Intersatellite link navigation method

Abstract

The present invention provides an intersatellite link navigation method, including: when a satellite state suddenly changes, determining the current satellite orbit information through celestial navigation, so that the satellite can restore the link with other stars; introducing a suboptimal fading factor λ in the celestial navigation extended Kalman filter, performing strong tracking filtering, so that the satellite orbit information determined by the celestial navigation quickly approaches the real satellite orbit, and speeding up the intersatellite restoration and link establishment; generating intersatellite observation information through the intersatellite restoration and link establishment; generating initial orbit information using celestial navigation; and restoring autonomous navigation of the intersatellite link using the intersatellite observation information and the initial orbit information. The present invention makes full use of the complete autonomy of celestial navigation and the rapid convergence of strong tracking, so that when the satellite state suddenly changes and the intersatellite link is interrupted, it has the ability to quickly restore the link. The present invention effectively improves the autonomous operation capability of intersatellite link navigation.

Description

Technical Field

[0001] The present invention relates to the field of satellite navigation technology, and in particular to an inter-satellite link navigation method. Background Art

[0002] The intersatellite link navigation algorithm is not a completely autonomous satellite navigation. In order to determine the satellite navigation information, intersatellite ranging and intersatellite information exchange must be completed first. Therefore, if one satellite in the constellation fails, the accuracy of the satellite link establishment and autonomous navigation of the entire network will be affected. The stability of the intersatellite link navigation algorithm is poor. Furthermore, the intersatellite link navigation algorithm also lacks the ability to recover from autonomous faults. When a satellite establishes a link with other satellites, it needs to obtain its own orbital information in real time to adjust the link direction. Once the satellite maneuvers or the satellite attitude is unstable, the on-board orbital information will deviate, and the satellite link will not be able to align with other satellites to establish a link. More seriously, since it is impossible to obtain observation information through link establishment to correct the orbital error, the satellite will not be able to autonomously restore the link, and the intersatellite link navigation algorithm will not be able to be used. Summary of the invention

[0003] The object of the present invention is to provide an inter-satellite link navigation method to solve the problem that the existing inter-satellite link cannot autonomously restore the link in the event of a failure.

[0004] In order to solve the above technical problems, the present invention provides an inter-satellite link navigation method, comprising:

[0005] A suboptimal fading factor λ is introduced into the celestial navigation extended Kalman filter to perform strong tracking filtering. The suboptimal fading factor λ satisfies

[0006]

[0007] Among them, H k is the observation matrix, which is expressed in celestial navigation as in, is the orbital state transfer matrix from time k-1 to time k, P k-1 is the covariance matrix at time k-1;

[0008] R k is the observation noise matrix, which is represented by R in celestial navigation k =(υ CEL ) 2 ;

[0009] Q k is the process noise matrix, which is expressed in celestial navigation as

[0010] V k is the innovation covariance matrix of the actual output of the filter, which is expressed in celestial navigation as:

[0011]

[0012] Among them, ρ is the forgetting factor, ρ = 0.95, ε is the filter output information, which is expressed in celestial navigation as:

[0013]

[0014] Among them, a s is the observed starlight angular distance, r sat is the satellite inertial system position vector, w rsat is the satellite position vector process noise information, w vsat is the satellite velocity vector process noise information, w rsat , w vsat are all regarded as zero-mean white noise vectors, is the unit sidereal vector representation in the inertial coordinate system.

[0015] Optionally, in the intersatellite link navigation method,

[0016] Introducing a suboptimal fading factor λ into the celestial navigation extended Kalman filter to perform strong tracking filtering, so that the satellite orbit information determined by the celestial navigation quickly approaches the real satellite orbit, thereby accelerating the inter-satellite recovery and link establishment;

[0017] The astronomical navigation algorithm is combined with the intersatellite link navigation algorithm to improve the stability and autonomy of the intersatellite link algorithm.

[0018] Optionally, in the intersatellite link navigation method, the method further includes:

[0019] The strong tracking filter algorithm introduces the suboptimal fading factor λ into the prediction covariance matrix P of the traditional EKF filter algorithm. k / k-1 In this way, the filtering algorithm can still maintain the ability to track the true state of the satellite when the satellite state changes suddenly; its introduction method is shown in the following formula;

[0020]

[0021] Since the strong tracking filter needs to meet the following conditions,

[0022]

[0023] Thus, the suboptimal fading factor λ satisfies the equation

[0024]

[0025] The suboptimal fading factor λ is regarded as a single factor. Therefore, according to the above formula, λ is calculated using the following formula:

[0026]

[0027] Optionally, in the intersatellite link navigation method, when the satellite state suddenly changes, the satellite orbit information is determined by astronomical navigation so that the satellite can restore the link with other stars;

[0028] Introducing a suboptimal fading factor λ into the celestial navigation extended Kalman filter to perform strong tracking filtering, so that the satellite orbit information determined by the celestial navigation quickly approaches the real satellite orbit, thereby accelerating the inter-satellite recovery and link establishment;

[0029] Through the inter-satellite recovery link establishment, inter-satellite ranging is achieved and inter-satellite observation information is generated;

[0030] Generate initial orbit information of intersatellite links using astronomical navigation;

[0031] The intersatellite observation information and the intersatellite link initial orbit information are used to restore the intersatellite link autonomous navigation.

