CN109269508A - A kind of satellite is with respect to asteroid vision autonomous navigation method - Google Patents

Abstract

The invention discloses a kind of satellites with respect to asteroid vision autonomous navigation method, can provide high precision position and velocity information for satellite formation flying, effectively solves the problems, such as that navigation accuracy caused by satellite formation flying observation information deficiency is lower.Star sensor of the present invention is the heavenly body sensor for observing fixed star, and need to meet specified conditions using star sensor observation asteroid, the present invention proposes that asteroid is observed the illumination condition and star sensor observation condition for needing to meet, solve the problems, such as traditional star sensor can only passive measurement, improve autonomous observation accuracy.On the basis of being observed between realizing star, the present invention proposes to calculate asteroid relative satellite orientation vector and azimuth and pitch angle method in real time, and asteroid is continuously tracked using universal axial adjustment star sensor optical axis direction, it solves the problems, such as that conventional observation can not be continuously tracked, improves continuous observation efficiency between star.

Description

A kind of satellite is with respect to asteroid vision autonomous navigation method

Technical field

The present invention relates to the space measurement fields of spacecraft deep space exploration, regard more particularly to a kind of satellite with respect to asteroid Feel autonomous navigation method.

Background technique

With the development of deep space exploration technology, asteroid detection has become one of the important content of 21 century deep space exploration, Carry out asteroid detection and not only facilitate the mystery for opening the solar system and origin of life, evolution, and can promote earth protection, sky Between science and space technology application development, verifying can be provided for farther deep space exploration key technology.

Currently, each main space power all attaches great importance to deep space exploration formulation work, specified from strategic height Respective developing direction and emphasis, but be still in the primary stage at present to the detection operations of asteroid, the relevant technologies are not yet mature, Need further verify and it is perfect.The main means of asteroid detection include: to leap detection, detection of being diversion, landing sampling detection Three kinds of modes, wherein requiring to carry out leaping detection and accompanying flying detection near asteroid through spaceborne monitoring device pair Asteroid carries out remote observation, with external informations such as the landform, the landforms that obtain asteroid.Further, since gravitation around asteroid The features such as field complexity, observing and controlling absence of information, big detector and ground control station communication delay, traditional radio tracking technology are difficult To meet the requirement of navigation real-time, so that the autonomous navigation technology in asteroid detection task becomes asteroid detection technology One of emphasis for needing to study.

In conclusion this allows for grinding since observation asteroid can not be continuously tracked in traditional Visible Light Camera always The novel optical measurement method studied carefully for asteroid seems urgent important, not only can provide technology for China's deep space exploration task Deposit, while the progress of other correlative study projects and engineering project can also be promoted, it is done for the progress of China's aeronautical and space technology It contributes out.

Summary of the invention

Goal of the invention: the present invention satellite can not be continuously tracked always observation asteroid lead to loss of learning aiming at the problem that, Deep space exploration towards asteroid proposes that a kind of satellite utilizes the autonomous continuous observation asteroid progress independent navigation of star sensor New method provides high-precision opposite observation information for deep space exploration satellite.

Technical solution: to reach this purpose, the invention adopts the following technical scheme:

Satellite of the present invention is with respect to asteroid vision autonomous navigation method, comprising the following steps:

S1: using target asteroid as tracking observation object, according to asteroid ephemeris, satellite theory flight track ginseng is designed Number；

S2: it according to the kinetic model of satellite relative target asteroid track under geocentric inertial coordinate system, establishes from leading Boat System State Model；

S3: calculating satellite and asteroid relative distance according to step S1, judges whether asteroid meets star sensor observation Required distance: if it is satisfied, then entering step S4；Otherwise, S12 is entered step；

S4: the positional relationship of the sun, the earth and asteroid three is resolved according to step S1, judges whether asteroid is in too Positive area of illumination: if it is, entering step S5；Otherwise, S12 is entered step；

