CN108958064A - Posture guidance law error judgement method, system and electronic equipment - Google Patents
Posture guidance law error judgement method, system and electronic equipment Download PDFInfo
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Abstract
The present invention provides a kind of posture guidance law error judgement method, system and electronic equipment, including setting with spacecraft and with reference to relative movement orbit of the spacecraft in orbital plane the lateral drift ellipse for being twice of semi-minor axis as major semiaxis in the LVLH coordinate system with reference to spacecraft, the relational expression of posture guidance law parameter and relative position is obtained according to C-W equation；It sets with the opposite relative movement orbit for referring to spacecraft of spacecraft as line of collimation and lateral drift ellipse from front to back, and obtains the emulation relative motion data of the corresponding adjoint spacecraft of two groups of motion profiles；Respectively according to two groups it is error free, have error emulation relative motion data and relative orbit forecast data, estimate posture guidance law parameter, and calculate the error of it is expected pitch angle with emulation pitch angle.Posture guidance law error judgement method, system and electronic equipment of the invention is carried out by calculating expectation pitching angle error, provides foundation for engineer application.
Description
Technical field
The present invention relates to the satellite relative movement technology fields in dynamics of orbits, more particularly to a kind of posture guidance law
Error judgement method, system and electronic equipment.
Background technique
With the continuous development of space technology and the continuous extension of spacecraft application field, distributed networked research center is by general
It reads and turns to practical application.The micro-nano satellite that more function distributions, information interconnect can replace large satellite to carry out by Collaborative Control
Space mission.Due to very strong flexibility and the anti-destruction of system, and have in terms of lead time and development cost aobvious
Work advantage, the micro-nano satellite that more function distributions, information interconnect have very wide application prospect.Each state of our times all exists
The relation technological researching and application mode for actively developing distributed networked research center are explored.
Relative distance is closer between the spacecraft of formation flight, can be analyzed under relative motion frame.Common phase
There is following two to sports immunology method:
(1) kinematic method based on two spacecraft orbit radicals
This method is input with two spacecraft absolute orbit radicals, and computational accuracy is higher, but needs high-precision numerical integration,
Calculation amount is larger, and disturbance rejection is poor.
(2) dynamic method based on two spacecraft relative position speed, also referred to as C-W equation method
When ignoring Perturbation Effect, with reference to spacecraft being circular orbit, C-W equation can be parsed by linearization process
Solution.C-W equation analytic solutions are input with the relative status of two spacecrafts, and physical meaning is clear, and calculation amount is small, and strong robustness is non-
It is often that relatively closely (the ratio between relative distance and nominal track radius are no more than ginseng for circular orbit, tail clearance suitable for spacecraft is referred to
Examine spacecraft eccentricity magnitude) Problem of Relative Movement analysis.
However, the spacecraft of practical in-orbit flight can not be the circular orbit of standard, distance is also impossible to keep constantly
In closer distance, need according to relative motion law, exploration is suitable for compared with the remote and lesser relative motion fitting side of error
Method obtains the relative motion related physical quantity of mission requirements, the foundation as engineering design.
Summary of the invention
In view of the foregoing deficiencies of prior art, the purpose of the present invention is to provide a kind of posture guidance law error judgments
Method, system and electronic equipment, it is optimal using least square according to short distance spacecraft high-resolution imaging pitching posture adjustment demand
The method for parameter estimation of estimation theory accurately estimates C-W equation parameter solution, derives the appearance of relative target spacecraft orientation
State guidance law, and prediction error empirical model and in-orbit actual orbit prediction precision for relative orbit, lead posture
Draw the expectation pitching angle error that rule is calculated to be analyzed, to provide foundation for engineer application.
In order to achieve the above objects and other related objects, the present invention provides a kind of suitable for the coplanar formation of short distance spacecraft
Posture guidance law error judgement method, comprising the following steps: set in orbital plane with the opposite spacecraft that refers to of spacecraft
Relative movement orbit be that major semiaxis is twice of semi-minor axis in the LVLH coordinate system with reference to spacecraft lateral drift is oval, according to
C-W equation obtains relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) it is initial
The elliptical center at moment, b are oval semi-minor axis, and Θ is the phase with reference to spacecraft on ellipse；It sets opposite with spacecraft
Relative movement orbit with reference to spacecraft is line of collimation and lateral drift ellipse from front to back, and obtains two groups of motion profiles pair
The emulation relative motion data for the adjoint spacecraft answered；Based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, y's
Relational expression and two groups of emulation relative motion datas estimate posture guidance law parameter x according to least-square principlec0,yc0,Θ,
B, and the expectation pitch angle at two groups of emulation relative motion datas corresponding each moment is calculated, and it is expected pitch angle and imitate
The difference of true pitch angle；Error is carried out to two groups of emulation relative motion datas according to relative orbit prediction error empirical model to repair
Just, it is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of revised emulation are transported relatively
Dynamic data estimate posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and after two groups of amendments are calculated
Emulation relative motion data corresponding each moment expectation pitch angle, and expectation pitch angle and the difference for emulating pitch angle
Value；Based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of motion profiles are corresponding opposite
Orbit prediction data estimate posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and it is calculated two groups
The expectation pitch angle at relative orbit forecast data corresponding each moment；According to the corresponding two groups of subsequent orbit determination of two groups of motion profiles
Data calculate the practical pitch angle at corresponding each moment, and calculate the difference of desired pitch angle Yu practical pitch angle.
In one embodiment of the invention, relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b are as follows:
Wherein, n is the mean angular velocity of satellite motion with reference to spacecraft absolute orbit, and t is the time started from initial time.