[0032] Optionally, in the intersatellite link navigation method, when the satellite state suddenly changes, determining the satellite orbit information by celestial navigation includes:

[0033] One-step prediction information of satellite orbits is calculated through satellite dynamics models;

[0034] Generate astronomical navigation observation information by collecting information from star sensors and earth sensors;

[0035] The one-step forecast information and observation information are introduced into the extended Kalman filter to optimize the real-time satellite orbit information.

[0036] Optionally, in the intersatellite link navigation method, collecting information through a star sensor and an earth sensor to generate celestial navigation observation information includes:

[0037] The star sensor collects the coordinates of the main star point on the charge-coupled device and calculates the unit vector of the star in the star-sensing coordinate system;

[0038] Using the attitude conversion matrix from the star-sensing coordinate system to the local system, the unit vector of the star in the star-sensing coordinate system is converted into the star unit vector in the local system;

[0039] The unit geocentric vector in the geosensitive coordinate system is obtained by observing the earth sensor, and the unit geocentric vector in the body coordinate system is calculated;

[0040] According to the obtained star unit vector and unit geocentric vector in this system, the observed starlight angular distance is calculated:

[0041]

[0042] in, is the unit star vector representation in the satellite body coordinate system, is the unit geocentric vector in the satellite coordinate system, a s It is the starlight angular distance observation information.

[0043] Optionally, in the intersatellite link navigation method, introducing a suboptimal fading factor λ into the extended Kalman filter to perform strong tracking filtering so that the satellite orbit information determined by celestial navigation quickly approaches the real satellite orbit, and accelerating the intersatellite recovery link establishment includes:

[0044] The suboptimal fading factor λ is introduced into the prediction covariance matrix P of the extended Kalman filter k / k-1 In, it is expressed as:

[0045]

[0046] in, is the orbital state transfer matrix from time k-1 to time k, P k-1 is the covariance matrix at time k-1.

[0047] Optionally, in the intersatellite link navigation method, the suboptimal fading factor λ satisfies

[0048]

[0049] Among them, H k is the observation matrix, which is expressed in celestial navigation as

[0050] R k is the observation noise matrix, which is represented by R in celestial navigation k =(υ CEL ) 2 ;

[0051] Q k is the process noise matrix, which is expressed in celestial navigation as

[0052] V k is the innovation covariance matrix of the actual output of the filter, which is expressed in celestial navigation as:

[0053]

[0054] Among them, ρ is the forgetting factor, ρ = 0.95, ε is the filter output information, which is expressed in celestial navigation as:

[0055]

[0056] Optionally, in the intersatellite link navigation method, the suboptimal fading factor λ is regarded as a single subfactor, and the calculation method is:

[0057]

[0058] Optionally, in the intersatellite link navigation method, implementing intersatellite ranging through the intersatellite link recovery and generation of intersatellite observation information includes:

[0059] The original information of satellite inter-satellite two-way ranging is:

[0060]

[0061] Among them, ρ AB , ρ BA are the two-way distance measurements between two stars A and B, is the theoretical value of intersatellite two-way ranging; δt A ,δt B is the clock difference between the two stars A and B; They are the inter-satellite two-way measurement error values, which include the transmit and receive delay error, antenna phase center deviation, relativistic effect error, and ionospheric delay error;

[0062] Adding the two equations in equation (7) together, we get the equation described in equation (8):

[0063]

[0064] In the formula, is the inter-satellite two-way measurement error; That is the intersatellite observation information.

[0065] Optionally, in the intersatellite link navigation method, using the intersatellite observation information and the intersatellite link initial orbit information to restore the intersatellite link autonomous navigation includes:

[0066] Using the orbit information generated by celestial navigation as the initial orbit estimate for intersatellite link navigation;

[0067] Use intersatellite ranging and error processing to obtain high-precision intersatellite observation information;

[0068] The extended Kalman filter is used to iteratively obtain high-precision satellite navigation information.

[0069] The present invention also provides an intersatellite link autonomous recovery system based on celestial navigation, comprising:

[0070] The strong tracking celestial navigation module is configured to determine the satellite orbit information through celestial navigation when the satellite state suddenly changes, so that the satellite can restore the link between the satellite and other stars; and introduce a suboptimal fading factor λ into the celestial navigation filter to perform strong tracking filtering, so that the satellite orbit information determined by the celestial navigation can quickly approach the real satellite orbit, thereby accelerating the restoration of the link between the satellites;

[0071] An intersatellite link establishment and recovery module is configured to determine satellite orbit information according to celestial navigation, adjust the intersatellite link pointing, and implement the intersatellite link recovery and establishment;

[0072] An observation information module is configured to achieve inter-satellite ranging and generate inter-satellite observation information through the inter-satellite recovery link establishment;

[0073] An initial orbit information module is configured to generate initial orbit information of an intersatellite link using celestial navigation;

[0074] The inter-satellite link autonomous navigation module is configured to use the inter-satellite observation information and the inter-satellite link initial orbit information to restore the inter-satellite link autonomous navigation.