S5: resolving the positional relationship of the earth, satellite and asteroid three according to step S1, judges that the earth is no and enters star sensitivity Device visual field: if it is, entering step S6；Otherwise, S12 is entered step；

S6: according to step S1 calculate asteroid can the apparent magnitude, judge asteroid can the apparent magnitude whether be less than star sensor can Observe threshold value: if it is, entering step S7；Otherwise, S12 is entered step；

S7: asteroid relative satellite direction vector is calculated according to step S1 and star sensor optical axis is directed toward angle, is judged small Whether planet is in star sensor field range: if it is, entering step S8；Otherwise, pass through universal axial adjustment star sensor Optical axis is directed toward, and judges that asteroid is to enter step S8, otherwise whether in star sensor field range again according to step S7 Into S12；

S8: asteroid is calculated in star sensor two dimension image planes battle array coordinate according to step S7, judges whether asteroid is quick in star In sensor two dimension image planes battle array: if it is, entering step S9；Otherwise, S12 is entered step；

S9: theory orientation vector, azimuth and the pitch angle of asteroid relative satellite are calculated；

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed asteroid, Asteroid relative satellite true directions vector is calculated, is established using unit direction vector and distance as the observational equation of observed quantity；

S11: the observational equation discretization established to the step S2 state model established and step S10 utilizes Unscented Kalman filtering algorithm estimates satellite position and speed；

S12: terminate observation.

Further, in the step S2, the process for establishing autonomous navigation system state model is as follows:

Under geocentric inertial coordinate system, when satellite position distance is greater than asteroid and satellite relative distance, establish Satellite relative target asteroid dynamics of orbits model:

Wherein, δ r^(Ast0)It is satellite with respect to asteroid direction vector, δ v^(Ast0)It is satellite with respect to asteroid velocity vector,For the first derivative of the relatively small planetary position of satellite at any time,For satellite with respect to asteroid speed at any time one Order derivative, r⁽⁰⁾For satellite position vectors, μ_eFor Gravitational coefficient of the Earth, a_fFor perturbation acceleration；

Definition status variable x=[(δ r^(Ast0))^T(δv^(Ast0))^T]^T, establish autonomous navigation system state model；

Wherein, f is mission nonlinear continuous state transfer function, w_tFor state-noise.

Further, in the step S3, judge that the no process for meeting star sensor observed range requirement of asteroid is as follows: meter Satellite is calculated with respect to asteroid distance δ r^(Ast0), judge δ r^(Ast0)Whether formula (3) shown in condition is met:

L_min≤δr^(Ast0)≤L_max (3)

Wherein, δ r^(Ast0)It is satellite with respect to asteroid distance, is represented by δ r^(Ast0)=| δ r^(Ast0)|=| r^(Ast)-r⁽⁰⁾|, r⁽⁰⁾For satellite position vectors, r^(Ast)For asteroid position vector, L_minMinimum range needed for being observed between star, L_maxIt is observed between star Required maximum distance.

Further, in the step S4, judge asteroid whether be in solar irradiation area process it is as follows: analysis the earth yin Shadow range and asteroid travel through the critical condition of the shaded region, if asteroid position vector r^(Ast)It is sweared with position of sun Measure r^(sun)Angle is ψ, and the critical angle that asteroid enters and leaves earth's shadow range is respectivelyWithThen asteroid Solar irradiation area is in need to meet condition:

Further, in the step S5, judge the earth whether enter star sensor visual field process it is as follows: set satellite position Vector r⁽⁰⁾With satellite with respect to asteroid direction vector δ r^(Ast0)Angle be θ, facing of causing background light excessively weak is blocked by the earth Boundary's condition is satellite with respect to asteroid direction vector δ r^(Ast0)Tangent with earth edge, defining this critical angle is θ_cri, then the earth Do not enter star sensor viewing conditions are as follows:

θ > θ_cri (5)。

Further, in the step S6, asteroid can the apparent magnitude be calculated according to formula (6):