In one embodiment of the invention, the expectation pitch angle Pitch is calculated according to the following formula:
Pitch=π+γ
γ=arctan (x/y) or γ=π+arctan (x/y)
Wherein, the value of γ byWithSymbol codetermine, xc0,yc0,
Θ, b take the estimated value obtained according to least-square principle.
In one embodiment of the invention, the relative orbit prediction error mode is linear plus trigonometric function model, institute
The period for stating trigonometric function is the orbital period with reference to spacecraft.
Meanwhile the present invention also provides a kind of posture guidance law error judgment systems suitable for the coplanar formation of short distance spacecraft
System, including expression formula obtain module, data acquisition module, first processing module, Second processing module and third processing module；
The expression formula obtains module for being set in orbital plane with the opposite opposite fortune for referring to spacecraft of spacecraft
Dynamic rail mark is that the lateral drift that major semiaxis is twice of semi-minor axis in the LVLH coordinate system with reference to spacecraft is oval, according to C-W equation
Obtain relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) it is the ellipse of initial time
Circle center, b are oval semi-minor axis, and Θ is the phase with reference to spacecraft on ellipse；
The data acquisition module is used to set with the opposite relative movement orbit for referring to spacecraft of spacecraft as before
Line of collimation and lateral drift backward is oval, and obtains the emulation relative motion number of the corresponding adjoint spacecraft of two groups of motion profiles
According to；
The first processing module is used to be based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relationship of y
Formula and two groups of emulation relative motion datas estimate posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and
The expectation pitch angle at two groups of emulation relative motion datas corresponding each moment is calculated, and expectation pitch angle is bowed with emulation
The difference at the elevation angle；
The Second processing module is used for according to relative orbit prediction error empirical model to two groups of emulation relative motion numbers
According to error correction is carried out, it is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, after the relational expression of y and two groups of amendments
Emulation relative motion data, according to least-square principle estimate posture guidance law parameter xc0,yc0, Θ, b, and calculate
To two groups of revised expectation pitch angles for emulating relative motion data corresponding each moment, and expectation pitch angle and emulation
The difference of pitch angle；
The third processing module is used to be based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relationship of y
Formula and the corresponding relative orbit forecast data of two groups of motion profiles estimate posture guidance law parameter according to least-square principle
xc0,yc0, Θ, b, and the expectation pitch angle at two groups of relative orbit forecast datas corresponding each moment is calculated；According to two groups
The corresponding two groups of subsequent orbit determination data of motion profile calculate the practical pitch angle at corresponding each moment, and calculate desired pitch angle
With the difference of practical pitch angle.
In one embodiment of the invention, relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b are as follows:
Wherein, n is the mean angular velocity of satellite motion with reference to spacecraft absolute orbit, and t is the time started from initial time.
In one embodiment of the invention, the first processing module, the Second processing module and the third handle mould
In block, the expectation pitch angle Pitch is calculated according to the following formula:
Pitch=π+γ
γ=arctan (x/y) or γ=π+arctan (x/y)
Wherein, the value of γ byWithSymbol codetermine, xc0,yc0,
Θ, b take the estimated value obtained according to least-square principle.
In one embodiment of the invention, in the Second processing module, the relative orbit prediction error mode is linear
Add the model of trigonometric function, the period of the trigonometric function is the orbital period with reference to spacecraft.
In addition, being total to the present invention also provides a kind of electronic equipment including any of the above-described short distance spacecraft that is suitable for
The posture guidance law error judgment system that face is formed into columns.
As described above, posture guidance law error judgement method, system and electronic equipment of the invention, have below beneficial to effect
Fruit:
(1) theoretical using least square optimal estimation according to short distance spacecraft high-resolution imaging pitching posture adjustment demand
Method for parameter estimation C-W equation parameter solution is accurately estimated, derive relative target spacecraft orientation posture guidance law；
(2) for the prediction error empirical model of relative orbit and in-orbit actual orbit prediction precision, posture is led
Draw the expectation pitching angle error that rule is calculated to be analyzed, to provide foundation for engineer application.
