CN105044745A - Circular orbit low orbit satellite zenith pass remaining visible duration prediction method - Google Patents

Abstract

The invention discloses a circular orbit low orbit satellite zenith pass remaining visible duration prediction method. According to the method, operation work characteristics of a satellite are fully used; the spatial position of the satellite and corresponding time information are constantly calculated and updated; the information is broadcasted to a user on the ground in the form of a simple navigation signal; after the user on the ground receives the navigation signal, clock calibration and user positioning are carried out through a GPS receiving device; the elevation angle and the track angle of the satellite relative to the user are calculated in real time; and satellite zenith pass remaining visible duration is solved; the method is simple and effective; less navigation signal content is needed; the degree of dependence on satellite ephemeris is low; and on the premise of accurate and timely prediction duration, the hardware cost of the user is not increased.

Description

A kind of circular orbit low orbit satellite crosses the visible duration prediction method of top residue

Technical field

The present invention relates to satellite communication field, particularly relate to a kind of circular orbit low orbit satellite and cross the visible duration prediction method of top residue.

Background technology

It take the earth's core as the circuit orbit in the center of circle that circular orbit low orbit satellite refers to that orbit is, and the satellite of orbit altitude between 500 kms to 2000 kms.This circular orbit low orbit satellite have propagation delay little, to terminal device antenna size and the advantage such as power consumption requirements is low, and satellite launch cost is lower, but satellite altitude is lower, its coverage is less, by the time in ground satellite terminal user overhead, namely Covering time is shorter, therefore often needs many low orbit satellite networkings to meet telex network demand.

When circular orbit low orbit satellite is for communicating, mainly deposit deficiency both ways: one is that Covering time is limited, and the communication time of user is likely greater than the Covering time of single satellite; Two is that user can select at switching instant between multi-satellite.The deficiency of these two aspects causes needing in telex network process to switch between many low orbit satellites, to ensure the continuity communicated.Therefore, the method of ground based terminal user switch satellite is extremely important, if user's Stochastic choice satellite provides communication service, likely choose and current there is the satellite of the shortest visible duration or the satellite of shorter visible duration, and then increase the number of times of satellite switching, increase the cutting off rate of user, and interrupt current talking and compare and block new call and can bring worse Consumer's Experience to user, from the angle reducing conversation loss rate, desirable the select of satellite mode is that user can select to provide the satellite of the longest service duration at switching instant from all visible satellites.Therefore, from the angle of user's observation, be necessary to make accurately predicting to satellite relative to the visible duration of the residue of user, for user selects satellite to provide technical support at switching instant.

In prior art, predicting that top satellite remained the method for visible duration mainly by downloading the ephemeris information of satellite, calculate satellite relative to the visible duration of the residue of user, but the method needing user to constantly update the ephemeris information of satellite.In addition, the people such as IrfanAli are at " PredictingtheVisibilityofLEOSatellites " (IEEETransactionsonAerospaceandElectronicSystems, vol.35, no.4,1999) a kind of method that top satellite remains visible duration of predicting is proposed in, but with the longitude residing for the intersection point of equatorial plane and corresponding moment the method needs input satellite orbit radius, orbit inclination and satellite from the Southern Hemisphere to Northern Hemisphere operational process, the method computational complexity is higher, real-time is not strong, is difficult to meet the demand selected fast satellite.

Summary of the invention

The technical matters that the present invention mainly solves is to provide a kind of circular orbit low orbit satellite and crosses the visible duration prediction method of top residue, solves in prior art and depends on the problems such as satellite ephemeris, computation complexity is high, real-time is not strong unduly.

For solving the problems of the technologies described above, the technical scheme that the present invention adopts is: provide a kind of circular orbit low orbit satellite to cross the visible duration prediction method of top residue, comprise the steps: the first step, user receives satellite navigation signals; Second step, calculates the elevation angle of satellite; 3rd step, calculates the mark angle of satellite; 4th step, calculates satellite and remains visible duration relative to the top of crossing of user.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, this user receives satellite navigation signals and comprises: (1) calculates satellite spatial position, by ground control centre by orbit prediction method, spatial positional information in this satellite of regular calculating following a period of time and the time information of correspondence thereof, and at this satellite through this overhead, ground control centre, upgrade the spatial positional information of this satellite; (2) satellite calibration clock, this satellite utilizes self GPS device to carry out clock alignment, and according to the clock information of calibration and the spatial positional information of this satellite of storage, the substar latitude and longitude information in this satellite current coverage range is determined in adjustment in real time; (3) satellite sends navigation signal, and this satellite regularly sends navigation signal, and this navigation signal comprises the substar latitude and longitude information of this satellite in its coverage and the time information of correspondence thereof; (4) user receives navigation signal, judge that this satellite is close to user or away from user, this user receives this navigation signal double that this satellite sends, and judges this satellite close to still away from this user according to the substar latitude and longitude information in this navigation signal relative to the distance change of customer location.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, this orbit prediction method is SGP4 method; The frequency of the spatial positional information in this regularly this satellite of calculating following a period of time and the time information of correspondence thereof is at interval of calculating 1 time 1 second; The frequency that this satellite regularly sends navigation signal is at interval of sending 1 time 1 second.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, the method at the mark angle of this calculating satellite comprises: (1) determines this customer position information, obtains the longitude λ of this user self _t, latitude η _twith elevation κ _tinformation; (2) clock alignment, according to the time information in received navigation signal, by consistent with satellite clock for user clock synchronous calibration; (3) calculating mark angle in real time, this user this navigation signal by receiving, calculating the mark angle γ of this user this satellite relative _m, computing method are as follows:

$\mathrm{\Ω}=\[\sqrt{{\sin }^2\left(\frac{\stackrel{\→}{{\mathrm{\λ}}_S}-{\mathrm{\λ}}_T}{2}\right)+cos\left(\stackrel{\→}{{\mathrm{\λ}}_S}\right)\×cos\left({\mathrm{\λ}}_T\right)\×{\sin }^2\left(\frac{\stackrel{\→}{{\mathrm{\η}}_S}-{\mathrm{\η}}_T}{2}\right)}\]$

γ _m＝min{2sin ^-1(Ω)}

Wherein, with be longitude and the latitude vector of the substar in this satellite coverage of comprising in this navigation signal, Ω also be vectorial.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, obtained the longitude λ of this user self by the GPS device measurement that this user is built-in _t, latitude η _twith elevation κ _tinformation.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, the method at the elevation angle of this calculating satellite comprises: (1) ECEF coordinate system is changed, by the longitude and latitude of this satellite current and user with elevation coordinate is unified is converted to ECEF coordinate system Satellite body-fixed coordinate system and user's body-fixed coordinate system; (2) local coordinate conversion, is converted to the satellite local coordinate system in local coordinate and user's origin by this satellite body-fixed coordinate system and this user's body-fixed coordinate system; (3) calculate the elevation angle in real time, utilize this satellite local coordinate system and this user's origin, in this local coordinate, calculate the elevation angle theta of this user this satellite relative.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, the spatial positional information of this satellite is (λ _s, η _s, κ _s), wherein λ _srepresent the longitude of this satellite, η _srepresent the latitude of this satellite, κ _srepresent the elevation of this satellite, this customer position information is (λ _t, η _t, κ _t), wherein λ _trepresent the longitude of this user, η _trepresent the latitude of this user, κ _trepresent the elevation of this user, this satellite body-fixed coordinate system is (X _s, Y _s, Z _s), this user's body-fixed coordinate system is (X _u, Y _u, Z _u), this ECEF coordinate system conversion method is: the earth is a spheroid, major axis a=6.3781370e+03km, semiaxis b=6.356752314245179e+03km, excentricity $\mathrm{\ϵ}=\frac{\sqrt{a^2-b^2}}{2},$ Definition intermediate value ${\mathrm{Normal}}_1=\frac{a}{\sqrt{1-{\mathrm{\ϵ}}^2{\sin }^2{\mathrm{\λ}}_S}},$ ${\mathrm{Normal}}_2=\frac{a}{\sqrt{1-{\mathrm{\ϵ}}^2{\sin }^2{\mathrm{\λ}}_T}},$ Then body-fixed coordinate system (the X of this satellite _s, Y _s, Z _s) and the body-fixed coordinate system (X of this user _u, Y _u, Z _u) be respectively:

$\left\{\begin{array}{c}{X}_S=({\mathrm{Normal}}_1+{\mathrm{\κ}}_S){\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S\\ Y_S=({\mathrm{Normal}}_1+{\mathrm{\κ}}_S){\mathrm{cos\λ}}_S{\mathrm{sin\η}}_S\\ Z_S=\left({\mathrm{Normal}}_1\right(1-{\mathrm{\ϵ}}^2)+{\mathrm{\κ}}_S){\mathrm{sin\λ}}_S\end{array}\right.,\left\{\begin{array}{c}{X}_U=({\mathrm{Normal}}_2+{\mathrm{\κ}}_T){\mathrm{cos\λ}}_T{\mathrm{cos\η}}_T\\ Y_U=({\mathrm{Normal}}_2+{\mathrm{\κ}}_T){\mathrm{cos\λ}}_T{\mathrm{sin\η}}_T\\ Z_U=\left({\mathrm{Normal}}_2\right(1-{\mathrm{\ϵ}}^2)+{\mathrm{\κ}}_T){\mathrm{sin\λ}}_T\end{array}\right.;$

This local coordinate conversion method is: this satellite take user as the satellite local coordinate system (x in this local coordinate of origin at this _s, y _s, z _s) be:

$\left(\begin{array}{c}{x}_S\\ y_S\\ z_S\end{array}\right)=\left(\begin{array}{ccc}-{\mathrm{sin\η}}_S& {\mathrm{cos\η}}_S& 0\\ -{\mathrm{sin\λ}}_S{\mathrm{cos\η}}_S& -{\mathrm{sin\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{cos\λ}}_S\\ {\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{sin\λ}}_S\end{array}\right)\left(\begin{array}{c}{X}_S-X_U\\ Y_S-Y_U\\ Z_S-Z_U\end{array}\right)$

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, the locus of this satellite is: longitude 25.257000 ° ,-84.298000 °, latitude, elevation 783.871776 (km), the locus of this user is: longitude 28.917655 ° ,-99.345589 °, latitude, elevation 0.0 (km), through the conversion of this ECEF coordinate system, the satellite body-fixed coordinate system of this satellite in this ECEF coordinate system is (X _s: 6.439025706247130e+05, Y _s:-6.448794085263775e+06, Z _s: 3.039310992962332e+06), the user body-fixed coordinate system of this user in this ECEF coordinate system is (X _u:-9.073090474340906e+05, Y _u:-5.513096509229902e+06, Z _u: 3.065915618754285e+06), unit is rice; Through the conversion of this local coordinate, this customer location is the initial point of this local coordinate, and the position of this satellite in this local coordinate is x _s: 1.682569089856995e+06, y _s:-3.479338007695247e+05, z _s: 5.748030724080029e+05, unit is rice; Calculate elevation angle theta in real time through this, obtain θ=18.497377364764110 °.