In formula (6), m be asteroid can the apparent magnitude, M be asteroid absolute magnitude, be calculated according to formula (7)；|r^(sunAst) | it is the distance between the sun and asteroid；ξ is satellite with respect to asteroid direction vector δ r^(Ast0)Asteroid opposite with sun direction Vector r^(sunAst)Angle, p (ξ) is phase integral, is calculated according to formula (8)；d₀It is the average departure between the earth and the sun From；

In formula (7), m_sunBe the sun can the apparent magnitude, r_dFor the radius for being observed celestial body, a is the reflection for being observed celestial body Rate；

Further, in the step S7, judging asteroid, whether process in star sensor field range is as follows: ifThen asteroid is in star sensor field range；Otherwise, then asteroid not in star sensor field range；Such as Shown in formula (9), FOV is star sensor field angle；

In formula (9),For star sensor direction vector,It is geocentric inertial coordinate system opposing body's coordinate system pose Transition matrix；

In the step S8, judging asteroid, whether process in star sensor two dimension image planes battle array is as follows: if met Formula (10a) and (10b), then asteroid is in star sensor two dimension image planes battle array；

Wherein,It is satellite with respect to asteroid direction vector δ r^(Ast0)It is projected in star sensor two dimension image planes battle array Coordinate, IP_longthFor the length of two-dimentional image planes battle array, IP_widthFor the width of two-dimentional image planes battle array.

Further, in the step S9, satellite distinguishes root with respect to the theory orientation vector of asteroid, azimuth and pitch angle It is calculated according to formula (11), (12a) and (12b):

In formula (11), δ r^(Ast0)Theory orientation vector for satellite with respect to asteroid, also referred to as satellite is with respect to asteroid side To vector；It is satellite with respect to asteroid unit direction vector；

Wherein, It is that geocentric inertial coordinate system opposing body sits Mark system pose transformation matrix, α are azimuth, and δ is pitch angle.

Further, in the step S10, the observational equation of foundation are as follows:

In formula (13), It is satellite with respect to asteroid unit The true measurement of direction vector,True measurement for satellite with respect to asteroid distance.

Further, detailed process is as follows by the step S11:

Discretization is carried out according to the state model that formula (14a) establishes step S2:

In formula (14a), x_k∈R^LFor k-th of state variable, x_k+1For+1 state variable of kth, w_k∈N(0,Q_k) it is kth A process noise, Q_kFor system noise intensity, f is mission nonlinear continuous state transfer function；

Discretization is carried out according to the observational equation that formula (14b) establishes step S10:

y_k=g (x_k)+v_k (14b)

In formula (14b), y_k∈R^MFor k-th of output vector, g is observational equation, v_k∈N(0,R_k) it is that k-th of measurement is made an uproar Sound, R_kFor observation noise intensity, w_kAnd v_kIt is uncorrelated；

Satellite position and speed are estimated using Unscented Kalman filtering algorithm, and steps are as follows:

S11.1: Unscented transformation is carried out according to formula (15a), (15b) and (15c):

Wherein,For x_kMean value,For x_kVariance, λ=α²(1+ κ) -1 is scalar parameter, and α determines that sigma point existsThe distribution of surrounding, κ are scalar parameter, and L is sigma point dimension, χ_i,k-1It is intermediate variable；

S11.2: it prediction process: is predicted according to formula (16a), (16b), (16c), (16d), (16e) and (16f)；

χ_i,k/k-1=f (χ_i,k-1) (16a)

Wherein, W_iWithRespectively weight coefficient used in U transformation calculations mean value and variance, χ_i,k/k-1、 Y_i,k/k-1WithIt is all intermediate variable；

S11.3: it renewal process: is updated according to formula (17a), (17b), (17c), (17d) and (17e)；

Wherein,And K_kIt is all intermediate variable；

S11.4: return step S11.1 carries out the filtering of next cycle.

