CN105572705A - Navigation device - Google Patents

Abstract

The invention discloses a navigation device comprising a capture tracking processing module, a message extraction module, an observation quantity extraction module, a positioning calculation module and an output module. Continuous tracking of satellite signals is performed through the capture tracking processing module so that navigation telegraph bit data of a satellite are obtained. Satellite navigation telegraphs and observation quantity are acquired through the message extraction module and the observation quantity extraction module, and then positioning calculation is performed through the positioning calculation module and outputting is performed through the output module according to a binary format so that the navigation telegraph bit data of the satellite can be automatically and continuously acquired for positioning, the situation that navigation is not timely can be avoided, and thus the navigation device has the technical effects of timeliness and accuracy.

Description

A kind of guider

Technical field

The application relates to technical field of satellite communication, particularly relates to a kind of guider.

Background technology

Along with the development of science and technology, communication technology of satellite have also been obtained development at full speed, and the kind of electronic product also gets more and more, and people have also enjoyed the various facilities that development in science and technology brings.Present people can pass through various types of electronic equipment, enjoy the comfortable life along with development in science and technology brings.

Such as navigating instrument is the product of modern science and technology, navigating instrument is used on automobile, for location, navigation and amusement, along with the universal of automobile and the construction of road, the economic interaction of intercity is more frequent, automatic navigator seems very important, and the navigation that accurately location, navigation, amusement function roll into one more can meet the demand of car owner, becomes the basic equipment on car.

Such as, but the navigation of current navigating instrument out of true, often there will be navigation way mistake or navigation situation not in time, occur that navigation postpones, and can not obtain follow-up course in time, make car owner be at a loss in the process of navigation; Or by car owner's misguidance in the sea, increase car owner dangerous.

Summary of the invention

For solving the problems of the technologies described above, the invention provides a kind of guider, comprising:

Acquisition and tracking processing module, continues to follow the tracks of for carrying out satellite-signal, and obtains the navigation message bit data of satellite;

Message extraction module, for receiving described navigation message bit data, and goes out satellite navigation message from described navigation message bit extracting data;

Observed quantity extraction module, for going out observed quantity from described navigation message bit extracting data;

Positioning calculation module, for carrying out co-located according to described observed quantity and described satellite navigation message, obtains navigation results;

Output module, for exporting described navigation results according to binary format.

Preferably, described satellite navigation message comprises ephemeris, satellite health, time device, the satellite almanac of satellite.

Preferably, described observed quantity extraction module, specifically for obtaining the pseudo range observed quantity of tracked satellite or Doppler measurements or carrier phase observed quantity according to Base-Band Processing data.

Preferably, described pseudo range observed quantity is utilize the transmission time of satellite-signal to be multiplied by transmission range that the light velocity obtains signal; The transmission time of described satellite-signal refers to satellite-signal transmitting time t _svdeduct receiver local zone time t _u.

Preferably, described Doppler measurements is after the frequency values obtaining carrier wave NCO, by the frequency values of described carrier wave NCO and the nominal intermediate frequency value f of radio-frequency front-end _csubtract each other acquisition.

Preferably, described positioning calculation module specifically comprises:

Computing module, for going out the position of satellite according to the ephemeris computation of described satellite;

Positioning calculation module, to position customer location for utilizing the position of described satellite and described observed quantity and resolves.

Preferably, described computing module specifically for:

The first step: the mean angular velocity n calculating satellite motion;

Second step: the mean anomaly M of calculating observation moment satellite _s;

M _s=M ₀+ n (t-t _oe); In formula, M ₀for reference moment t _oetime mean anomaly, provided by broadcast ephemeris;

3rd step: calculate eccentric anomaly: E _s=M _s+ e _ssinE _s;

4th step: calculate anomaly f _s: $f_s=srctan\frac{\sqrt{1-{e_s}^2}{\mathrm{sinE}}_s}{{\mathrm{cosE}}_s-e_s};$

5th step: calculate to rise and hand over elongation in formula, ω is the argument of perigee, is provided by broadcast ephemeris;

6th step: calculate perturbation correction member;

7th step: to a liter friendship elongation satellite radius vector r, orbit inclination i carry out perturbation and correct;

8th step: calculate the position of satellite in track areal coordinate system;

9th step: the longitude L of calculating observation moment ascending node;

Tenth step: calculate the position of satellite in spherical coordinate system instantaneously;

11 step: calculate the position of satellite in conventional terrestrial coordinate system.

