Space debris approaching sensing method based on space-based angle measurement information

Abstract

The invention discloses a space debris approaching sensing method based on space-based angle measurement information, and belongs to the field of space situation sensing. According to the regularity of relative movement between a high-orbit satellite and high-orbit fragments, space-based angle observed quantity of the fragments is obtained through visible light imaging detection equipment carried on the satellite, after observed quantity accumulation and observed quantity correlation are successfully achieved, the complete trend that the angle observed quantity changes along with time is fitted, and the high-orbit fragments are obtained. And the approaching perception of the high-orbit space debris without depending on distance information is realized by taking the curve trend as a criterion.

Description

Technical Field

[0001] The present invention relates to the field of space situational awareness, and in particular to a method for sensing the approach of space debris based on space-based angle measurement information. Background Art

[0002] The increasingly crowded space environment places higher and higher demands on the satellite's self-protection capabilities. Debris approach sensing is a method for determining whether debris is approaching a satellite.

[0003] At present, the related research on evaluating the motion state of debris based on visible light imaging detection includes: debris motion detection, collision warning and relative motion state estimation. The above three methods are all of good reference significance for debris approach perception, but they are not completely applicable. Motion detection is suitable for debris orbit cataloging and cataloging database maintenance. It is not easy to determine the relative motion state between the debris and the satellite for the approach perception problem; collision warning is suitable for the assessment of the collision risk between debris and satellites with known cataloging information, but not for debris with unknown cataloging information. In addition, collision warning focuses on the state at the moment of collision rather than the approach process; relative motion state estimation can effectively analyze the motion state between the debris and the satellite, but the accuracy of the relative motion state evaluation depends on the accuracy of the state estimation at the initial moment. For visible light imaging detection, the lack of distance information makes it impossible to accurately determine the approach state of the debris. Summary of the invention

[0004] The technical problem solved by the present invention is to overcome the shortcomings of the prior art and provide a space debris approach perception method based on space-based angle measurement information, which only uses space-based angle measurement information to perceive the debris approach status.

[0005] The technical solution of the present invention is: a method for sensing the approach of space debris based on space-based angle measurement information, the method comprising the following steps:

[0006] Obtain the azimuth angle Az of the space debris in the satellite orbit coordinate system;

[0007] When the amplitude of the fluctuation of the tangent value of the azimuth angle Az of the space debris in the satellite orbit coordinate system gradually increases with time, it is determined that the space debris is approaching the satellite;

[0008] When the amplitude of the fluctuation of the tangent value of the azimuth angle Az of the space debris in the satellite orbit coordinate system gradually decreases with time, it is determined that the space debris is moving away from the satellite.

[0009] Another technical solution provided by the present invention is: a method for sensing approaching space debris based on space-based angle measurement information, the method comprising the following steps:

[0010] S1. Obtaining the coordinates of space debris on the image plane within multiple observation periods through space-based detection to form observation trajectory information of space debris; the observation period is equal to the orbital period of the satellite;

[0011] S2, correlating the observed quantities of the same space debris in multiple observation periods;

[0012] S3, extracting multiple observation quantity information from the observation trajectory of the same space debris in multiple observation periods;

[0013] S4, converting the observed information of space debris into azimuth information;

[0014] S5. Substituting the azimuth angle sequence of the same space debris in multiple observation periods into the function of the azimuth angle tangent value of the debris changing with time, performing least squares parameter fitting, and obtaining a curve of the azimuth angle tangent value of the debris changing with time;

[0015] S6. According to the curve of the tangent value of the azimuth angle of the debris changing with time, the extreme value within three consecutive orbital periods is obtained to judge the approaching state of the space debris.

[0016] Preferably, the function of the azimuth tangent value of the fragments varying with time is:

[0017]

[0018] Among them, n2 is the average angular velocity of the satellite, and p1~p6 are parameters to be fitted.