Detailed description of the invention
Fig. 1 is shown as the posture guidance law error judgement method suitable for the coplanar formation of short distance spacecraft of the invention
Flow chart；
Fig. 2 is shown as in the present invention showing with spacecraft and with reference to spacecraft elliptical orbit of relative motion in orbital plane
It is intended to；
Fig. 3 is shown as adjoint spacecraft of the invention and illustrates with reference to the phase of spacecraft relative motion in orbital plane
Figure；
Fig. 4 is shown as the schematic diagram of relative motion coordinate system LVLH system；
Fig. 5 is shown as the schematic diagram of expectation pitch angle；
Fig. 6 is shown as the curve graph that error free line of collimation in the embodiment of the present invention moves corresponding expectation pitch angle；
Fig. 7 be shown as error free line of collimation in the embodiment of the present invention move corresponding parameter Estimation relative movement orbit with
Emulate the comparison schematic diagram of relative movement orbit；
Fig. 8 is shown as error free line of collimation in the embodiment of the present invention and moves corresponding parameter Estimation expectation pitch angle and imitate
The difference curve figure of true pitch angle；
Fig. 9 be shown as error free line of collimation in the embodiment of the present invention move corresponding parameter Estimation relative movement orbit with
Emulate the curve graph of the laterally and radially error of relative movement orbit；
Figure 10 is shown as the corresponding expectation of relative motion that error free ellipse semi-minor axis is 160m in the embodiment of the present invention
The curve graph of pitch angle；
Figure 11 is shown as the corresponding parameter of relative motion that error free ellipse semi-minor axis is 160m in the embodiment of the present invention
Estimate relative movement orbit and emulates the comparison schematic diagram of relative movement orbit；
Figure 12 is shown as the corresponding parameter of relative motion that error free ellipse semi-minor axis is 160m in the embodiment of the present invention
Estimation expectation pitch angle and the difference curve figure for emulating pitch angle；
Figure 13 is shown as the corresponding parameter of relative motion that error free ellipse semi-minor axis is 160m in the embodiment of the present invention
Estimate relative movement orbit and emulates the curve graph of the laterally and radially error of relative movement orbit；
Figure 14 is shown as in the embodiment of the present invention the curve graph for having error line of collimation to move corresponding expectation pitch angle；
Figure 15 is shown as having error line of collimation to move corresponding parameter Estimation relative movement orbit in the embodiment of the present invention
With the comparison schematic diagram of true relative movement orbit；
Figure 16 be shown as having in the embodiment of the present invention error line of collimation move corresponding parameter Estimation expectation pitch angle with
Emulate the difference curve figure of pitch angle；
Figure 17 is shown as having error line of collimation to move corresponding parameter Estimation relative movement orbit in the embodiment of the present invention
With the curve graph of the laterally and radially error of emulation relative movement orbit；
Figure 18 is shown as in the embodiment of the present invention the corresponding expectation of relative motion for having error ellipse semi-minor axis for 160m
The curve graph of pitch angle；
Figure 19 is shown as in the embodiment of the present invention the corresponding parameter of relative motion for having error ellipse semi-minor axis for 160m
Estimate relative movement orbit and emulates the comparison schematic diagram of relative movement orbit；
Figure 20 is shown as in the embodiment of the present invention the corresponding parameter of relative motion for having error ellipse semi-minor axis for 160m
Estimation expectation pitch angle and the difference curve figure for emulating pitch angle；
Figure 21 is shown as in the embodiment of the present invention the corresponding parameter of relative motion for having error ellipse semi-minor axis for 160m
Estimate relative movement orbit and emulates the curve graph of the laterally and radially error of relative movement orbit；
Figure 22 is shown as the curve graph of the corresponding true pitch angle of subsequent precise orbit determination data in the embodiment of the present invention；
Figure 23 is shown as relative orbit forecast data and the corresponding ginseng of subsequent precise orbit determination data in the embodiment of the present invention
The comparison schematic diagram of number estimation relative movement orbit and true relative movement orbit；
Figure 24 is shown as relative orbit forecast data and the corresponding ginseng of subsequent precise orbit determination data in the embodiment of the present invention
The difference curve figure of number estimation expectation pitch angle and true pitch angle；
Figure 25 is shown as relative orbit forecast data and the corresponding ginseng of subsequent precise orbit determination data in the embodiment of the present invention
The curve graph of the laterally and radially error of number estimation relative movement orbit and true relative movement orbit；
The posture guidance law parameter suitable for the coplanar formation of short distance spacecraft that Figure 26 is shown as of the invention generates system
Structural schematic diagram；
Figure 27 is shown as the structural schematic diagram of electronic equipment of the invention.
Component label instructions
1 expression formula obtains module
2 data acquisition modules
3 first processing modules
4 Second processing modules
5 third processing modules
Specific embodiment
Illustrate embodiments of the present invention below by way of specific specific example, those skilled in the art can be by this specification
Other advantages and efficacy of the present invention can be easily understood for disclosed content.The present invention can also pass through in addition different specific realities
The mode of applying is embodied or practiced, the various details in this specification can also based on different viewpoints and application, without departing from
Various modifications or alterations are carried out under spirit of the invention.It should be noted that in the absence of conflict, following embodiment and implementation
Feature in example can be combined with each other.
It should be noted that illustrating the basic structure that only the invention is illustrated in a schematic way provided in following embodiment
Think, only shown in diagram then with related component in the present invention rather than component count, shape and size when according to actual implementation
Draw, when actual implementation kenel, quantity and the ratio of each component can arbitrarily change for one kind, and its assembly layout kenel
It is likely more complexity.
As shown in Figure 1, the posture guidance law error judgement method suitable for the coplanar formation of short distance spacecraft of the invention
The following steps are included:
Step S1, it is set in orbital plane and refers to the relative movement orbit of spacecraft as with reference to space flight with spacecraft is opposite
The lateral drift that major semiaxis is twice of semi-minor axis in the LVLH coordinate system of device is oval, obtains relative position x according to C-W equation, y with
Posture guidance law parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) be initial time elliptical center, b is oval short
Semiaxis, Θ are the phase with reference to spacecraft on ellipse.
It as shown in Figures 2 and 3, can be by following formula with the opposite elliptical orbit for referring to spacecraft movement of spacecraft in orbital plane
It indicates:
Wherein, (xc,yc) it is elliptical center, b is oval semi-minor axis.xc=xc0, yc=yc0-1.5nxc0T, n are with reference to boat
Its deviceAbsolute orbitMean angular velocity of satellite motion, (xc0,yc0) be initial time elliptical center, t is to start from initial time
Time.
LVLH coordinate system, that is, Local Vertical Local Horizontal coordinate system, with the matter of space spacecraft
The heart is origin, and it is x-axis that spacecraft direction is directed toward in the earth's core, and orbit plane normal direction is that z-axis, y-axis and x-axis and z-axis constitute right-handed scale (R.H.scale)
System.As shown in figure 4, the x-axis in LVLH coordinate system is radially towards day, wherein being directed toward the direction with reference to spacecraft centroid by the earth's core is diameter
To；Y-axis is in orbital plane perpendicular to x-axis and along heading for laterally；Z-axis meets the right-hand rule, is orbital plane normal direction.
