and the number of orbits of the orbit control mode; if so, determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits; wherein the track control strategy comprises: at least one control moment and a track control speed increment corresponding to each control moment; and adjusting the running speed of the spacecraft based on the corresponding orbit control speed increment at the target control moment.
The invention provides a spacecraft orbit maneuver control method, and an orbit control mode comprises at least one of the following steps: the method comprises a semi-long axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity maintaining control mode and a sub-satellite point track maintaining control mode, wherein each maintaining control mode has a corresponding trigger condition, and in the orbit process of the spacecraft, whether the spacecraft needs to be subjected to orbit control and an orbit control strategy when the spacecraft needs to be controlled can be judged through the number of orbits and an actually selected orbit control mode, so that the operating speed of the spacecraft is adjusted according to a corresponding orbit control speed increment at a target control moment. Therefore, the method belongs to a self-adaptive adjusting method of the spacecraft orbit, and compared with a method for controlling the spacecraft orbit by using control parameters which are injected in advance, the method can effectively solve the technical problem of poor spacecraft orbit maintaining accuracy.
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In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
Fig. 1 is a flowchart of a spacecraft orbit maneuver control method according to an embodiment of the present invention;
fig. 2 is a schematic flow chart of a spacecraft orbit maneuver control method according to an embodiment of the present invention;
fig. 3 is a functional block diagram of a spacecraft orbit maneuver control device provided by the embodiment of the invention;
fig. 4 is a schematic diagram of an electronic device according to an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. The components of embodiments of the present invention generally described and illustrated in the figures herein may be arranged and designed in a wide variety of different configurations.
Thus, the following detailed description of the embodiments of the present invention, presented in the figures, is not intended to limit the scope of the invention, as claimed, but is merely representative of selected embodiments of the invention. All other embodiments, which can be obtained by a person skilled in the art without inventive step based on the embodiments of the present invention, are within the scope of protection of the present invention.
Some embodiments of the invention are described in detail below with reference to the accompanying drawings. The embodiments and features of the embodiments described below can be combined with each other without conflict.
In the actual engineering business execution process, a large part of spacecraft orbit maneuvering control is calculated by a ground system according to measurement data and a control target by a ground computing system in advance, and is processed into control parameters which are injected to a spacecraft from a ground station to execute control operation. However, in the actual operation process of the spacecraft, some situations which cannot be expected in advance may be encountered, and then the orbit deviation may occur when the spacecraft is controlled by using the control parameters calculated in advance, so that the orbit of the spacecraft is not maintained accurately enough. In view of the above, the embodiments of the present invention provide a spacecraft orbit maneuver control method to alleviate the above technical problems.
Example one
Fig. 1 is a flowchart of a spacecraft orbit maneuver control method provided in an embodiment of the present invention, and as shown in fig. 1, the method specifically includes the following steps:
and S102, acquiring the orbit control mode and the number of orbits of the spacecraft.
Specifically, as can be seen from the above description of the conventional method, a method for forecasting the orbit control of the spacecraft according to the control target type and the autonomy according to the threshold is lacked in the prior art, and therefore, the embodiment of the present invention aims to provide a method capable of performing the autonomous orbit control of the spacecraft according to the control target and the corresponding threshold condition.
Specifically, to control the orbit maneuver of the spacecraft, the embodiment of the invention first needs to acquire an orbit control mode and the number of orbits of the spacecraft, wherein the orbit control mode includes at least one of the following: a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition. That is, to make the spacecraft perform adaptive adjustment of the orbit, the orbit control mode thereof may be one or more of the above maintenance control modes, which are compatible with each other. Moreover, each maintenance control mode has a corresponding trigger condition, and if the orbit control mode selects 3 maintenance control modes, the orbit control of the spacecraft has 3 corresponding trigger conditions, and the 3 trigger conditions perform trigger judgment at the same time at each moment.
And step S104, judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition and the number of the orbits of the orbit control mode.
If yes, executing the following step S106; if not, judging whether the track control strategy to be executed exists at the current moment, if not, performing the next track forecast through track recursion, and if so, executing the following step S108.
And S106, determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits.
After the orbit control mode and the number of orbits of the spacecraft are obtained, triggering judgment can be performed by using the number of orbits of the spacecraft and at least one triggering condition in the determined orbit control mode, if the number of orbits can meet a target triggering condition (any one of at least one triggering condition set currently), an orbit control strategy of the spacecraft is determined according to a maintaining control mode and the number of orbits corresponding to the target triggering condition, that is, if the number of orbits simultaneously meets a plurality of triggering conditions set currently, a plurality of orbit control strategies of the spacecraft need to be determined according to the maintaining control mode and the number of orbits corresponding to each met triggering condition, respectively, wherein the orbit control strategies include: at least one control moment and a track control velocity increment corresponding to each control moment.
If the number of the orbits does not meet any one of the currently set triggering conditions, whether an orbit control strategy to be executed exists at the current moment or not needs to be judged, namely whether the current moment is the control moment corresponding to a certain orbit control strategy determined before or not needs to be judged, if not, the orbit of the spacecraft needs to be forecasted and calculated according to the existing method without the control strategy, namely, the orbit is forecasted by analyzing according to the space stress of the spacecraft only without considering the control strategy through orbit recursion. If so, the following step S108 is performed.
And S108, adjusting the running speed of the spacecraft based on the corresponding track control speed increment at the target control time.
In the embodiment of the invention, the target control time is any control time included in any orbit control strategy, when the time reaches the target control time, the orbit control of the spacecraft is required, that is, the operation speed (vector) of the spacecraft is adjusted by using the orbit control speed increment corresponding to the target control time, so that the aim of the orbit control is fulfilled. After the orbit control is completed once, the orbit of the spacecraft can be continuously forecasted by an orbit recursion method, and the steps S102-S108 are repeatedly executed until a preset ending moment is reached.