The utility model has the advantages that can be compiled the invention discloses a kind of opposite asteroid vision autonomous navigation method of satellite for satellite Team's flight provides high precision position and velocity information, effectively solves navigation essence caused by satellite formation flying observation information deficiency Spend lower problem.Compared with the existing technology, the invention has the following advantages:

1) star sensor is the heavenly body sensor for observing fixed star, and needs to meet using star sensor observation asteroid Specified conditions, the present invention propose that asteroid is observed the illumination condition and star sensor observation condition for needing to meet, solve tradition Star sensor can only passive measurement problem, improve autonomous observation accuracy；

2) on the basis of observing between realizing star, the present invention proposes to calculate asteroid relative satellite orientation vector and orientation in real time Angle and pitch angle method, and asteroid is continuously tracked using universal axial adjustment star sensor optical axis direction, solve conventional observation Problem can not be continuously tracked, improve continuous observation efficiency between star.

Detailed description of the invention

Fig. 1 is the method flow diagram in the specific embodiment of the invention；

Fig. 2 is that specific embodiment of the invention Satellite star sensor observes asteroid flow chart；

Fig. 3 is specific embodiment of the invention Satellite with respect to specific distance range schematic diagram between asteroid star；

Fig. 4 is asteroid illumination condition schematic diagram in the specific embodiment of the invention；

Fig. 5 is star sensor visual field and position of the earth relation schematic diagram in the specific embodiment of the invention；

Fig. 6 is that the visual magnitude of asteroid calculates schematic diagram in the specific embodiment of the invention；

Fig. 7 is asteroid in the specific embodiment of the invention in star sensor two-dimensional image area array projection schematic diagram；

Fig. 8 is the small phase of planets in the specific embodiment of the invention to satellite direction vector and azimuth schematic diagram.

Specific embodiment

Technical solution of the present invention is further introduced with attached drawing With reference to embodiment.

Present embodiment discloses a kind of satellite with respect to asteroid vision autonomous navigation method, such as Fig. 1 and Fig. 2 institute Show, be directed to the asteroid deep space exploration stage, satellite carries out independent navigation using the autonomous continuous observation asteroid of star sensor, is One kind being very suitable for deep space exploration autonomous navigation of satellite method.Specifically includes the following steps:

S1: using target asteroid as tracking observation object, according to asteroid ephemeris, satellite theory flight track ginseng is designed Number；Wherein, satellite theory flight track parameter includes semi-major axis of orbit a, orbital eccentricity e, orbit inclination angle i, right ascension of ascending node Ω, argument of perigee ω, time of perigee passage t_p；

S2: it according to the kinetic model of satellite relative target asteroid track under geocentric inertial coordinate system, establishes from leading Boat System State Model；

S3: calculating satellite and asteroid relative distance according to step S1, judges whether asteroid meets star sensor observation Required distance: if it is satisfied, then entering step S4；Otherwise, S12 is entered step；

S4: the positional relationship of the sun, the earth and asteroid three is resolved according to step S1, judges whether asteroid is in too Positive area of illumination: if it is, entering step S5；Otherwise, S12 is entered step；

S5: resolving the positional relationship of the earth, satellite and asteroid three according to step S1, judges that the earth is no and enters star sensitivity Device visual field: if it is, entering step S6；Otherwise, S12 is entered step；

S6: according to step S1 calculate asteroid can the apparent magnitude, judge asteroid can the apparent magnitude whether be less than star sensor can Observe threshold value: if it is, entering step S7；Otherwise, S12 is entered step；

S7: asteroid relative satellite direction vector is calculated according to step S1 and star sensor optical axis is directed toward angle, is judged small Whether planet is in star sensor field range: if it is, entering step S8；Otherwise, pass through universal axial adjustment star sensor Optical axis is directed toward, and judges that asteroid is to enter step S8, otherwise whether in star sensor field range again according to step S7 Into S12；

S8: asteroid is calculated in star sensor two dimension image planes battle array coordinate according to step S7, judges whether asteroid is quick in star In sensor two dimension image planes battle array: if it is, entering step S9；Otherwise, S12 is entered step；

S9: theory orientation vector, azimuth and the pitch angle of asteroid relative satellite are calculated；

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed asteroid, Asteroid relative satellite true directions vector is calculated, is established using unit direction vector and distance as the observational equation of observed quantity；

S11: the observational equation discretization established to the step S2 state model established and step S10 utilizes Unscented Kalman filtering algorithm estimates satellite position and speed；

S12: terminate observation.