[0075] In the intersatellite link navigation method provided by the present invention, inventive point 1 utilizes the completely autonomous characteristics of celestial navigation, and introduces the satellite orbit information determined by celestial navigation into the link pointing algorithm, so that the link between the satellite and other stars is restored, thereby enabling the intersatellite link autonomous navigation algorithm to have the ability to autonomously restore operation; inventive point 2 introduces a strong tracking suboptimal fading factor λ into the celestial navigation extended Kalman filter, so that the satellite orbit information determined by celestial navigation quickly approaches the real satellite orbit, accelerates the recovery of the intersatellite link navigation algorithm, and further improves the autonomous recovery performance of the intersatellite link navigation algorithm. BRIEF DESCRIPTION OF THE DRAWINGS

[0076] Figure 1 It is a schematic diagram of an intersatellite link autonomous navigation method based on celestial navigation according to an embodiment of the present invention;

[0077] Figure 2 is a schematic diagram of an intersatellite link navigation method according to an embodiment of the present invention;

[0078] FIG3( a ) is a diagram of the celestial navigation position error accuracy during stable operation of an embodiment of the present invention;

[0079] FIG3( b ) is a diagram of the navigation velocity error accuracy of celestial navigation during stable operation according to an embodiment of the present invention;

[0080] FIG4( a ) is a schematic diagram of the azimuth error accuracy of the satellite inter-satellite link during stable operation according to an embodiment of the present invention;

[0081] FIG4( b ) is a schematic diagram of the elevation angle error accuracy of the satellite inter-satellite link during stable operation according to an embodiment of the present invention;

[0082] FIG5( a ) is a schematic diagram of the navigation position error accuracy of celestial navigation when the state changes suddenly according to an embodiment of the present invention;

[0083] FIG5( b ) is a schematic diagram of the navigation speed error accuracy of celestial navigation when the state changes suddenly according to an embodiment of the present invention;

[0084] FIG6( a ) is a schematic diagram of the azimuth error accuracy of the satellite inter-satellite link when the state changes suddenly according to an embodiment of the present invention;

[0085] FIG6( b ) is a schematic diagram of the elevation angle error accuracy of the satellite inter-satellite link when the state changes suddenly according to an embodiment of the present invention;

[0086] FIG. 7( a ) is a diagram of the URE error of the PRN01 inter-satellite link autonomous navigation during initial convergence according to an embodiment of the present invention.

[0087] FIG. 7( b ) is a diagram of the autonomous navigation URE error of the PRN01 inter-satellite link during stable operation according to an embodiment of the present invention. DETAILED DESCRIPTION

[0088] The intersatellite link navigation method proposed in the present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments. The advantages and features of the present invention will become more apparent from the following description and claims. It should be noted that the accompanying drawings are all in a very simplified form and are not in precise proportions, and are only used to conveniently and clearly assist in explaining the purpose of the embodiments of the present invention.

[0089] In addition, unless otherwise specified, the features in different embodiments of the present invention may be combined with each other. For example, a feature in the second embodiment may be used to replace a corresponding feature in the first embodiment or a feature with the same or similar function, and the resulting embodiment also falls within the disclosure or description scope of the present application.

[0090] The core idea of ​​the present invention is to provide an inter-satellite link navigation method to solve the problem that the existing inter-satellite link cannot be restored autonomously in the event of a failure.

[0091] To realize the above idea, the present invention provides an inter-satellite link navigation method, including: when a satellite state suddenly changes, satellite orbit information is determined by astronomical navigation, so that the satellite can restore the link with other stars; a suboptimal fading factor λ is introduced into the astronomical navigation extended Kalman filter to perform strong tracking filtering, so that the satellite orbit information determined by the astronomical navigation quickly approaches the real satellite orbit, thereby accelerating the inter-satellite restoration and link establishment; through the inter-satellite restoration and link establishment, inter-satellite ranging is realized and inter-satellite observation information is generated; astronomical navigation is used to generate initial orbit information of the inter-satellite link; and the inter-satellite observation information and the initial orbit information of the inter-satellite link are used to restore autonomous navigation of the inter-satellite link.

[0092] The intersatellite link navigation algorithm is a semi-autonomous navigation algorithm, which has the disadvantages of poor stability and inability to recover from faults. A celestial navigation algorithm is proposed to improve the stable operation capability of the intersatellite link algorithm. The algorithm uses the extremely high stability and complete autonomy of celestial navigation to determine the direction of intersatellite link establishment to ensure stable intersatellite link establishment. In addition, the algorithm adds a strong tracking filter link to the celestial navigation, so that when the orbit information deviates and the intersatellite link is interrupted, it has the ability to quickly determine the autonomous orbit and restore the link. Simulation tests show that when the satellite is running smoothly, the elevation and azimuth pointing errors of the intersatellite link are less than 0.1° using the algorithm of the present invention. After the link is interrupted due to orbital maneuvers, the algorithm can enable the satellite to quickly restore the link establishment. The simulation test verifies the effectiveness of the celestial navigation algorithm.

[0093] In order to improve navigation accuracy and autonomous operation capability in wartime, all satellite navigation systems have carried out research on satellite autonomous navigation algorithms. The US GPS satellite system was the first to initiate such research. In 1984, Ananda first proposed a navigation satellite autonomous navigation technology that does not rely on the support of the ground monitoring system and only uses inter-satellite ranging information. This autonomous navigation technology was successfully applied to GPS BLOCK IIR satellites. GPS BLOCK IIR satellites correct the long-term forecast ephemeris through inter-satellite two-way ranging information. Within 75 days, the user ranging error is less than 3m.

[0094] The new generation Beidou global navigation satellite system has also introduced an intersatellite link measurement system and carried out intersatellite link autonomous navigation tests. Unlike the UHF band intersatellite links used by GPS BLOCK IIR satellites, Beidou global navigation satellites use Ka-band intersatellite links with higher ranging accuracy and stronger communication capabilities. According to on-orbit test evaluation, the intersatellite ranging accuracy of the Ka link is better than 10cm, and the communication rate can reach 50-100kbps. And through phased array technology, satellites can achieve rapid intersatellite Ka link switching and link establishment, and can generally complete the establishment of links with 14 other satellites within a 5-minute cycle. At present, Beidou satellite autonomous navigation technology is still in the whole network test stage, and its autonomous navigation accuracy has not yet been announced. At the same time, Beidou satellites have also carried out relevant research on laser intersatellite links and completed the demonstration of Beidou satellite networking solutions based on laser intersatellite links, striving to further improve the orbit and time synchronization accuracy of navigation satellites.

[0095] In order to make navigation satellites truly autonomous, in addition to improving the accuracy of autonomous navigation and enabling satellites to autonomously broadcast high-precision time and space references, satellites must also have the ability to operate stably and recover from failures over a long period of time. Since the intersatellite link algorithm is a semi-autonomous navigation algorithm with poor stability, it is urgent to conduct research on methods to improve the algorithm's ability to operate stably. However, there is currently no mature and effective solution.

[0096] Specifically, the intersatellite link algorithm is not a completely autonomous satellite navigation algorithm. In order to determine the satellite navigation information, the algorithm must first complete the intersatellite ranging and intersatellite information exchange. Therefore, if a satellite in the constellation fails, the accuracy of the satellite link establishment and the autonomous navigation accuracy of the entire network will be affected. The stability of the intersatellite link navigation algorithm is poor. Furthermore, the intersatellite link algorithm also lacks the ability to recover from faults autonomously. When a satellite establishes a link with other satellites, it is necessary to obtain its own orbital information in real time to adjust the link direction. Once the satellite maneuvers or the satellite attitude becomes unstable, the on-board orbital information will deviate, and the satellite link will not be able to align with the other satellite to establish a link. More seriously, since it is impossible to obtain observation information through link establishment to correct the orbital error, the satellite will not be able to autonomously restore the link, and the intersatellite link algorithm will not be able to be used.

[0097] To solve the above problems, the present invention proposes an intersatellite link autonomous recovery method based on celestial navigation to improve the stable operation capability of the intersatellite link navigation algorithm. First, the celestial navigation algorithm does not need to exchange information with the outside world and has extremely high stability and complete autonomy. Therefore, the output orbit information of the celestial navigation algorithm can be used to determine the link direction, so that the satellite can establish a long-term stable link with other satellites, ensuring the long-term stable operation of the intersatellite link navigation algorithm. In addition, the celestial navigation algorithm can collect the required astronomical information at any direction to obtain the current satellite orbit information. Therefore, when the satellite attitude is unstable or the orbit information deviation on the satellite is generated due to orbital maneuvers, resulting in link interruption, the celestial navigation can determine the current satellite orbit, so that the satellite can re-establish the link with other satellites. The present invention adds a strong tracking filter algorithm to the celestial navigation algorithm, so that the celestial navigation algorithm can quickly converge the on-satellite orbit deviation, so that the intersatellite link algorithm can quickly restore the link. Finally, the celestial navigation algorithm can provide the initial orbit reference information of the intersatellite link navigation algorithm, help the navigation information to converge quickly when the intersatellite link navigation algorithm restarts, and get rid of the dependence on the ground orbit.

[0098] This embodiment describes the principles of celestial navigation and intersatellite link navigation algorithms. The intersatellite link navigation algorithm based on celestial navigation combines the celestial navigation algorithm with the intersatellite link navigation algorithm to improve the stability and autonomy of the intersatellite link algorithm. In the design, the celestial navigation algorithm and the intersatellite link navigation algorithm are different except for the observation model. The satellite dynamics orbit prediction model and the filter model selection method are the same. The principles of the two navigation algorithms are introduced together below.

[0099] The satellite dynamics orbit prediction model is shown in formula (1), and the satellite dynamics orbit prediction model can be obtained by Newton's law.

[0100] (1)

[0102] In the formula, r sat is the satellite inertial system position vector, v satis the satellite inertial system velocity vector, a sat is the satellite inertial acceleration vector. rsat is the satellite position vector process noise information, w vsat is the satellite velocity vector process noise information. rsat , w vsat can all be regarded as zero-mean white noise vectors.

[0103] Since the navigation satellite is a medium-to-high orbit satellite, the satellite acceleration vector a in the prediction model sat The calculation mainly considers the changes in satellite acceleration caused by the 4×4 order non-spherical gravity of the earth, the gravity of the sun and the moon, and the pressure perturbation force of sunlight.

[0104] This embodiment proposes an observation model for a celestial navigation algorithm. The celestial navigation algorithm collects sensitive information from star sensors and earth sensors to form starlight angular distance observation information, thereby forming celestial navigation algorithm observation quantities and observation equations. The specific steps are as follows.

[0105] First, the star sensor collects the coordinates of the main star point (x s ,y s ), using formula (2) to calculate the unit vector of the star in the star-sensitive coordinate system

[0106]

[0107] In the formula, f s is the star-sensing optical focal length.

[0108] Then, the attitude transformation matrix R from the star-sensitive coordinate system to the satellite system is used bs , convert the unit vector of the star in the sensor coordinate system into the unit vector of the star expressed in this system As shown in formula (3).

[0109]

[0110] Similarly, the geocentric vector of the geosensitive coordinate system It can be obtained by observing the earth sensor. In addition, the representation of the unit geocentric vector in the body coordinate system can be calculated by formula (4).

[0111]

[0112] In the formula, R bh It is the transformation matrix from the ground-sensitive coordinate system to the satellite's own system.

[0113] According to the obtained stellar unit vector of this system and the unit geocentric vector in this system Observed starlight angular distance a s It can be calculated by formula (5).

[0114]

[0115] Furthermore, the starlight angular distance observation a s It can also be expressed by formula (6)

[0116]

[0117] In the formula, is the unit star vector representation in the inertial coordinate system, which can be calculated by formula (7). is the representation of the unit geocentric vector in the inertial system, which can be obtained by formula (8). υ is the observation noise, which can usually be regarded as Gaussian white noise. The design of the observation noise standard deviation can be obtained from the observation errors of the earth sensor and the star sensor.

[0118]

[0119] In formula (7), It is the celestial coordinate of the star of the inertial coordinate system satellite. Through star map recognition and the built-in star ephemeris in the star sensor, the celestial coordinate of the star can be accurately obtained.

[0120]

[0121] In formula (8), r sat is the satellite position vector, |r sat | is the satellite's distance from the Earth's center.

[0122] Substituting equation (8) into equation (6), the observation equation of the autonomous navigation algorithm of the star sensor and the earth sensor can be expressed by equation (9):

[0123]

[0124] This embodiment proposes an observation model for an inter-satellite link navigation algorithm. The inter-satellite link navigation algorithm uses inter-satellite bidirectional ranging information as an observation quantity and constructs an observation equation. The inter-satellite link navigation algorithm can be divided into two processing methods: distributed and centralized. The present invention uses a centralized processing method to introduce the inter-satellite link algorithm observation quantity and observation equation implementation method.

[0125] The original information of inter-satellite two-way ranging can be expressed by formula (10).

[0126]

[0127] In the formula, ρ AB , ρ BAare the two-way distance measurements between two stars A and B, is the theoretical value of intersatellite two-way ranging. A ,δt B is the clock difference between stars A and B. They are the inter-satellite two-way measurement error values, which include the transmit and receive delay error, antenna phase center deviation, relativistic effect error, and ionospheric delay error.

[0128] Adding the two equations in equation (10) together, we can get the equation described in equation (11). Therefore, the left-hand variable It can be observed in real time as an algorithm observation quantity.

[0129]

[0130] Furthermore, based on the relationship between the algorithm observation quantity and the satellite position and velocity information to be determined by the algorithm, the observation equation shown in formula (12) can be obtained.

[0131]

[0132] In the formula, are the position vectors of satellites A and B respectively.

[0133] This embodiment proposes a filtering model, as shown in equations (9) and (12), the observation equations of the celestial navigation algorithm and the observation equations of the intersatellite link navigation algorithm are both nonlinear equations. Therefore, in the design, the extended Kalman filter algorithm (EKF) is used to estimate the two algorithms.

[0134] According to the principle of EKF algorithm, the state vectors of the two navigation algorithms are adjusted to the first-order error of the satellite position and velocity information. At the same time, the observation equations of the two navigation algorithms are also adjusted to the forms shown in equations (13) and (14) according to the state vectors. Among them, equations (13) and (14) are respectively equations (9), and equation (12) is obtained by first-order Taylor expansion at the predicted position information.

[0135]

[0136]

[0137] The guidance method based on astronomical navigation information proposed in the present invention enables the intersatellite link navigation algorithm to have the ability of autonomous stable link establishment and fault recovery, thereby greatly improving the stable operation capability of the intersatellite link autonomous navigation algorithm. The following is an introduction to the autonomous link establishment and fault recovery implementation method based on astronomical information.

[0138] Among them, the autonomous link building method based on astronomical navigation information includes: The intersatellite link autonomous navigation algorithm is a semi-autonomous navigation algorithm. Each satellite in the constellation needs to exchange intersatellite ranging information and navigation messages with multiple other satellites to update the navigation information of each satellite. If there is an error in the navigation information of a satellite, the autonomous navigation accuracy of the satellites linked to it and even the entire network of satellites will be affected. Therefore, the navigation performance of a single satellite depends on the accuracy of the information exchanged by the entire network of satellites, and the algorithm stability is poor.