[0019] Preferably, if three consecutive orbital periods satisfy EP k -EP k-1 >Δep, the debris is considered to be approaching; if three consecutive orbital periods meet EP k-1 -EP k >Δep means the fragments are far away; k ,EP k-1 are the extreme values ​​of the curve of the azimuth tangent of the debris varying with time in two adjacent orbital periods, and Δep is the approach tolerance.

[0020] Preferably, the approach tolerance Δep has a value range of [0.5, 1.5].

[0021] Preferably, the steps of associating the observed quantities of the same space debris in multiple observation periods are as follows:

[0022] S3.1. Using the satellite's orbital period as the time difference criterion for the association of multi-period observations of the same debris, preliminarily judge whether the observations in two adjacent observation periods belong to the same debris, preliminarily pair the space debris in two adjacent observation periods to obtain space debris pairs. The space debris pair of space debris i' is represented by: (space debris i', space debris j'), i'∈[1, I], j'∈[1, J];

[0023] S3.2, if there is only one set of space debris pairs for space debris i, record this set of space debris pairs as: the optimal space debris pair of space debris i', and go to step S3.3; otherwise, use the association method based on trajectory consistency to select the optimal space debris from multiple space debris that are initially paired with space debris i, and form the optimal space debris pair with space debris i';

[0024] S3.3. According to the distance residual s between the observed trajectories of two space debris in the optimal debris pair of space debris i' in adjacent measurement cycles, it is confirmed whether the two space debris in the optimal debris pair belong to the same space debris.

[0025] Preferably, the method for preliminarily judging whether the observed quantities in two observation periods belong to the same fragment is:

[0026] The time when space debris A in the first observation period reaches intersection point a is T1, and the time when space debris B in the second observation period reaches intersection point a is T2, if:

[0027] TS-ΔT<|T1-T2| <TS+ΔT

[0028] It is preliminarily determined that the fragment A in the first observation period and the fragment B in the second observation period belong to the same fragment, where ΔT is the time tolerance.

[0029] Preferably, the time tolerance ΔT has a value range of [60, 180], in seconds.

[0030] Preferably, the association method based on trajectory consistency is:

[0031] The observed trajectory of space debris i in the kth observation period is recorded as the first observed trajectory;

[0032] The observed trajectory of space debris j' that is initially paired with space debris i' in the k-1th observation period is recorded as the second observed trajectory;

[0033] Solving the trajectory consistency criterion of the first observation trajectory and all second observation trajectories;

[0034] The space debris corresponding to the second observed trajectory with the smallest trajectory consistency criterion is taken as the optimal paired space debris of space debris i.

[0035] Preferably, the trajectory consistency criterion is calculated by the following method:

[0036] Calculate the Euclidean distance Dis between each scattered point between the first observation trajectory and the second observation trajectory i,j ,i=1~m,j=1~n, the Euclidean distances between the scattered points form a matrix of size m×n; m is the number of scattered points of the real trajectory, and n is the number of scattered points of the predicted trajectory;

[0037] The trajectory consistency criterion D(m,n) between the first trajectory and the second trajectory is calculated according to the following formula:

[0038] D(m,n)=|Dis m,n |+min{|Dis m-1,n |,|Dis m-1,n-1 |,|Dis m,n-1 |}.

[0039] Preferably, the Euclidean distance between the i-th scattered point in the first observation trajectory and the j-th scattered point in the second observation trajectory is:

[0040] Dis i,j =(U i -U j ′)+(V i -V j ′),0 <i≤m,0<j≤n

[0041] Among them, (U i , V i ) is the coordinate of the i-th scattered point in the real trajectory on the image plane;

[0042] (U j ′,V j ′) is the image plane coordinate of the jth scattered point in the prediction trajectory.

[0043] The beneficial results of the present invention compared with the prior art are as follows:

[0044] (1) The present invention proposes a debris approach perception method based on multi-period observation quantity correlation and extreme point fitting. The method is applicable to the visible light imaging detection of debris by high-orbit satellites and can effectively solve the problem of limited accuracy in relative motion state assessment of visible light imaging detection equipment.