In step S2, the process for establishing autonomous navigation system state model is as follows:

Under geocentric inertial coordinate system, when satellite position distance is greater than asteroid and satellite relative distance, establish Satellite relative target asteroid dynamics of orbits model:

Wherein, δ r^(Ast0)It is satellite with respect to asteroid direction vector, δ v^(Ast0)It is satellite with respect to asteroid velocity vector,For the first derivative of the relatively small planetary position of satellite at any time,For satellite with respect to asteroid speed at any time one Order derivative, r⁽⁰⁾For satellite position vectors, μ_eFor Gravitational coefficient of the Earth, a_fFor perturbation acceleration；

Definition status variable x=[(δ r^(Ast0))^T(δv^(Ast0))^T]^T, establish autonomous navigation system state model；

Wherein, f is mission nonlinear continuous state transfer function, w_tFor state-noise.

In step S3, judge that the no process for meeting star sensor observed range requirement of asteroid is as follows: it is opposite to calculate satellite Asteroid distance δ r^(Ast0), as shown in figure 3, judging δ r^(Ast0)Whether formula (3) shown in condition is met:

L_min≤δr^(Ast0)≤L_max (3)

Wherein, δ r^(Ast0)It is satellite with respect to asteroid distance, is represented by δ r^(Ast0)=| δ r^(Ast0)|=| r^(Ast)-r⁽⁰⁾|, r⁽⁰⁾For satellite position vectors, r^(Ast)For asteroid position vector, L_minMinimum range needed for being observed between star, L_maxIt is observed between star Required maximum distance.

When moonscope asteroid, asteroid needs are sufficiently irradiated by sunlight.When asteroid is at global illumination area, Asteroid can sufficiently be irradiated by sunlight；Conversely, when asteroid enters earth's shadow area, since the earth blocks, sunlight without Method is irradiated to asteroid, it is therefore desirable to judge asteroid illumination condition.In step S4, judge whether asteroid is in too The process of positive area of illumination is as follows: as shown in figure 4, analysis earth shaded region and asteroid travel through facing for the shaded region Boundary's condition, if asteroid position vector r^(Ast)With position of sun vector r^(sun)Angle is ψ, and asteroid enters and leaves earth yin The critical angle of shadow range is respectivelyWithThen asteroid is in solar irradiation area and needs to meet condition:

Wherein,Re is Earth radius, r⁽⁺⁾And r^(-)Critical localisation when earth's shadow is entered and left for asteroid.

In step S5, judge the earth whether enter star sensor visual field process it is as follows: as shown in figure 5, setting satellite position Vector r⁽⁰⁾With satellite with respect to asteroid direction vector δ r^(Ast0)Angle be θ, facing of causing background light excessively weak is blocked by the earth Boundary's condition is satellite with respect to asteroid direction vector δ r^(Ast0)Tangent with earth edge, defining this critical angle is θ_cri, then the earth Do not enter star sensor viewing conditions are as follows:

θ > θ_cri (5)。

In step S6, asteroid can the apparent magnitude be calculated according to formula (6):

In formula (6), m be asteroid can the apparent magnitude, M be asteroid absolute magnitude, be calculated according to formula (7)；|r^(sunAst) | it is the distance between the sun and asteroid；ξ is satellite with respect to asteroid direction vector δ r^(Ast0)Asteroid opposite with sun direction Vector r^(sunAst)Angle be calculated as shown in fig. 6, p (ξ) is phase integral according to formula (8)；d₀Be the earth and the sun it Between average distance；

In formula (7), m_sunBe the sun can the apparent magnitude, r_dFor the radius for being observed celestial body, a is the reflection for being observed celestial body Rate；

