CN109269509A - In face of the space multi-robot self-determination air navigation aid of geostationary orbit object run - Google Patents

Abstract

The invention discloses a kind of space multi-robot self-determination air navigation aids in face of geostationary orbit object run, high precision position and velocity information can be provided for satellite formation flying, effectively solve the problems, such as that navigation accuracy caused by satellite formation flying observation information deficiency is lower.Star sensor is the heavenly body sensor for observing fixed star, and Inter-satellite relative measure is carried out using star sensor and needs to meet specified conditions, the invention proposes the illumination conditions that needs are observed between star and star sensor observation condition, solve the problems, such as traditional star sensor can only passive measurement, improve and independently select star accuracy.On the basis of being observed between realizing star, the invention proposes the real-time primarys that calculates with respect to component orientation vector and azimuth and pitch angle method, and component 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

In face of the space multi-robot self-determination air navigation aid of geostationary orbit object run

Technical field

The present invention relates to spacecraft in-orbit service space measurement fields, face geostationary orbit mesh more particularly to one kind Mark the space multi-robot self-determination air navigation aid of operation.

Background technique

Geostationary orbit (Geostationary orbit, GEO) is the unique track resources of the mankind, and being located at should The satellite (abbreviation GEO satellite) of track, area coverage is big, and relative to ground be it is static, in communication, navigation, early warning, meteorology Equal civil and militaries field just increasingly plays increasingly important role.For certain tasks, need multiple Satellite Networkings, It forms constellation and has remained 5 (3 works on GEO track such as DSP (Defense support program) ballistic missile early warning satellite in the U.S. Make, 2 are spare) satellite；The high rail section of its space-based infrared system (SBIRS) also includes 4 GEO satellites and 2 highly elliptic orbits Satellite.The dipper system built is China's independent development capability, independently operated Global Satellite Navigation System, by 5 static rails Road satellite and 30 other types satellite compositions.Compared with developed countries, the on-orbit fault rate of China's satellite is higher, loses in recent years The important GEO satellite of effect includes No. two satellites of prosperous promise (2006), 04 star of Beidou No.1 (2007), Nigeria's star (2007 Year transmitting, fails for 2008), Beidou G2 star (2009) etc., seriously affected the development of China's Aerospace Technology.Especially Beidou The failure of G2 star (one of 5 GEO satellites in Beidou satellite navigation system), affects the networking process of entire navigation system, makes China have to transmit again in 2012 substitution star (G2R, also known as G6) pinpoint with the position of the failure astrology away from 0.2 ° It sets.

To ensure that in-orbit spacecraft is run steadily in the long term, and protect GEO track resources, it is necessary to which development is with robot for space In-orbit service technology for the purpose of means, satellite maintenence and space trash removing.Due to orbit altitude is high, third body gravitation can not Ignore, transmitting, management and the maintenance cost of the robot for space itself for GEO service are also very high.Therefore space multimachine is needed Device people system carries out in-orbit maintenance to more GEO satellites in a certain segmental arc, can greatly save the cost of satellite maintenance, improves Rail efficiency of service.Current Space Robot System, ETS-VII, Orbital Express including emitted and in-orbit demonstration, and The systems such as FREND, the DEOS carried out are using single satellite as service object, and service content is more single, Bu Nengman The sufficient purpose in more spacecraft maintainable technology on-orbits of GEO track, therefore study the space multirobot towards GEO track in-orbit service System is imperative.

In order to which space multirobot carries out formation flight near GEO satellite, fixed point is kept and position control altogether, it is necessary to first Position and the posture information of satellite can be first obtained in real time, and cannot generate interference to star is faced, since GEO satellite usually exists There is observation signal deficiency of navigating in 36000km height, space multirobot: 1. commonly using GNSS navigation mode and there is navigation Signal is weak, the earth blocks and the serious problems such as visible satellite is few；2. other independent navigation modes: earth's magnetic field is not available, radar Altimeter and celestial navigation can not provide high-precision navigation information, all be difficult to meet navigation request, therefore this as observation information Allowing for the research new observation method of space multirobot seems urgent important.

Summary of the invention

Goal of the invention: the present invention is directed to the problem that observation information deficiency causes navigation accuracy lower, proposes a kind of in face of ground The space multi-robot self-determination air navigation aid of ball stationary orbit object run utilizes the autonomous continuous observation relative direction of star sensor The method of vector provides high-precision opposite observation information for the space multirobot of in-orbit service.