[0045] (1) The present invention realizes the perception of the approach of space debris by fitting the trend of azimuth angle change over time. Compared with the prior art, for visible light imaging detection, it effectively makes up for the problem that the lack of distance information makes it impossible to accurately determine the approach status of debris.

[0046] (2) The present invention uses trajectory consistency as the criterion to achieve the association between space-based observations of space debris and cataloging information. It is not affected by factors such as target motion characteristics and observation environment, and can effectively reduce the probability of false association. BRIEF DESCRIPTION OF THE DRAWINGS

[0047] Figure 1 The present invention is a flow chart of a method for sensing approaching space debris based on space-based angle measurement information. DETAILED DESCRIPTION

[0048] The present invention is described in detail below with reference to the accompanying drawings and specific embodiments.

[0049] The present invention first analyzes the variation law of the relative position of debris and derives the criterion for the space-based angular approach perception of space debris; then, the proposed debris approach perception method is described in detail, including two steps: correlation of observation quantities within multiple observation periods of the same debris and judgment of the debris approach behavior, such as Figure 1 shown.

[0050] S1. Criteria for determining the space-based approach sensing only by angle measurement

[0051] The relative motion equation between the high-orbit satellite and the debris can be expressed as:

[0052]

[0053] Where n1 and n2 are the average angular velocities of the debris and the satellite respectively; (x, y, z) is the position vector of the debris in the satellite orbit coordinate system.

[0054] Solving the above relative motion equation, the position vector of the debris in the satellite orbit coordinate system can be expressed as follows:

[0055]

[0056] Where C' xi , C' yi , C' zi are constant coefficients, and τ is time. From the above formula, we can see that only the x-direction has a component that changes with time τ.

[0057] Visible light imaging detection equipment can only obtain the angle measurement information of the debris. The relationship between the azimuth angle Az and elevation angle El of the debris in the satellite orbit coordinate system and the position vector (x, y, z) of the debris in the satellite orbit coordinate system is as follows:

[0058]

[0059] Since only the x-direction has a component that varies with time τ, the present invention derives the criterion for space-based angular proximity perception based on the solution of the relative motion equation between the space debris and the satellite and the relationship between the azimuth angle Az and x, y, that is, when the debris approaches the satellite, the amplitude of the fluctuation of the tangent value of the azimuth angle Az in the satellite orbit coordinate system over time gradually increases; when the debris moves away from the satellite, the amplitude of the fluctuation of the tangent value of the azimuth angle Az over time gradually decreases.

[0060] The following space debris approach perception method based on space-based angle measurement information is proposed in the project. The specific steps are as follows:

[0061] S1. Obtaining the coordinates of space debris on the image plane within multiple observation periods through space-based detection to form observation trajectory information of space debris; the observation period is equal to the orbital period of the satellite;

[0062] S2, correlating the observed quantities of the same space debris in multiple observation periods;

[0063] S3, extracting multiple observation quantity information from the observation trajectory of the same space debris in multiple observation periods;

[0064] S4, converting the observed information of space debris into azimuth information;

[0065] First, according to the installation relationship between the optical system and the detector in the visible light imaging detection equipment, the direct observation quantity (U, V) is converted into a vector (Vec) in the satellite orbit coordinate system. x ,Vec y ,Vec z ), the azimuth is calculated as follows

[0066]

[0067] S5. Substituting the azimuth angle sequence of the same space debris in multiple observation periods into the function of the azimuth angle tangent value of the debris changing with time, performing least squares parameter fitting, and obtaining a curve of the azimuth angle tangent value of the debris changing with time;

[0068] S6. According to the curve of the tangent value of the azimuth angle of the debris changing with time, the extreme value within three consecutive orbital periods is obtained to judge the approaching state of the space debris.

[0069] S2. Correlation of multi-period observations of the same fragment

[0070] Space-based angular proximity sensing of debris requires the accumulation of multi-cycle observations, but due to the influence of the external environment of stray light, the observations obtained by the detection equipment are intermittent. The correlation of the observations of the same debris in multiple observation cycles is the prerequisite for proximity sensing.