In step S7, judging asteroid, whether process in star sensor field range is as follows: ifThen Asteroid is in star sensor field range；Otherwise, then asteroid not in star sensor field range；As shown in formula (9), FOV is star sensor field angle；

In formula (9),For star sensor direction vector,It is geocentric inertial coordinate system opposing body's coordinate system pose Transition matrix；

In step S8, judging asteroid, whether process in star sensor two dimension image planes battle array is as follows: if meeting formula (10a) and (10b), then asteroid is in star sensor two dimension image planes battle array；

Wherein,It is satellite with respect to asteroid direction vector δ r^(Ast0)It is projected in star sensor two dimension image planes battle array Coordinate, as shown in fig. 7, IP_longthFor the length of two-dimentional image planes battle array, IP_widthFor the width of two-dimentional image planes battle array.

In step S9, satellite with respect to asteroid theory orientation vector, azimuth and pitch angle respectively according to formula (11), (12a) and (12b) is calculated, as shown in Figure 8:

In formula (11), δ r^(Ast0)Theory orientation vector for satellite with respect to asteroid, also referred to as satellite is with respect to asteroid side To vector；It is satellite with respect to asteroid unit direction vector；

Wherein, It is that geocentric inertial coordinate system opposing body sits Mark system pose transformation matrix, α are azimuth, and δ is pitch angle.

In step S10, the observational equation of foundation are as follows:

In formula (13), It is satellite with respect to asteroid unit The true measurement of direction vector,True measurement for satellite with respect to asteroid distance.

Detailed process is as follows by step S11:

Discretization is carried out according to the state model that formula (14a) establishes step S2:

In formula (14a), x_k∈R^LFor k-th of state variable, x_k+1For+1 state variable of kth, w_k∈N(0,Q_k) it is kth A process noise, Q_kFor system noise intensity, f is mission nonlinear continuous state transfer function；

Discretization is carried out according to the observational equation that formula (14b) establishes step S10:

y_k=g (x_k)+v_k (14b)

In formula (14b), y_k∈R^MFor k-th of output vector, g is observational equation, v_k∈N(0,R_k) it is that k-th of measurement is made an uproar Sound, R_kFor observation noise intensity, w_kAnd v_kIt is uncorrelated；

Satellite position and speed are estimated using Unscented Kalman filtering algorithm, and steps are as follows:

S11.1: Unscented transformation is carried out according to formula (15a), (15b) and (15c):

Wherein,For x_kMean value,For x_kVariance, λ=α²(1+ κ) -1 is scalar parameter, and α determines sigma point ?The distribution of surrounding, κ are scalar parameter, and L is sigma point dimension, χ_i,k-1It is intermediate variable；

S11.2: it prediction process: is predicted according to formula (16a), (16b), (16c), (16d), (16e) and (16f)；

χ_i,k/k-1=f (χ_i,k-1) (16a)

Wherein, W_iWithRespectively weight coefficient used in U transformation calculations mean value and variance, χ_i,k/k-1、 Y_i,k/k-1WithIt is all intermediate variable；

S11.3: it renewal process: is updated according to formula (17a), (17b), (17c), (17d) and (17e)；

Wherein,And K_kIt is all intermediate variable；

S11.4: return step S11.1 carries out the filtering of next cycle.

Claims (10)

1. a kind of satellite is with respect to asteroid vision autonomous navigation method, it is characterised in that: the following steps are included:

S1: using target asteroid as tracking observation object, according to asteroid ephemeris, satellite theory flight track parameter is designed；

S2: according to the kinetic model of satellite relative target asteroid track under geocentric inertial coordinate system, independent navigation system is established System state model；

S3: satellite and asteroid relative distance are calculated according to step S1, judge whether asteroid meets star sensor observed range It is required that: if it is satisfied, then entering step S4；Otherwise, S12 is entered step；

S4: the positional relationship of the sun, the earth and asteroid three is resolved according to step S1, judges whether asteroid is in sunlight According to area: if it is, entering step S5；Otherwise, S12 is entered step；