Since the track control speed increment is typically defined in the track coordinate system, the track coordinate system is defined as:
Figure BDA0003617806840000091
wherein i represents an x-direction unit vector, j represents a y-direction unit vector, k represents a z-direction unit vector, r represents a spacecraft position vector, and v represents a spacecraft velocity vector.
The spacecraft orbit maneuver control method provided by the embodiment of the invention belongs to a pulse type maneuver control mode, and assumes that the applied orbit control speed increment at a certain moment is [ delta v [ ] x ,Δv y ,Δv z ]Then the inertia system velocity increment vector is Δ v = Δ v x i+Δv y j+Δv z k, the position of the spacecraft after maneuvering is unchanged, the speed is v = Δ v + v 0 Wherein v is 0 Representing the velocity vector of the spacecraft before the maneuver and v representing the velocity vector of the spacecraft after the maneuver. The orbit integral dynamic model is as follows:
Figure BDA0003617806840000092
wherein,
Figure BDA0003617806840000093
denotes the spacecraft acceleration, mu denotes the Earth's gravitational constant, a p The method is characterized by comprising the following steps of (1) representing the sum of various perturbation forces suffered by the spacecraft, wherein the perturbation forces comprise: the method comprises the steps of earth non-spherical perturbation, atmospheric resistance perturbation, light pressure perturbation, tide perturbation and the like, and the long-term forecasting and calculating method is used in the calculating process and the non-control strategy. That is, the orbit control of the spacecraft is performed only by adding the corresponding orbit control speed increment to the speed vector at a given control moment. And the track forecast between the two maneuvering moments is calculated according to the long-term forecast without a control strategy.
Fig. 2 is a schematic flow chart of a spacecraft orbit maneuver control method according to an embodiment of the present invention, and it can be known from the above description that a basic flow of orbit control on a spacecraft is to calculate a spacecraft orbit according to a given time step length and a non-control strategy forecast, after each step of forecast, according to a set orbit control mode and a corresponding trigger condition, determine whether a maneuver is required (i.e., determine whether an orbit control is required), and if a control is required, calculate an orbit control speed increment according to the set mode, and generate a set of control strategies (which may include multiple control times and multiple orbit control speed increments) to be stored in a memory. And when the forecasted time is equal to the time of the first control strategy which is not implemented, adding the corresponding orbit control speed increment into the spacecraft running state, continuing to forecast next time, marking the control as implemented, and outputting the number, position, speed and orbit control speed increment of the spacecraft in the current step.
The invention provides a spacecraft orbit maneuver control method, and an orbit control mode comprises at least one of the following steps: the method comprises a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity maintaining control mode and a satellite point track maintaining control mode, wherein each maintaining control mode has a corresponding trigger condition, in the orbit process of the spacecraft, whether the spacecraft needs to be subjected to orbit control or not and an orbit control strategy when the spacecraft needs to be controlled can be judged through the number of orbits and an actually selected orbit control mode, and then the running speed of the spacecraft is adjusted according to a corresponding orbit control speed increment in the target control moment. Therefore, the method belongs to a self-adaptive adjusting method of the spacecraft orbit, and compared with a method for controlling the spacecraft orbit by using the control parameters which are injected in advance, the method can effectively solve the technical problem of poor spacecraft orbit maintaining accuracy.
In the above description, the orbit control mode of the spacecraft is introduced to be at least one of the semi-major axis maintaining control mode, the inclination angle maintaining control mode, the eccentricity maintaining control mode and the sub-satellite point trajectory maintaining control mode, and for convenience of description, a detailed method for orbit control of the spacecraft by each maintaining control mode will be described below.
In an alternative embodiment, the trigger condition for the semi-major axis maintenance control mode includes: the orbit flat semi-major axis of the spacecraft reaches the preset semi-major axis limit value.
Specifically, the semimajor axis maintenance control is directed to the case where the semimajor axis damping is caused by atmospheric resistance. In the embodiment of the present invention, the predetermined semimajor axis limit is a min . In atmospheric resistanceUnder the influence of perturbation, when the orbit mean semi-major axis of the spacecraft reaches a min In the process, the semimajor axis needs to be raised, and the single-pulse raising time is set to be at a far place so as to ensure that the eccentricity is not continuously increased.
Therefore, under the condition that the orbit control mode selects the semimajor axis maintaining control mode, attention needs to be paid to the size of the orbit semimajor axis of the spacecraft all the time, and if the orbit semimajor axis of the spacecraft reaches the preset semimajor axis limit value, the orbit control strategy of the spacecraft needs to be updated.
If the track control mode includes: if the semi-major axis maintains the control mode, the step S106 determines the orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, including:
step S1061, obtaining a nominal semi-major axis value, a nominal track average speed and a maximum value of the track semi-major axis at the current moment.
Step S1062 determines a first adjustment amount of the track semimajor axis based on the maximum value of the track semimajor axis and a preset semimajor axis limit value.
In the semi-major axis maintaining control mode, before the orbit control strategy of the spacecraft is prepared, a nominal semi-major axis value is firstly obtained
Figure BDA0003617806840000113
Nominal track average velocity V and maximum value a of track semimajor axis at current moment of cutoff max Then, using the maximum a of the track semimajor axis max And a preset semi-major axis limit a min Calculating a first adjustment delta a of the semi-major axis of the track 1 Wherein, Δ a 1 ＝a max -a min 。
Step S1063, determining a first track control speed increment based on the first adjustment for the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed.
