CN103424116B - A kind of geostationary satellite precise orbit determination method adapting to orbit maneuver - Google Patents

Abstract

The invention discloses a kind of geostationary satellite precise orbit determination method adapting to orbit maneuver, belong to aerospace measurement and control field.First the method calculates Maneuver Acceleration; Calculating observation amount ρ _i,jand observation residual sequence Δ ρ _i,j; To public system error ρ _bcarry out valuation; By public system error ρ _bfrom observation residual sequence Δ ρ _i,jmiddle deduction, obtains each survey station system deviation ; By ρ _b, deduct from observed quantity, then carry out improvement of orbit, the original state after being improved by again secondary for the observation data after deduction System level gray correlation

Description

A kind of geostationary satellite precise orbit determination method adapting to orbit maneuver

Technical field

Patent of the present invention belongs to aerospace measurement and control field, relates to a kind of geostationary satellite (GEO) precise orbit determination method adapting to orbit maneuver.

Background technology

According to National Airspace development plan, the Big Dipper No. two satellite navigation systems of China adopt GEO/IGSO/MEO hybrid constellation to carry out passive navigation location.For passive location, orbit maneuver just means that satellite temporarily lost efficacy, and because GEO satellite control system regularly carries out momenttum wheel unloading, frequent orbit maneuver, GEO satellite becomes the crucial satellite of influential system navigation performance.The systematic errors such as GEO Satellite orbital maneuver is comparatively frequent, ground observation geometry difference and satellite clock correction and equipment delay are difficult to features such as being separated, make GEO precision orbit determination and cross over one of rail control period forecasting technique gordian technique becoming system.During orbit maneuver, on star, the thrust of the jet generation of engine is determined to bring difficulty to Precise Orbit, utilizes conventional method to carry out track and determines to be difficult to orbit prediction ensure navigation accuracy requirement.In order to provide Navsat continuous print precise ephemeris during orbit maneuver, need study track mechanomotive force model, the track explored during orbit maneuver is determined and forecasting procedure, to improve the navigation performance of this system.

Summary of the invention

The object of the invention is: in order to solve geostationary orbits maintenance period, utilize conventional method to carry out problem that track determines to be difficult to orbit prediction to ensure trajectory accuracy, the present invention provides the geostationary satellite precise orbit determination method of the adaptation orbit maneuver of a set of practicality.

Technical scheme of the present invention is: the geostationary satellite precise orbit determination method adapting to orbit maneuver, comprises the steps:

Step one: calculate Maneuver Acceleration:

Suppose that on star, the engine jet pipe initial working time is T ₀, finish time is T _e, F _ijthrust size when being i-th jet pipe jth task, work pulsewidth is t _j, then have the acceleration produced during jth task can be expressed as in J2000 inertial coordinates system:

Wherein, m is satellite quality; α _i, β _i, γ _ibe three the vector cosine angles of i-th nozzle axis orientation under whole star mechanical coordinate system; Matrix M _jfor jth task moment inertial coordinates system is to the transition matrix of Centroid orbit coordinate system, matrix P _jfor jth task moment Centroid orbit coordinate system is to the rotation matrix of whole star mechanical coordinate system.

Step 2: from t initial epoch ₀with original state amount set out, utilize accurate numerical method to carry out Orbit extrapolation, calculate the observed quantity ρ of i-th survey station jth time observational record _i,j, obtain the observation residual sequence Δ ρ of i-th survey station jth time observational record _i,j, Δ ρ _i,jin, public system error is designated as ρ _b, each station system deviation is designated as

Step 3: to public system error ρ _bcarry out valuation,

ρ _bcomprise a constant component and a linear change part, namely

ρ _b＝ρ _b0+ρ _b1T

Wherein, T is the time apart from initial observation epoch, namely apart from initial orbit epoch time, and ρ _b0and ρ _b1be respectively constant component and a linear change part.If ρ _i,jcorresponding true value is then have,