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

Space multi-robot self-determination air navigation aid of the present invention in face of geostationary orbit object run, including with Lower step:

S1: using GEO target satellite as in-orbit service object, being set to primary and component for two spaces robot, if Count primary and component formation flight configuration and orbit parameter；Orbit parameter includes that semi-major axis of orbit a, orbital eccentricity e, track incline Angle i, right ascension of ascending node Ω, argument of perigee ω, time of perigee passage t_p；

S2: according to satellite relative motion dynamics model under geocentric inertial coordinate system, autonomous navigation system state mould is established Type；

S3: primary and component relative distance are calculated according to step S1, judge whether component 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 component three is resolved according to step S1, judges whether component 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, primary and component three according to step S1, judges that the earth is no and enters star sensor Visual field: if it is, entering step S6；Otherwise, S12 is entered step；

S6: according to step S1 calculate component can the apparent magnitude, judge component 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: primary is calculated according to step S1 and is directed toward angle with respect to component direction vector and star sensor optical axis, judges component 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 component whether in star sensor field range, is to enter step S8, otherwise enters again according to step S7 S12；

S8: component is calculated in star sensor two dimension image planes battle array coordinate according to step S7, judges component 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 pitch angle of the primary with respect to component are calculated；

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed component, counts Operator star 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.

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

Under geocentric inertial coordinate system, when primary positional distance is greater than component and primary relative distance, foundation is defended Star relative target component dynamics of orbits model:

Wherein, δ r⁽¹⁰⁾It is main astrology to component direction vector, δ v⁽¹⁰⁾It is main astrology to component velocity vector,To defend Championship sets first derivative at any time,For the first derivative of satellite velocities at any time, r⁽⁰⁾For main star position vector, μ_eFor Gravitational coefficient of the Earth, a_fFor perturbation acceleration；

Definition status variable x=[(δ r⁽¹⁰⁾)^T (δv⁽¹⁰⁾)^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 in component is as follows: calculating Primary is with respect to component distance δ r⁽¹⁰⁾, judge δ r⁽¹⁰⁾Whether formula (3) shown in condition is met:

L_min≤δr⁽¹⁰⁾≤L_max (3)

Wherein, δ r⁽¹⁰⁾=| δ r⁽¹⁰⁾|=| r⁽¹⁾-r⁽⁰⁾|, r⁽⁰⁾For main star position vector, r⁽¹⁾For component position vector, L_min Minimum range needed for being observed between star, L_maxMaximum distance needed for being observed between star.

Further, in the step S4, judge component whether be in solar irradiation area process it is as follows: analysis earth's shadow Range and component travel through the critical condition of the shaded region, if component position vector r⁽¹⁾With position of sun vector r^(sun) Angle is ψ, and the critical angle that component enters and leaves earth's shadow range is ψ_cri, then component is in solar irradiation area and needs completely Sufficient condition:

ψ < ψ_cri (4)。

Further, in the step S5, judge the earth whether enter star sensor visual field process it is as follows: set primary position Vector r⁽⁰⁾With primary with respect to component direction vector δ r⁽¹⁰⁾Angle be θ, cause background light excessively weak critical is blocked by the earth Condition is primary with respect to component direction vector δ r⁽¹⁰⁾Tangent with earth edge, defining this critical angle is θ_cri, then the earth not into Enter star sensor viewing conditions are as follows:

θ > θ_cri (5)。

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

In formula (6), m be component can the apparent magnitude, M be component absolute magnitude, be calculated according to formula (7)；|r^(sun0)| it is too The distance between sun and component；ξ is primary with respect to component direction vector δ r⁽¹⁰⁾Component direction vector r opposite with the sun^(sun1)'s 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 reflection for being observed celestial body Rate；

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

In formula (9), δ r⁽¹⁰⁾It is main astrology to component direction vector,For star sensor direction vector,It is that the earth's core is used Property coordinate system opposing body's coordinate system pose transition matrix；

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

Wherein,It is main astrology to component direction vector δ r⁽¹⁰⁾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, theory orientation vector, azimuth and pitch angle respectively basis of the primary with respect to component Formula (11), (12a) and (12b) is calculated:

In formula (11), δ r⁽¹⁰⁾It is main astrology to the theory orientation vector of component, also referred to as primary is sweared with respect to component direction Amount；It is main astrology to component unit direction vector；

Wherein, It is geocentric inertial coordinate system opposing body's coordinate It is pose transformation matrix, α is azimuth, and δ is pitch angle.

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

In formula (13), It is main astrology to the true of component unit direction vector Measured value,It is main astrology to the true measurement of component 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_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.