[0071] The present invention realizes the association of observations in multiple observation periods of the fragments by combining time difference and trajectory consistency. The steps are as follows:

[0072] S3.1. Using the satellite's orbital period as the time difference criterion for the association of multi-period observations of the same debris, preliminarily judge whether the observations in two adjacent observation periods belong to the same debris, preliminarily pair the space debris in two adjacent observation periods to obtain space debris pairs. The space debris pair of space debris i' is represented by: (space debris i', space debris j'), i'∈[1, I], j'∈[1, J];

[0073] S3.2, if there is only one set of space debris pairs for space debris i, record this set of space debris pairs as: the optimal space debris pair of space debris i', and go to step S3.3; otherwise, use the association method based on trajectory consistency to select the optimal space debris from multiple space debris that are initially paired with space debris i, and form the optimal space debris pair with space debris i';

[0074] S3.3. According to the distance residual s between the observed trajectories of two space debris in the optimal debris pair of space debris i' in adjacent measurement cycles, it is confirmed whether the two space debris in the optimal debris pair belong to the same space debris.

[0075] From the expression of the position vector (x, y, z) of the debris in the satellite orbit coordinate system, it can be seen that the time interval between two consecutive times when the same debris reaches the same intersection point with the satellite orbit plane is theoretically the orbital period of the debris. However, since the orbital period of the debris is unknown, the space debris close to the satellite is generally close to the satellite. Therefore, the orbital period TS of the satellite is used as the time difference criterion for the association of multi-period observations of the same debris.

[0076] The method to preliminarily determine whether the observed quantities in two observation periods belong to the same fragment is:

[0077] The time when space debris A in the first observation period reaches intersection point a is T1, and the time when space debris B in the second observation period reaches intersection point a is T2, if:

[0078] TS-ΔT<|T1-T2| <TS+ΔT

[0079] It is preliminarily determined that the fragment A in the first observation period and the fragment B in the second observation period belong to the same fragment, where ΔT is the time tolerance. The value range of the time tolerance ΔT is [60, 180], unit: second.

[0080] The observations over a period of time can form an observation trajectory. The observation trajectory of the same debris has strong consistency in adjacent orbital periods, while the observation trajectories of different debris are quite different.

[0081] The correlation method based on trajectory consistency is used to confirm the correlation of multi-period observations of the same fragment. The correlation method based on trajectory consistency is:

[0082] The observed trajectory of space debris i in the kth observation period is recorded as the first observed trajectory;

[0083] The observed trajectory of space debris j' that is initially paired with space debris i' in the k-1th observation period is recorded as the second observed trajectory;

[0084] Solving the trajectory consistency criterion of the first observation trajectory and all second observation trajectories;

[0085] The space debris corresponding to the second observed trajectory with the smallest trajectory consistency criterion is taken as the optimal paired space debris of space debris i.

[0086] The trajectory consistency criterion is calculated by the following method:

[0087] Calculate the Euclidean distance Dis between each scattered point between the first observation trajectory and the second observation trajectory i,j ,i=1~m,j=1~n, the Euclidean distances between the scattered points form a matrix of size m×n; m is the number of scattered points of the real trajectory, and n is the number of scattered points of the predicted trajectory;

[0088] The trajectory consistency criterion D(m,n) between the first trajectory and the second trajectory is calculated according to the following formula:

[0089] D(m,n)=|Dis m,n |+min{|Dis m-1,n |,|Dis m-1,n-1 |,|Dis m,n-1 |}.

[0090] The Euclidean distance between the i-th scattered point in the first observation trajectory and the j-th scattered point in the second observation trajectory is:

[0091] Dis i,j =(U i -U j ′)+(V i -V j ′),0 <i≤m,0<j≤n

[0092] Among them, (U i , V i ) is the coordinate of the i-th scattered point in the real trajectory on the image plane;

[0093] (U j ′,V j ′) is the image plane coordinate of the jth scattered point in the prediction trajectory.

