CN111959828B

CN111959828B - Spacecraft orbit maneuver detection method and device based on nonlinear deviation evolution

Info

Publication number: CN111959828B
Application number: CN202011128409.3A
Authority: CN
Inventors: 杨震; 罗亚中; 舒鹏; 张进; 李泽越
Original assignee: National University of Defense Technology
Current assignee: National University of Defense Technology
Priority date: 2020-10-21
Filing date: 2020-10-21
Publication date: 2020-12-29
Anticipated expiration: 2040-10-21
Also published as: CN111959828A

Abstract

The application relates to a spacecraft orbit maneuver detection method and device based on nonlinear deviation evolution. The method comprises the following steps: calculating the statistic value vector of the measuring orbit determination state of the target spacecraft at the two moments before and after the orbit measurement data, predicting the state vector of the target spacecraft at the later measurement moment according to the previous measurement data, and comparing the predicted state vector with the statistic value vector of the measuring orbit determination state at the later measurement moment. When the deviation is larger than a preset value, through nonlinear deviation evolution, forward prediction is carried out based on the measurement orbit determination state statistic vector at the previous moment, backward prediction is carried out based on the measurement orbit determination state statistic vector at the next moment, two orbit determination state statistic vectors are obtained at each prediction moment, and when the deviation of the two measured data at a certain moment is smaller than the preset value, the two measured data are judged to correspond to the same target spacecraft before and after maneuvering. By adopting the method, the target spacecraft data obtained by two times of detection can be matched, and the orbital maneuver condition of the target spacecraft can be identified in near real time.

Description

Spacecraft orbit maneuver detection method and device based on nonlinear deviation evolution

Technical Field

The application relates to the field of spatial target situation awareness and orbit anomaly detection, in particular to a spacecraft orbit maneuver detection method and device based on nonlinear deviation evolution.

Background

With the increasing frequency of space activities, the number of targets in the on-orbit space is increasing. The trajectory of these in-orbit space objects is not constant for various reasons. For example, the spacecraft in a normal working state has an orbit control capability, and in order to complete a corresponding task, an orbit maneuver needs to be performed to maintain an orbit or meet a specific target. In another example, the low orbit spacecraft has no capability of orbit mobility after failure, and the orbit height can be continuously attenuated. In addition, various space events including spacecraft maneuvering, on-orbit collisions, disintegration, abrupt changes in the space environment, etc., can cause orbital anomalies in space targets. In order to monitor the operating state of the spacecraft and ensure that the spacecraft operates safely without interference, the orbit detection of the concerned target spacecraft is required, and particularly, the orbit change of the target spacecraft is found, so that on one hand, whether the spacecraft is in a normal working state or not can be judged, on the other hand, whether the target spacecraft is maneuvered or not can be judged, and the intention and the purpose of the maneuvering of the target spacecraft can be further judged according to the change.

For non-cooperative spacecraft, the current approach is to perform continuous orbit detection. Firstly, an A target is determined according to the first track measuring data, a B target is determined according to the latter track measuring data, if the track maneuvering is not executed by the targets between the first track measuring data and the second track measuring data, A, B two targets can be matched and associated to be the same target, otherwise, the B target is temporarily cataloged to be a new target. The existing track maneuver detection method is mainly a post maneuver detection method based on historical track data, and comprises a moving window curve fitting method, a track forecast error fitting method and a cluster analysis method.

Disclosure of Invention

Based on this, it is necessary to provide a spacecraft orbit maneuver detection method and device based on nonlinear deviation evolution, which can match the target spacecraft obtained by two detections and realize the in-orbit spacecraft near-real-time orbit maneuver abnormity warning, for solving the technical problems.

A method for spacecraft orbital maneuver detection based on nonlinear bias evolution, the method comprising:

and according to the measured orbit determination state statistic value vectors of the target spacecraft at the 0 th moment and the n th moment. The measuring of the orbit state statistic vector comprises measuring of an orbit state mean vector and a state covariance matrix, and the orbit state mean vector comprises a position statistic component and a velocity statistic component.

And according to the measured orbit determination state statistic value vector at the 0 th moment, obtaining a corresponding forward prediction orbit determination state statistic value vector at the nth moment through nonlinear deviation evolution.

When the deviation between the forward predicted orbit determination state statistic vector at the nth moment and the measured orbit determination state statistic vector at the nth moment is larger than a preset value, obtaining the forward predicted orbit determination state statistic vector at the ith moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the 0 th moment, and obtaining the backward predicted orbit determination state statistic vector at the ith moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the nth moment, wherein i is more than or equal to 0 and less than n.

And when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is smaller than a preset value, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as maneuvering.

In one embodiment, the step of obtaining the corresponding forward prediction tracking state statistic vector at the nth time through nonlinear bias evolution according to the measured tracking state statistic vector at the 0 th time further includes:

and when the deviation between the forward prediction orbit determination state statistic value vector at the nth moment and the measurement orbit determination state statistic value vector at the nth moment is smaller than a preset value, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as the unmoved state.

In one embodiment, when the deviation between the forward predicted tracking state statistic vector at the nth time and the measured tracking state statistic vector at the nth time is greater than a preset value, the step of obtaining the forward predicted tracking state statistic vector at the ith time through nonlinear deviation evolution according to the measured tracking state statistic vector at the 0 th time, and the step of obtaining the backward predicted tracking state statistic vector at the ith time through nonlinear deviation evolution according to the measured tracking state statistic vector at the nth time comprises:

and calculating the forward prediction orbit determination state mean vector and the state covariance matrix at the nth moment, and calculating the measurement orbit determination state mean vector and the state covariance matrix at the nth moment.