The utility model has the advantages that the invention discloses a kind of space multi-robot self-determinations in face of geostationary orbit object run to lead Boat method can provide high precision position and velocity information for satellite formation flying, effectively solve satellite formation flying observation letter The lower problem of navigation accuracy caused by breath deficiency.Compared with the existing technology, the present invention have it is following the utility model has the advantages that

1) star sensor is the heavenly body sensor for observing fixed star, and carries out Inter-satellite relative measure using star sensor and need completely Sufficient specified conditions, the invention proposes the illumination conditions that needs are observed between star and star sensor observation condition, solve traditional star Sensor can only passive measurement problem, improve and independently select star accuracy；

2) on the basis of observing between realizing star, the invention proposes the real-time primarys that calculates with respect to component orientation vector and orientation Angle and pitch angle method, and component 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 of the specific embodiment of the invention；

Fig. 2 is that primary star sensor observes component flow chart in the specific embodiment of the invention；

Fig. 3 is primary in the specific embodiment of the invention with respect to specific distance range schematic diagram between the star of component；

Fig. 4 is specific embodiment of the invention neutron star illumination condition schematic diagram；

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 specific embodiment of the invention neutron star can apparent magnitude calculating schematic diagram；

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

Fig. 8 is primary in the specific embodiment of the invention with respect to component 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 space multi-robot self-determination in face of geostationary orbit object run and leads Boat method, as depicted in figs. 1 and 2, comprising the following steps:

S1: using GEO target satellite as in-orbit service object, being set to primary and component for two spaces robot, if Count primary and component formation flight configuration and orbit parameter；Orbit parameter includes that semi-major axis of orbit a, orbital eccentricity e, track incline Angle i, right ascension of ascending node Ω, argument of perigee ω, time of perigee passage t_p；

S2: according to satellite relative motion dynamics model under geocentric inertial coordinate system, autonomous navigation system state mould is established Type.

S3: calculating primary and component relative distance according to step S1, as shown in figure 3, judging whether component meets star sensitivity Device observed range requirement: if it is satisfied, then entering step S4；Otherwise, S12 is entered step.

S4: when primary observes component, component needs sufficiently to be irradiated by sunlight.When component is at global illumination area, son Star can sufficiently be irradiated by sunlight；Conversely, since the earth blocks, sunlight can not irradiate when component enters earth's shadow area To component, it is therefore desirable to judge sub- star illumination condition.The position of the sun, the earth and component three is resolved according to step S1 Relationship, judges whether component is in solar irradiation area: if it is, entering step S5；Otherwise, S12 is entered step.

S5: resolving the positional relationship of the earth, primary and component three according to step S1, judges that the earth is no and enters star sensor Visual field: if it is, entering step S6；Otherwise, S12 is entered step.

S6: according to step S1 calculate component can the apparent magnitude, judge component 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: primary is calculated according to step S1 and is directed toward angle with respect to component direction vector and star sensor optical axis, judges component 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 component whether in star sensor field range, is to enter step S8, otherwise enters again according to step S7 S12；

S8: component is calculated in star sensor two dimension image planes battle array coordinate according to step S7, judges component 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 pitch angle of the primary with respect to component are calculated.

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed component, counts Operator star 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.

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

Under geocentric inertial coordinate system, when primary positional distance is greater than component and primary relative distance, foundation is defended Star relative target component dynamics of orbits model:

Wherein, δ r⁽¹⁰⁾It is main astrology to component direction vector, δ v⁽¹⁰⁾It is main astrology to component velocity vector,To defend Championship sets first derivative at any time,For the first derivative of satellite velocities at any time, r⁽⁰⁾For main star position vector, μ_eFor Gravitational coefficient of the Earth, a_fFor perturbation acceleration；

Definition status variable x=[(δ r⁽¹⁰⁾)^T (δv⁽¹⁰⁾)^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 in component is as follows: it is relatively sub to calculate primary Star distance δ r⁽¹⁰⁾, as shown in figure 3, judging δ r⁽¹⁰⁾Whether formula (3) shown in condition is met:

L_min≤δr⁽¹⁰⁾≤L_max(3)

Wherein, δ r⁽¹⁰⁾=| δ r⁽¹⁰⁾|=| r⁽¹⁾-r⁽⁰⁾|, r⁽⁰⁾For main star position vector, r⁽¹⁾For component position vector, L_min Minimum range needed for being observed between star, L_maxMaximum distance needed for being observed between star.δr⁽¹⁰⁾With δ r⁽¹⁰⁾Difference be: δ r⁽¹⁰⁾With Normal font r indicates scalar, and δ r⁽¹⁰⁾Vector r is indicated with overstriking font, this is the common representation of this field, His parameter also similarly, repeats no more.