Further, a first adjustment amount delta a according to the semi-major axis of the track 1 Nominal semi-major axis value
Figure BDA0003617806840000111
And the nominal track average velocity V can be calculatedOne track control velocity increment Δ v 1 Specifically, the embodiments of the present invention utilize equations
Figure BDA0003617806840000112
A first orbital control speed increment is calculated.
Step S1064, determining a first control time corresponding to the first orbit control speed increment based on the mean anomaly angle and the nominal semi-major axis value of the spacecraft.
While determining the first orbit control speed increment, a first control time corresponding to the orbit control speed increment is also determined, in the embodiment of the invention, the first control time needs to be calculated based on the mean anomaly of the spacecraft and the nominal semi-major axis value, and the mean anomaly of the spacecraft is a parameter contained in the orbit number of the spacecraft.
Optionally, in step S1064, the determining a first control time corresponding to the first orbit control speed increment based on the mean anomaly angle and the nominal semi-major axis value of the spacecraft specifically includes the following steps:
step S10641, judging whether the mean anomaly angle of the spacecraft is larger than 180 degrees.
If the mean anomaly of the spacecraft is not more than pi, executing the following step S10642; if the mean anomaly angle of the spacecraft is greater than pi, the following step S10643 is performed.
Step S10642, using the equation
Figure BDA0003617806840000121
Calculating a first control moment; wherein, t c1 Representing a first control moment, t representing the current moment, M representing the mean anomaly of the spacecraft,
Figure BDA0003617806840000122
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000123
the nominal half-length axis value is indicated.
Step S10643, using the equation
Figure BDA0003617806840000124
The first control timing is calculated.
In step S1065, a first trajectory control strategy is determined based on the first control time and the first trajectory control speed increment.
In an alternative embodiment, the triggering condition of the tilt angle maintenance control mode includes: the orbit inclination angle of the spacecraft exceeds the preset inclination angle range.
In the embodiment of the present invention, if the nominal inclination angle value is
Figure BDA0003617806840000125
The limit value of the change of the inclination angle is delta i max If the current track tilt value is i, the trigger condition of the tilt maintaining control mode is:
Figure BDA0003617806840000126
or
Figure BDA0003617806840000127
That is, the orbit inclination of the spacecraft exceeds the preset inclination range.
Therefore, under the condition that the inclination angle maintaining control mode is selected in the orbit control mode, attention needs to be paid to the inclination angle range where the current orbit inclination angle value is located all the time, and if the orbit inclination angle of the spacecraft exceeds the preset inclination angle range, the orbit control strategy of the spacecraft needs to be updated.
If the track control mode includes: if the inclination angle is maintained in the control mode, the step S106 determines the orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, and specifically includes the following steps:
step S201, a nominal inclination angle value, a nominal track average speed and a nominal half-major axis value are obtained.
Step S202, determining an orbit inclination angle adjustment amount based on the nominal inclination angle value and the current orbit inclination angle value.
In the inclination angle maintaining control mode, before an orbit control strategy of a spacecraft is prepared, a nominal inclination angle value is firstly obtained
Figure BDA0003617806840000128
Nominal rail mean velocity V and nominal semi-major axis value
Figure BDA0003617806840000129
Next, using the nominal tilt angle value
Figure BDA00036178068400001210
And a current track inclination value i, wherein,
Figure BDA00036178068400001211
in step S203, a second track control speed increment is determined based on the track pitch adjustment and the nominal track average speed.
Further, after determining the track pitch angle adjustment Δ i, a second track control velocity increment Δ ν is calculated using the track pitch angle adjustment Δ i and the nominal track average velocity V 2 Specifically, using the formula Δ v 2 = Δ iV calculates the second orbit control speed increment.
And S204, determining a second control moment corresponding to the second orbit control speed increment based on the nominal half-major axis value, the orbit ascension point of the spacecraft and the mean-anomaly angle of the spacecraft.
At the same time as the second track control speed increment is determined, a second control moment corresponding to the speed increment should also be determined and selected as the next closest point of ascent or descent. According to the embodiment of the invention, the second control moment is calculated by using the nominal half-length axis value, the orbit rising point right ascension of the spacecraft and the flat near point angle of the spacecraft, and the orbit rising point right ascension of the spacecraft and the flat near point angle of the spacecraft are parameters contained in the orbit number of the spacecraft.
Optionally, in step S204, the second control time corresponding to the second orbit control speed increment is determined based on the nominal half-length axis value, the right ascension of the orbit intersection of the spacecraft, and the mean-anomaly angle of the spacecraft, and specifically includes the following contents:
if the spacecraft moves to the lifting intersection point earlier than the lowering intersection pointAt that time, the formula is used
Figure BDA0003617806840000131
Calculating a second control moment; wherein, t c2 Represents a second control moment, t represents the current moment, omega represents the orbit ascent point right ascension of the spacecraft, M represents the mean anomaly angle of the spacecraft,
Figure BDA0003617806840000132
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000133
the nominal half-length axis value is indicated. mod (ω + M,2 π) represents the current latitude argument modulo 2 π.
If the time when the spacecraft moves to the ascending intersection point is later than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure BDA0003617806840000134
A second control time is calculated.
In step S205, a second trajectory control strategy is determined based on the second control time and the second trajectory control speed increment.
In an alternative embodiment, the triggering conditions for the eccentricity maintenance control mode include: the variation of the orbit eccentricity of the spacecraft exceeds a preset eccentricity variation limit value.