Preferably, described mean angular velocity n is obtained by following steps:

First according to the parameter provided in broadcast ephemeris calculate with reference to moment t _oemean angular velocity n ₀: in formula, GM is the product of gravitational constant G and earth gross mass M, GM=3.9860047 × 1014m ³/ s ², then according to the mean angular velocity n:n=n of perturbation parameter Δ n calculating observation moment satellite given in broadcast ephemeris ₀+ Δ n.

Preferably, described calculating perturbation correction member comprises the following steps:

By the C provided in broadcast ephemeris _uc, C _us, C _rc, C _rs, C _ic, C _is6 perturbation parameters, obtain the liter caused due to the second order zonal harmonic coefficient item of earth gravitational field bit function and hand over elongation perturbation correction member δ _μ, satellite radius vector r perturbation correction member δ _r, and the perturbation correction member δ of inclination of satellite orbit i _i, computing formula is as follows:

${\mathrm{\δ}}_{\mathrm{\μ}}=C_{uc}cos2\widetilde{\mathrm{\μ}}+C_{us}sin2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_r=C_{rc}cos2\widetilde{\mathrm{\μ}}+C_{rs}sin2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_i=C_{ic}cos2\widetilde{\mathrm{\μ}}+C_{is}sin2\widetilde{\mathrm{\μ}}.$

Preferably, described to liter handing over an elongation μ, satellite radius vector r, orbit inclination i to carry out to perturb the step corrected be:

$\widetilde{\mathrm{\μ}}=\widetilde{\mathrm{\μ}}+{\mathrm{\δ}}_{\mathrm{\μ}}$

r＝a _s(1-e _scosE _s)+δ _r，

$i=i_0+{\mathrm{\δ}}_i+\frac{di}{dt}(t-t_{oe})$

In formula: a _sfor the major radius of satellite orbit, i ₀for t _oethe orbit inclination in moment, is provided by the kepler parameters in broadcast ephemeris; for the rate of change of i, provided by the perturbation parameter in broadcast ephemeris.

By one or more technical scheme of the present invention, the present invention has following beneficial effect or advantage:

The invention provides a kind of guider, comprising: acquisition and tracking processing module, message extraction module, observed quantity extraction module, positioning calculation module, output module.Carry out continuing to follow the tracks of the navigation message bit data obtaining satellite to satellite-signal by acquisition and tracking processing module; By message extraction module, observed quantity extraction module, obtain satellite navigation message and observed quantity, then positioned by positioning calculation module and resolve, and exported according to binary format by output module, the navigation message bit data of acquisition satellite that can be automatic continuous position, and then avoid occurring that navigation situation not in time occurs, and have technique effect promptly and accurately.

Accompanying drawing explanation

Fig. 1 is the structural representation of guider in the embodiment of the present invention.

Embodiment

In order to make the application the technical staff in the technical field more clearly understand the application, below in conjunction with accompanying drawing, by specific embodiment, technical scheme is described in detail.

Guider is mainly divided into four parts principle: acquisition and tracking processing section, message and observed quantity Extraction parts, positioning calculation and system export.

Based on above-mentioned principle, the structure of guider as shown in Figure 1.Guider mainly comprises: acquisition and tracking processing module 11, message extraction module 12, observed quantity extraction module 13, positioning calculation module 14, output module 15.