[0139] The establishment of intersatellite links requires the acquisition of satellite orbit information. If the navigation results obtained by the intersatellite link algorithm are directly introduced, the stability of the link establishment will be greatly restricted. Therefore, the present invention proposes an autonomous link establishment method based on astronomical information. Since the astronomical navigation algorithm has high reliability, the autonomous link establishment method based on astronomical navigation information can make the intersatellite link establishment have extremely high stability.

[0140] Combined with the principle of intersatellite link establishment, the proposed autonomous link establishment design method is as follows: Figure 1 shown.

[0141] 1) Determine the constellation satellites to be linked with the local satellite based on the intersatellite link planning table, and use astronomical navigation information to determine the satellite inertial system position Use the ephemeris of other stars to predict the position of the inertial system of other stars Using formula (15), we can get the satellite link inertial system pointing vector P: Ai .

[0142]

[0143] 2) Using celestial navigation information, the conversion matrix from the inertial system to the satellite orbit system is obtained as described in equation (16). And combined with the satellite attitude information, the conversion matrix from the satellite inertial system to the present system is obtained as described in equation (17).

[0144]

[0145]

[0146] In the formula, θ, ψ are the satellite three-axis attitude angle information, R oi , R bi They are the transformation matrices from the inertial system to the orbital system and from the inertial system to this system respectively.

[0147] 3) According to the link phase center local coordinates P Lb , link installation matrix R under this system Lb And the transformation matrix R from inertial system to this system bi , point the inertial system downward to vector P Ai The link establishment pointing vector P converted to the link coordinate system is represented by AL, the method is shown in formula (18).

[0148] P AL =R Lb (R bi P Ai -P Lb ) (18)

[0149] 4) Link building pointing vector representation P based on link coordinate system AL , the target rotation angle in the link coordinate system can be finally calculated and used for link pointing attitude adjustment. The method is shown in equations (19) and (20).

[0150]

[0151]

[0152] Where P AL =(x AL ,y AL ,z AL ) is the three-axis coordinate of the link-building pointing vector, (E i ,A z ) are the link pointing elevation angle and azimuth angle.

[0153] According to the above four steps, the autonomous establishment of satellite inter-satellite links based on astronomical navigation information can be completed.

[0154] This embodiment also provides a strong tracking autonomous link recovery method based on celestial navigation information, such as Figure 2 As shown in the figure, when a satellite uses an intersatellite link navigation algorithm to operate autonomously, if the satellite's orbit maneuvers or its attitude becomes unstable, the intersatellite link between the satellite and other satellites will be interrupted. When the satellite orbit maneuver ends or the satellite's attitude becomes stable again, the intersatellite link cannot be re-established because the real-time satellite orbit information cannot be obtained, and thus the intersatellite link autonomous navigation algorithm cannot resume operation. This section proposes a strong tracking autonomous link recovery method based on astronomical navigation information. By utilizing the complete autonomy of astronomical navigation and the rapid convergence of the strong tracking algorithm, the satellite can quickly obtain the current real-time orbit information after the link is interrupted, thereby restoring the link and resuming the operation of the intersatellite link algorithm. At the same time, at the beginning of the restart of the intersatellite link algorithm, astronomical navigation provides the link navigation algorithm with initial orbit information, so that the link navigation algorithm obtains the initial benchmark and quickly and stably outputs it. The process of the designed autonomous link recovery method is as follows: Figure 2 shown.

[0155] The present invention adds a strong tracking filter processing link to the traditional celestial navigation algorithm, so that the satellite can quickly determine its own orbit when the orbit information is lost. The strong tracking filter algorithm introduces the suboptimal fading factor λ into the prediction covariance matrix P of the traditional EKF filter algorithm. k / k-1In this way, the filtering algorithm can still maintain the ability to track the true state of the satellite when the satellite state changes suddenly. The introduction method is shown in formula (21).

[0156]

[0157] In the formula, is the orbital state transfer matrix from time k-1 to time k, P k-1 is the covariance matrix at time k-1.

[0158] Since the strong tracking filter needs to satisfy the condition shown in equation (22), the equation shown in equation (23) is obtained.

[0159]

[0160] Where ε is the new information output by the filter, which can be expressed as

[0161]

[0162] In the formula, H k is the observation matrix, which is expressed in the celestial navigation algorithm as V k Calculated by formula (24); R k is the observation noise matrix, which is represented by R in the celestial navigation algorithm. k =(υ CEL ) 2 ;Q k is the process noise matrix, expressed as

[0163]

[0164] In the formula, ρ is the forgetting factor, and in this embodiment, ρ=0.95 is selected.

[0165] The present invention regards the suboptimal fading factor λ as a single subfactor. Therefore, according to formula (23), formula (25) can be used to calculate λ

[0166]

[0167] According to the designed intersatellite link autonomous navigation algorithm based on celestial navigation, the autonomous link establishment performance based on celestial navigation information and the link establishment recovery capability based on celestial navigation information are simulated and analyzed. The simulation scenario and simulation results are described as follows.

[0168] A simulation scenario of 24 MEO BeiDou constellation satellites is established, and the satellite numbers in the constellation are simplified to PRN01 to PRN24. Each satellite flies in a yaw attitude. The standard orbit and standard attitude of each satellite are generated by Satellite Kit Tools (STK) software. According to the evaluation results of BeiDou satellite on-orbit test, the predicted satellite orbit is obtained by adding 10% light pressure error based on the satellite prediction dynamics model. The predicted satellite attitude is based on the standard attitude, and the roll, pitch, and yaw attitudes are added with 0.02°, 0.02°, and 0.04° errors, respectively.

[0169] In the simulation scenario, the PRN01 satellite is used to verify the intersatellite link navigation algorithm of the astronomical navigation designed by the present invention. For the astronomical navigation algorithm, the algorithm simulation cycle is set to 4s, the three-axis random noise of the star sensor is 5" (3σ), the random noise of the earth sensor is 0.015° (3σ), and the system noise is 0.01°. For the intersatellite ranging navigation algorithm, the algorithm simulation cycle is set to 5min, the Ka link coordinate system installation matrix is ​​set to the unit matrix, the installation position coordinates are (0, 0, 0.5)m, and the earth shielding, link elevation angle, azimuth angle dead zone and other factors are considered. A link planning table is established, and each satellite completes the intersatellite two-way ranging according to the planning table. Each satellite establishes a link with an average of 10 other satellites per cycle, and the intersatellite ranging error is 0.1m.

[0170] Firstly, the stability of intersatellite link establishment based on astronomical navigation information is evaluated according to the simulation scenario.

[0171] According to the simulation conditions of the celestial navigation algorithm, after 2 days of celestial navigation algorithm simulation, its navigation accuracy is shown in Figure 3 (a) and (b). After the algorithm is stabilized and converged, its maximum position accuracy of the three axes is 7270.1m, 6936.9m, and 6753.9m, and the maximum speed errors of the three axes are 1.6205m / s, 1.0982m / s, and 1.4526m / s, respectively. The algorithm error is basically consistent with the on-orbit Beidou satellite celestial navigation evaluation accuracy of 6000m and the speed accuracy of 1.5m / s, which verifies the effectiveness of the simulation results of the celestial navigation algorithm of the present invention.

[0172] The orbit information obtained by the celestial navigation algorithm is introduced into the autonomous link establishment method designed by the present invention to evaluate the satellite link establishment accuracy based on the celestial navigation algorithm. After 2 days of evaluation, the link establishment pointing accuracy is shown in Figures 4(a) and (b). After introducing the attitude error [0.02°, 0.02°, 0.04°], the elevation pointing error of the link establishment is 0.042°, and the azimuth pointing accuracy error is 0.082°. Therefore, through simulation analysis, it can be seen that by using the designed autonomous link establishment method based on astronomical information, the inter-satellite link establishment can maintain high precision and stability.

[0173] Then, the strong tracking and autonomous link recovery effect based on astronomical navigation information is evaluated.

[0174] After a period of stable operation, the satellite was subjected to a 30-minute orbital maneuver with a thrust acceleration of 0.03 m / s2 to evaluate the strong tracking and link-building recovery capability of the celestial navigation algorithm when attitude instability and satellite maneuvers occur. After two days of simulation, the algorithm's autonomous link-building recovery effect is described in Figures 5 (a) and (b) of the satellite navigation position accuracy and Figures 6 (a) and (b) of the link-building pointing accuracy.

[0175] When the satellite orbit maneuvers, the onboard orbit determination accuracy decreases rapidly. As shown in Figure 5(a) and (b), after the orbit maneuver, the three-axis orbit position error can reach [3.74, 0.832, 1.67] × 10 5 m, the velocity error can reach [113.05,9.23,7.45]m / s. At the same time, affected by this orbital accuracy, the link establishment pointing accuracy also drops rapidly. As shown in Figure 6(a) and (b), the maximum elevation angle pointing error of the intersatellite link establishment can reach 1.04°. Based on this link establishment pointing accuracy, the satellite intersatellite link can no longer establish a link with other satellites, and the intersatellite link autonomous navigation algorithm will not be able to run.

[0176] By using the strong tracking autonomous link establishment and recovery method based on celestial navigation information designed by the present invention, as shown in Figures 5(a) and (b) and Figures 6(a) and (b), the satellite orbit determination error and link establishment pointing accuracy after orbital maneuvers converge quickly. The elevation angle accuracy and azimuth angle accuracy of link establishment pointing are restored to within 0.1° after 5000s of orbital maneuvering, so that the intersatellite link can be restored.

[0177] The converged celestial navigation algorithm orbit information is used as the initial orbit in the intersatellite link autonomous navigation algorithm to evaluate the autonomous navigation accuracy of the intersatellite link navigation algorithm after link establishment is restored. When the link establishment pointing accuracy is restored to 0.1°, the celestial navigation orbit information is used as the initial orbit of the intersatellite link algorithm (the position error is [1.9643, 0.1114, 0.4478] × 10 4 m, the velocity error is [-5.2834, -1.6846, 1.3775] m / s), and the intersatellite link navigation algorithm is restarted for a 30-day algorithm simulation. As shown in Figure 7(a) and (b), after 15 iterations of the PRN01 intersatellite link navigation algorithm, the URE error converges to within 3 m, and the final error stabilizes at 0.2 m.