In the embodiments of the present invention, if the nominal eccentricity is expressed as
Figure BDA0003617806840000141
The preset eccentricity change limit value is delta e max The current track eccentricity is (e) x ，e y ) Then the triggering condition of the eccentricity maintenance control mode is expressed as: Δ e > Δ e max Wherein, in the process,
Figure BDA0003617806840000142
that is, the variation Δ e of the orbital eccentricity of the spacecraft exceeds the preset eccentricity variation limit Δ e max . I.e. selecting bias in the track control modeIn the case of the heart rate maintenance control mode, attention needs to be paid to the orbital eccentricity (e) of the spacecraft all the time x ，e y ) Whether or not to exceed the nominal eccentricity
Figure BDA0003617806840000143
Is a circle with a center, if the center of the circle exceeds the center of the circle, the orbit control strategy of the spacecraft needs to be updated.
If the track control mode includes: if the eccentricity maintains the control mode, in the step S106, the orbit control strategy of the spacecraft is determined based on the orbit control mode and the number of orbits, which specifically includes the following steps:
step S301, a nominal eccentricity and a nominal track average speed are obtained.
And step S302, determining the variation of the orbital eccentricity of the spacecraft based on the nominal eccentricity and the current orbital eccentricity.
In the eccentricity maintaining control mode, before the orbit control strategy of the spacecraft is determined, the nominal eccentricity is acquired firstly
Figure BDA0003617806840000144
And a nominal track average velocity V, followed by first using the nominal eccentricity
Figure BDA0003617806840000145
And current track eccentricity (e) x ，e y ) The variation deltae in the orbital eccentricity of the spacecraft is determined. The calculation method of Δ e has already been described above, and will not be described here.
Step S303 determines a third track control speed increment and a fourth track control speed increment based on the variation of the track eccentricity and the nominal track average speed.
In the eccentricity maintaining control mode, once the maintaining control condition is triggered, two track control speed increments are adopted, namely the latitude argument and the tangential pulse speed increment of the eccentricity control double pulse are respectively adopted, namely, the spacecraft needs to be ignited twice. In the embodiment of the invention, after the variation delta e of the orbital eccentricity of the spacecraft is determined, the variation delta e of the orbital eccentricity and the nominal value are utilizedThe average track velocity V may be used to determine a third track control velocity increment Δ V, respectively 3 And a fourth track control speed increment Δ v 4 . Specifically, the formula of utilization
Figure BDA0003617806840000151
Calculating a third track control speed increment; equation of utilization
Figure BDA0003617806840000152
A fourth orbit control speed increment is calculated.
Step S304, determining a third control time corresponding to a third orbit control speed increment and a fourth control time corresponding to a fourth orbit control speed increment based on the right ascension of the orbit intersection point of the spacecraft, the mean-near point angle of the spacecraft and the variation of the orbit eccentricity.
And when determining the third and fourth orbit control speed increments, determining a third and fourth control time corresponding to the speed increment, in the embodiment of the present invention, a third control time corresponding to the third orbit control speed increment is calculated by using the orbit ascent point right ascent of the spacecraft, the mean-near point angle of the spacecraft, and the variation of the orbit eccentricity, and a fourth control time corresponding to the fourth orbit control speed increment is calculated by using the orbit ascent point right ascent of the spacecraft, the mean-near point angle of the spacecraft, and the variation of the orbit eccentricity. The right ascension point of the orbit of the spacecraft and the mean and near point angle of the spacecraft are parameters contained in the number of orbits of the spacecraft.
Specifically, the third control timing and the fourth control timing are calculated using the following equations:
Figure BDA0003617806840000153
wherein, t c3 Denotes a third control time, t c4 Represents a fourth control moment, omega represents the orbit ascent point right ascension of the spacecraft, M represents the mean anomaly angle of the spacecraft,
Figure BDA0003617806840000154
an x-direction component representing the variation deltae in the orbital eccentricity of the spacecraft,
Figure BDA0003617806840000155
a y-direction component representing the variation Δ e of the orbital eccentricity of the spacecraft.
Step S305, a third track control strategy is determined based on the third control time, the third track control speed increment, the fourth control time and the fourth track control speed increment.
In an optional embodiment, the triggering condition of the sub-satellite point trajectory maintenance control mode includes:
Figure BDA0003617806840000156
where lambda denotes the current longitude of the spacecraft,
Figure BDA0003617806840000157
nominal longitude, representing the target latitude circle, N represents the number of intervals that the regressive orbit trajectory network will divide one latitude circle equally,
Figure BDA0003617806840000158
representing the inter-adjacent tracings, Δ λ max Representing a preset longitude boundary value.
Specifically, the main function of the undersatellite point track maintenance control is to offset the long-term drift of the undersatellite point track caused by the change of atmospheric resistance to the semi-major axis. In an embodiment of the invention, the nominal longitude through a certain latitude circle (target latitude circle) is
Figure BDA0003617806840000161
Presetting a longitude boundary value as delta lambda max And the current longitude of the spacecraft is λ, the triggering condition of the substellar point trajectory maintenance control mode can be expressed as:
Figure BDA0003617806840000162
therefore, when the orbit control mode selects the intersatellite point trajectory maintenance control mode, the current longitude of the spacecraft needs to be concerned all the time, and once the above conditions are triggered, the orbit control strategy of the spacecraft needs to be updated.
If the track control mode includes: if the track of the sub-satellite point maintains the control mode, the step S106 determines the orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, and specifically includes the following steps:
step S401, acquiring a nominal track average speed, a nominal longitude, a preset longitude boundary value, a nominal half-major axis value and a half-major axis attenuation speed caused by atmospheric resistance.
Step S402, determining the semimajor axis deviation value based on the nominal longitude, the preset longitude boundary value, the nominal semimajor axis value and the semimajor axis attenuation speed.