[0094] Take the two trajectories with the smallest criterion value to calculate the residual s of the distance between the trajectories. According to the statistical distribution law, the residual s obeys a normal distribution with a mean of 0 and a standard deviation of σ, and a confidence interval of ±εσ. If it satisfies:

[0095]

[0096] It is confirmed that the two observed tracks belong to the same debris.

[0097] The residual distance s between trajectories is calculated as follows

[0098]

[0099] Where, K = min (m, n), E t =(U t -U t ′)+(V t -V t ′),(0 <t≤K)

[0100] S3. Debris approach behavior judgment

[0101] After successfully correlating the multi-period observations of the same debris, the complete curve of the fluctuation of the tangent value of the azimuth angle Az over time is fitted, and then the criterion for space-based angle-only approach perception is derived based on S1 to perform debris approach perception.

[0102] From the expression of the three-axis position vector (x, y, z) of the debris in the satellite orbit coordinate system and the relationship between the tangent value of the azimuth angle and x, y, we can get:

[0103] The function of the azimuth tangent of the fragment changing with time is:

[0104]

[0105] Among them, n2 is the average angular velocity of the satellite, and p1~p6 are parameters to be fitted.

[0106] Substituting the observed quantity sequence of the same fragment in multiple observation periods into the above formula for least squares parameter fitting, the parameter to be fitted can be obtained: r ,r=1~6.

[0107] The p r , r = 1 ~ 6 into the expression of tan Az to obtain the fitting curve of the azimuth tangent value changing with time. If three consecutive orbital periods satisfy EP k -EP k-1 >Δep, the debris is considered to be approaching; if three consecutive orbital periods meet EP k-1 -EP k >Δep means the fragments are far away; k,EP k-1 are the extreme values ​​of the curve of the azimuth tangent of the debris varying with time in two adjacent orbital periods, and Δep is the approach tolerance.

[0108] The value range of the approach tolerance Δep is [0.5, 1.5].

[0109] In summary, in order to solve the problem of only using angle measurement information to perceive the approaching state of debris, the present invention proposes a debris approaching perception method based on multi-period observation quantity correlation and extreme point fitting. The method is applicable to debris and satellites in high orbits, and can effectively solve the problem of limited accuracy in relative motion state assessment of visible light imaging detection equipment. It mainly involves the derivation of space-based angle measurement-only space debris approaching perception criteria, the correlation of space-based observation quantities of the same space debris in multiple periods, observation quantity curve fitting and extreme point prediction.

[0110] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A space debris approach perception method based on space-based angle measurement information, characterized in that The steps include: Obtain the azimuth angle Az of the space debris in the satellite orbit coordinate system; When the amplitude of the fluctuation of the tangent value of the azimuth angle Az of the space debris in the satellite orbit coordinate system gradually increases with time, it is determined that the space debris is approaching the satellite; When the amplitude of the fluctuation of the tangent value of the azimuth angle Az of the space debris in the satellite orbit coordinate system gradually decreases with time, it is determined that the space debris is moving away from the satellite.

2. A space debris approach perception method based on space-based angle measurement information, characterized in that: The steps include: S1. Obtaining the coordinates of space debris on the image plane within multiple observation periods through space-based detection to form observation trajectory information of space debris; the observation period is equal to the orbital period of the satellite; S2, correlating the observed quantities of the same space debris in multiple observation periods; S3, extracting multiple observation quantity information from the observation trajectory of the same space debris in multiple observation periods; S4, converting the observed information of space debris into azimuth information; S5. Substituting the azimuth angle sequence of the same space debris in multiple observation periods into the function of the azimuth angle tangent value of the debris changing with time, performing least squares parameter fitting, and obtaining a curve of the azimuth angle tangent value of the debris changing with time; S6. According to the curve of the tangent value of the azimuth angle of the debris changing with time, the extreme value within three consecutive orbital periods is obtained to judge the approaching state of the space debris.

3. The space debris approach perception method based on space-based angle measurement information according to claim 2 is characterized in that: The function of the azimuth tangent of the fragment changing with time is: Among them, n2 is the average angular velocity of the satellite, and p1~p6 are parameters to be fitted.