And obtaining the Mahalanobis distance of the position mean component of the forward predicted orbit determination state mean vector and the measured orbit determination state mean vector at the nth time through nonlinear covariance analysis according to the forward predicted orbit determination state mean vector, the measured orbit determination state mean vector and the corresponding covariance at the nth time.

And when the Mahalanobis distance is larger than the preset value, obtaining a forward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the 0 th moment, and obtaining a backward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the nth moment.

In one embodiment, when a deviation between the backward predicted orbit determination state statistic vector at the ith time and the forward predicted orbit determination state statistic vector at the ith time is smaller than a preset value, the step of associating the measured orbit determination state statistic vectors at the 0 th time and the nth time to the same target spacecraft and setting the state value of the corresponding target spacecraft to be maneuvering further includes:

and acquiring a state value which is the minimum deviation value between a backward prediction orbit determination state statistic vector and a forward prediction orbit determination state statistic vector of the maneuvering target spacecraft at the ith moment, and setting the moment corresponding to the minimum deviation value as maneuvering time corresponding to the target spacecraft.

In one embodiment, after the step of obtaining a minimum deviation value between a backward prediction orbit determination state statistic vector and a forward prediction orbit determination state statistic vector of a maneuvering target spacecraft at the ith time and setting the time corresponding to the minimum deviation value as the maneuvering time corresponding to the target spacecraft, the method further includes:

and calculating the maneuvering speed increment of the corresponding target spacecraft according to the speed statistic component difference between the backward prediction orbit determination state statistic vector and the forward prediction orbit determination state statistic vector of the maneuvering time.

In one embodiment, when a deviation between the forward predicted tracking state statistic vector at the nth time and the measured tracking state statistic vector at the nth time is greater than a preset value, the step of obtaining the forward predicted tracking state statistic vector at the ith time through nonlinear deviation evolution according to the measured tracking state statistic vector at the 0 th time, and obtaining the backward predicted tracking state statistic vector at the ith time through nonlinear deviation evolution according to the measured tracking state statistic vector at the nth time further includes:

and when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is larger than a preset value, respectively associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to different target spacecrafts.

In one embodiment, n +1 time points from the 0 th time point to the n th time point are uniformly distributed, and the interval between adjacent time points is less than 10 seconds.

A spacecraft orbit maneuver detection apparatus based on nonlinear bias evolution, the apparatus comprising:

and the measurement orbit determination state statistic value vector calculation module is used for calculating the measurement orbit determination state statistic value vector of the target spacecraft at the 0 th moment and the n th moment. The measuring of the orbit state statistic vector comprises measuring of an orbit state mean vector and a state covariance matrix, and the orbit state mean vector comprises a position statistic component and a velocity statistic component.

And the forward orbit determination state forecasting module is used for obtaining a corresponding forward prediction orbit determination state statistic value vector at the nth time through nonlinear deviation evolution according to the measurement orbit determination state statistic value vector at the 0 th time.

And the bidirectional orbit determination state forecasting module is used for obtaining a forward predicted orbit determination state statistic vector at the ith moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the 0 th moment and obtaining a backward predicted orbit determination state statistic vector at the ith moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the nth moment, wherein i is more than or equal to 0 and less than n.

And the target spacecraft maneuvering detection module is used for associating the measured orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft and setting the state value of the corresponding target spacecraft as maneuvering when the deviation between the backward predicted orbit determination state statistic value vector at the ith moment and the forward predicted orbit determination state statistic value vector at the ith moment is smaller than a preset value.

A computer device comprising a memory and a processor, the memory storing a computer program, the processor implementing the following steps when executing the computer program:

A computer-readable storage medium, on which a computer program is stored which, when executed by a processor, carries out the steps of:

According to the spacecraft orbit maneuver detection method, the spacecraft orbit maneuver detection device, the computer equipment and the storage medium based on the nonlinear deviation evolution, the forward predicted orbit determination state statistic vector at the later measurement moment is predicted according to the previous measurement data, the forward predicted orbit determination state statistic vector is compared with the measured orbit determination state statistic vector at the later measurement moment, and when the deviation between the two is larger than a preset value, whether the same target spacecraft which is subjected to orbit change or different target spacecrafts corresponds to the two measurement data is continuously judged. Through nonlinear deviation evolution, forward prediction is carried out on the basis of the measured orbit determination state statistic vector at the previous moment at the preset prediction moment, backward prediction is carried out on the basis of the measured orbit determination state statistic vector at the next moment, two orbit determination state statistic vectors are obtained at each prediction moment, and when the deviation of the two measured data at a certain prediction moment is smaller than a preset value, the same target spacecraft which is subjected to orbit transfer is judged to correspond to the measured data twice. The method and the device can match the target spacecraft data obtained by two times of detection, and recognize the orbital maneuver condition of the target spacecraft in near real time.

Drawings

FIG. 1 is a diagram of an application scenario of a spacecraft orbit maneuver detection method based on nonlinear deviation evolution in an embodiment;

FIG. 2 is a diagram of the steps of a method for spacecraft orbital maneuver detection based on nonlinear bias evolution, under an embodiment;

FIG. 3 is a schematic flow chart of a method for detecting spacecraft orbital maneuver based on nonlinear bias evolution in another embodiment;

FIG. 4 is a diagram of a data distribution of a forward predicted tracking state statistic vector and a backward predicted tracking state statistic vector in one embodiment;

FIG. 5 is a diagram illustrating an internal structure of a computer device according to an embodiment.