In the method for maintaining and controlling the track of the sub-satellite point, before the orbit control strategy of the spacecraft is determined, the nominal orbit average speed V and the nominal longitude are acquired
Figure BDA0003617806840000163
Default longitude boundary value Δ λ max Nominal semi-major axis value
Figure BDA0003617806840000164
And half-major axis decay rate caused by atmospheric resistance
Figure BDA0003617806840000165
The known average angular velocity rate of change of the orbit is:
Figure BDA0003617806840000166
the change in longitude of the sub-satellite points for each pass through the same latitude is accumulated as:
Figure BDA0003617806840000167
wherein, Δ a 0 Represents the relative nominal semi-major axis value of the semi-major axis value a of the current track
Figure BDA0003617806840000168
I.e. the semimajor axis deviation. As can be seen from the above expression of Δ λ, Δ λ is a quadratic function of time, and there is an inflection point
Figure BDA0003617806840000171
And the inflection point t' corresponds to
Figure BDA0003617806840000172
Setting the semi-major axis of the control target as
Figure BDA0003617806840000173
Then the offset deltaa 0 Should be such that the inflection point of the sub-satellite longitude is exactly equal to
Figure BDA0003617806840000174
In this case the subsatellite point bias will drift back towards the nominal value after reaching the maximum value, forming a parabola. Namely, it is
Figure BDA0003617806840000175
Substitution into
Figure BDA0003617806840000176
The following can be obtained:
Figure BDA0003617806840000177
by solving the expression, the semimajor axis deviation value delta a can be obtained 0 。
And S403, determining a second adjustment quantity of the orbit semi-major axis based on the semi-major axis deviation quantity, the nominal semi-major axis value and the current orbit semi-major axis value of the spacecraft.
Obtaining the semimajor axis deviation amount Delta a 0 Then, using the equation
Figure BDA0003617806840000178
The second adjustment amount delta a of the semi-major axis of the track can be calculated 2 Wherein, Δ a 0 The amount of semi-major axis deviation is shown,
Figure BDA0003617806840000179
the nominal half-major axis value is represented, and a represents the current orbit half-major axis value of the spacecraft.
And S404, determining the second adjustment amount based on the semi-major axis of the track, the nominal semi-major axis value and the nominal track average speed.
And S405, determining that the fifth orbit control speed increment corresponds to the nominal semi-major axis value and the mean anomaly angle of the spacecraft.
After determining the second adjustment amount for the track semi-major axis, the method for determining the fifth track control speed increment and its corresponding fifth control time may refer to steps S1063-S1064, discussed above, that is,
Figure BDA00036178068400001710
if the mean-near point angle of the spacecraft is not more than pi, the formula is utilized
Figure BDA00036178068400001711
Calculating a fifth control moment; otherwise, using the equation
Figure BDA00036178068400001712
The fifth control timing is calculated.
In step S406, a fourth trajectory control strategy is determined based on the fifth control time and the fifth trajectory control speed increment.
Example two
The embodiment of the invention also provides a spacecraft orbit maneuver control device, which is mainly used for executing the spacecraft orbit maneuver control method provided by the first embodiment of the invention, and the spacecraft orbit maneuver control device provided by the embodiment of the invention is specifically introduced below.
Fig. 3 is a functional block diagram of a spacecraft orbit maneuver control device provided in an embodiment of the present invention, and as shown in fig. 3, the device mainly includes: the device comprises an acquisition module 10, a judgment module 20, a determination module 30 and an adjustment module 40, wherein:
the acquisition module 10 is configured to acquire an orbit control mode and an orbit number of the spacecraft; wherein the track control mode comprises at least one of: a semi-long shaft maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition.
And the judging module 20 is configured to judge whether the spacecraft needs to be subjected to orbit control based on the triggering condition of the orbit control mode and the number of the orbits.
The determining module 30 is configured to determine an orbit control strategy of the spacecraft based on an orbit control mode and the number of orbits under the condition that the spacecraft needs to be controlled in orbit; wherein the trajectory control strategy comprises: at least one control moment and a track control velocity increment corresponding to each control moment.
And the adjusting module 40 is used for adjusting the operating speed of the spacecraft based on the corresponding orbit control speed increment at the target control moment.
In the spacecraft orbit maneuver control method executed by the spacecraft orbit maneuver control device provided by the invention, the orbit control mode comprises at least one of the following modes: the method comprises a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity maintaining control mode and a satellite point track maintaining control mode, wherein each maintaining control mode has a corresponding trigger condition, in the orbit process of the spacecraft, whether the spacecraft needs to be subjected to orbit control or not and an orbit control strategy when the spacecraft needs to be controlled can be judged through the number of orbits and an actually selected orbit control mode, and then the running speed of the spacecraft is adjusted according to a corresponding orbit control speed increment in the target control moment. Therefore, the device can be used for self-adaptive adjustment of the spacecraft orbit, and compared with a method for controlling the spacecraft orbit by using control parameters which are injected in advance, the device can effectively solve the technical problem of poor spacecraft orbit maintaining accuracy existing in the conventional spacecraft orbit control method.
Optionally, the trigger condition of the semi-major axis maintaining control mode includes: the orbit flat semi-major axis of the spacecraft reaches a preset semi-major axis limit value; if the track control mode includes: the semi-major axis maintaining control mode, the determining module 30 includes:
and the first acquisition unit is used for acquiring the nominal semi-major axis value, the nominal track average speed and the maximum value of the track semi-major axis at the current moment.
And the first determining unit is used for determining a first adjustment amount of the track semimajor axis based on the maximum value of the track semimajor axis and a preset semimajor axis limit value.