According to the sun, the earth and component three's geometry site, as shown in figure 4, determining shadow of the sun and component fortune Row track passes through the critical condition in the shadow region.In step S4, judge component whether be in solar irradiation area process it is as follows: point Analysis earth's shadow range and component travel through the critical condition of the shaded region, if component position vector r⁽¹⁾With sun position Set vector r^(sun)Angle is ψ, and the critical angle that component enters and leaves earth's shadow range is ψ_cri, then component is in sunlight It needs to meet condition according to area:

ψ < ψ_cri (4)。

Wherein,R_eIt is earth radius.

In step S5, judge the earth whether enter star sensor visual field process it is as follows: as shown in figure 5, setting primary position Vector r⁽⁰⁾With primary with respect to component direction vector δ r⁽¹⁰⁾Angle be θ, cause background light excessively weak critical is blocked by the earth Condition is primary with respect to component direction vector δ r⁽¹⁰⁾Tangent with earth edge, defining this critical angle is θ_cri, then the earth not into Enter star sensor viewing conditions are as follows:

θ > θ_cri (5)。

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

In formula (6), m be component can the apparent magnitude, M be component absolute magnitude, be calculated according to formula (7)；|r^(sun0)| it is too The distance between sun and component；As shown in fig. 6, ξ is primary with respect to component direction vector δ r⁽¹⁰⁾Component opposite with sun direction arrow Measure r^(sun1)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 reflection for being observed celestial body Rate；

In step S7, judge whether process in star sensor field range is as follows for component: ifIt is then sub Star is in star sensor field range；Otherwise, then component not in star sensor field range；As shown in formula (9), FOV is Star sensor field angle；

In formula (9), δ r⁽¹⁰⁾It is main astrology to component direction vector,For star sensor direction vector,It is that the earth's core is used Property coordinate system opposing body's coordinate system pose transition matrix；

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

Wherein,It is main astrology to component direction vector δ r⁽¹⁰⁾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, primary with respect to component theory orientation vector, azimuth and pitch angle respectively according to formula (11), (12a) and (12b) is calculated:

In formula (11), δ r⁽¹⁰⁾It is main astrology to the theory orientation vector of component, also referred to as primary is sweared with respect to component direction Amount；It is main astrology to component unit direction vector, primary can be by azimuth and pitch angle with respect to component theory direction vector Description, as shown in figure 8, in satellite body coordinate system o_b-x_by_bz_bIn, definition azimuth angle alpha is δ r⁽¹⁰⁾In o_b-y_bz_bThe projection of plane With y_bAxle clamp angle, pitch angle δ are δ r⁽¹⁰⁾With x_bAxle clamp angle, is represented by；

Wherein, It is geocentric inertial coordinate system opposing body's coordinate It is pose transformation matrix, α is azimuth, and δ is pitch angle.

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

In formula (13), It is main astrology to the true of component unit direction vector Measured value,It is main astrology to the true measurement of component 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_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.

To sum up, present embodiment designs two spaces robot first using GEO target satellite as in-orbit service object (being set as primary and component) formation flight configuration and orbit parameter, it is then dynamic according to satellite relative orbit under geocentric inertial coordinate system Mechanical model establishes autonomous navigation system state model；Secondly the theory met needed for primary star sensor observation component is proposed Illumination condition and image-forming condition.Primary is calculated with respect to component theory azimuth and pitch angle, adjust true star sensor optical axis and Theory orientation is consistent, is really observed component, establishes using relative unit direction vector and distance as the observation side of observed quantity Journey；Unscented Kalman Filter Estimation primary relative position and speed are finally used, the invention belongs to aerospace navigation technology necks Domain not only can provide high-precision navigation information in GEO formation flight for satellite, but also can design for its autonomous navigation system Reference is provided.