[0178] Therefore, through simulation analysis, the effectiveness of the strong tracking autonomous link recovery method based on celestial navigation information has also been verified.

[0179] In view of the shortcomings of the intersatellite link navigation algorithm, such as poor stability and inability to recover from faults, an intersatellite link autonomous navigation algorithm for celestial navigation is proposed. The algorithm introduces the satellite orbit information obtained by celestial navigation into the satellite link establishment method, so that the satellite can autonomously obtain high-precision intersatellite link establishment pointing. A strong tracking filter processing method is added to the celestial navigation algorithm, so that when there is an orbit information deviation of the satellite, the satellite can quickly determine the orbit and restore the link establishment. The simulation results show that when the satellite is running smoothly, the intersatellite elevation and azimuth pointing errors are less than 0.1° using the algorithm of the present invention. When the satellite link is interrupted due to a 30-minute orbital maneuver, the satellite link pointing accuracy can be restored to 0.1° within 5000 seconds using the algorithm of the present invention. The algorithm of the present invention greatly improves the autonomy and stable operation capability of the satellite intersatellite link autonomous navigation.

[0180] An embodiment of the present invention also provides an intersatellite link autonomous recovery system based on celestial navigation, comprising: a strong tracking celestial navigation module, configured to determine satellite orbit information through celestial navigation when a satellite state suddenly changes, so that the satellite can restore the link with other stars; and introduce a suboptimal fading factor λ into the celestial navigation filter to perform strong tracking filtering, so that the satellite orbit information determined by the celestial navigation quickly approaches the real satellite orbit, thereby accelerating the intersatellite recovery and link establishment; an intersatellite link establishment recovery module, configured to use celestial navigation to determine satellite orbit information, adjust the intersatellite link pointing, and realize the intersatellite recovery and link establishment; an observation information module, configured to realize intersatellite ranging and generate intersatellite observation information through the intersatellite recovery and link establishment; an initial orbit information module, configured to use celestial navigation to generate initial orbit information of the intersatellite link; and an intersatellite link autonomous navigation module, configured to use the intersatellite observation information and the intersatellite link initial orbit information to restore and perform intersatellite link autonomous navigation.

[0181] In summary, the above embodiments describe in detail different configurations of the intersatellite link autonomous navigation method and system based on celestial navigation. Of course, the present invention includes but is not limited to the configurations listed in the above embodiments, and any content that is transformed based on the configurations provided in the above embodiments belongs to the scope of protection of the present invention. Those skilled in the art can draw inferences from the contents of the above embodiments.

[0182] In this specification, each embodiment is described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the embodiments can be referred to each other. For the system disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant parts can be referred to the method part.

[0183] The above description is only a description of the preferred embodiments of the present invention, and is not intended to limit the scope of the present invention. Any changes or modifications made by a person skilled in the art in the field of the present invention based on the above disclosure shall fall within the scope of protection of the claims.

Claims

1. An intersatellite link navigation method, characterized in that: include: A suboptimal fading factor λ is introduced into the celestial navigation extended Kalman filter to perform strong tracking filtering. The suboptimal fading factor λ satisfies Among them, H k is the observation matrix, which is expressed in celestial navigation as in, is the orbital state transfer matrix from time k-1 to time k, P k-1 is the covariance matrix at time k-1; R k is the observation noise matrix, which is represented by R in celestial navigation k =(υ CEL ) 2 ; Q k is the process noise matrix, which is expressed in celestial navigation as V k is the innovation covariance matrix of the actual output of the filter, which is expressed in celestial navigation as: Among them, ρ is the forgetting factor, ρ = 0.95, ε is the filter output information, which is expressed in celestial navigation as: Among them, a s is the observed starlight angular distance, r sat is the satellite inertial system position vector, w rsat is the satellite position vector process noise information, w vsat is the satellite velocity vector process noise information, w rsat , w vsat are all regarded as zero-mean white noise vectors, is the unit sidereal vector representation in the inertial coordinate system.

2. The intersatellite link navigation method according to claim 1, characterized in that: Introducing a suboptimal fading factor λ into the celestial navigation extended Kalman filter to perform strong tracking filtering, so that the satellite orbit information determined by the celestial navigation quickly approaches the real satellite orbit, thereby accelerating the inter-satellite recovery and link establishment; The astronomical navigation algorithm is combined with the intersatellite link navigation algorithm to improve the stability and autonomy of the intersatellite link algorithm.

3. The intersatellite link navigation method according to claim 1, characterized in that: Also includes: The strong tracking filter algorithm introduces the suboptimal fading factor λ into the prediction covariance matrix P of the traditional EKF filter algorithm. k / k-1 In this way, the filtering algorithm can still maintain the ability to track the true state of the satellite when the satellite state changes suddenly; its introduction method is shown in the following formula; Since the strong tracking filter needs to meet the following conditions, Thus, the suboptimal fading factor λ satisfies the equation The suboptimal fading factor λ is regarded as a single factor. Therefore, according to the above formula, λ is calculated using the following formula:

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