A second determination unit to determine a first track control speed increment based on the first adjustment of the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed.
And the third determining unit is used for determining a first control moment corresponding to the first orbit control speed increment based on the mean anomaly angle and the nominal semi-major axis value of the spacecraft.
A fourth determination unit for determining the first trajectory control strategy based on the first control instant and the first trajectory control speed increment.
Optionally, the third determining unit is specifically configured to:
and judging whether the mean anomaly angle of the spacecraft is larger than 180 degrees.
If not, using the formula
Figure BDA0003617806840000191
Calculating a first control moment; wherein, t c1 Representing a first control moment, t representing the current moment, M representing the mean anomaly of the spacecraft,
Figure BDA0003617806840000192
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000193
the nominal half-length axis value is indicated.
If it is greater than, the formula is used
Figure BDA0003617806840000194
The first control timing is calculated.
Optionally, the triggering condition of the tilt angle maintaining control mode includes: the orbit inclination angle of the spacecraft exceeds a preset inclination angle range; if the track control mode includes: the tilt angle maintaining control mode, the determining module 30 includes:
and the second acquisition unit is used for acquiring a nominal inclination angle value, a nominal track average speed and a nominal half-length axis value.
A fifth determining unit for determining an orbital inclination adjustment amount based on the nominal inclination value and the current orbital inclination value.
A sixth determining unit for determining a second track control speed increment based on the track pitch adjustment and the nominal track average speed.
And the seventh determining unit is used for determining a second control moment corresponding to the second orbit control speed increment based on the nominal half-long axis value, the right ascension of the orbit intersection of the spacecraft and the mean-near-point angle of the spacecraft.
An eighth determining unit for determining the second trajectory control strategy based on the second control timing and the second trajectory control speed increment.
Optionally, the seventh determining unit is specifically configured to:
if the time when the spacecraft moves to the ascending intersection point is earlier than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure BDA0003617806840000201
Calculating a second control moment; wherein, t c2 Represents a second control moment, t represents the current moment, omega represents the orbit ascent point right ascension of the spacecraft, M represents the mean anomaly angle of the spacecraft,
Figure BDA0003617806840000202
represents the orbital angular velocity, mu represents the earth's gravitational constant,
Figure BDA0003617806840000203
the nominal half-length axis value is indicated.
If the time when the spacecraft moves to the ascending intersection point is later than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure BDA0003617806840000204
The second control timing is calculated.
Optionally, the triggering condition of the eccentricity maintenance control mode includes: the variation of the orbital eccentricity of the spacecraft exceeds a preset eccentricity variation limit value; if the track control mode includes: the eccentricity maintains the control mode, the determining module 30 includes:
a third obtaining unit for obtaining the nominal eccentricity and the nominal track average velocity.
A ninth determining unit for determining a variation of the orbital eccentricity of the spacecraft on the basis of the nominal eccentricity and the current orbital eccentricity.
A tenth determining unit for determining a third track control speed increment and a fourth track control speed increment based on the variation of the track eccentricity and the nominal track average speed.
And the eleventh determining unit is used for determining a third control moment corresponding to the third orbit control speed increment and a fourth control moment corresponding to the fourth orbit control speed increment based on the right ascension point of the orbit of the spacecraft, the mean-near point angle of the spacecraft and the variation of the orbit eccentricity.
And the twelfth determining unit is used for determining a third track control strategy based on the third control time, the third track control speed increment, the fourth control time and the fourth track control speed increment.
Optionally, the triggering condition of the sub-satellite trajectory maintenance control mode includes:
Figure BDA0003617806840000211
wherein λ represents the current longitude of the spacecraft,
Figure BDA0003617806840000212
nominal longitude representing the target latitude circle, N representing the number of intervals that the regression orbit trajectory net divides a latitude circle equally, Δ λ max Representing a preset longitude boundary value; if the track control mode includes: if the sub-satellite trajectory maintains the control mode, the determining module 30 includes:
and the fourth acquisition unit is used for acquiring the nominal orbit average speed, the nominal longitude, a preset longitude boundary value, a nominal semi-major axis value and the semi-major axis attenuation speed caused by the atmospheric resistance.
A thirteenth determining unit for determining the semimajor axis deviation amount based on the nominal longitude, a preset longitude boundary value, the nominal semimajor axis value, and the semimajor axis fading speed.
And the fourteenth determining unit is used for determining a second adjustment quantity of the orbit semi-major axis based on the semi-major axis deviation quantity, the nominal semi-major axis value and the current orbit semi-major axis value of the spacecraft.
A fifteenth determining unit for determining a fifth track control speed increment based on the second adjustment for the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed.
And the sixteenth determining unit is used for determining a fifth control moment corresponding to the fifth orbit control speed increment based on the mean anomaly angle and the nominal semi-major axis value of the spacecraft.
A seventeenth determining unit for determining a fourth trajectory control strategy based on the fifth control timing and the fifth trajectory control speed increment.
EXAMPLE III
Referring to fig. 4, an embodiment of the present invention provides an electronic device, including: a processor 60, a memory 61, a bus 62 and a communication interface 63, wherein the processor 60, the communication interface 63 and the memory 61 are connected through the bus 62; the processor 60 is arranged to execute executable modules, such as computer programs, stored in the memory 61.
The Memory 61 may include a high-speed Random Access Memory (RAM) and may also include a non-volatile Memory (non-volatile Memory), such as at least one disk Memory. The communication connection between the network element of the system and at least one other network element is realized through at least one communication interface 63 (which may be wired or wireless), and the internet, a wide area network, a local network, a metropolitan area network, and the like can be used.
The bus 62 may be an ISA bus, PCI bus, EISA bus, or the like. The bus may be divided into an address bus, a data bus, a control bus, etc. For ease of illustration, only one double-headed arrow is shown in FIG. 4, but that does not indicate only one bus or one type of bus.
The memory 61 is used for storing a program, the processor 60 executes the program after receiving an execution instruction, and the method executed by the apparatus defined by the process disclosed in any of the foregoing embodiments of the present invention may be applied to the processor 60, or implemented by the processor 60.
The processor 60 may be an integrated circuit chip having signal processing capabilities. In implementation, the steps of the above method may be performed by integrated logic circuits of hardware or instructions in the form of software in the processor 60. The Processor 60 may be a general-purpose Processor, including a Central Processing Unit (CPU), a Network Processor (NP), and the like; the Integrated Circuit may also be a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field-Programmable Gate Array (FPGA) or other Programmable logic device, discrete Gate or transistor logic device, or discrete hardware components. The various methods, steps and logic blocks disclosed in the embodiments of the present invention may be implemented or performed. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like. The steps of the method disclosed in connection with the embodiments of the present invention may be directly implemented by a hardware decoding processor, or implemented by a combination of hardware and software modules in the decoding processor. The software modules may be located in ram, flash, rom, prom, or eprom, registers, etc. as is well known in the art. The storage medium is located in a memory 61, and the processor 60 reads the information in the memory 61 and, in combination with its hardware, performs the steps of the above method.
The method, the apparatus, and the computer program product for controlling a spacecraft orbit maneuver provided in the embodiments of the present invention include a computer-readable storage medium storing a non-volatile program code executable by a processor, where instructions included in the program code may be used to execute the method described in the foregoing method embodiments, and specific implementation may refer to the method embodiments, and thus are not described herein again.
In addition, functional units in the embodiments of the present invention may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit.
The functions, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a non-volatile computer-readable storage medium executable by a processor. Based on such understanding, the technical solution of the present invention or a part thereof which substantially contributes to the prior art may be embodied in the form of a software product, which is stored in a storage medium and includes several instructions for causing a computer device (which may be a personal computer, a server, or a network device) to execute all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a magnetic disk, or an optical disk, and various media capable of storing program codes.
It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, it need not be further defined and explained in subsequent figures.
In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings or the orientations or positional relationships that the products of the present invention are conventionally placed in use, and are only used for convenience in describing the present invention and simplifying the description, but do not indicate or imply that the devices or elements referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," "third," and the like are used solely to distinguish one from another and are not to be construed as indicating or implying relative importance.
Furthermore, the terms "horizontal", "vertical", "overhang" and the like do not imply that the components are required to be absolutely horizontal or overhang, but may be slightly inclined. For example, "horizontal" merely means that the direction is more horizontal than "vertical" and does not mean that the structure must be perfectly horizontal, but may be slightly inclined.
In the description of the present invention, it should also be noted that, unless otherwise explicitly stated or limited, the terms "disposed," "mounted," "connected," and "connected" are to be construed broadly and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit the same; while the invention has been described in detail and with reference to the foregoing embodiments, it will be understood by those skilled in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present invention.

Claims (9)

1. A spacecraft orbital maneuver control method, comprising:
acquiring an orbit control mode and the number of orbits of a spacecraft; wherein the trajectory control mode comprises at least one of: a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition;
judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition of the orbit control mode and the number of the orbits;
if so, determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits; wherein the trajectory control strategy comprises: at least one control moment and a track control speed increment corresponding to each control moment;
adjusting the operating speed of the spacecraft based on the corresponding orbit control speed increment at a target control time;
wherein the trigger condition of the semi-major axis maintaining control mode comprises: the orbit flat semi-major axis of the spacecraft reaches a preset semi-major axis limit value;
if the track control mode comprises: if the semi-major axis maintains the control mode, determining the orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits comprises:
acquiring a nominal half-length axis value, a nominal track average speed and a maximum value of a track half-length axis at the current moment;
determining a first adjustment amount of the track semimajor axis based on the maximum value of the track semimajor axis and the preset semimajor axis limit value;
determining a first track control speed increment based on a first adjustment to the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed;
determining a first control moment corresponding to the first orbit control speed increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value;
a first trajectory control strategy is determined based on the first control time and the first trajectory control speed increment.
2. A spacecraft orbital maneuver control method according to claim 1, wherein the determining a first control instant corresponding to the first orbital control velocity increment based on the mean anomaly angle and the nominal semi-major axis value of the spacecraft comprises:
judging whether the mean anomaly angle of the spacecraft is larger than 180 degrees;
if not, using the formula
Figure FDA0003833859210000021
Calculating the first control moment; wherein, t c1 Representing the first control moment, t representing the current moment, M representing the mean anomaly of the spacecraft,
Figure FDA0003833859210000022
represents the orbital angular velocity, m represents the earth's gravitational constant,
Figure FDA0003833859210000023
representing the nominal half-major axis value;
if it is greater than the above value, the formula is used
Figure FDA0003833859210000024
And calculating the first control moment.
3. A spacecraft orbit maneuver control method according to claim 1, wherein the triggering conditions of the tilt angle maintenance control mode include: the orbit inclination angle of the spacecraft exceeds a preset inclination angle range;
if the track control mode comprises: and if the inclination angle is maintained in the control mode, determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, wherein the determining comprises the following steps:
acquiring a nominal inclination angle value, a nominal track average speed and a nominal half-major axis value;
determining an orbit inclination angle adjustment amount based on the nominal inclination angle value and the current orbit inclination angle value;
determining a second track control speed increment based on the track pitch adjustment and the nominal track average speed;
determining a second control moment corresponding to the second orbit control speed increment based on the nominal half-length axis value, the orbit ascension point of the spacecraft and the mean-anomaly angle of the spacecraft;
a second trajectory control strategy is determined based on the second control time and the second trajectory control speed increment.
4. A spacecraft orbit maneuver control method according to claim 3, wherein the determining a second control time instant corresponding to the second orbit control speed increment based on the nominal semi-major axis value, the spacecraft orbit ascension point and the spacecraft mean anomaly angle comprises:
if the time when the spacecraft moves to the ascending intersection point is earlier than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure FDA0003833859210000031
Calculating the second control moment; wherein, t c2 Representing the second control moment, t representing the current moment, ω representing the orbital ascent crossing right ascension of the spacecraft, M representing the mean-anomaly angle of the spacecraft,
Figure FDA0003833859210000032
represents the orbital angular velocity, m represents the earth's gravitational constant,
Figure FDA0003833859210000033
representing the nominal half-major axis value;
if the time when the spacecraft moves to the ascending intersection point is later than the time when the spacecraft moves to the descending intersection point, the formula is utilized
Figure FDA0003833859210000034
And calculating the second control moment.
5. A spacecraft orbit maneuver control method according to claim 1, wherein the triggering condition of the eccentricity maintenance control mode comprises: the variation of the orbital eccentricity of the spacecraft exceeds a preset eccentricity variation limit value;
if the track control mode comprises: an eccentricity maintaining control mode, and then determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits, including:
acquiring a nominal eccentricity and a nominal track average speed;
determining a variation in orbital eccentricity of the spacecraft based on the nominal eccentricity and a current orbital eccentricity;
determining a third track control speed increment and a fourth track control speed increment based on the amount of change in track eccentricity and the nominal track average speed;
determining a third control time corresponding to the third orbit control speed increment and a fourth control time corresponding to the fourth orbit control speed increment based on the orbit ascension point right ascension of the spacecraft, the mean-anomaly angle of the spacecraft and the variation of the orbit eccentricity;
determining a third trajectory control strategy based on the third control time, the third trajectory control speed increment, the fourth control time, and the fourth trajectory control speed increment.
6. A spacecraft orbit maneuver control method according to claim 1, wherein the triggering conditions of the substellar point trajectory maintenance control mode include:
Figure FDA0003833859210000041
wherein λ represents a current longitude of the spacecraft,
Figure FDA0003833859210000042
nominal longitude representing the target latitude circle, N representing the number of intervals that the regression orbit trajectory net divides a latitude circle equally, Δ λ max Representing a preset longitude boundary value;
if the track control mode comprises: if the track of the sub-satellite points maintains the control mode, determining the orbit control strategy of the spacecraft based on the orbit control mode and the number of the orbits comprises the following steps:
acquiring a nominal orbit average speed, the nominal longitude, the preset longitude boundary value, a nominal semi-major axis value and a semi-major axis attenuation speed caused by atmospheric resistance;
determining a semi-major axis deviation amount based on the nominal longitude, the preset longitude boundary value, the nominal semi-major axis value, and the semi-major axis decay rate;
determining a second adjustment quantity of the orbit semi-major axis based on the semi-major axis deviation quantity, the nominal semi-major axis value and the current orbit semi-major axis value of the spacecraft;
determining a fifth track control speed increment based on a second adjustment to the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed;
determining a fifth control moment corresponding to the fifth orbit control speed increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value;
determining a fourth trajectory control strategy based on the fifth control time and the fifth trajectory control speed increment.
7. A spacecraft orbital maneuver control device, comprising:
the acquisition module is used for acquiring the orbit control mode and the number of orbits of the spacecraft; wherein the trajectory control mode comprises at least one of: a semi-major axis maintaining control mode, an inclination angle maintaining control mode, an eccentricity rate maintaining control mode and a sub-satellite point track maintaining control mode; each maintenance control mode has a corresponding trigger condition;
the judging module is used for judging whether the spacecraft needs to be subjected to orbit control or not based on the triggering condition of the orbit control mode and the number of the orbits;
the determining module is used for determining an orbit control strategy of the spacecraft based on the orbit control mode and the number of orbits under the condition that the spacecraft needs to be subjected to orbit control; wherein the trajectory control strategy comprises: at least one control time and a track control speed increment corresponding to each control time;
the adjusting module is used for adjusting the running speed of the spacecraft based on the corresponding track control speed increment at the target control moment;
wherein the trigger condition of the semi-major axis maintaining control mode comprises: the orbit flat semi-major axis of the spacecraft reaches a preset semi-major axis limit value;
if the track control mode comprises: and the semi-major axis maintaining control mode, the determining module comprises:
the first acquisition unit is used for acquiring a nominal semi-major axis value, a nominal track average speed and a maximum value of a track semi-major axis at the current moment;
a first determining unit, configured to determine a first adjustment amount of the track semi-major axis based on the maximum value of the track semi-major axis and the preset semi-major axis limit;
a second determining unit for determining a first track control speed increment based on the first adjustment of the track semi-major axis, the nominal semi-major axis value, and the nominal track average speed;
a third determining unit, configured to determine a first control time corresponding to the first orbit control speed increment based on the mean anomaly angle of the spacecraft and the nominal semi-major axis value;
a fourth determination unit for determining a first trajectory control strategy based on the first control instant and the first trajectory control speed increment.
8. An electronic device comprising a memory, a processor, and a computer program stored on the memory and executable on the processor, wherein the processor implements the steps of the method of any of claims 1 to 6 when executing the computer program.
9. A computer-readable medium having non-volatile program code executable by a processor, the program code causing the processor to perform the method of any of claims 1 to 6.
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