Detailed Description

In order to make the objects, technical solutions and advantages of the present application more apparent, the present application is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the present application and are not intended to limit the present application.

The spacecraft orbit maneuver detection method based on the nonlinear deviation evolution can be applied to the application environment shown in figure 1. Wherein the spacecraft orbit detection device 102 communicates with the space monitoring device 104 over a network. The spacecraft orbit detection equipment 102 and the space monitoring equipment 104 may be implemented by a single server or a server cluster composed of a plurality of servers, and may be, but not limited to, various personal computers, notebook computers, and the like which are sufficient to provide the computing power and storage space required by the method.

In one embodiment, as shown in fig. 2, a method for detecting spacecraft orbit maneuver based on nonlinear deviation evolution is provided, which is illustrated by applying the method to the spacecraft orbit detection device 102 in fig. 1, and includes the following steps:

and 202, according to the measurement orbit determination state statistic value vectors of the target spacecraft at the 0 th moment and the n th moment. The measuring of the orbit state statistic vector comprises measuring of an orbit state mean vector and a state covariance matrix, and the orbit state mean vector comprises a position statistic component and a velocity statistic component.

Specifically, according to orbit determination data of the ground surface to the target spacecraft, measurement orbit determination state statistic vectors of the target spacecraft at the current time (set as the nth time) and the previous time (set as the 0 th time) are obtained, and the measurement orbit determination state statistic vectors can be represented under the geocentric inertial system and comprise actual measurement position coordinates and speed information of the target spacecraft at the corresponding time. Each component in the vector is the statistic value of the position component and the velocity component of the target spacecraft,

and 204, obtaining a corresponding forward prediction orbit determination state statistic vector at the nth time through nonlinear deviation evolution according to the measurement orbit determination state statistic vector at the 0 th time.

And step 206, when the deviation between the forward predicted orbit determination state statistic vector at the nth time and the measured orbit determination state statistic vector at the nth time is larger than a preset value, obtaining the forward predicted orbit determination state statistic vector at the ith time through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the 0 th time, and obtaining the backward predicted orbit determination state statistic vector at the ith time through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the nth time, wherein i is more than or equal to 0 and less than n.

And predicting a forward predicted orbit determination state vector of the corresponding target spacecraft at the current moment according to the measured orbit determination state statistic vector at the previous moment, comparing the predicted orbit determination state vector with the measured orbit determination state statistic vector at the current moment, and continuously judging whether the two measured data correspond to the same target spacecraft which is subjected to the orbit change or not when the deviation between the predicted orbit determination state vector and the measured orbit determination state statistic vector is greater than a preset value. The specific way of judging is as follows: through nonlinear deviation evolution, forward prediction is carried out based on the measured orbit determination state statistic vector at the previous moment, backward prediction is carried out based on the measured orbit determination state statistic vector at the next moment, and two predicted orbit determination state statistic vectors are obtained at the preset prediction moment. And respectively comparing the two corresponding predicted orbit determination state vector values at the previous measuring time and the preset predicting time.

And 208, when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is smaller than a preset value, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as maneuvering.

When the deviation of the two predicted orbit determination state vectors at a certain moment is smaller than a preset value, the situation that the same target spacecraft which is subjected to orbit change corresponds to the two measurement data at the previous moment and the current moment is judged.

The spacecraft orbit maneuver detection method based on the nonlinear deviation evolution carries out forward and reverse cross arc prediction on twice orbit measurement data based on the nonlinear deviation evolution, matches target spacecrafts obtained by twice detection, can identify the target spacecrafts subjected to orbit change in near real time, and sends out corresponding abnormal alarm information on the basis.

In one embodiment, as shown in fig. 3, a method for detecting spacecraft orbital maneuver based on nonlinear bias evolution is provided, which includes the following steps:

step 302, respectively calculating according to the orbit measurement data of the target spacecraft

Time of day and

and measuring orbit determination state mean vector and state covariance matrix of the target spacecraft at the moment. The components of the measured orbit state mean vector include a position mean component and a velocity mean component.

In particular, at the current moment

At the nth time, in

The time is the 0 th time and is set as the time

The target spacecraft corresponding to the orbit measurement data is a target B, and the time is set

The target spacecraft corresponding to the orbit measurement data is a target A. Obtaining the current time according to the ground track measurement data

Mean value of orbit determination state of target B under geocentric inertial system

And covariance matrix

Wherein

For the state location vector component of the target spacecraft,

is a component of the spacecraft state velocity vector,

is shown in

Constructing a square matrix for diagonal lines, wherein off-diagonal elements of the square matrix are all 0,

，

for corresponding orbit state component

Standard deviation of (2). In the same way, obtain

Mean orbit determination state of time target A

And covariance matrix

。

Step 304, according to

Measuring orbit determination state mean vector and state covariance matrix at moment, and obtaining corresponding state through nonlinear deviation evolution

Forward prediction orbit determination state mean vector and state covariance matrix at a time.

Specifically, the target A orbit state mean value

And covariance matrix

Forecast to

Time of day, get

Time target A state mean

And covariance matrix

In the position space calculation

Mean orbit determination state of time target B

Mahalanobis distance to a state distribution

Wherein

Is a B target

The time-of-day position vector component,

is an A target

The time-of-day position vector component,

is composed of

Is determined by the symmetric positive definite matrix of (a),

is composed of

Formed by the first 3 rows and the first 3 columns

A symmetric positive definite matrix.