Wherein, ε _i,jfor observation random noise.ρ _i,jto ρ _b0and ρ _b1partial derivative be:

Complete public system error ρ based on above-mentioned partial derivative _bvaluation;

Step 4: by public system error ρ _bfrom observation residual sequence Δ ρ _i,jmiddle deduction, obtains Δ ρ ' _i,j=Δ ρ _i,j-ρ _b, now Δ ρ ' _i,jin only comprise the system deviation of i-th survey station and random noise.All observation moment Δ ρ ' are calculated to i-th survey station _i,jmean value, obtain each survey station system deviation

Step 5: by ρ _b, deduct from observed quantity, then carry out improvement of orbit, the original state after being improved by again secondary for the observation data after deduction System level gray correlation

Step 6: if do not meet the condition of convergence, this condition of convergence is improvement be less than certain setting value, for those skilled in the art are known, if return step 2, repeat above-mentioned steps; Otherwise end loop, completes geostationary satellite precise orbit determination.

The invention has the beneficial effects as follows: GEO satellite can be solved during orbit maneuver, on star, the thrust of the jet generation of engine determines the difficulty brought to Precise Orbit, simultaneously, the problem adopting public system error estimator and each station deviation adaptive iteration removing method to solve the systematic errors such as satellite clock correction and each survey station equipment delay to be difficult to be separated, effectively can improve the Orbit Determination and Orbit Forecast precision of GEO satellite during orbit maneuver.

Embodiment

Be treated to example with BeiDou-I satellite multistation ranging data 12 to 13 Dec in 2006 below, use step and the effect of this method are described.Wherein, data on the 13rd thing that contains satellite is motor-driven.

The geostationary satellite precise orbit determination method of the adaptation orbit maneuver in the present embodiment, comprises the steps:

Step one: obtain BeiDou-I satellite star lifting force device installation parameter and motor-driven period the attitude of satellite, star lifting force device array mode, thrust size, the working pulse cycle, pulsewidth and jet pipe work times etc., and get out satellite original state amount and orbit determination data file, these data are control precontract one day and control time about 20 minute data, calculate the orbit maneuver acceleration in control time according to above-mentioned information;

Step 2: from t initial epoch ₀with original state amount set out, utilize accurate numerical method to carry out Orbit extrapolation, calculating observation amount ρ _i,jwith each station range finding observation residual sequence Δ ρ _i,j;

Step 3: to public system error ρ _bcarry out valuation, and adopt adaptive iteration method to eliminate each station range measurement system skew component

Step 4: adopt batch processing least square method to carry out improvement of orbit, and by the Orbit extrapolation after improving to controlling latter 4 hours, prediction orbit is used for comparing with Precise Orbit afterwards, to assess the orbit determination accuracy of this method;

Step 5: by ρ _b, deduct from observed quantity, then carry out improvement of orbit, the original state after being improved by again secondary for the observation data after deduction System level gray correlation

Step 6: if do not meet the condition of convergence, this condition of convergence is improvement be less than certain setting value, for those skilled in the art are known, if return step 2, repeat above-mentioned steps; Otherwise end loop, completes geostationary satellite precise orbit determination.

After adopting the method in the present embodiment to carry out geostationary satellite precise orbit determination, after adopting control, about one day data carries out orbit determination, as benchmark, assessment is checked to above-mentioned prediction orbit, visible in table 4-1, the control of forecast rail terminates the track in latter 1 hour, and satellite position forecast maximum error is within 20m.

Precise Orbit comparative result (adding telemetry parameter) after data Orbit Determination and Orbit Forecast result and control before the control of table 4-1 rail and during rail control

Directly forecast according to data orbit determination before rail control, do not add track machine dynamic model, then forecast that rail control terminates the track in latter 1 hour, satellite position forecast maximum error is about 175m.Visible, before adopting control, data orbit determination directly forecasts the rear track of control, and prediction error is larger.