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, this calculating satellite comprises relative to the method crossing the visible duration of top residue of user: (1) calculates satellite and crosses top duration, calculates from the minimum visible elevation angle theta of this user relative to this satellite ₀the moment occurred is θ (t to the current elevation angle _eduring)=θ the duration t of process _e, computing method are as follows:

${\mathrm{\γ}}_0={\cos }^{-1}\left(\frac{R_E}{R_S}{\mathrm{cos\θ}}_0\right)-{\mathrm{\θ}}_0$

$\mathrm{\ψ}\left(0\right)={\mathrm{\ψ}}_0={\cos }^{-1}\left(\frac{{\mathrm{cos\γ}}_0}{{\mathrm{cos\γ}}_m}\right)$

$\mathrm{\θ}\left(t_e\right)={\tan }^{-1}\left(\frac{cos\left(\right({\mathrm{\ω}}_E\cos i-{\mathrm{\ω}}_S)t_e+{\mathrm{\ψ}}_0)\×{\mathrm{cos\γ}}_m-\frac{R_E}{R_S}}{sin\left(cos\right(\left({\mathrm{\ω}}_E\cos i-{\mathrm{\ω}}_S\right)t_e+{\mathrm{\ψ}}_0)\×{\mathrm{cos\γ}}_m)}\right)$

Wherein, R _eand R _sbe respectively the radius of the earth and the orbit radius of this satellite, ω _eand ω _sbe respectively the earth and the angular velocity of this satellite under geocentric inertial coordinate system, γ ₀for the elevation angle of this user this satellite relative is θ ₀time this sub-satellite point to the angular distance of this user, calculate t _ethere are two solution values with ;

(2) calculate satellite and cross the visible duration t of top residue _r, computing method are as follows:

$t_c=\frac{1}{\mathrm{\ω}}{\cos }^{-1}\left(\frac{{\mathrm{cos\γ}}_0}{{\mathrm{cos\γ}}_m}\right),\mathrm{\ω}=({\mathrm{\ω}}_S-{\mathrm{\ω}}_E\cos i)/2$

Cross in visible another embodiment of duration prediction method of top residue at circular orbit low orbit satellite of the present invention, this mark angle γ _m=0.229891909587292 (rad), θ=18.497377364764110, R _e=6378.137km, ω _e=0.00007292115, κ _s=780km, i=86.4 °, θ ₀=8.2 °, μ=398601.58 (km ³/ s ²), r _s=R _e+ κ _s, calculate:

γ ₀＝0.347752358714960，ψ ₀＝0.264117321639204，

t _c＝5.089429154649187e+02，

$t_e^1=3.289250321990000e+02$ With $t_e^2=1.800178832660000e+02$ ，

If this satellite is away from this user, then this residue Covering time is:

$t_r=t_c-t_e^1=1.800178832659187e+02;$

If this satellite is near this user, then this residue Covering time is: $t_r=t_c-t_e^2=3.289250321989186e+02.$

The invention has the beneficial effects as follows: circular orbit low orbit satellite provided by the invention crosses the visible duration prediction method of top residue, satellite transit work characteristics can be made full use of, the continuous time information calculating renewal satellite spatial position and substar position and correspondence, and with the user's broadcast earthward of the form of easy navigation signal, after terrestrial user receives this navigation signal, the GPS receiving equipment calibration clock of recycling itself, determine customer location, the elevation angle of real-time calculating satellite and mark angle, and then solve satellite cross top residue visible duration, there is implementation method simply effective, few to the content requirements of navigation signal, ensure that accuracy and the high-timeliness of prediction duration, not adding users hardware cost again, application value is high.

Accompanying drawing explanation

Fig. 1 is the process flow diagram that circular orbit low orbit satellite of the present invention crosses a top residue visible duration prediction method embodiment;

Fig. 2 crosses user in the visible duration prediction method of top residue according to circular orbit low orbit satellite of the present invention to receive the process flow diagram of an embodiment of satellite navigation signals;

Fig. 3 is the composition diagram of the embodiment crossing navigation signal message in the visible duration prediction method of top residue according to circular orbit low orbit satellite of the present invention;

Fig. 4 crosses according to circular orbit low orbit satellite of the present invention the process flow diagram calculating an embodiment of satellite elevation angle in the visible duration prediction method of top residue;

Fig. 5 is the schematic diagram crossing the top visible duration prediction method Satellite of residue and a terrestrial user position relationship embodiment according to circular orbit low orbit satellite of the present invention;

Fig. 6 crosses according to circular orbit low orbit satellite of the present invention the process flow diagram calculating an embodiment at satellite mark angle in the visible duration prediction method of top residue;

Fig. 7 crosses in the visible duration prediction method of top residue according to circular orbit low orbit satellite of the present invention to calculate the process flow diagram that satellite remains an embodiment of visible duration.