Step 306, calculate and

forward predicted orbit determination state statistics vector sum of time

The Mahalanobis distance k of the position mean value component of the statistic value vector of the orbit determination state at the moment is measured, and when the Mahalanobis distance is smaller than a preset value, the Mahalanobis distance k is used for measuring the mean value component of the position mean value component of the orbit determination state statistic value vector of the orbit determination state at the moment

Time of day and

and the measurement orbit determination state statistic value vector at the moment is associated to the same target spacecraft, and the state value of the corresponding target spacecraft is set as the unmoved state.

Specifically, the preset mahalanobis distance is

When is coming into contact with

When the target A and the target B are the same target, the target A is judged to be not in track abnormality, and the state value of the target A is set as non-maneuvering.

Step 308, when the Mahalanobis distance is larger than the preset value, according to the second step

Measuring the mean vector and the state covariance matrix of the orbit determination state at the moment, obtaining the mean vector and the state covariance matrix of the forward prediction orbit determination state at the ith moment through nonlinear covariance analysis, and obtaining the mean vector and the state covariance matrix of the forward prediction orbit determination state at the ith moment according to the ith moment

Measuring orbit determination state mean vector and state covariance matrix at moment, and obtaining ith moment through nonlinear covariance analysisAnd (3) backward predicting the orbit determination state mean vector and the state covariance matrix, wherein i is more than or equal to 0 and is less than n.

In particular, when

Then, it is determined that target a and target B may be: the same target spacecraft which is motorized or different target spacecraft needs to be further judged. The specific judgment method comprises the following steps:

mean orbit determination state of target A

And covariance matrix

According to step length

Forecast to

Time of day, get

A moment of time

Target A orbital state mean

And covariance matrix

The forecast data of (1). Similarly, the mean value of the orbit determination state of the target B

And covariance matrix

According to step length

Forecast back to

Time of day, get

A moment of time

Target B orbital state mean

And covariance matrix

The forecast data of (1).

Target A, B at

A moment of time

State mean of

、

And covariance matrix

、

Merging to obtain the product with A as reference point,

A moment of time

Relative state of

And total covariance matrix

And satisfies the following conditions:

in position space calculation

Mahalanobis distance of orbit determination state mean value of target B at each moment relative to orbit determination state mean value of target A in combined total covariance matrix

Namely:

wherein the content of the first and second substances,

，

the first 3 components of the first group of components,

is composed of

A symmetrical positive definite matrix composed of the first 3 rows and the first 3 columns.

Further, from the second

Time to

N +1 of the moments are evenly distributed, the interval between adjacent moments (i.e. step size)

) Less than 10 seconds.

310, when the Mahalanobis distance between the backward predicted orbit determination state average value vector at the ith moment and the position average value component of the forward predicted orbit determination state average value vector at the ith moment is larger than the preset value, the Mahalanobis distance between the backward predicted orbit determination state average value vector at the ith moment and the position average value component of the forward predicted orbit determination state average value vector at the ith moment is used for calculating

Time of day and

and respectively associating the measured orbit determination state mean value vectors at the moment to different target spacecrafts.

If for all

Are all provided with

Then target A and target B are different target spacecrafts

Time of day and

and the measurement orbit determination state mean value vector at the moment and the corresponding orbit measurement data are associated to different target spacecrafts.

Step 312, when the deviation between the backward predicted orbit determination state statistic vector at the ith time and the forward predicted orbit determination state statistic vector at the ith time is less than the preset value, the second step is executed

Time of day and

and the measurement orbit determination state mean value vector at the moment is associated to the same target spacecraft, and the state value of the corresponding target spacecraft is set as maneuvering.

If for one

Is provided with

Target A and target B are the same target spacecraft which has undergone orbital transfer

Time of day and

and the measurement orbit determination state mean vector at the moment and the corresponding orbit measurement data are associated to a target A, and the state value of the target A is set as maneuvering.

And step 314, acquiring the minimum value of the Mahalanobis distance of the position components of the backward prediction orbit determination state mean vector and the forward prediction orbit determination state mean vector of the maneuvering target spacecraft at the ith moment, and setting the moment when the minimum value appears as the maneuvering time of the corresponding target spacecraft.

And step 316, calculating the maneuvering speed increment of the corresponding target spacecraft according to the difference value of the speed statistic value component between the backward prediction orbit determination state mean value vector and the forward prediction orbit determination state mean value vector of the maneuvering time.

Specifically, for the object A whose state value is maneuvering, at which

Minimum value is extracted by searching in Mahalanobis distance

And the time at which the minimum occurs

. Is provided with

，

Then, then

The moment is the maneuvering time of the target A for maneuvering, and the speed increment of the target A

And amount of mobility

The size is as follows:

wherein the content of the first and second substances,

is composed of

The last three (4 th to 6 th) components,

is composed of

The last three (4 th to 6 th) components of (a).

The spacecraft orbit maneuver detection method based on the nonlinear deviation evolution provided by the embodiment can judge whether the target spacecraft is maneuvered in near real time, and can judge the maneuvering time, the speed increment and the maneuvering amount of the maneuvered target spacecraft, so that abundant target state information is provided.