Therefore, relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b is expressed from the next:
Wherein, Θ=nt+ θ is the phase with reference to spacecraft on ellipse, and θ is the initial phase with reference to spacecraft on ellipse
Position.
Step S2, set with the opposite relative movement orbit for referring to spacecraft of spacecraft as from front to back line of collimation and
Lateral drift is oval, and obtains the emulation relative motion data of the corresponding adjoint spacecraft of two groups of motion profiles.
Specifically, two groups of relative movement orbits are set as line of collimation and lateral drift ellipse from front to back, and are obtained
Two groups of emulation relative motion positions speed of the corresponding adjoint spacecraft of two kinds of relative movement orbits in relative motion certain period of time
Degree evidence, abbreviation relative motion data), using the foundation data analyzed as subsequent simulation.
Step S3, it is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of emulation phases
To exercise data, posture guidance law parameter x is estimated according to least-square principlec0,yc0, Θ, b, and be calculated two groups and imitate
The expectation pitch angle at true relative motion data corresponding each moment, and expectation pitch angle and the difference for emulating pitch angle.
Specifically, it according to the relationship and emulation relative motion data of posture guidance law parameter and relative motion data, adopts
Posture guidance law parameter x is obtained according to residual error minimum principle with least-squares parameter estimation methodc0,yc0,Θ,b；Further according to
The estimated value of relative motion data is calculated in posture guidance law parameter, and then obtains the expectation pitch angle of gesture stability needs.
According to the difference of desired pitch angle and emulation pitch angle, least-squares estimation theory can be analyzed in posture guidance law parameter Estimation
In effect.Wherein, emulation pitch angle is calculated using emulation relative motion data and is obtained.
As shown in figure 5, expectation pitch angle be directed toward with the load optical axis of spacecraft ontology-angle of x-axis, under
Formula is calculated:
Pitch=π+γ
Wherein, the value of γ=arctan (x/y) or γ=π+arctan (x/y), γ byWithSymbol codetermine.
Step S4, error is carried out to two groups of emulation relative motion datas according to relative orbit prediction error empirical model to repair
Just, it is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of revised emulation are transported relatively
Dynamic data estimate posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and after two groups of amendments are calculated
Emulation relative motion data corresponding each moment expectation pitch angle, and expectation pitch angle and the difference for emulating pitch angle
Value.
Specifically, according to relative orbit prediction error empirical model, error is added on two groups of emulation relative motion datas,
Expectation pitch angle is then calculated in the same way and emulates the difference between pitch angle.By being carried out with error-free result
Compare, the robustness of posture guidance law parameter estimation theories can be analyzed.
Wherein, according to engineering experience, the relationship of relative orbit prediction error and time are provided, relative orbit forecast is established and misses
Differential mode type.The relative orbit prediction error mode is " linear+trigonometric function " empirical model, and wherein the period of trigonometric function is rail
The road period.
Step S5, it is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of movement rails
The corresponding relative orbit forecast data of mark estimates posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and
The expectation pitch angle at two groups of relative orbit forecast datas corresponding each moment is calculated；It is corresponding according to two groups of motion profiles
Two groups of subsequent orbit determination data calculate the practical pitch angle at corresponding each moment, and calculate desired pitch angle and practical pitch angle
Difference.
Wherein, by comparing the difference of desired pitch angle and practical pitch angle, error caused by posture guidance law has been obtained.
Therefore, the posture guidance law error judgement method suitable for the coplanar formation of short distance spacecraft of the invention by with
Lower three classes input calculates expectation pitch angle and emulation/practical pitch angle difference, so as to provide the conclusion of engineer application:
(1) line of collimation and the elliptical two groups of emulation relative motion number of lateral drift moved from front to back with spacecraft
According to；
(2) according to relative orbit prediction error empirical model, error is added to obtain containing mistake two groups of emulation relative motion datas
The relative motion data of difference；
(3) two groups of relative orbit forecast datas and subsequent precise orbit determination data.
The posture suitable for the coplanar formation of short distance spacecraft that present invention will be further explained below with reference to specific examples
Guidance law error judgement method.
13 divide 47 seconds on September 15,22 2016, and TG-2 companion star enters the orbit with TG-2；After in-orbit storage 38 days, in 2016
31 divide success in 0 second to discharge from TG-2 when 23 days 7 October, have started the space wedding photographer's of its TG-2 and SZ-11 assembly
Missions.54 divide 16 seconds when 30 days 5 October in 2016, and TG-2 companion star is realized on assembly by 7 accurate orbits controllings
The success of square 1.4km or so is leapt, and is completed the high definition imaging of assembly, is had taken 900 several " wedding photos ".
It is with spacecraft with TG-2 companion star, assembly is with reference to spacecraft, then for TG-2 companion star, to assembly
It leaps and needs to calculate corresponding posture guidance law parameter during taking pictures, to obtain final in-orbit expectation pitching angle error, with
Foundation is provided for engineer application.The demand that assembly is leapt for TG-2 companion star, according to the movement of TG-2 companion star's relative combinations body
For track in-plane moving, TG-2 companion star orbit altitude 1.46km high compared with assembly or so sets two groups of relative motions as from forward direction
Line of collimation and relative motion ellipse semi-minor axis afterwards be 160m the curve of cyclical fluctuations, and provide leap the moment before 120min~after
Two groups of emulation relative motion datas of 90min, the foundation data as subsequent simulation analysis.