Acquisition and tracking processing module 11, continues to follow the tracks of for carrying out satellite-signal, and obtains the navigation message bit data of satellite;

Message extraction module 12, for receiving described navigation message bit data, and goes out satellite navigation message from described navigation message bit extracting data;

Observed quantity extraction module 13, for going out observed quantity from described navigation message bit extracting data;

Positioning calculation module 14, for carrying out co-located according to described observed quantity and described satellite navigation message, obtains navigation results;

Output module 15, for exporting described navigation results according to binary format.

Introduce the concrete function of modules below respectively.

In acquisition and tracking processing module 11, acquisition and tracking data processing comprises ALT-CH alternate channel state, adjustment is followed the tracks of carrier loop and code loop circuit state, maintains correct tracking mode, drawn the information such as correct Doppler, code phase and carrier phase, also can carry out the navigation message bit data of bit synchronous and satellite.

In the measurement moment, acquisition and tracking data processing module can encapsulate the observation value obtained, and passes to next unit (message extraction module 12, observed quantity extraction module 13) by system interface, carries out message and resolves and extract with observed quantity.

Message extraction module 12 for receiving described navigation message bit data, and goes out the navigation messages such as satellite ephemeris, almanac, UTC parameter, Ionospheric Parameters from described navigation message bit extracting data.And observed quantity extraction module 13, then mainly try to achieve the observed quantities such as the pseudorange of satellite, Doppler, carrier phase.

Lower mask body introduces message extraction module 12 and observed quantity extraction module 13.

Message extraction module 12.

Message extraction module 12, mainly decodes to bit stream information according to GPS and BD2 (a kind of positioning system) interface control document, obtains satellite navigation message.Satellite navigation message provides the data such as ephemeris, satellite health, time system, satellite almanac of concerned satellites for user, is that user positions and navigate necessary basic data.For the structure of different global position system navigation message, need to adopt different decoding processes, correctly to obtain navigation message efficiently.

On the L1 wave band of GPS, the transfer rate of people's code navigation message is 50bit/s, with " frame " for unit outwards sends.The length of every frame is 1500bit, has broadcast a frame and has needed 30s.Every frame comprises 5 subframes, and each subframe all comprises 300bit, and the time of broadcasting is 6s.Each subframe can be divided into again 10 words, and each word is made up of 30bit.First, second and third subframe of navigation message is the ephemeris parameter of satellite, and every 30s repeats once, and its content upgrades once every 2h.Fourth, fifth subframe is the almanac data of satellite, and every 30s overturns one page, and its content includes 25 pages, so the navigation message sending complete set needs 12.5min, and then repeats.

The Z counting of leading character, subframe is comprised in gps navigation message bit stream information, and the content such as the ephemeris of satellite, almanac data.Bit stream is synchronously searched for, until find leading character " 10001011b ".Then according to checking algorithm, information is verified, if not by verification, abandon this data; If by verification, then current data is effective, according to navigation message form, can parse Ionospheric Parameters, ephemeris and almanac data etc.

Extraction and the GPS of BD2 navigation message are similar, also comprise and separate frame structure and information analysis.

Observed quantity extraction module 13.

Observed quantity extraction module 13, obtains the observed quantity such as pseudorange, Doppler, carrier phase of tracked satellite according to Base-Band Processing data.

Pseudo range observed quantity is utilize the transmission time of satellite-signal to be multiplied by transmission range that the light velocity obtains signal.Here transmission time refers to satellite-signal transmitting time t _svdeduct receiver local zone time t _u.Local zone time t _udirectly provided by receiver local clock, and satellite-signal transmitting time t _svneed the counting by subframe or word string, and could accurately calculate after providing local pseudo-code phase by code tracking loop.

The extraction of Doppler measurements is after the frequency values obtaining carrier wave NCO, by the nominal intermediate frequency value f of itself and radio-frequency front-end _ccompare, its difference is exactly Doppler shift observed quantity.

Positioning calculation module 14.