Embodiment

For the ease of understanding the present invention, below in conjunction with the drawings and specific embodiments, the present invention will be described in more detail.Preferred embodiment of the present invention is given in accompanying drawing.But the present invention can realize in many different forms, is not limited to the embodiment described by this instructions.On the contrary, provide the object of these embodiments be make the understanding of disclosure of the present invention more comprehensively thorough.

It should be noted that, unless otherwise defined, all technology that this instructions uses and scientific terminology are identical with belonging to the implication that those skilled in the art of the present invention understand usually.The object of term used in the description of the invention just in order to describe specific embodiment is not for limiting the present invention.The term "and/or" that this instructions uses comprises arbitrary and all combinations of one or more relevant Listed Items.

Fig. 1 shows the process flow diagram that circular orbit low orbit satellite of the present invention crosses an embodiment of the visible duration prediction method of top residue.First, user receives satellite navigation signals S101, mainly terrestrial user received the navigation signal that top satellite sends, here satellite is not only the signal of communication as telstar transmission user, also to send navigation signal at a certain time interval, include the positional information of this satellite within following a period of time and the time information of correspondence in this navigation signal, pass through received satellite navigation signals and can judge that this satellite is satellite away from user or the satellite near user.Then, calculate the mark angle S103 of the elevation angle S102 of satellite relative to user user relative to calculating satellite, these two steps are that the satellite position information that the satellite navigation signals received based on the positional information residing for user and user comprises calculates, these two steps both can walk abreast and also can complete successively, and the elevation angle obtained and mark angle remain visible duration for the top of crossing calculating this satellite further.Finally, remain visible duration S104 by calculating satellite relative to the top of crossing of user, obtain satellite and push up the visible duration of residue relative to crossing of user.

Below by way of specific embodiment, above-mentioned steps S101, S102, S103 and S104 are described further.

Fig. 2 shows circular orbit low orbit satellite of the present invention and crosses user in the visible duration prediction method of top residue and receive the process flow diagram of an embodiment of satellite navigation signals.First, calculate and upgrade satellite spatial positional information S1011, wherein, calculate satellite spatial positional information to be completed by ground control centre, mainly calculate the positional information of satellite within following a period of time and (comprise longitude and latitude, elevation) and the time information of correspondence, the method calculated adopts SGP4 (SimpleGeneralPerturbationversion4) method, U.S. Department of Defense disclosed mathematical model and the code thereof of SGP4 in 1980, the method is predicted for the position of the flyer being less than to 255 minutes the cycle of operation in near-earth orbit, moving object particularly for height above sea level 1000 km has good forecast precision.Here following a period of time mainly refers to satellite after overhead, ground control centre to again through overhead, ground control centre during this period of time, and the frequency that ground control centre calculates is 1 time/second, namely calculated the locus at 1 satellite place every 1 second.Upgrading satellite spatial positional information is then when satellite arrives overhead, ground control centre, by ground control centre to the satellite spatial positional information in this satellite injection next following a period of time; Calculate satellite spatial position, by ground control centre by orbit prediction method, spatial positional information in the described satellite of regular calculating following a period of time and the time information of correspondence thereof, and at described satellite through overhead, described ground control centre, upgrade the spatial positional information of described satellite; Then, satellite calibration clock S1012, here, satellite utilizes the GPS device of self to carry out clock alignment, obtain clock information accurately, and the substar positional information of this satellite is determined in real time according to clock information and the satellite spatial positional information that stores in step S1011, this substar positional information refers to the substar latitude and longitude information in satellite current coverage range, and this is the positional information that the spatial positional information of satellite projects at the earth's surface; Then, satellite sends navigation signal S1013, mainly refer to the latitude and longitude information of track and the clock information of correspondence thereof that include this satellite each substar composition in its coverage in this navigation signal, and satellite sends navigation signal with the frequency of 1 time/second.Fig. 3 shows embodiment, wherein (a λ of the message structure of this navigation signal ₁, η ₁, κ ₁) represent at moment timestamp ₁longitude, latitude and elevation that satellite spatial position is corresponding, (λ ₂, η ₂, κ ₂) represent at moment timestamp ₂longitude, latitude and elevation that satellite spatial position is corresponding, by that analogy, (λ _n, η _n, κ _n) represent at moment timestamp _nlongitude, latitude and elevation that satellite spatial position is corresponding, and moment timestamp ₁, timestamp ₂timestamp _ncan be differ 1 second successively, represent that satellite sends navigation signal with the frequency of 1 time/second; User accepts navigation signal and judges satellite away from still close to S1014, terrestrial user constantly receives the navigation signal that satellite sends, wherein can utilize twice satellite navigation signals of continuous adjacent, calculate and judge that this satellite is close to user or away from user, this computation process is mainly judged by the distance of calculating two substar position relative user position, if second substar position is nearer than first substar positional distance customer location, then show that satellite is close to user, otherwise satellite is away from user.

By navigating shown in process flow diagram shown in Fig. 2 and Fig. 3, message structure figure can find out, the satellite that circular orbit low orbit satellite of the present invention is crossed in the visible duration prediction method of top residue can send navigation signal, but this navigation signal has simplification, only need comprise the longitude of satellite, the time information of latitude and elevation information and correspondence, can broadcast as a kind of broadcast singal of this satellite, can be mutually compatible with the communication system of this telstar, and do not need as the location navigation satellites such as GPS, send complicated satellite ephemeris information, therefore the more applicable this satellite for communicating.