4. The space debris approach perception method based on space-based angle measurement information according to claim 2 is characterized in that: If three consecutive orbital periods satisfy EP k -EP k-1 >Δep, the debris is considered to be approaching; if three consecutive orbital periods meet EP k-1 -EP k >Δep means the fragments are far away; k ,EP k-1 are the extreme values ​​of the curve of the azimuth tangent of the debris varying with time in two adjacent orbital periods, and Δep is the approach tolerance.

5. The space debris approach perception method based on space-based angle measurement information according to claim 2, characterized in that: The steps to correlate the observations of the same space debris over multiple observation periods are as follows: S3.

1. Using the satellite's orbital period as the time difference criterion for the association of multi-period observations of the same debris, preliminarily judge whether the observations in two adjacent observation periods belong to the same debris, preliminarily pair the space debris in two adjacent observation periods to obtain space debris pairs. The space debris pair of space debris i' is represented by: (space debris i', space debris j'), i'∈[1, I], j'∈[1, J]; S3.2, if there is only one set of space debris pairs for space debris i, record this set of space debris pairs as: the optimal space debris pair of space debris i', and go to step S3.3; otherwise, use the association method based on trajectory consistency to select the optimal space debris from multiple space debris that are initially paired with space debris i, and form the optimal space debris pair with space debris i'; S3.

3. According to the distance residual s between the observed trajectories of two space debris in the optimal debris pair of space debris i' in adjacent measurement cycles, it is confirmed whether the two space debris in the optimal debris pair belong to the same space debris.

6. The space debris approach perception method based on space-based angle measurement information according to claim 1, characterized in that: The method to preliminarily determine whether the observed quantities in two observation periods belong to the same fragment is: The time when space debris A in the first observation period reaches intersection point a is T1, and the time when space debris B in the second observation period reaches intersection point a is T2, if: TS-ΔT<|T1-T2| <TS+ΔT It is preliminarily determined that the fragment A in the first observation period and the fragment B in the second observation period belong to the same fragment, where ΔT is the time tolerance.

7. The method for sensing approaching space debris based on space-based angle measurement information according to claim 6, characterized in that: The value range of the time tolerance ΔT is [60,180], unit: second.

8. The method for sensing approaching space debris based on space-based angle measurement information according to claim 2, characterized in that: The association method based on trajectory consistency is: The observed trajectory of space debris i in the kth observation period is recorded as the first observed trajectory; The observed trajectory of space debris j' that is initially paired with space debris i' in the k-1th observation period is recorded as the second observed trajectory; Solving the trajectory consistency criterion of the first observation trajectory and all second observation trajectories; The space debris corresponding to the second observed trajectory with the smallest trajectory consistency criterion is taken as the optimal paired space debris of space debris i.

9. The space debris approach perception method based on space-based angle measurement information according to claim 8, characterized in that: The trajectory consistency criterion is calculated by the following method: Calculate the Euclidean distance Dis between each scattered point between the first observation trajectory and the second observation trajectory i,j ,i=1~m,j=1~n, the Euclidean distances between the scattered points form a matrix of size m×n; m is the number of scattered points of the real trajectory, and n is the number of scattered points of the predicted trajectory; The trajectory consistency criterion D(m,n) between the first trajectory and the second trajectory is calculated according to the following formula: D(m,n)=|Dis m,n |+min{|Dis m-1,n |,|Dis m-1,n-1 |,|Dis m,n-1 |}。 10. A high-precision space debris collision warning method assisted by space-based observation information according to claim 9, characterized in that The Euclidean distance between the i-th scattered point in the first observation trajectory and the j-th scattered point in the second observation trajectory is: Dis i,j =(U i -U j ′)+(V i -V j ′),0 <i≤m,0<j≤n Among them, (U i , V i ) is the coordinate of the i-th scattered point in the real trajectory on the image plane; (U j ′,V j ′) is the image plane coordinate of the jth scattered point in the prediction trajectory.

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