S5: resolving the positional relationship of the earth, satellite and asteroid three according to step S1, judges that the no star sensor that enters of the earth regards : if it is, entering step S6；Otherwise, S12 is entered step；

S6: according to step S1 calculate asteroid can the apparent magnitude, judge asteroid can the apparent magnitude whether be less than star sensor Observable Threshold value: if it is, entering step S7；Otherwise, S12 is entered step；

S7: asteroid relative satellite direction vector is calculated according to step S1 and star sensor optical axis is directed toward angle, judges asteroid Whether in star sensor field range: if it is, entering step S8；Otherwise, pass through universal axial adjustment star sensor optical axis It is directed toward, judges that asteroid whether in star sensor field range, is to enter step S8, otherwise enters again according to step S7 S12；

S8: asteroid is calculated in star sensor two dimension image planes battle array coordinate according to step S7, judges asteroid whether in star sensor In two-dimentional image planes battle array: if it is, entering step S9；Otherwise, S12 is entered step；

S9: theory orientation vector, azimuth and the pitch angle of asteroid relative satellite are calculated；

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed asteroid, calculates Asteroid relative satellite true directions vector is established using unit direction vector and distance as the observational equation of observed quantity；

S11: the observational equation discretization established to the step S2 state model established and step S10 utilizes Unscented Kalman filtering algorithm estimates satellite position and speed；

S12: terminate observation.

2. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S2 In, the process for establishing autonomous navigation system state model is as follows:

Under geocentric inertial coordinate system, when satellite position distance is greater than asteroid and satellite relative distance, satellite is established Relative target asteroid dynamics of orbits model:

Wherein, δ r^(Ast0)It is satellite with respect to asteroid direction vector, δ v^(Ast0)It is satellite with respect to asteroid velocity vector, For the first derivative of the relatively small planetary position of satellite at any time,It is led for satellite with respect to the single order of asteroid speed at any time Number, r⁽⁰⁾For satellite position vectors, μ_eFor Gravitational coefficient of the Earth, a_fFor perturbation acceleration；

Definition status variable x=[(δ r^(Ast0))^T (δv^(Ast0))^T]^T, establish autonomous navigation system state model；

Wherein, f is mission nonlinear continuous state transfer function, w_tFor state-noise.

3. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S3 In, judge that the no process for meeting star sensor observed range requirement of asteroid is as follows: calculating satellite with respect to asteroid distance δ r^(Ast0), judge δ r^(Ast0)Whether formula (3) shown in condition is met:

L_min≤δr^(Ast⁰⁾≤L_max (3)

Wherein, δ r^(Ast0)It is satellite with respect to asteroid distance, is represented by δ r^(Ast0)=| δ r^(Ast0)|=| r^(Ast)-r⁽⁰⁾|, r⁽⁰⁾ For satellite position vectors, r^(Ast)For asteroid position vector, L_minMinimum range needed for being observed between star, L_maxNeeded for being observed between star Maximum distance.

4. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S4 In, judge asteroid whether be in solar irradiation area process it is as follows: analysis earth shaded region and asteroid travel through The critical condition of the shaded region, if asteroid position vector r^(Ast)With position of sun vector r^(sun)Angle is ψ, asteroid into The critical angle for entering and leaving earth's shadow range is respectivelyWithThen asteroid is in solar irradiation area and needs to meet item Part:

5. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S5 In, judge the earth whether enter star sensor visual field process it is as follows: set satellite position vectors r⁽⁰) and satellite with respect to asteroid side To vector delta r^(Ast0)Angle be θ, the critical condition that being blocked by the earth causes background light excessively weak is satellite with respect to asteroid side To vector delta r^(Ast0)Tangent with earth edge, defining this critical angle is θ_cri, then the earth does not enter star sensor viewing conditions Are as follows:

θ > θ_cri (5)。

6. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S6 In, asteroid can the apparent magnitude be calculated according to formula (6):