Fig. 4 shows circular orbit low orbit satellite of the present invention and crosses the process flow diagram calculating an embodiment of satellite elevation angle in the visible duration prediction method of top residue.First, carry out ECEF coordinate system conversion S1021, exactly by spatial positional information (mainly the longitude and latitude of user and elevation information) the unified positional information be converted in solid (ECEF) coordinate system of ground heart according to the spatial positional information of present satellites and user, normally represent with 3 d space coordinate.The concrete conversion method of ECEF coordinate system is: the spatial positional information of satellite is (λ _s, η _s, κ _s), wherein λ _srepresent the longitude of this satellite, η _srepresent the latitude of this satellite, κ _srepresent the elevation of this satellite, the spatial positional information of user or customer position information are (λ _t, η _t, κ _t), wherein λ _trepresent the longitude of this user, η _trepresent the latitude of this user, κ _trepresent the elevation of this user, this satellite body-fixed coordinate system is (X _s, Y _s, Z _s), this user's body-fixed coordinate system is (X _u, Y _u, Z _u), the earth is a spheroid, major axis a=6.3781370e+03km, semiaxis b=6.356752314245179e+03km, excentricity $\mathrm{\ϵ}=\frac{\sqrt{a^2-b^2}}{2},$ Definition intermediate value ${\mathrm{Normal}}_1=\frac{a}{\sqrt{1-{\mathrm{\ϵ}}^2{\sin }^2{\mathrm{\λ}}_S}},{\mathrm{Normal}}_2=\frac{a}{\sqrt{1-{\mathrm{\ϵ}}^2{\sin }^2{\mathrm{\λ}}_T}},$ Then body-fixed coordinate system (the X of this satellite _s, Y _s, Z _s) and the body-fixed coordinate system (X of this user _u, Y _u, Z _u) be respectively:

$\left\{\begin{array}{c}{X}_S=({\mathrm{Normal}}_1+{\mathrm{\κ}}_S){\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S\\ Y_S=({\mathrm{Normal}}_1+{\mathrm{\κ}}_S){\mathrm{cos\λ}}_S{\mathrm{sin\η}}_S\\ Z_S=\left({\mathrm{Normal}}_1\right(1-{\mathrm{\ϵ}}^2)+{\mathrm{\κ}}_S){\mathrm{sin\λ}}_S\end{array}\right.,\left\{\begin{array}{c}{X}_U=({\mathrm{Normal}}_2+{\mathrm{\κ}}_T){\mathrm{cos\λ}}_T{\mathrm{cos\η}}_T\\ Y_U=({\mathrm{Normal}}_2+{\mathrm{\κ}}_T){\mathrm{cos\λ}}_T{\mathrm{sin\η}}_T\\ Z_U=\left({\mathrm{Normal}}_2\right(1-{\mathrm{\ϵ}}^2)+{\mathrm{\κ}}_T){\mathrm{sin\λ}}_T\end{array}\right.$

Such as, according to above-mentioned ECEF coordinate system conversion method, the locus of a certain moment satellite is: longitude 25.257000 °,-84.298000 °, latitude, elevation 783.871776 (km), the satellite body-fixed coordinate system be transformed in solid (ECEF) coordinate system of ground heart is (X _s: 6.439025706247130e+05, Y _s:-6.448794085263775e+06, Z _s: 3.039310992962332e+06), unit is rice.And for example, the locus of user is: longitude 28.917655 ° ,-99.345589 °, latitude, elevation 0.0 (km), is transformed into the user's body-fixed coordinate system (X in solid (ECEF) coordinate system of ground heart _u:-9.073090474340906e+05, Y _u:-5.513096509229902e+06, Z _u: 3.065915618754285e+06), unit is rice.

Then, carry out local coordinate conversion S1022, it is exactly the local coordinate situation according to the actual use of user, satellite in solid for ground heart in S1021 (ECEF) coordinate system and customer position information are converted to the satellite in local coordinate (ENU) and customer position information, this is actually the conversion of solid (ECEF) coordinate system of ground heart and local coordinate, during conversion, using the initial point of customer location as local coordinate, therefore in local coordinate user at the initial point of this coordinate system, be called user's origin, determine satellite position coordinates relative to user in local coordinate accordingly.Conversion method is: if the longitude of satellite is λ _s, latitude is η _s, this satellite is expressed as (X at the satellite body-fixed coordinate system of ECEF coordinate system _s, Y _s, Z _s), user's body-fixed coordinate system is (X _u, Y _u, Z _u), then the coordinate of this satellite in the local coordinate taking user as initial point is expressed as (x _s, y _s, z _s), then have:

$\left(\begin{array}{c}{x}_S\\ y_S\\ z_S\end{array}\right)=\left(\begin{array}{ccc}-{\mathrm{sin\η}}_S& {\mathrm{cos\η}}_S& 0\\ -{\mathrm{sin\λ}}_S{\mathrm{cos\η}}_S& -{\mathrm{sin\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{cos\λ}}_S\\ {\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{cos\λ}}_S{\mathrm{cos\η}}_S& {\mathrm{sin\λ}}_S\end{array}\right)\left(\begin{array}{c}{X}_S-X_U\\ Y_S-Y_U\\ Z_S-Z_U\end{array}\right)$