Claims (10)

1. facing the space multi-robot self-determination air navigation aid of geostationary orbit object run, it is characterised in that: including following Step:

S1: using GEO target satellite as in-orbit service object, two spaces robot is set to primary and component, design master Star and component formation flight configuration and orbit parameter；Orbit parameter include 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: according to satellite relative motion dynamics model under geocentric inertial coordinate system, autonomous navigation system state model is established；

S3: primary and component relative distance are calculated according to step S1, judge whether component meets star sensor observed range requirement: If it is satisfied, then entering step S4；Otherwise, S12 is entered step；

S4: the positional relationship of the sun, the earth and component three is resolved according to step S1, judges whether component is in solar irradiation Area: if it is, entering step S5；Otherwise, S12 is entered step；

S5: resolving the positional relationship of the earth, primary and component 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 component can the apparent magnitude, judge component 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: primary is calculated according to step S1 and is directed toward angle with respect to component direction vector and star sensor optical axis, whether judges component In star sensor field range: if it is, entering step S8；Otherwise, it is directed toward by universal axial adjustment star sensor optical axis, Judge that component is to enter step S8 whether in star sensor field range again according to step S7, otherwise enters S12；

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

S9: theory orientation vector, azimuth and pitch angle of the primary with respect to component are calculated；

S10: adjustment star sensor true optical axis direction is consistent with theory orientation vector, is really observed component, and son is calculated Star 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. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: the process for establishing autonomous navigation system state model is as follows in the step S2:

Under geocentric inertial coordinate system, when primary positional distance is greater than component and primary relative distance, satellite phase is established To target component dynamics of orbits model:

Wherein, δ r⁽¹⁰⁾It is main astrology to component direction vector, δ v⁽¹⁰⁾It is main astrology to component velocity vector,For satellite position First derivative at any time is set,For the first derivative of satellite velocities at any time, r⁽⁰⁾For main star position vector, μ_eFor the earth Gravitational constant, a_fFor perturbation acceleration；

Definition status variable x=[(δ r⁽¹⁰⁾)^T(δv⁽¹⁰⁾)^T]^T, establish autonomous navigation system state model；

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

3. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: judging that the no process for meeting star sensor observed range requirement in component is as follows in the step S3: calculating primary Opposite component distance δ r⁽¹⁰⁾, judge δ r⁽¹⁰⁾Whether formula (3) shown in condition is met:

L_min≤δr⁽¹⁰⁾≤L_max (3)

Wherein, δ r⁽¹⁰⁾=| δ r⁽¹⁰⁾|=| r⁽¹⁾-r⁽⁰⁾|, r⁽⁰⁾For main star position vector, r⁽¹⁾For component position vector, L_minBetween star Minimum range needed for observing, L_maxMaximum distance needed for being observed between star.

4. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S4, judge component whether be in solar irradiation area process it is as follows: analysis earth shaded region And component travels through the critical condition of the shaded region, if component position vector r⁽¹⁾With position of sun vector r^(sun)Angle For ψ, the critical angle that component enters and leaves earth's shadow range is ψ_cri, then component is in solar irradiation area and needs to meet item Part:

ψ < ψ_cri (4)。

5. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S5, judge the earth whether enter star sensor visual field process it is as follows: set primary position vector r⁽⁰⁾With primary with respect to component direction vector δ r⁽¹⁰⁾Angle be θ, the critical condition for causing background light excessively weak is blocked by the earth It is primary with respect to component direction vector δ r⁽¹⁰⁾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. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S6, component can the apparent magnitude be calculated according to formula (6):

In formula (6), m be component can the apparent magnitude, M be component absolute magnitude, be calculated according to formula (7)；|r^(sun0)| be the sun with The distance between component；ξ is primary with respect to component direction vector δ r⁽¹⁰⁾Component direction vector r opposite with the sun^(sun1)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. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S7, judge whether process in star sensor field range is as follows for component: ifThen component is in star sensor field range；Otherwise, then component not in star sensor field range；Such as formula (9) Shown, FOV is star sensor field angle；

In formula (9), δ r⁽¹⁰⁾It is main astrology to component direction vector,For star sensor direction vector,It is that Earth central inertial is sat Mark system opposing body coordinate system pose transition matrix；

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

Wherein,It is main astrology to component direction vector δ r⁽¹⁰⁾It is projected in the coordinate of star sensor two dimension image planes battle array, IP_longthFor the length of two-dimentional image planes battle array, IP_widthFor the width of two-dimentional image planes battle array.

8. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S9, primary with respect to component theory orientation vector, azimuth and pitch angle respectively according to formula (11), (12a) and (12b) is calculated:

In formula (11), δ r⁽¹⁰⁾It is main astrology to the theory orientation vector of component, also referred to as primary is with respect to component direction vector；It is main astrology to component 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. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: in the step S10, the observational equation of foundation are as follows:

In formula (13), True measurement for main astrology to component unit direction vector Value,It is main astrology to the true measurement of component distance.

10. the space multi-robot self-determination air navigation aid according to claim 1 in face of geostationary orbit object run, It is characterized by: 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 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.