Relative motion position using the HPOP Orbit extrapolation model (considering all perturbations) of STK software, before export control, after control
It sets, speed data, the foundation as posture guidance law parameter Estimation.
Relative orbit prediction error empirical model is shown below:
1, it is based on error free emulation relative motion data
A, error free line of collimation motion simulation data posture guidance law parameter estimation result is
xc0=-1463.246218127131m
yc0=-17900.84783337577m
Θ=4.238784462530427rad
B=54.59450495536964m
Wherein, it is expected that pitch angle as shown in fig. 6, parameter Estimation relevant path and true relevant path as shown in fig. 7, parameter
Estimate expectation pitch angle and emulates the difference of pitch angle as shown in figure 8, parameter Estimation relevant path and the transverse direction for emulating relevant path
It is as shown in Figure 9 with radial error.
B, the relative motion emulation data posture guidance law parameter estimation result that error free oval semi-minor axis is 160m is
xc0=-1463.673145480557m
yc0=-18108.07121033077m
Θ=2.029183483729073rad
B=161.2360679716028m
Wherein, it is expected that pitch angle is as shown in Figure 10, the difference of parameter Estimation relevant path and emulation relevant path is such as Figure 11 institute
Show, parameter Estimation it is expected that pitch angle and the difference for emulating pitch angle are as shown in figure 12, parameter Estimation relevant path and the opposite rail of emulation
The laterally and radially error of mark is as shown in figure 13.
2, error emulates relative motion data
A, the error line of collimation motion simulation data posture guidance law parameter estimation result is
xc0=-1464.490714356118m
yc0=-17900.67669131771m
Θ=4.234450195813123rad
B=53.68400543336465m
Wherein, it is expected that pitch angle is as shown in figure 14, the difference of parameter Estimation relevant path and emulation relevant path is such as Figure 15 institute
Show, parameter Estimation it is expected pitch angle and emulates the poor as shown in figure 16 of pitch angle, parameter Estimation relevant path and true opposite rail
The laterally and radially error of mark is as shown in figure 17.
B, relative motion that error ellipse semi-minor axis the is 160m emulation data posture guidance law parameter estimation result is
xc0=-1464.917641617445m
yc0=-18107.90010651827m
Θ=2.025524539809740rad
B=161.9671171786372m
Wherein, it is expected that pitch angle is as shown in figure 18, the difference of parameter Estimation relevant path and emulation relevant path is such as Figure 19 institute
Show, parameter Estimation it is expected that pitch angle and the difference for emulating pitch angle are as shown in figure 20, parameter Estimation relevant path and the opposite rail of emulation
The laterally and radially error of mark is as shown in figure 21.
3, in-orbit to leap Process Forecasting track and subsequent precise orbit determination data
In the present embodiment, two groups of relative motion datas of 120min before leaping the moment~rear 90min are given, according to in-orbit
Guidance law calculates and the experience of upper note, can calculate that the orbit prediction time (to 90min after the moment is leapt) is about 5 hours.According to
Practical in-orbit orbit prediction error experience, corresponding relative orbit forecast worst error is about 21m within 5 hours, and start
(about 1 rail) interior indiffusion in 1.5 hours, corresponding prediction error mode parameter are
Δxσ=Δ r × 8%=21 × 8%=1.7m
Δyσ=Δ xσ
Δy0=0
Therefore, the in-orbit Process Forecasting track profile guidance law parameter estimation result that leaps is
xc0=-1407.093235535373m
yc0=-17148.77536192946m
Θ=2.249255219230809rad
B=20.91089797730896m
Wherein, it is expected that pitch angle is as shown in figure 22.Parameter Estimation relevant path and true relevant path are as shown in figure 23.Ginseng
The difference of number estimation expectation pitch angle and true pitch angle is as shown in figure 24.The cross of parameter Estimation relevant path and true relevant path
To as shown in figure 25 with radial error.
As shown in figure 26, the posture guidance law error judgment system suitable for the coplanar formation of short distance spacecraft of the invention
Module 1, data acquisition module 2, first processing module 3, Second processing module 4 and third processing module 5 are obtained including expression formula.
Expression formula obtains module 1 for being set in orbital plane with the opposite relative motion rail for referring to spacecraft of spacecraft
Mark is that the lateral drift that major semiaxis is twice of semi-minor axis in the LVLH coordinate system with reference to spacecraft is oval, is obtained according to C-W equation
Relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) it is in the ellipse of initial time
The heart, b are oval semi-minor axis, and Θ is the phase with reference to spacecraft on ellipse.
It as shown in Figures 2 and 3, can be by following formula with the opposite elliptical orbit for referring to spacecraft movement of spacecraft in orbital plane
It indicates:
Wherein, (xc,yc) it is elliptical center, b is oval semi-minor axis.xc=xc0, yc=yc0-1.5nxc0T, n are with reference to boat
The mean angular velocity of satellite motion of its device absolute orbit, (xc0,yc0) be initial time elliptical center, t is to start from initial time
Time.
LVLH coordinate system, that is, Local Vertical Local Horizontal coordinate system, with the matter of space spacecraft
The heart is origin, and it is x-axis that spacecraft direction is directed toward in the earth's core, and orbit plane normal direction is that z-axis, y-axis and x-axis and z-axis constitute right-handed scale (R.H.scale)
System.As shown in figure 4, the x-axis in LVLH coordinate system is radially towards day, wherein being directed toward the direction with reference to spacecraft centroid by the earth's core is diameter
To；Y-axis is in orbital plane perpendicular to x-axis and along heading for laterally；Z-axis meets the right-hand rule, is orbital plane normal direction.