Such as, the locus of above-mentioned a certain moment satellite is: longitude λ _s25.257000 °, latitude η _sbe-84.298000 °, the satellite body-fixed coordinate system of this satellite in solid (ECEF) coordinate system of ground heart is (X _s: 6.439025706247130e+05, Y _s:-6.448794085263775e+06, Z _s: 3.039310992962332e+06), the user's body-fixed coordinate system in solid (ECEF) coordinate system of ground heart of user is (X _u:-9.073090474340906e+05, Y _u:-5.513096509229902e+06, Z _u: 3.065915618754285e+06), after being transformed into local coordinate, the coordinate of this satellite in the local coordinate taking user as initial point is: x _s: 1.682569089856995e+06, y _s:-3.479338007695247e+05, z _s: 5.748030724080029e+05, unit is rice; Real-time calculating elevation angle S1023, the elevation angle theta of current time user observation satellite is calculated in local coordinate, this elevation angle is real-time change, such as, can calculate current time θ=18.497377364764110 ° according to the local coordinate system position of above-mentioned satellite and user.

As can be seen here, in order to the elevation angle of accurate Calculation user observation satellite, by carrying out coordinate conversion to satellite spatial position and user's space position, the two is converted in same coordinate system and calculates the elevation angle, contribute to like this eliminating the error brought because coordinate system is different, and have fixing geometric transformation relation between selected coordinate system after, computing machine can be utilized automatically to change, improve conversion counting yield.

Fig. 5 shows the schematic diagram that circular orbit low orbit satellite of the present invention crosses the top visible duration prediction method Satellite of residue and a terrestrial user position relationship embodiment.Wherein, the locus of the substar in satellite S current coverage range is represent with vector form, illustrate that satellite is in the state being in constantly motion, its substar H position is (λ _s, η _s, κ ' _s), the locus of user T is (λ _t, η _t, κ _t), γ (t) represents the shortest arc distance of user T to present satellites S substar H, and Q point represents when satellite S is through user T overhead, this satellite at the sub-satellite track of the earth to the shortest arc distance of user T and the intersection point of this sub-satellite track, γ _mfor the mark angle of user T, ψ (t) represents that present satellites S substar H is with the shortest arc distance between Q point, and θ is the elevation angle of active user's T-phase to satellite S.

Fig. 6 shows circular orbit low orbit satellite of the present invention and crosses the process flow diagram calculating an embodiment at satellite mark angle in the visible duration prediction method of top residue.First, determine the positional information S1031 of user, longitude and latitude and the elevation information of this user can be calculated by the GPS that subscriber equipment is built-in, be expressed as user's longitude λ _t, user's latitude η _twith user's elevation κ _t; Clock alignment S1032, obtains the higher satellite clock of precision, by consistent with satellite clock for user clock synchronous calibration from the satellite navigation signals received; Calculate mark angle S1033, from the satellite navigation signals received, obtain the substar positional information of this satellite in real time, i.e. substar latitude and longitude information, is expressed as sub-satellite point longitude with dynamic vector form sub-satellite point latitude show that this substar longitude and latitude is in time in continuous change, the mark angle of user's relative satellite is expressed as γ _m, first calculate:

$\mathrm{\Ω}=\[\sqrt{{\sin }^2\left(\frac{\stackrel{\→}{{\mathrm{\λ}}_S}-{\mathrm{\λ}}_T}{2}\right)+cos\left(\stackrel{\→}{{\mathrm{\λ}}_S}\right)\×cos\left({\mathrm{\λ}}_T\right)\×{\sin }^2\left(\frac{\stackrel{\→}{{\mathrm{\η}}_S}-{\mathrm{\η}}_T}{2}\right)}\]$

Calculate further again:

γ _m＝min{2sin ^-1(Ω)}

Such as, as the longitude λ of user _t=28.917655 °, latitude η _t=-99.345589 ° constant when, and the substar longitude of satellite and latitude cross top motion along with satellite and constantly change, the Ω value that the longitude and latitude of substar changes situation and correspondence and calculates can be represented, as shown in table 1 part by the form of list.Then, then its arcsin function is asked to all Ω values, find minimum value wherein, then be multiplied by 2, finally can obtain γ _m=0.229891909587292 (rad).

Substar longitude and latitude and Ω value situation of change when table 1 satellite crosses top

Fig. 7 show circular orbit low orbit satellite of the present invention cross top residue visible duration prediction method in calculate the process flow diagram that satellite remains an embodiment of visible duration.First, calculate and cross top duration S1034, calculate that to see that the moment of satellite plays the current elevation angle from user be θ (t exactly _e)=θ time institute elapsed time t _e, user sees that the moment of satellite refers to the minimum visible elevation angle theta of user's relative satellite here ₀in the moment occurred, computing method are as follows:

${\mathrm{\γ}}_0={\cos }^{-1}\left(\frac{R_E}{R_S}{\mathrm{cos\θ}}_0\right)-{\mathrm{\θ}}_0$

$\mathrm{\ψ}\left(0\right)={\mathrm{\ψ}}_0={\cos }^{-1}\left(\frac{{\mathrm{cos\γ}}_0}{{\mathrm{cos\γ}}_m}\right)$