Positioning calculation module 14, carries out co-located according to observed quantity information and the navigation message obtained, and utilizes filtering algorithm to make positioning result more reliable and more stable, meet high sensitivity demand.

Positioning calculation module 14 specifically comprises:

Computing module, for going out the position of satellite according to the ephemeris computation of described satellite;

Positioning calculation module 14, to position customer location for utilizing the position of described satellite and described observed quantity and resolves.

Lower mask body introduces the concrete enforcement principle of computing module, and its distance resolving satellite is carried out step by step, in addition, and can the distance of resolving GPS satellite or the distance of BD2 (a kind of positioning system) satellite.

1) calculating of gps satellite locus

The first step: the mean angular velocity n calculating satellite motion.

First, according to the parameter provided in broadcast ephemeris calculate with reference to moment t _oemean angular velocity n ₀: in formula, GM is the product of gravitational constant G and earth gross mass M, and its value is GM=3.9860047 × 1014m ³/ s ².Then according to the mean angular velocity n:n=n of perturbation parameter Δ n calculating observation moment satellite given in broadcast ephemeris ₀+ Δ n.

Second step: the mean anomaly M of calculating observation moment satellite _s.

The cycle of operation due to satellite is 12 hours, adopts time of satellite perigee passing t ₀when calculating mean anomaly M, extrapolating, interval is maximum reaches 6 hours.And broadcast ephemeris upgrades once for every 2 hours, with reference to when being constantly located at the central moment, extrapolation interval is less than or equal to 1 hour.So use t _oereplace time of satellite perigee passing t ₀after, extrapolation interval will reduce greatly, also can obtain the higher result of precision with better simply model.

M _s＝M ₀+n(t-t _oe)

In formula, M ₀for reference moment t _oetime mean anomaly, provided by broadcast ephemeris.

3rd step: calculate eccentric anomaly.

The Kepler's equation represented with radian is: E _s=M _s+ e _ssinE _s.

4th step: calculate true anomaly.

According to Keplerian orbit equation, anomaly f can be obtained _swith eccentric anomaly E _sbetween relation:

${\mathrm{cosf}}_s=\frac{{\mathrm{cosE}}_s-e_s}{1-e_s{\mathrm{cosE}}_s}$

in formula, e _sfor the excentricity of satellite orbit, provided by broadcast ephemeris.True anomaly can be obtained thus and calculate common-used formula:

5th step: calculate to rise and hand over elongation

Utilize formula calculate to rise and hand over elongation, in formula, ω is the argument of perigee, is provided by broadcast ephemeris.

6th step: calculate perturbation correction member.

C is given in broadcast ephemeris _uc, C _us, C _rc, C _rs, C _ic, C _is6 perturbation parameters, can obtain the liter caused due to the second order zonal harmonic coefficient item of earth gravitational field bit function accordingly and hand over elongation perturbation correction member δ _μ, satellite radius vector r perturbation correction member δ _r, and the perturbation correction member δ of inclination of satellite orbit i _i.Computing formula is as follows:

${\mathrm{\δ}}_{\mathrm{\μ}}=C_{uc}cos2\widetilde{\mathrm{\μ}}+C_{us}sin2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_r=C_{rc}cos2\widetilde{\mathrm{\μ}}+C_{rs}sin2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_i=C_{ic}cos2\widetilde{\mathrm{\μ}}+C_{is}sin2\widetilde{\mathrm{\μ}},$

7th step: to a liter friendship elongation satellite radius vector r, orbit inclination i carry out perturbation and correct.

$\widetilde{\mathrm{\μ}}=\widetilde{\mathrm{\μ}}+{\mathrm{\δ}}_{\mathrm{\μ}},$

r＝a _s(1-e _scosE _s)+δ _r，

$i=i_0+{\mathrm{\δ}}_i+\frac{di}{dt}(t-t_{oe}),$

In formula: a _sfor the major radius of satellite orbit, i ₀for t _oethe orbit inclination in moment, is provided by the kepler parameters in broadcast ephemeris. for the rate of change of i, provided by the perturbation parameter in broadcast ephemeris.

8th step: calculate the position of satellite in track areal coordinate system.

In track plane right-angle coordinate, the planimetric rectangular coordinates of (true origin is positioned at the earth's core, and X-axis points to ascending node) satellite is:

x _p＝rcosμ

y _p＝rsinμ

9th step: the longitude L of calculating observation moment ascending node.

If with reference to moment t _oetime ascending node right ascension be ascending node is Δ Ω to the rate of change of time, and so the right ascension of ascending node Ω of observation moment t should be:

$\mathrm{\Ω}={\mathrm{\Ω}}_{t_{oe}}+\mathrm{\Δ}\mathrm{\Ω}(t-t_{oe}),$

In formula, Δ Ω can provide from the perturbation parameter of broadcast ephemeris.If start time this week (when Sun. 0) Greenwich sidereal time is GASTweek, then observe moment Greenwich sidereal time be:

GAST＝GAST _week+ω _et，

ω in formula _efor rotational-angular velocity of the earth, its value is ω _e=7.292115 × l0-5rad/s, t are the time (s) in this week.The longitude so just can trying to achieve observation moment ascending node is:

$L=\mathrm{\Ω}-GAST={\mathrm{\Ω}}_{t_{oe}}-{\mathrm{GAST}}_{week}+\mathrm{\Δ}\mathrm{\Ω}(t-t_{oe})-{\mathrm{\ω}}_{e}t,$

Order ${\mathrm{\Ω}}_0={\mathrm{\Ω}}_{t_{oe}}-{\mathrm{GAST}}_{week},$ Then have:

L＝Ω ₀+ΔΩ(t-t _oe)-ω _et＝Ω ₀+(ΔΩ-ω _e)t-ΔΩt _oe。

Tenth step: calculate the position of satellite in spherical coordinate system instantaneously.

After the geodetic longitude L of known ascending node and the inclination angle i of orbit plane, just try to achieve the position of satellite in body-fixed coordinate system by twice rotation:

$\left[\begin{array}{c}X\\ Y\\ Z\end{array}\right]=R_Z(-L)R_X(-i)\left[\begin{array}{c}x\\ y\\ z\end{array}\right]=\left[\begin{array}{c}x\cos L-y\cos i\sin L\\ x\cos L+y\cos i\sin L\\ y\sin L\end{array}\right].$

11 step: calculate the position of satellite in conventional terrestrial coordinate system.

The observation moment position of satellite in conventional terrestrial coordinate system

${\left[\begin{array}{c}x\\ y\\ z\end{array}\right]}_{CTS}=R_Y(-x_p)R_X(-y_p)\left[\begin{array}{c}X\\ Y\\ Z\end{array}\right]=\left[\begin{array}{ccc}1& 0& x_p\\ 0& 1& -y\\ -x_p& y_p& 1\end{array}\right]\left[\begin{array}{c}X\\ Y\\ Z\end{array}\right],$

The form of BD2 satellite broadcasting ephemeris and the broadcast ephemeris of gps satellite substantially similar, therefore the calculating of its satellite spatial position is also similar with upper joint, repeats no more here.

Positioning calculation module 14.

Positioning calculation module 14, uses satellite position and pseudo range observed quantity, is calculated the position of receiver current time by solving equation.If the position of known satellite i is r _i(x _i, y _i, z _i), the pseudorange of this satellite is p _iif receiver location is r _u(x _u, y _u, z _u), receiver clock-offsets is b _u.For bimodulus positioning system, system time difference Δ T also as a unknown quantity, then can obtain measuring equation:

$p_i=\sqrt{{(x_i-x_u)}^2+{(y_i-y_u)}^2+{(z_i-z_u)}^2}+b_u+\mathrm{\Δ}T,$

For bimodulus positioning system, the position of known at least five satellites and pseudorange, substitute into above formula and do linearization, can combine and solve, obtain r _u(x _u, y _u, z _u) and the value of receiver clock-offsets.

Based on above-mentioned location ultimate principle, in software program, least square and kalman filter method is used to carry out solving of receiver location and speed.Then, positioning result is detected.If by this detection, then export this positioning result, witness marker puts 1, otherwise does not export this positioning result, and witness marker sets to 0.

Least square method is the basic skills of positioning calculation, and it can try to achieve optimum solution containing between error and each measured value of noise, makes the sum of square of deviations of all measured values minimum.

Meet at least four stars under single system station-keeping mode, under dual system station-keeping mode when at least five stars, according to the method that GPS and BD2 interface control document provides, utilize ephemeris parameter to calculate position and the speed of participation position location satellite.Owing to comprising satellite clock correction and ionosphere, tropospheric delay error in pseudorange, ephemeris parameter before positioning calculation, is first utilized to calculate satellite clock correction; For GPS and BD2 satellite, all ionosphere delay can be calculated according to ionosphere delay parameter; And estimate according to correlation model tropospheric delay.These correcting values are utilized to correct pseudorange afterwards.After being rejected by pseudorange larger for measuring error, Newton iteration method is used to solve, until iteration terminates.

Kalman filtering algorithm, it is a kind of Linear Minimum Variance Filter method, it adopts state equation and measurement equation to describe stochastic linear system, according to the criterion that the variance of estimated state error is minimum, the measuring value from measured noise pollution, estimate each state in each moment of system in real time.It comprises prediction and revises two processes.Kalman filter adopts recursive form in computing method, is easy to realize on computers.

If the system state equation after discretize and measurement equation are respectively:

$\left\{\begin{array}{c}{X}_k={\mathrm{\Φ}}_{k|k-1}{X}_{k-1}+{\mathrm{\Γ}}_{k-1}{W}_{k-1}\\ Z_k=H_k{X}_k+V_k\end{array}\right.,$

In formula:

X _krepresent k moment (n × 1) dimension system state vector; X _k-1it is then the system state vector in k-1 moment; Φ _k|k-1for the transition matrix of etching system during k-1; W _k-1the k-1 moment (n × 1) maintain system noise vector; Γ _k-1that system noise drives matrix; Z _krepresent that (m × 1) dimension in k moment measures vector; H _kfor (m × n) in k moment ties up observing matrix; V _kit is k moment (m × 1) dimension measurement noise vector.

Meanwhile, W _k-1and V _kmutual incoherent zero-mean white noise sequence, that is:

E[W _k-1]＝0， $Cov\[W_{k-1},W_j\]=E\[W_{k-1}{W}_j^T\]=Q_{k-1}{\mathrm{\δ}}_{(k-1)j},$

E[V _k]＝0， $Cov\[V_k,V_j\]=E\[V_k{V}_j^T\]=R_k{\mathrm{\δ}}_{kj},$

$Cov\[W_{k-1},V_j\]=E\[W_{k-1}{V}_j^T\]=0,$

Wherein, Q _kfor system noise variance matrix, be nonnegative definite, R _kfor measurement noises variance matrix, it is positively definite matrix; δ _kjkroneker function:

${\mathrm{\δ}}_{kj}=\left\{\begin{array}{cc}0& k\≠j\\ 1& k=j\end{array}\right.,$

If discrete system meets above-mentioned requirements, the state optimization of discrete system is estimated to adopt following steps to try to achieve:

State one-step prediction: ${\hat{X}}_{k|k-1}={\mathrm{\Φ}}_{k|k-1}{\hat{X}}_{k-1},$

State estimation: ${\hat{X}}_k={\hat{X}}_{k|k-1}+K_k\[Z_k-{\hat{Z}}_{k|k-1}\],$

Filter gain: $K_k=P_{k|k-1}{H}_k^T{\[H_k{P}_{k|k-1}{H}_k^T+R_k\]}^{-1},$

One-step prediction square error: $P_{k/k-1}={\mathrm{\Φ}}_{k|k-1}{P}_{k-1}{\mathrm{\Φ}}_{k|k-1}^T+{\mathrm{\Γ}}_{k-1}{Q}_{k-1}{\mathrm{\Γ}}_{k-1}^T,$

Estimate square error: $P_k=\[I-K_k{H}_k\]P_{k/k-1}{\[I-K_k{H}_k\]}^T+K_k{R}_k{K}_k^T,$

Output module 15, for exporting described navigation results according to binary format.