Therefore, relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b is expressed from the next:
Wherein, Θ=nt+ θ is the phase with reference to spacecraft on ellipse, and θ is the initial phase with reference to spacecraft on ellipse
Position.
Data acquisition module 2 refers to the relative movement orbit of spacecraft as from front to back with spacecraft is opposite for setting
Line of collimation and lateral drift it is oval, and obtain the emulation relative motion data of the corresponding adjoint spacecraft of two groups of motion profiles.
Specifically, two groups of relative movement orbits are set as line of collimation and lateral drift ellipse from front to back, and are obtained
Two groups of emulation relative motion positions speed of the corresponding adjoint spacecraft of two kinds of relative movement orbits in relative motion certain period of time
Degree evidence, abbreviation relative motion data), using the foundation data analyzed as subsequent simulation.
First processing module 3 obtains module 1 with expression formula and data acquisition module 2 is connected, for being based on posture guidance law
Parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of emulation relative motion datas, according to least-squares estimation original
Reason estimation posture guidance law parameter xc0,yc0, Θ, b, and two groups of emulation relative motion datas corresponding each moment is calculated
Expectation pitch angle, and expectation pitch angle and emulate pitch angle difference.
Specifically, it according to the relationship and emulation relative motion data of posture guidance law parameter and relative motion data, adopts
Posture guidance law parameter x is obtained according to residual error minimum principle with least-squares parameter estimation methodc0,yc0,Θ,b；Further according to
The estimated value of relative motion data is calculated in posture guidance law parameter, and then obtains the expectation pitch angle of gesture stability needs.
According to the difference of desired pitch angle and emulation pitch angle, least-squares estimation theory can be analyzed in posture guidance law parameter Estimation
In effect.Wherein, emulation pitch angle is calculated using emulation relative motion data and is obtained.
As shown in figure 5, expectation pitch angle be directed toward with the load optical axis of spacecraft ontology-angle of x-axis, under
Formula is calculated:
Pitch=π+γ
Wherein, the value of γ=arctan (x/y) or γ=π+arctan (x/y), γ byWithSymbol codetermine.
Second processing module 4 obtains module 1 with expression formula and data acquisition module 2 is connected, for pre- according to relative orbit
It reports error empirical model to carry out error correction to two groups of emulation relative motion datas, is based on posture guidance law parameter xc0,yc0,Θ,
B and relative position x, the relational expression of y and two groups of revised emulation relative motion datas, estimate according to least-square principle
Posture guidance law parameter xc0,yc0, Θ, b, and be calculated two groups of revised emulation relative motion datas it is corresponding each when
The expectation pitch angle at quarter, and expectation pitch angle and the difference for emulating pitch angle.
Specifically, according to relative orbit prediction error empirical model, error is added on two groups of emulation relative motion datas,
Expectation pitch angle is then calculated in the same way and emulates the difference between pitch angle.By being carried out with error-free result
Compare, the robustness of posture guidance law parameter estimation theories can be analyzed.
Wherein, according to engineering experience, the relationship of relative orbit prediction error and time are provided, relative orbit forecast is established and misses
Differential mode type.The relative orbit prediction error mode is " linear+trigonometric function " empirical model, and wherein the period of trigonometric function is rail
The road period.
Third processing module 5 obtains module 1 with expression formula and data acquisition module 2 is connected, for being based on posture guidance law
Parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and the corresponding relative orbit forecast data of two groups of motion profiles, according to
Least-square principle estimates posture guidance law parameter xc0,yc0, Θ, b, and two groups of relative orbit forecast datas are calculated
The expectation pitch angle at corresponding each moment；It is calculated according to the corresponding two groups of subsequent orbit determination data of two groups of motion profiles corresponding each
The practical pitch angle at a moment, and calculate the difference of desired pitch angle Yu practical pitch angle.
Wherein, by comparing the difference of desired pitch angle and practical pitch angle, error caused by posture guidance law has been obtained.
Therefore, the posture guidance law error judgment system suitable for the coplanar formation of short distance spacecraft of the invention by with
Lower three classes input calculates expectation pitch angle and emulation/practical pitch angle difference, so as to provide the conclusion of engineer application:
(1) line of collimation and the elliptical two groups of emulation relative motion number of lateral drift moved from front to back with spacecraft
According to；
(2) according to relative orbit prediction error empirical model, error is added to obtain containing mistake two groups of emulation relative motion datas
The relative motion data of difference；
(3) two groups of relative orbit forecast datas and subsequent precise orbit determination data.
As shown in figure 27, electronic equipment of the invention includes the above-mentioned posture suitable for the coplanar formation of short distance spacecraft
Guidance law error judgment system.
In conclusion posture guidance law error judgement method, system and electronic equipment of the invention.So the present invention has
Effect overcomes various shortcoming in the prior art and has high industrial utilization value.
The above-described embodiments merely illustrate the principles and effects of the present invention, and is not intended to limit the present invention.It is any ripe
The personage for knowing this technology all without departing from the spirit and scope of the present invention, carries out modifications and changes to above-described embodiment.Cause
This, institute is complete without departing from the spirit and technical ideas disclosed in the present invention by those of ordinary skill in the art such as
At all equivalent modifications or change, should be covered by the claims of the present invention.