$\mathrm{\θ}\left(t_e\right)={\tan }^{-1}\left(\frac{cos\left(\right({\mathrm{\ω}}_E\cos i-{\mathrm{\ω}}_S)t_e+{\mathrm{\ψ}}_0)\×{\mathrm{cos\γ}}_m-\frac{R_E}{R_S}}{sin\left(cos\right(\left({\mathrm{\ω}}_E\cos i-{\mathrm{\ω}}_S\right)t_e+{\mathrm{\ψ}}_0)\×{\mathrm{cos\γ}}_m)}\right)$

Wherein, R _efor earth radius, R _sfor the circuit orbit radius of satellite, ω _ethe angular velocity of the earth, ω _sthe angular velocity of satellite, γ ₀for user is θ at the elevation angle of relative satellite ₀time sub-satellite point to the angular distance of user.

Can find out, t _eby negating, tangent trigonometric function obtains, and angle value corresponding to same tan has two, therefore corresponding t _ethere are two solutions with .

Further, calculated top residual time length S1034, residual time length is expressed as t _r, computing method are as follows:

$t_c=\frac{1}{\mathrm{\ω}}{\cos }^{-1}\left(\frac{{\mathrm{cos\γ}}_0}{{\mathrm{cos\γ}}_m}\right),\mathrm{\ω}=({\mathrm{\ω}}_S-{\mathrm{\ω}}_E\cos i)/2$

Such as, at mark angle γ _mwhen=0.229891909587292 (rad) and θ=18.497377364764110, R _e=6378.137km, ω _e=0.00007292115, R _s=R _e+ κ _s, κ _s=780km, i=86.4 °, θ ₀=8.2 °, ω=(ω _s-ω _ecosi)/2, μ=398601.58 (km ³/ s ²), then can obtain:

γ ₀＝0.347752358714960，ψ ₀＝0.264117321639204，

t _c＝5.089429154649187e+02，

$t_e^1=3.289250321990000e+02$ With $t_e^2=1.800178832660000e+02$ , unit is second, if satellite is away from user, then remaining Covering time is:

$t_r=t_c-t_e^1=1.800178832659187e+02;$

If satellite is near user, then remaining Covering time is:

$t_r=t_c-t_e^2=3.289250321989186e+02.$

The visible duration prediction method of top residue is crossed by circular orbit low orbit satellite of the present invention, can according to satellite transit work characteristics, by constantly calculating and upgrading satellite spatial position, the time information of substar position and correspondence, and with the form of easy navigation signal earthward user broadcast these information, after terrestrial user receives this navigation signal, the GPS receiving equipment calibration clock of recycling itself, determine customer location, the elevation angle of the relative user of real-time calculating satellite and mark angle, and then solve satellite cross top residue visible duration, there is implementation method simply effective, few to the content requirements of navigation signal, low to satellite ephemeris dependency degree, under the accuracy guaranteeing to predict duration and high-timeliness prerequisite, not adding users hardware cost again.

The foregoing is only embodiments of the invention; not thereby the scope of the claims of the present invention is limited; every equivalent structure transformation utilizing instructions of the present invention and accompanying drawing content to do, or be directly or indirectly used in other relevant technical fields, include in scope of patent protection of the present invention.

Claims (10)

1. circular orbit low orbit satellite crosses the visible duration prediction method of top residue, it is characterized in that, comprises the steps:

The first step, user receives satellite navigation signals;

Second step, calculates the elevation angle of satellite;

3rd step, calculates the mark angle of satellite;

4th step, calculates satellite and remains visible duration relative to the top of crossing of user.

2. circular orbit low orbit satellite according to claim 1 crosses the visible duration prediction method of top residue, and it is characterized in that, described user receives satellite navigation signals and comprises:

(1) satellite spatial position is calculated, by ground control centre by orbit prediction method, spatial positional information in the described satellite of regular calculating following a period of time and the time information of correspondence thereof, and at described satellite through overhead, described ground control centre, upgrade the spatial positional information of described satellite;

(2) satellite calibration clock, described satellite utilizes self GPS device to carry out clock alignment, and according to the clock information of calibration and the spatial positional information of described satellite of storage, the substar latitude and longitude information in described satellite current coverage range is determined in adjustment in real time;

(3) satellite sends navigation signal, and described satellite regularly sends navigation signal, and described navigation signal comprises the substar latitude and longitude information of described satellite in its coverage and the time information of correspondence thereof;

(4) user receives navigation signal, judge that described satellite is close to user or away from user, described user receives the double described navigation signal that described satellite sends, and judges described satellite close to still away from described user according to the substar latitude and longitude information in described navigation signal relative to the distance change of customer location.

3. circular orbit low orbit satellite according to claim 2 crosses the visible duration prediction method of top residue, and it is characterized in that, described orbit prediction method is SGP4 method; The frequency of the spatial positional information in the described satellite of described regular calculating following a period of time and the time information of correspondence thereof is at interval of calculating 1 time 1 second; The frequency that described satellite regularly sends navigation signal is at interval of sending 1 time 1 second.