In formula (6), m be asteroid can the apparent magnitude, M be asteroid absolute magnitude, be calculated according to formula (7)；|r^(sunAst)| it is The distance between the sun and asteroid；ξ is satellite with respect to asteroid direction vector δ r^(Ast0)Asteroid opposite with sun direction arrow Measure r^(sunAst)Angle, p (ξ) is phase integral, is calculated according to formula (8)；d₀It is the average distance between the earth and the sun；

In formula (7), m_sunBe the sun can the apparent magnitude, r_dFor the radius for being observed celestial body, a is the reflectivity for being observed celestial body；

7. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S7 In, judging asteroid, whether process in star sensor field range is as follows: if

Then asteroid is in star sensor field range；Otherwise, then asteroid not in star sensor field range；As shown in formula (9), FOV is star sensor field angle；

In formula (9),For star sensor direction vector,It is geocentric inertial coordinate system opposing body's coordinate system pose conversion square Battle array；

In the step S8, judging asteroid, whether process in star sensor two dimension image planes battle array is as follows: if meeting formula (10a) and (10b), then asteroid is in star sensor two dimension image planes battle array；

Wherein,It is satellite with respect to asteroid direction vector δ r^(Ast0)It is projected in the seat of star sensor two dimension image planes battle array Mark, IP_longthFor the length of two-dimentional image planes battle array, IP_widthFor the width of two-dimentional image planes battle array.

8. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step S9 In, satellite is calculated according to formula (11), (12a) and (12b) respectively with respect to theory orientation vector, azimuth and the pitch angle of asteroid It obtains:

In formula (11), δ r^(Ast0)Theory orientation vector for satellite with respect to asteroid, also referred to as satellite are sweared with respect to asteroid direction Amount；It is satellite with respect to asteroid unit direction vector；

Wherein, It is geocentric inertial coordinate system opposing body's coordinate system appearance State transition matrix, α are azimuth, and δ is pitch angle.

9. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step In S10, the observational equation of foundation are as follows:

In formula (13), It is satellite with respect to asteroid unit direction The true measurement of vector,True measurement for satellite with respect to asteroid distance.

10. satellite according to claim 1 is with respect to asteroid vision autonomous navigation method, it is characterised in that: the step Detailed process is as follows by S11:

Discretization is carried out according to the state model that formula (14a) establishes step S2:

In formula (14a), x_k∈R^LFor k-th of state variable, x_k+1For+1 state variable of kth, w_k∈N(0,Q_k) it is k-th of process Noise, Q_kFor system noise intensity, f is mission nonlinear continuous state transfer function；

Discretization is carried out according to the observational equation that formula (14b) establishes step S10:

y_k=g (x_k)+v_k (14b)

In formula (14b), y_k∈R^MFor k-th of output vector, g is observational equation, v_k∈N(0,R_k) it is k-th of measurement noise, R_kFor Observation noise intensity, w_kAnd v_kIt is uncorrelated；

Satellite position and speed are estimated using Unscented Kalman filtering algorithm, and steps are as follows:

S11.1: Unscented transformation is carried out according to formula (15a), (15b) and (15c):

Wherein,For x_kMean value,For x_kVariance, λ=α²(1+ κ) -1 is scalar parameter, and α determines that sigma point existsWeek The distribution enclosed, κ are scalar parameter, and L is sigma point dimension, χ_i,k-1It is intermediate variable；

S11.2: it prediction process: is predicted according to formula (16a), (16b), (16c), (16d), (16e) and (16f)；

χ_i,k/k-1=f (χ_i,k-1) (16a)

Wherein, W_iAnd W_i ^*Respectively weight coefficient used in U transformation calculations mean value and variance, χ_i,k/k-1、 Y_i,k/k-1WithIt is all intermediate variable；

S11.3: it renewal process: is updated according to formula (17a), (17b), (17c), (17d) and (17e)；

Wherein,And K_kIt is all intermediate variable；

S11.4: return step S11.1 carries out the filtering of next cycle.