Claims (9)
1. a kind of posture guidance law error judgement method suitable for the coplanar formation of short distance spacecraft, it is characterised in that: including
Following steps:
It is set in orbital plane with the opposite relative movement orbit for referring to spacecraft of spacecraft as the LVLH seat with reference to spacecraft
The lateral drift that major semiaxis is twice of semi-minor axis in mark system is oval, obtains relative position x, y and posture guidance law according to C-W equation
Parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) be initial time elliptical center, b is oval semi-minor axis, and Θ is ginseng
Examine phase of the spacecraft on ellipse；
Set with the opposite relative movement orbit for referring to spacecraft of spacecraft as from front to back line of collimation and lateral drift it is ellipse
Circle, and obtain the emulation relative motion data of the corresponding adjoint spacecraft of two groups of motion profiles；
Based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of emulation relative motion datas,
Posture guidance law parameter x is estimated according to least-square principlec0,yc0, Θ, b, and two groups of emulation relative motions are calculated
The expectation pitch angle at data corresponding each moment, and expectation pitch angle and the difference for emulating pitch angle；
Error correction is carried out to two groups of emulation relative motion datas according to relative orbit prediction error empirical model, is led based on posture
Draw rule parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of revised emulation relative motion datas, according to most
Small two multiply estimation principle estimation posture guidance law parameter xc0,yc0, Θ, b, and two groups of revised emulation fortune relatively is calculated
The expectation pitch angle at dynamic data corresponding each moment, and expectation pitch angle and the difference for emulating pitch angle；
Based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups of motion profiles are corresponding opposite
Orbit prediction data estimate posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and it is calculated two groups
The expectation pitch angle at relative orbit forecast data corresponding each moment；According to the corresponding two groups of subsequent orbit determination of two groups of motion profiles
Data calculate the practical pitch angle at corresponding each moment, and calculate the difference of desired pitch angle Yu practical pitch angle.
2. the posture guidance law error judgement method according to claim 1 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b are as follows:
Wherein, n is the mean angular velocity of satellite motion with reference to spacecraft absolute orbit, and t is the time started from initial time.
3. the posture guidance law error judgement method according to claim 2 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: the expectation pitch angle Pitch is calculated according to the following formula:
Pitch=π+γ
γ=arctan (x/y) or γ=π+arctan (x/y)
Wherein, the value of γ byWithSymbol codetermine, xc0,yc0,Θ,b
Take the estimated value obtained according to least-square principle.
4. the posture guidance law error judgement method according to claim 1 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: the relative orbit prediction error mode is linear plus trigonometric function model, the period of the trigonometric function
For the orbital period with reference to spacecraft.
5. a kind of posture guidance law error judgment system suitable for the coplanar formation of short distance spacecraft, it is characterised in that: including
Expression formula obtains module, data acquisition module, first processing module, Second processing module and third processing module；
The expression formula obtains module for being set in orbital plane with the opposite relative motion rail for referring to spacecraft of spacecraft
Mark is that the lateral drift that major semiaxis is twice of semi-minor axis in the LVLH coordinate system with reference to spacecraft is oval, is obtained according to C-W equation
Relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b, wherein (xc0,yc0) it is in the ellipse of initial time
The heart, b are oval semi-minor axis, and Θ is the phase with reference to spacecraft on ellipse；
The data acquisition module refers to the relative movement orbit of spacecraft as from front to back with spacecraft is opposite for setting
Line of collimation and lateral drift it is oval, and obtain the emulation relative motion data of the corresponding adjoint spacecraft of two groups of motion profiles；
The first processing module is used to be based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two
Group emulation relative motion data, estimates posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and calculate
To the expectation pitch angle at two groups of emulation relative motion datas corresponding each moment, and expectation pitch angle and emulation pitch angle
Difference；
The Second processing module be used for according to relative orbit prediction error empirical model to two groups of emulation relative motion datas into
Row error correction is based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two groups it is revised imitative
True relative motion data estimates posture guidance law parameter x according to least-square principlec0,yc0, Θ, b, and it is calculated two
The expectation pitch angle at the revised emulation relative motion data of group corresponding each moment, and expectation pitch angle and emulation pitching
The difference at angle；
The third processing module is used to be based on posture guidance law parameter xc0,yc0, Θ, b and relative position x, the relational expression of y and two
The corresponding relative orbit forecast data of group motion profile estimates posture guidance law parameter x according to least-square principlec0,
yc0, Θ, b, and the expectation pitch angle at two groups of relative orbit forecast datas corresponding each moment is calculated；According to two groups of movements
The corresponding two groups of subsequent orbit determination data in track calculate the practical pitch angle at corresponding each moment, and calculate desired pitch angle and reality
The difference of border pitch angle.
6. the posture guidance law error judgment system according to claim 5 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: relative position x, y and posture guidance law parameter xc0,yc0, the relational expression of Θ, b are as follows:
Wherein, n is the mean angular velocity of satellite motion with reference to spacecraft absolute orbit, and t is the time started from initial time.
7. the posture guidance law error judgment system according to claim 6 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: in the first processing module, the Second processing module and the third processing module, the expectation pitching
Angle Pitch is calculated according to the following formula:
Pitch=π+γ
γ=arctan (x/y) or γ=π+arctan (x/y)
Wherein, the value of γ byWithSymbol codetermine, xc0,yc0,Θ,b
Take the estimated value obtained according to least-square principle.
8. the posture guidance law error judgment system according to claim 5 suitable for the coplanar formation of short distance spacecraft,
It is characterized by: the relative orbit prediction error mode is linear plus trigonometric function model in the Second processing module,
The period of the trigonometric function is the orbital period with reference to spacecraft.
9. a kind of electronic equipment, it is characterised in that: coplanar including being suitable for short distance spacecraft described in one of claim 5-8
The posture guidance law error judgment system of formation.