Binary format:

Binary format according to client's needs, can export different forms, and the pseudorange of time of formatted output receiver below, position and each satellite and Doppler are comparatively conventional binary format.

Table 1

N is observed reading number.

Note: 1, the low byte of each data is front, and high byte is rear.

2, double-precision quantity structure (64bits)

MSB(bit63)＝Sign

Bit62-52＝Exponent(exp)

Bit51-00＝Mantissa＝2exp(-1)*bit51+2exp(-2)*bit50...

Value＝Sign*1.mantissa*2exp(EXP-1023)

In serial data, by first to after order be: Bit7-0, Bit15-8, Bit23-16, Bit31-24, Bit39-32, Bit47-40, Bit55-48, Bit6354

3,16 bit check and be without symbol short number, it is all information bytes (comprising frame head) summation, ignores the carry and spilling that exceed 16.Low byte is first sent out, and sends out after high byte.

4, in BG, the satellite number of BD is its real satellite numbers+33, to distinguish gps satellite.

By one or more embodiment of the present invention, the present invention has following beneficial effect or advantage:

The invention provides a kind of guider, comprising: acquisition and tracking processing module, message extraction module, observed quantity extraction module, positioning calculation module, output module.Carry out continuing to follow the tracks of the navigation message bit data obtaining satellite to satellite-signal by acquisition and tracking processing module; By message extraction module, observed quantity extraction module, obtain satellite navigation message and observed quantity, then positioned by positioning calculation module and resolve, and exported according to binary format by output module, the navigation message bit data of acquisition satellite that can be automatic continuous position, and then avoid occurring that navigation situation not in time occurs, and have technique effect promptly and accurately.

Although described the preferred embodiment of the application, one of ordinary skilled in the art once obtain the basic creative concept of cicada, then can make other change and amendment to these embodiments.So claims are intended to be interpreted as comprising preferred embodiment and falling into all changes and the amendment of the application's scope.

Obviously, those skilled in the art can carry out various change and modification to the application and not depart from the spirit and scope of the application.Like this, if these amendments of the application and modification belong within the scope of the application's claim and equivalent technologies thereof, then the application is also intended to comprise these change and modification.

Claims (10)

1. a guider, is characterized in that, comprising:

Acquisition and tracking processing module, continues to follow the tracks of for carrying out satellite-signal, and obtains the navigation message bit data of satellite;

Message extraction module, for receiving described navigation message bit data, and goes out satellite navigation message from described navigation message bit extracting data;

Observed quantity extraction module, for going out observed quantity from described navigation message bit extracting data;

Positioning calculation module, for carrying out co-located according to described observed quantity and described satellite navigation message, obtains navigation results;

Output module, for exporting described navigation results according to binary format.

2. device as claimed in claim 1, it is characterized in that, described satellite navigation message comprises ephemeris, satellite health, time device, the satellite almanac of satellite.

3. device as claimed in claim 1, is characterized in that, described observed quantity extraction module, specifically for obtaining the pseudo range observed quantity of tracked satellite or Doppler measurements or carrier phase observed quantity according to Base-Band Processing data.

4. device as claimed in claim 3, is characterized in that, described pseudo range observed quantity is utilize the transmission time of satellite-signal to be multiplied by transmission range that the light velocity obtains signal; The transmission time of described satellite-signal refers to satellite-signal transmitting time t _svdeduct receiver local zone time t _u.

5. device as claimed in claim 3, it is characterized in that, described Doppler measurements is after the frequency values obtaining carrier wave NCO, by the frequency values of described carrier wave NCO and the nominal intermediate frequency value f of radio-frequency front-end _csubtract each other acquisition.

6. device as claimed in claim 2, it is characterized in that, described positioning calculation module specifically comprises:

Computing module, for going out the position of satellite according to the ephemeris computation of described satellite;

Positioning calculation module, to position customer location for utilizing the position of described satellite and described observed quantity and resolves.

7. device as claimed in claim 6, is characterized in that, described computing module specifically for:

The first step: the mean angular velocity n calculating satellite motion;

Second step: the mean anomaly M of calculating observation moment satellite _s;

M _s=M ₀+ n (t-t _oe); In formula, M ₀for reference moment t _oetime mean anomaly, provided by broadcast ephemeris;

3rd step: calculate eccentric anomaly: E _s=M _s+ e _ssinE _s;

4th step: calculate anomaly f _s: $f_s=src\tan \frac{\sqrt{1-{e_s}^2}\sin E_s}{\cos E_s-e_s};$

5th step: calculate to rise and hand over elongation in formula, ω is the argument of perigee, is provided by broadcast ephemeris;

6th step: calculate perturbation correction member;

7th step: to a liter friendship elongation satellite radius vector r, orbit inclination i carry out perturbation and correct;

8th step: calculate the position of satellite in track areal coordinate system;

9th step: the longitude L of calculating observation moment ascending node;

Tenth step: calculate the position of satellite in spherical coordinate system instantaneously;

11 step: calculate the position of satellite in conventional terrestrial coordinate system.

8. device as claimed in claim 7, it is characterized in that, described mean angular velocity n is obtained by following steps:

First according to the parameter provided in broadcast ephemeris calculate with reference to moment t _oemean angular velocity n ₀: in formula, GM is the product of gravitational constant G and earth gross mass M, GM=3.9860047 × 1014m ³/ s ², then according to the mean angular velocity n:n=n of perturbation parameter △ n calculating observation moment satellite given in broadcast ephemeris ₀+ △ n.

9. device as claimed in claim 7, is characterized in that, described calculating perturbation correction member comprises the following steps:

By the C provided in broadcast ephemeris _uc, C _us, C _rc, C _rs, C _ic, C _is6 perturbation parameters, obtain the liter caused due to the second order zonal harmonic coefficient item of earth gravitational field bit function and hand over elongation perturbation correction member δ _μ, satellite radius vector r perturbation correction member δ _r, and the perturbation correction member δ of inclination of satellite orbit i _i, computing formula is as follows:

${\mathrm{\δ}}_{\mathrm{\μ}}=C_{uc}\cos 2\widetilde{\mathrm{\μ}}+C_{us}\sin 2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_r=C_{rc}\cos 2\widetilde{\mathrm{\μ}}+C_{rs}\sin 2\widetilde{\mathrm{\μ}}$

${\mathrm{\δ}}_i=C_{ic}\cos 2\widetilde{\mathrm{\μ}}+C_{is}\sin 2\widetilde{\mathrm{\μ}}.$

10. device as claimed in claim 7, is characterized in that, described to liter handing over an elongation μ, satellite radius vector r, the orbit inclination i step corrected of carrying out perturbing is:

$\widetilde{\mathrm{\μ}}=\widetilde{\mathrm{\μ}}+{\mathrm{\δ}}_{\mathrm{\μ}}$

r＝a _s(1-e _scosE _s)+δ _r，

$i=i_0+{\mathrm{\δ}}_i+\frac{di}{dt}\left(t-t_{oe}\right)$

In formula: a _sfor the major radius of satellite orbit, i ₀for t _oethe orbit inclination in moment, is provided by the kepler parameters in broadcast ephemeris; for the rate of change of i, provided by the perturbation parameter in broadcast ephemeris.