In one embodiment, a spacecraft orbit maneuver detection method based on nonlinear deviation evolution is provided, and comprises the following steps:

step 402, respectively calculating according to the orbit measurement data of the target spacecraft

Time of day and

Specifically, in

The time of day orbital measurement data is shown in table 1. Setting the current fixed rail

At the moment of time of

=86400s, spacecraft in

=43200s, the maneuvering impulse under the local orbit coordinate LVLH system of the spacecraft (the origin o is in the center of mass of the spacecraft, ox is along the radial direction of the center of the spacecraft, oz is along the normal direction of the orbit surface, oy forms a right-hand system) is

m/s, the magnitude of the maneuvering quantity is

Setting mahalanobis distance threshold for rail maneuvering detection

=4, the forecast time step number of the crossed arc track is set as n =8640, and the forecast time step length is determined

The orbit forecast uses a two-body model.

TABLE 1 initial number of orbits of in-orbit spacecraft

Semi-major axis/m	Eccentricity ratio	Orbital inclination angle/°	Ascending crossing point Chin Jing/°	Angular distance between near points/° c	True angle/degree of approach
						7181727.864	0.0005	45	50	60	30

The orbit is determined according to the ground, and the current state of the in-orbit spacecraft after the maneuver can be calculated by the data in the table 1

The mean orbit determination state of the time target B under the geocentric inertial system is as follows:

covariance matrix

Wherein

For the spacecraft state position vector components,

is a component of the spacecraft state velocity vector,

is shown in

，

. Similarly, the previous t0 of the storage record is set (t 0)<tf) the mean value of the orbit determination state of the target A at the moment is:

covariance matrix

，

，

，

。

Step 404, according to

In particular, obtained from a measuring rail in the centroidal inertial system

Mean orbit determination state of time target A

And covariance matrix

Generating (2 n + 1) sigma sample points with certain weights

I.e. by

Wherein each sample point corresponds to a weight of

Wherein the content of the first and second substances,

，

and

for free parameters, the suggested value is

，

. For a gaussian distribution, the distribution of the power,

。

is a covariance matrix

The square root of (i), i.e.

；

Is composed of

Column i.

Forecasting all initial sigma points to the terminal by using a given orbit forecasting algorithm

Time of day, the forecasting process using non-linear mapping

Expressed, the sigma sample point of the terminal can be expressed as

，i = 0, 1, …, 12。

Computing A target by using terminal sigma sample point

State mean of time

And covariance matrix

Namely:

in position space calculation

Mean orbit determination state of time target B

Mahalanobis distance to a state distribution, i.e.:

it is obvious that

A and B are different targets or the same target that has been motorized.

Step 406, when the Mahalanobis distance is greater than the preset value, according to the second step

And measuring the mean vector and the state covariance matrix of the orbit determination state at the moment, and obtaining the mean vector and the state covariance matrix of the backward prediction orbit determination state at the ith moment through nonlinear covariance analysis, wherein i is more than or equal to 0 and is less than n.

Specifically, the target A orbit state mean value

And covariance matrix

According to step length

Forecast to

Time of day, get

A moment of time

Target A orbital state mean

And covariance matrix

And covariance matrix

According to step length

Forecast back to

Time of day, get

A moment of time

Target B orbital state mean

And covariance matrix

The forecast data of (1).

Step 408, at

Minimum value is extracted by searching in Mahalanobis distance

And the time at which the minimum occurs

Comparison of

And

and (4) judging the maneuvering state of the target spacecraft. And judging the maneuvering time, the speed increment and the maneuvering quantity of the maneuvering target.

Specifically, in

Minimum value is extracted by searching in Mahalanobis distance

And the time at which the minimum occurs

Is calculated to

，

(ii) a It is obvious that

A and B are the same object of implementing the maneuver, A is

A maneuver is performed at a time.

For target A maneuver size

An estimation is performed. Obtaining target A maneuver time

Then, it must satisfy

Is provided with

Is calculated to

Then can obtain

Time of day, velocity increment of object A

And amount of mobility

The size is as follows:

wherein

Is composed of

The last three (4 th to 6 th) components,

is composed of

The last three (4 th to 6 th) components of (a). The maneuvering impulse under the spacecraft local orbit coordinate LVLH system is calculated to be

The magnitude of the maneuvering quantity is

。

In the method for detecting orbital maneuver of spacecraft based on nonlinear deviation evolution provided by this embodiment, when a computer program implementing the method is run on a notebook computer of intel core i7-5500U CPU @2.4GHz, the orbital maneuver anomaly detection of the in-orbit spacecraft can be realized only in 16 seconds, the orbital maneuver of the spacecraft is successfully detected based on the orbit determination data of the current time and the previous time, and the relative error of the maneuver time estimation is

The relative error of the magnitude estimate of the maneuver is

It can be seen that the invention has higher calculation accuracy and efficiency, and has low requirements for calculation resources. In the embodiment, the two times of orbit determination data are distributed and forecasted to the maneuvering time of the orbit, and the distribution result is obtainedAs shown in fig. 4, marked with an "x" symbol are state distribution points of the target a obtained on the basis of the orbit determination state mean vector and the state covariance matrix obtained by forward prediction of the target a; the symbol of good quality indicates the state distribution point of the target B obtained on the basis of the orbit determination state mean vector and the state covariance matrix obtained by backward prediction of the target B. Obviously, the distribution of the two groups of orbit determination state data at the maneuvering moment is obviously overlapped, which indicates that the spacecraft executes maneuvering at the maneuvering moment. In addition, the stepless transformation theory is adopted for forward and backward orbit deviation propagation, forward and backward high-precision nonlinear prediction of the orbit determination state and the covariance matrix thereof can be realized only by 26 sigma sample points, the calculation efficiency is high, and the prediction process of the stepless transformation theory treats the power system as a black box and is suitable for any high-precision orbit dynamics model.