Data Orbit Determination and Orbit Forecast result and the rear Precise Orbit comparative result of control (directly forecasting) before the control of table 4-2 rail

In sum, the geostationary satellite precise orbit determination method of answering orbit maneuver that the present invention proposes, based on the Orbit Determination and Orbit Forecast of track machine dynamic model and telemetry parameter, effectively improves Orbit Determination and Orbit Forecast precision, is conducive to the fast quick-recovery controlling rear track.

Claims (1)

1. adapt to the geostationary satellite precise orbit determination method of orbit maneuver, comprise the steps:

Step one: calculate Maneuver Acceleration:

Suppose that on star, the engine jet pipe initial working time is T ₀, finish time is T _e, F _ijthrust size when being i-th jet pipe jth task, work pulsewidth is t _j, then have the acceleration produced during jth task can be expressed as in J2000 inertial coordinates system:

Wherein, m is satellite quality; α _i, β _i, γ _ibe three the vector cosine angles of i-th nozzle axis orientation under whole star mechanical coordinate system; Matrix M _jfor jth task moment inertial coordinates system is to the transition matrix of Centroid orbit coordinate system, matrix P _jfor jth task moment Centroid orbit coordinate system is to the rotation matrix of whole star mechanical coordinate system;

Step 2: from t initial epoch ₀with original state amount set out, utilize accurate numerical method to carry out Orbit extrapolation, calculate the observed quantity ρ of i-th survey station jth time observational record _i,j, obtain the observation residual sequence Δ ρ of i-th survey station jth time observational record _i,j, Δ ρ _i,jin, public system error is designated as ρ _b, each station system deviation is designated as ρ _bsi;

Step 3: to public system error ρ _bcarry out valuation,

ρ _bcomprise a constant component and a linear change part, namely

ρ _b＝ρ _b0+ρ _b1T

Wherein, T is the time apart from initial observation epoch, namely apart from initial orbit epoch time, and ρ _b0and ρ _b1be respectively constant component and a linear change part; If ρ _i,jcorresponding true value is then have,

$\begin{array}{c}{\mathrm{\ρ}}_{i,j}={\widehat{\mathrm{\ρ}}}_{i,j}+{\mathrm{\ρ}}_b+{\mathrm{\ρ}}_{{\mathrm{bs}}_i}+{\mathrm{\ϵ}}_{i,j}\\ ={\widehat{\mathrm{\ρ}}}_{i,j}+{\mathrm{\ρ}}_{b0}+{\mathrm{\ρ}}_{{\mathrm{bs}}_i}+{\mathrm{\ρ}}_{b1}T+{\mathrm{\ϵ}}_{i,j}\end{array}$

Wherein, ε _i,jfor observation random noise; ρ _i,jto ρ _b0and ρ _b1partial derivative be:

$\frac{\∂{\mathrm{\ρ}}_{i,j}}{\∂{\mathrm{\ρ}}_{b0}}=1$

$\frac{\∂{\mathrm{\ρ}}_{i,j}}{\∂{\mathrm{\ρ}}_{b1}}=T$

Complete public system error ρ based on above-mentioned partial derivative _bvaluation;

Step 4: by public system error ρ _bfrom observation residual sequence Δ ρ _i,jmiddle deduction, obtains Δ ρ ' _i,j=Δ ρ _i,j-ρ _b, now Δ ρ ' _i,jin only comprise the system deviation of i-th survey station and random noise; All observation moment Δ ρ ' are calculated to i-th survey station _i,jmean value, obtain each survey station system deviation

Step 5: by ρ _b, deduct from observed quantity, then carry out improvement of orbit, the original state after being improved by again secondary for the observation data after deduction System level gray correlation

Step 6: if do not meet the condition of convergence, if return step 2, repeat above-mentioned steps; Otherwise end loop, completes geostationary satellite precise orbit determination.