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Citations (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04283615A (en) * | 1991-03-13 | 1992-10-08 | Tokimec Inc | Gyroscope apparatus |
| EP0999128A1 (en) * | 1998-11-06 | 2000-05-10 | Société Européenne des Satellites | Method for determining the yaw angle of a satellite |
| US20120104150A1 (en) * | 2010-11-01 | 2012-05-03 | Honeywell International Inc. | Projectile 3d attitude from 3-axis magnetometer and single-axis accelerometer |
| US20120221244A1 (en) * | 2011-02-28 | 2012-08-30 | Trusted Positioning Inc. | Method and apparatus for improved navigation of a moving platform |
| EP2685214A2 (en) * | 2012-07-10 | 2014-01-15 | Honeywell International Inc. | Multiple truth reference system and method |
| CN103728980A (en) * | 2014-01-08 | 2014-04-16 | 哈尔滨工业大学 | Spacecraft relative orbit control method |
| CN103852082A (en) * | 2012-11-30 | 2014-06-11 | 上海航天控制工程研究所 | Inter-satellite measurement and gyro attitude orbit integrated smoothing estimation method |
| CN104765373A (en) * | 2015-01-30 | 2015-07-08 | 北京控制工程研究所 | On-satellite relative motion state acquiring method |
| CN105094139A (en) * | 2015-07-24 | 2015-11-25 | 上海微小卫星工程中心 | Elliptical short half shaft control method of close-range spacecraft coplanar ellipse formation |
| CN105159304A (en) * | 2015-06-26 | 2015-12-16 | 哈尔滨工业大学 | Finite time fault-tolerant control method for approaching and tracking space non-cooperative target |
| CN105222788A (en) * | 2015-09-30 | 2016-01-06 | 清华大学 | The automatic correcting method of the aircraft course deviation shift error of feature based coupling |
| CN105678061A (en) * | 2015-12-31 | 2016-06-15 | 中国人民解放军国防科学技术大学 | Space station safe region design method capable of satisfying passive safety requirement |
| CN106227225A (en) * | 2016-08-25 | 2016-12-14 | 上海微小卫星工程中心 | Spacecraft coplanar formation accompanying flying configuration control method |
| CN106297422A (en) * | 2016-10-09 | 2017-01-04 | 中国人民解放军国防科学技术大学 | A kind of non-linear relevant path forecasting procedure based on least square |
| CN106384152A (en) * | 2016-09-09 | 2017-02-08 | 哈尔滨工业大学 | PF space non-cooperative target track prediction method based on firefly group optimization |
-
2017
- 2017-05-17 CN CN201710345816.1A patent/CN108958064B/en active Active
Patent Citations (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04283615A (en) * | 1991-03-13 | 1992-10-08 | Tokimec Inc | Gyroscope apparatus |
| EP0999128A1 (en) * | 1998-11-06 | 2000-05-10 | Société Européenne des Satellites | Method for determining the yaw angle of a satellite |
| US20120104150A1 (en) * | 2010-11-01 | 2012-05-03 | Honeywell International Inc. | Projectile 3d attitude from 3-axis magnetometer and single-axis accelerometer |
| US20120221244A1 (en) * | 2011-02-28 | 2012-08-30 | Trusted Positioning Inc. | Method and apparatus for improved navigation of a moving platform |
| EP2685214A2 (en) * | 2012-07-10 | 2014-01-15 | Honeywell International Inc. | Multiple truth reference system and method |
| CN103852082A (en) * | 2012-11-30 | 2014-06-11 | 上海航天控制工程研究所 | Inter-satellite measurement and gyro attitude orbit integrated smoothing estimation method |
| CN103728980A (en) * | 2014-01-08 | 2014-04-16 | 哈尔滨工业大学 | Spacecraft relative orbit control method |
| CN104765373A (en) * | 2015-01-30 | 2015-07-08 | 北京控制工程研究所 | On-satellite relative motion state acquiring method |
| CN105159304A (en) * | 2015-06-26 | 2015-12-16 | 哈尔滨工业大学 | Finite time fault-tolerant control method for approaching and tracking space non-cooperative target |
| CN105094139A (en) * | 2015-07-24 | 2015-11-25 | 上海微小卫星工程中心 | Elliptical short half shaft control method of close-range spacecraft coplanar ellipse formation |
| CN105222788A (en) * | 2015-09-30 | 2016-01-06 | 清华大学 | The automatic correcting method of the aircraft course deviation shift error of feature based coupling |
| CN105678061A (en) * | 2015-12-31 | 2016-06-15 | 中国人民解放军国防科学技术大学 | Space station safe region design method capable of satisfying passive safety requirement |
| CN106227225A (en) * | 2016-08-25 | 2016-12-14 | 上海微小卫星工程中心 | Spacecraft coplanar formation accompanying flying configuration control method |
| CN106384152A (en) * | 2016-09-09 | 2017-02-08 | 哈尔滨工业大学 | PF space non-cooperative target track prediction method based on firefly group optimization |
| CN106297422A (en) * | 2016-10-09 | 2017-01-04 | 中国人民解放军国防科学技术大学 | A kind of non-linear relevant path forecasting procedure based on least square |
Non-Patent Citations (3)
| Title |
|---|
| ZHENG JUN: "Analysis of Manned Spacecraft Attitude Variation to Relay Terminal Antenna Pointing Error", 《2015 7TH INTERNATIONAL CONFERENCE ON INTELLIGENT HUMAN-MACHINE SYSTEMS AND CYBERNETICS》 * |
| 刘胜: "基于C-W方程的近程导引制导与控制方法", 《上海航天》 * |
| 白显宗: "空间目标轨道预报误差与碰撞概率问题研究", 《中国博士学位论文全文数据库 工程科技Ⅱ辑》 * |
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