It should be understood that although the various steps in the flow charts of fig. 2-3 are shown in order as indicated by the arrows, the steps are not necessarily performed in order as indicated by the arrows. The steps are not performed in the exact order shown and described, and may be performed in other orders, unless explicitly stated otherwise. Moreover, at least some of the steps in fig. 2-3 may include multiple sub-steps or multiple stages that are not necessarily performed at the same time, but may be performed at different times, and the order of performance of the sub-steps or stages is not necessarily sequential, but may be performed in turn or alternating with other steps or at least some of the sub-steps or stages of other steps.

In one embodiment, there is provided a spacecraft orbital maneuver detection apparatus based on nonlinear bias evolution, comprising:

In one embodiment, the system further includes an unmoving target spacecraft identification module, configured to associate the measured orbit state statistic vector at the 0 th time and the measured orbit state statistic vector at the nth time with the same target spacecraft when a deviation between the forward predicted orbit state statistic vector at the nth time and the measured orbit state statistic vector at the nth time is smaller than a preset value, and set a state value of the corresponding target spacecraft to be unmoving.

In one embodiment, the bidirectional orbit determination state prediction module is configured to calculate a forward prediction orbit determination state mean vector and a state covariance matrix at a time n, and calculate a measurement orbit determination state mean vector and a state covariance matrix at the time n. And obtaining the Mahalanobis distance of the position mean component of the forward predicted orbit determination state mean vector and the measured orbit determination state mean vector at the nth time through nonlinear covariance analysis according to the forward predicted orbit determination state mean vector, the measured orbit determination state mean vector and the corresponding covariance at the nth time. And when the Mahalanobis distance is larger than the preset value, obtaining a forward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the 0 th moment, and obtaining a backward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the nth moment.

In one embodiment, the system further includes a maneuvering time obtaining module, configured to obtain a minimum deviation value between a backward prediction orbit determination state statistic vector and a forward prediction orbit determination state statistic vector of the maneuverable target spacecraft at the ith time, and set a time corresponding to the minimum deviation value as the maneuvering time corresponding to the target spacecraft.

In one embodiment, the system further comprises a maneuvering speed increment calculation module, configured to calculate a maneuvering speed increment of the corresponding target spacecraft according to a speed statistic component difference between the backward predicted orbit state statistic vector and the forward predicted orbit state statistic vector of the maneuvering time.

In one embodiment, the system further comprises different target spacecraft identification modules, and the different target spacecraft identification modules are used for respectively associating the measured orbit determination state statistic value vectors at the 0 th moment and the n th moment to different target spacecraft when the deviation between the backward predicted orbit determination state statistic value vector at the ith moment and the forward predicted orbit determination state statistic value vector at the ith moment is greater than a preset value.

For specific limitations of the spacecraft orbit maneuver detection device based on the nonlinear deviation evolution, reference may be made to the above limitations of the spacecraft orbit maneuver detection method based on the nonlinear deviation evolution, and details are not repeated here. The modules in the spacecraft orbit maneuver detection device based on the nonlinear deviation evolution can be wholly or partially realized by software, hardware and a combination thereof. The modules can be embedded in a hardware form or independent from a processor in the computer device, and can also be stored in a memory in the computer device in a software form, so that the processor can call and execute operations corresponding to the modules.

In one embodiment, a computer device is provided, which may be a server, the internal structure of which may be as shown in fig. 5. The computer device includes a processor, a memory, a network interface, and a database connected by a system bus. Wherein the processor of the computer device is configured to provide computing and control capabilities. The memory of the computer device comprises a nonvolatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program, and a database. The internal memory provides an environment for the operation of an operating system and computer programs in the non-volatile storage medium. The database of the computer equipment is used for storing target spacecraft orbit measurement data and implementing data generated by a spacecraft orbit maneuver detection method based on nonlinear deviation evolution. The network interface of the computer device is used for communicating with an external terminal through a network connection. The computer program is executed by a processor to implement a method for spacecraft orbital maneuver detection based on nonlinear bias evolution.

Those skilled in the art will appreciate that the architecture shown in fig. 5 is merely a block diagram of some of the structures associated with the disclosed aspects and is not intended to limit the computing devices to which the disclosed aspects apply, as particular computing devices may include more or less components than those shown, or may combine certain components, or have a different arrangement of components.

In one embodiment, there is provided a computer device comprising a memory storing a computer program and a processor implementing the following steps when the processor executes the computer program:

In one embodiment, the processor, when executing the computer program, further performs the steps of: and when the deviation between the forward prediction orbit determination state statistic value vector at the nth moment and the measurement orbit determination state statistic value vector at the nth moment is smaller than a preset value, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as the unmoved state.

In one embodiment, the processor, when executing the computer program, further performs the steps of: and calculating the forward prediction orbit determination state mean vector and the state covariance matrix at the nth moment, and calculating the measurement orbit determination state mean vector and the state covariance matrix at the nth moment. And obtaining the Mahalanobis distance of the position mean component of the forward predicted orbit determination state mean vector and the measured orbit determination state mean vector at the nth time through nonlinear covariance analysis according to the forward predicted orbit determination state mean vector, the measured orbit determination state mean vector and the corresponding covariance at the nth time. And when the Mahalanobis distance is larger than the preset value, obtaining a forward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the 0 th moment, and obtaining a backward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the nth moment.

In one embodiment, the processor, when executing the computer program, further performs the steps of: and acquiring a state value which is the minimum deviation value between a backward prediction orbit determination state statistic vector and a forward prediction orbit determination state statistic vector of the maneuvering target spacecraft at the ith moment, and setting the moment corresponding to the minimum deviation value as maneuvering time corresponding to the target spacecraft.

In one embodiment, the processor, when executing the computer program, further performs the steps of: and calculating the maneuvering speed increment of the corresponding target spacecraft according to the speed statistic component difference between the backward prediction orbit determination state statistic vector and the forward prediction orbit determination state statistic vector of the maneuvering time.

In one embodiment, the processor, when executing the computer program, further performs the steps of: and when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is larger than a preset value, respectively associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to different target spacecrafts.

In one embodiment, a computer-readable storage medium is provided, having a computer program stored thereon, which when executed by a processor, performs the steps of:

In one embodiment, the computer program when executed by the processor further performs the steps of: and when the deviation between the forward prediction orbit determination state statistic value vector at the nth moment and the measurement orbit determination state statistic value vector at the nth moment is smaller than a preset value, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as the unmoved state.

In one embodiment, the computer program when executed by the processor further performs the steps of: and calculating the forward prediction orbit determination state mean vector and the state covariance matrix at the nth moment, and calculating the measurement orbit determination state mean vector and the state covariance matrix at the nth moment. And obtaining the Mahalanobis distance of the position mean component of the forward predicted orbit determination state mean vector and the measured orbit determination state mean vector at the nth time through nonlinear covariance analysis according to the forward predicted orbit determination state mean vector, the measured orbit determination state mean vector and the corresponding covariance at the nth time. And when the Mahalanobis distance is larger than the preset value, obtaining a forward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the 0 th moment, and obtaining a backward prediction orbit determination state mean vector at the ith moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the nth moment.

In one embodiment, the computer program when executed by the processor further performs the steps of: and acquiring a state value which is the minimum deviation value between a backward prediction orbit determination state statistic vector and a forward prediction orbit determination state statistic vector of the maneuvering target spacecraft at the ith moment, and setting the moment corresponding to the minimum deviation value as maneuvering time corresponding to the target spacecraft.

In one embodiment, the computer program when executed by the processor further performs the steps of: and calculating the maneuvering speed increment of the corresponding target spacecraft according to the speed statistic component difference between the backward prediction orbit determination state statistic vector and the forward prediction orbit determination state statistic vector of the maneuvering time.

In one embodiment, the computer program when executed by the processor further performs the steps of: and when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is larger than a preset value, respectively associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to different target spacecrafts.

It will be understood by those skilled in the art that all or part of the processes of the methods of the embodiments described above can be implemented by hardware instructions of a computer program, which can be stored in a non-volatile computer-readable storage medium, and when executed, can include the processes of the embodiments of the methods described above. Any reference to memory, storage, database, or other medium used in the embodiments provided herein may include non-volatile and/or volatile memory, among others. Non-volatile memory can include read-only memory (ROM), Programmable ROM (PROM), Electrically Programmable ROM (EPROM), Electrically Erasable Programmable ROM (EEPROM), or flash memory. Volatile memory can include Random Access Memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in a variety of forms such as Static RAM (SRAM), Dynamic RAM (DRAM), Synchronous DRAM (SDRAM), Double Data Rate SDRAM (DDRSDRAM), Enhanced SDRAM (ESDRAM), Synchronous Link DRAM (SLDRAM), Rambus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM).

The technical features of the above embodiments can be arbitrarily combined, and for the sake of brevity, all possible combinations of the technical features in the above embodiments are not described, but should be considered as the scope of the present specification as long as there is no contradiction between the combinations of the technical features.

The above-mentioned embodiments only express several embodiments of the present application, and the description thereof is more specific and detailed, but not construed as limiting the scope of the invention. It should be noted that, for a person skilled in the art, several variations and modifications can be made without departing from the concept of the present application, which falls within the scope of protection of the present application. Therefore, the protection scope of the present patent shall be subject to the appended claims.

Claims

1. A spacecraft orbit maneuver detection method based on nonlinear deviation evolution is characterized by comprising the following steps:

acquiring measurement orbit determination state statistic value vectors of the target spacecraft at the 0 th moment and the nth moment; the measurement orbit determination state statistic value vector comprises a measurement orbit determination state mean value vector and a state covariance matrix, and the orbit determination state mean value vector comprises a position statistic value component and a speed statistic value component;

according to the measured orbit determination state statistic vector at the 0 th moment, obtaining a corresponding forward prediction orbit determination state statistic vector at the nth moment through nonlinear deviation evolution;

when the deviation between the forward prediction orbit determination state statistic vector at the nth moment and the measurement orbit determination state statistic vector at the nth moment is larger than a preset value, obtaining a forward prediction orbit determination state statistic vector at the p th moment through nonlinear deviation evolution according to the measurement orbit determination state statistic vector at the 0 th moment, and obtaining a backward prediction orbit determination state statistic vector at the q th moment through nonlinear deviation evolution according to the measurement orbit determination state statistic vector at the nth moment; wherein p =1, 2, …, n-1, q =0, 1, …, n-1;

when the deviation between the backward prediction orbit determination state statistic value vector at the ith moment and the forward prediction orbit determination state statistic value vector at the ith moment is smaller than a preset value at the ith moment, associating the measurement orbit determination state statistic value vectors at the 0 th moment and the nth moment to the same target spacecraft, and setting the state value of the corresponding target spacecraft as maneuvering; wherein i is more than or equal to 0 and less than n.

2. The method according to claim 1, wherein said step of obtaining a corresponding forward predicted tracking state statistic vector at time n by nonlinear bias evolution according to said measured tracking state statistic vector at time 0 further comprises:

3. The method of claim 1, wherein the forward predicted tracking state statistic vector comprises a forward predicted tracking state mean vector and a state covariance matrix;

when the deviation between the forward prediction orbit determination state statistic vector at the nth moment and the measurement orbit determination state statistic vector at the nth moment is larger than a preset value, the step of obtaining the forward prediction orbit determination state statistic vector at the p th moment through nonlinear deviation evolution according to the measurement orbit determination state statistic vector at the 0 th moment, and obtaining the backward prediction orbit determination state statistic vector at the q th moment through nonlinear deviation evolution according to the measurement orbit determination state statistic vector at the nth moment comprises the following steps:

calculating the forward prediction orbit determination state mean vector and the state covariance matrix at the nth moment, and calculating the measurement orbit determination state mean vector and the state covariance matrix at the nth moment;

obtaining the Mahalanobis distance of the position mean component of the forward prediction orbit determination state mean vector and the measurement orbit determination state mean vector at the nth moment through nonlinear covariance analysis according to the forward prediction orbit determination state mean vector, the measurement orbit determination state mean vector and the corresponding covariance at the nth moment;

and when the Mahalanobis distance is larger than a preset value, obtaining a forward prediction orbit determination state mean vector at a p moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the 0 moment, and obtaining a backward prediction orbit determination state mean vector at a q moment through nonlinear covariance analysis according to the measurement orbit determination state mean vector at the n moment.

4. The method according to claim 1, wherein, when there is a time i, and the deviation between the backward predicted orbiting state statistic vector at the time i and the forward predicted orbiting state statistic vector at the time i is smaller than a preset value, the step of associating the measured orbiting state statistic vectors at the time 0 and the time n to the same target spacecraft and setting the state value of the corresponding target spacecraft to the maneuver further comprises:

acquiring a deviation minimum value between the backward prediction orbit determination state statistic vector and the forward prediction orbit determination state statistic vector of the maneuvering target spacecraft from the 0 th moment to the n th moment, and setting the moment corresponding to the deviation minimum value as the maneuvering time corresponding to the target spacecraft.

5. The method according to claim 4, wherein the obtaining the target spacecraft whose state value is maneuvering further comprises, after the step of setting the time corresponding to the minimum deviation value between the backward predicted orbiting state statistic vector and the forward predicted orbiting state statistic vector from the 0 th time to the n th time as the maneuvering time of the corresponding target spacecraft, the step of obtaining the minimum deviation value between the backward predicted orbiting state statistic vector and the forward predicted orbiting state statistic vector:

6. The method according to claim 1, wherein said step of obtaining a forward predicted tracking state statistic vector at a time p by nonlinear bias evolution based on said measured tracking state statistic vector at a time 0 when a deviation between said forward predicted tracking state statistic vector at a time n and said measured tracking state statistic vector at a time n is greater than a predetermined value, and obtaining a backward predicted tracking state statistic vector at a time q by nonlinear bias evolution based on said measured tracking state statistic vector at a time n further comprises:

and when the ith moment does not exist, enabling the deviation between the backward prediction orbit determination state statistic vector at the ith moment and the forward prediction orbit determination state statistic vector at the ith moment to be smaller than a preset value, and respectively associating the measurement orbit determination state statistic vector at the 0 th moment and the measurement orbit determination state statistic vector at the nth moment to different target spacecrafts.

7. Method according to any of claims 1 to 6, characterized in that n +1 moments from the 0 th moment to the n th moment are evenly distributed, with adjacent moments being separated by less than 10 seconds.

8. A spacecraft orbit maneuver detection device based on nonlinear bias evolution, the device comprising:

the measurement orbit determination state statistic value vector calculation module is used for acquiring measurement orbit determination state statistic value vectors of the target spacecraft at the 0 th moment and the n th moment; the measurement orbit determination state statistic value vector comprises a measurement orbit determination state mean value vector and a state covariance matrix, and the orbit determination state mean value vector comprises a position statistic value component and a speed statistic value component;

the forward orbit determination state forecasting module is used for obtaining a corresponding forward prediction orbit determination state statistic value vector at the nth moment through nonlinear deviation evolution according to the measurement orbit determination state statistic value vector at the 0 th moment;

the bidirectional orbit determination state forecasting module is used for obtaining a forward predicted orbit determination state statistic vector at a p moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at a 0 th moment when the deviation between the forward predicted orbit determination state statistic vector at the n moment and the measured orbit determination state statistic vector at the n moment is larger than a preset value, and obtaining a backward predicted orbit determination state statistic vector at a q moment through nonlinear deviation evolution according to the measured orbit determination state statistic vector at the n moment; wherein p =1, 2, …, n-1, q =0, 1, …, n-1;

a target spacecraft maneuver detection module, configured to, when there is a time i and a deviation between the backward predicted orbit determination state statistic vector at the time i and the forward predicted orbit determination state statistic vector at the time i is smaller than a preset value, associate the measured orbit determination state statistic vectors at the time 0 and the time n to the same target spacecraft, and set a state value of the corresponding target spacecraft as a maneuver; wherein i is more than or equal to 0 and less than n.

9. A computer device comprising a memory and a processor, the memory storing a computer program, wherein the processor implements the steps of the method of any one of claims 1 to 7 when executing the computer program.

10. A computer-readable storage medium, on which a computer program is stored, which, when being executed by a processor, carries out the steps of the method according to any one of claims 1 to 7.