4. the circular orbit low orbit satellite according to any one of claims 1 to 3 crosses the visible duration prediction method of top residue, and it is characterized in that, the method at the mark angle of described calculating satellite comprises:

(1) determine described customer position information, obtain the longitude λ of described user self _t, latitude η _twith elevation κ _tinformation;

(2) clock alignment, according to time information described in received navigation signal, by consistent with satellite clock for user clock synchronous calibration;

(3) calculate mark angle in real time, described user, by the described navigation signal received, calculates the mark angle γ of the relatively described satellite of described user _m, computing method are as follows:

γ _m＝min{2sin ^-1(Ω)}

Wherein, with be longitude and the latitude vector of the substar in the described satellite coverage that comprises in described navigation signal, Ω also be vectorial.

5. circular orbit low orbit satellite according to claim 4 crosses the visible duration prediction method of top residue, it is characterized in that, is obtained the longitude λ of described user self by the GPS device measurement that described user is built-in _t, latitude η _twith elevation κ _tinformation.

6. circular orbit low orbit satellite according to claim 5 crosses the visible duration prediction method of top residue, and it is characterized in that, the method at the elevation angle of described calculating satellite comprises:

(1) ECEF coordinate system conversion, is converted to ECEF coordinate system Satellite body-fixed coordinate system and user's body-fixed coordinate system by the locus of current described satellite and customer position information unification;

(2) local coordinate conversion, is converted to the satellite local coordinate system in local coordinate and user's origin by described satellite body-fixed coordinate system and described user's body-fixed coordinate system;

(3) calculate the elevation angle in real time, utilize described satellite local coordinate system and described user's origin, in described local coordinate, calculate the elevation angle theta of the relatively described satellite of described user.

7. circular orbit low orbit satellite according to claim 6 crosses the visible duration prediction method of top residue, and the spatial positional information of described satellite is (λ _s, η _s, κ _s), wherein λ _srepresent the longitude of described satellite, η _srepresent the latitude of described satellite, κ _srepresent the elevation of described satellite, described customer position information is (λ _t, η _t, κ _t), wherein λ _trepresent the longitude of described user, η _trepresent the latitude of described user, κ _trepresent the elevation of described user, described satellite body-fixed coordinate system is (X _s, Y _s, Z _s), described user's body-fixed coordinate system is (X _u, Y _u, Z _u), it is characterized in that, described ECEF coordinate system conversion method is:

The earth is a spheroid, major axis a=6.3781370e+03km, semiaxis b=6.356752314245179e+03km, excentricity definition intermediate value then body-fixed coordinate system (the X of described satellite _s, Y _s, Z _s) and the body-fixed coordinate system (X of described user _u, Y _u, Z _u) be respectively:

Described local coordinate conversion method is: the satellite local coordinate system (x of described satellite in the described described local coordinate taking user as origin _s, y _s, z _s) be:

。

8. circular orbit low orbit satellite according to claim 7 crosses the visible duration prediction method of top residue, it is characterized in that, the locus of described satellite is: longitude 25.257000 ° ,-84.298000 °, latitude, elevation 783.871776 (km), the locus of described user is: longitude 28.917655 ° ,-99.345589 °, latitude, elevation 0.0 (km), change through described ECEF coordinate system, described satellite body-fixed coordinate system is (X _s: 6.439025706247130e+05, Y _s:-6.448794085263775e+06, Z _s: 3.039310992962332e+06), described user's body-fixed coordinate system is (X _u:-9.073090474340906e+05, Y _u:-5.513096509229902e+06, Z _u: 3.065915618754285e+06), unit is rice; Change through described local coordinate, described customer location is the initial point of described local coordinate, and the position of described satellite in described local coordinate is x _s: 1.682569089856995e+06, y _s:-3.479338007695247e+05, z _s: 5.748030724080029e+05, unit is rice; Through described real-time calculating elevation angle theta, obtain θ=18.497377364764110 °.

9. the circular orbit low orbit satellite according to claim 6 or 7 crosses the visible duration prediction method of top residue, it is characterized in that, described calculating satellite comprises relative to the method crossing the visible duration of top residue of user:

(1) calculate satellite and cross top duration, calculate the minimum visible elevation angle theta from the relatively described satellite of described user ₀the moment occurred is θ (t to the current elevation angle _eduring)=θ the duration t of process _e, computing method are as follows:

Wherein, R _eand R _sbe respectively the radius of the earth and the orbit radius of described satellite, ω _eand ω _sbe respectively the earth and the angular velocity of described satellite under geocentric inertial coordinate system, γ ₀for the elevation angle of the relatively described satellite of described user is θ ₀shi Suoshu sub-satellite point, to the angular distance of described user, calculates t _ethere are two solution values with

(2) calculate satellite and cross the visible duration t of top residue _r, computing method are as follows:

ω＝(ω _S-ω _Ecosi)/2

。

10. circular orbit low orbit satellite according to claim 9 crosses the visible duration prediction method of top residue, it is characterized in that, described mark angle γ _m=0.229891909587292 (rad), θ=18.497377364764110, R _e=6378.137km, ω _e=0.00007292115, κ _s=780km, i=86.4 °, θ ₀=8.2 °, μ=398601.58 (km ³/ s ²),

R _s=R _e+ κ _s, calculate:

γ ₀＝0.347752358714960，ψ ₀＝0.264117321639204，

t _c＝5.089429154649187e+02，

with unit is second, if described satellite is away from described user, then described residue Covering time is:

If described satellite is near described user, then described residue Covering time is: