CN106153051B - A kind of spacecraft shading device combined navigation methods - Google Patents

Abstract

The present invention relates to a kind of spacecraft shading device combined navigation methods, the star based on PI filtering is quick/the autonomous celestial combined navigation of horizon instrument/inertia, solve the problems, such as due to add table exist drift cannot the location/velocity for a long time based on inertial navigation output calculate orbit parameter.The position and speed of drift correction inertial navigation output is directed toward based on the quick the earth's core determined with horizon instrument of star; inertia-celestial combined navigation can inhibit inertial navigation accumulated error; present invention GNSS compatible on star is abnormal; and without under the ground in real time abnormal conditions such as upper note orbit parameter; utilize the quick orbit parameter corresponding with horizon instrument inertia-location/velocity calculating of celestial combined navigation output of star, it can be ensured that based on the posture of the quick determining ontology relative orbit system of star.Compared with prior art, it is simple, effective to be algorithm for advantages and beneficial effects, and is easy to Project Realization.

Description

Spacecraft combined navigation method

Technical Field

The invention relates to a spacecraft integrated navigation method, in particular to a star sensor/horizon/inertia autonomous astronomical integrated navigation method based on PI filtering.

Background

Orbit parameters are needed for determining the posture of the body relative to an orbit system based on the satellite sensitivity, drift exists in the adding table, navigation accumulated errors are caused by inertial navigation integration, and the orbit parameters cannot be calculated based on the position/speed output by the inertial navigation for a long time.

The ground generally carries out inertial navigation solution based on a WGS-84 non-inertial system, and is not suitable for high-latitude areas; the on-orbit estimation is generally carried out based on the flat root of the orbit noted on the ground, and the on-orbit estimation cannot be used for a long time and is not suitable for the condition of orbit change.

The position and the speed of inertial navigation output are corrected based on geocentric pointing deviation determined by the star sensor and the horizon sensor, inertial-astronomical combined navigation is carried out based on a J2000 inertial system without limitation of latitude, and inertial navigation accumulated errors can be restrained.

Under the abnormal conditions that the on-satellite GNSS compatible machine is abnormal and no ground real-time orbit parameters are injected, the corresponding orbit parameters are calculated by utilizing the position/speed output by the inertia-astronomical combined navigation of the satellite sensitivity and the horizon, and the attitude of the body relative to the orbit system is ensured to be determined based on the satellite sensitivity.

Disclosure of Invention

In order to solve the problems in the prior art, the invention aims to provide a spacecraft combination navigation method, which is used for realizing star sensitivity/horizon/inertia autonomous astronomical combination navigation based on PI filtering, estimating position/speed error correction quantity based on star sensitivity and horizon geocentric pointing deviation, and determining orbit parameters based on inertia-astronomical combination navigation.

In order to achieve the above object, the technical solution of the present invention is to provide a spacecraft combination navigation method, wherein:

calculating an inertial navigation position error based on the star sensor and the horizon instrument geocentric pointing deviation, and estimating a position/speed error correction amount based on PI filtering;

performing inertial navigation calculation based on a J2000 inertial system, and performing inertial-astronomical combined navigation based on position/speed error correction;

orbit parameters are calculated based on the position/velocity of the inertial-astronomical combined navigation output, and the attitude of the star relative to the orbital system is determined based on the star sensitivity.

Compared with the prior art, the star sensor/horizon sensor/inertia autonomous astronomical combined navigation method based on PI filtering has the advantages and beneficial effects that:

the invention deduces the inertia-astronomical combined navigation based on the star sensor and horizon geocentric pointing deviation and PI filtering estimation position/speed error correction quantity, and the algorithm is simple and effective and is easy to realize in engineering.

The invention carries out inertia-astronomical combined navigation based on the J2000 inertial system, obtains the position and the speed relative to the J2000 inertial system, converts the position and the speed into orbit parameters required by star sensitive attitude calculation, and obtains the attitude of the star sensitive relative to the orbit system.

Under the abnormal conditions that the on-satellite GNSS compatible machine is abnormal and no ground real-time orbit parameters are annotated, the invention utilizes the position/speed output by the inertia-astronomical combined navigation of the satellite sensitivity and the horizon to calculate the corresponding orbit parameters, and can ensure that the attitude of the body relative to the orbit system is determined based on the satellite sensitivity.

Drawings

Fig. 1 is a schematic diagram of a spacecraft integrated navigation method according to the invention.

Detailed Description

As shown in fig. 1, the present invention provides a spacecraft integrated navigation method, which implements a satellite sensitive/horizon/inertial autonomous astronomical integrated navigation based on PI filtering, estimates a position/speed error correction amount based on a satellite sensitive and horizon geocentric pointing deviation, and determines an orbit parameter based on the inertial-astronomical integrated navigation, and specifically includes:

calculating an inertial navigation position error based on the star sensor and the horizon instrument geocentric pointing deviation, and estimating a position/speed error correction amount based on PI filtering;

performing inertial navigation calculation based on a J2000 inertial system, and performing inertial-astronomical combined navigation based on position/speed error correction;

orbit parameters are calculated based on the position/velocity of the inertial-astronomical combined navigation output, and the attitude of the star relative to the orbital system is determined based on the star sensitivity.

The calculation process of the spacecraft combination navigation method comprises the following steps:

s1 inertial navigation solution based on J2000 inertial system

Let the position vector of the satellite under J2000 inertia be r ═ x y z]^TThen, the three-axis component of the gravitational acceleration of the satellite under the J2000 inertial system is as follows:

wherein:is the distance from the satellite to the Earth's center, μ is the Earth's gravitational constant, R_eIs the radius of the earth, J₂Is the global non-spherical perturbation coefficient.

Setting the attitude transformation matrix from the adding table coordinate system to the J2000 inertial system as A_iaTo find the J2000 inertiaNon-inertial acceleration a of satellite under system_a,iThe following were used:

a_a,i＝A_ia·a_a,a

wherein: a is_a,aThe acceleration measured by the accelerometer under the accelerometer coordinate system is used.

Let a_g,i＝[a_gx,i a_gy,i a_gz,i]^TThe acceleration a of the satellite under the J2000 inertial system is obtained_iThe following were used:

a_i＝a_g,i+a_a,i

performing inertial navigation calculation in a J2000 inertial system to obtain a position r_iVelocity v_iThe following were used:

v_i＝∫a_i·dt

r_i＝∫v_i·dt。

s2, calculating q from position/speed output by inertial navigation_oi

Let position r under J2000 inertial system_i＝[x y z]^TSpeed, velocityFrom this, a position/velocity scalar r/v and a track plane normal vector are obtainedThe following were used:

order toAnd (3) solving the ascension angle omega of the ascending intersection point, the track inclination angle i and the latitude argument u as follows:

i＝cos^-1(σ₃/σ)

wherein: k- σ₁/σ，h＝-σ₂/σ，Quadrant discrimination is required in solving omega, i and u.

Obtaining J2000 quaternion q from inertia system to orbital system attitude from omega, i and u_oiThe following were used:

wherein:

s3, calculating the earth center pointing deviation angle based on star sensor and horizon

Setting the star sensitivity, considering the exposure time difference and data acquisition delay correction, and deducting the installation matrix to obtain a quaternion q of the attitude of the star relative to the J2000 inertial system_biCombining the quaternion q of the attitude of the orbital relative to the J2000 inertial system_oiObtaining the quaternion q of the attitude of the body relative to the orbital system_boThe following were used:

let q be_bo＝[q₀ q₁ q₂ q₃]^TThe transformation matrix expressed in quaternion form is as follows:

the transformation matrix expressed in Euler angle 3-1-2 rotation order is as follows:

wherein:is the roll attitude angle, theta is the pitch attitude angle, and psi is the yaw attitude angle.

The attitude of the star relative to the orbital system is obtained as follows:

the earth-centered pointing deviation angle can be approximately solved as follows:

dθ＝θ-θ_H

wherein,roll angle, theta, output for horizon finder_HThe pitch angle output by the horizon sensor.

S4 estimating position error based on geocentric pointing error angle

Pointing the roll deviation angle from the centroid of the opposite orbital systemAnd (3) solving the position errors in the X direction and the Y direction under the track system by the aid of the pitching deviation angle d theta and the ground center distance r output by inertial navigation:

dx＝r·dθ

considering the coupling relationship of dz and dx, dz is estimated using the ratio based on dx as follows:

dz＝k_pxz·dx

wherein: k is a radical of_pxzAnd (4) scaling the estimation coefficient.

Let the rail under position error Deltar_o＝[dx dy dz]^TConversion to J2000 position error Δ r under inertial system_iThe following were used:

Δr_i＝A_oi ^T·Δr_o

wherein: a. the_oiFor the inertial system to orbital system transformation matrix, from q_oiAnd (6) obtaining.

S5 estimating position/speed error correction amount based on PI filtering

Let the k-th step position error under the J2000 inertial system estimated based on the star sensor and the horizon instrument be delta r_i,kLet is given by_i,kPosition/velocity error after clipping ofK step inertial navigation solution position error correction dr_i,kThe following were used:

wherein:

k_p,restimating a scaling factor, k, for the position error correction_p,rFor a 3 x 3 diagonal array, the position error correction can be estimated independently.

Let dr_i,k＝[dx_i,k dy_i,k dz_i,k]^TAndthe triaxial independent estimation position error correction is in the form of:

wherein:

k_p,x、k_p,y、k_p,zscaling factors are estimated for the three axis position error correction.

The correction amount dr is based on the position error in consideration of the fact that the position error also reflects the speed error_ikA speed error correction is estimated. Let pair dr_i,kThe position error correction amount after the amplitude limiting isK-th step inertial navigation resolving speed error correction dv_i,kThe following were used:

wherein:

k_p,vestimating a scaling factor, k, for a speed error correction_p,vFor a 3 x 3 diagonal matrix, the speed error correction can be estimated independently.

Let dv_i,k＝[dvx_i,k dvy_i,k dvz_i,k]^TAndthe triaxial independent estimation speed error correction is in the form as follows:

wherein:

k_p,dx、k_p,dy、k_p,dza scaling factor is estimated for the three axis velocity error correction.

S6, position/speed error correction based integrated navigation

The position/speed error correction amount is gradually increased in the inertial navigation integration process, so that the stability of error correction can be ensured. And (3) deducting the position/speed error correction quantity in the J2000 inertial system, and applying a simplified integral algorithm to carry out inertial navigation solution as follows:

v_i,k＝v_i,k-1+[a_i,k-1+(a_i,k-a_i,k-1)/2]·T-dv_i,k-1

r_i,k＝r_i,k-1+[v_i,k-1+(v_i,k-v_i,k-1)/2]·T-dr_i,k-1

wherein:

a_i,k-1acceleration of the k-1 step;

a_i,kacceleration of the kth step;

dv_i,k-1the speed error correction amount estimated for the k-1 step;

v_i,k-1deducting the speed of the speed error correction quantity for the k-1 step;

v_i,kdeducting the speed of the speed error correction quantity for the k step;

dr_i,k-1the position error correction amount estimated for the (k-1) th step;

r_i,k-1deducting the position of the position error correction quantity for the k-1 step;

r_i,kdeducting the position of the position error correction quantity for the k step;

t is the navigation period.

S7, determining the attitude of the star relative to the orbital system based on the inertia-astronomical combined navigation:

calculating corresponding orbit parameters based on the position/speed of the inertia-astronomical combined navigation output, calculating attitude algorithms of the stars relative to the orbit system based on the orbit parameters, and calculating attitude angle correlation algorithms with reference to the star sensor. If the position and the velocity under the J2000 inertial system in S2 are replaced by the values of the velocity and the position after the error correction is subtracted in S6, the attitude quaternion of the star body relative to the orbital system can be calculated through the processes of S2 and S3.

In conclusion, the accelerometer acceleration drift is estimated based on PI filtering, the algorithm is simple and effective, and the engineering implementation is easy; the inertial navigation solution is not limited by latitude degree based on the J2000 inertial system, and can be used in all day regions; the star sensor and horizon instrument inertia-astronomical combined navigation based on PI filtering can be used for a long time and is still suitable for orbital transfer.

While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.

Claims (7)

1. A spacecraft combination navigation method is characterized by comprising the following processes:

calculating an inertial navigation position error based on the star sensor and the horizon instrument geocentric pointing deviation, and estimating a position/speed error correction amount based on PI filtering;

performing inertial navigation calculation based on a J2000 inertial system, and performing inertial-astronomical combined navigation based on position/speed error correction;

calculating orbit parameters based on the position/speed output by the inertia-astronomical combined navigation, and determining the attitude of a star body relative to an orbit system based on the star sensor;

correcting the star sensor according to the exposure time difference and the data acquisition delay, and deducting the installation matrix to obtain a quaternion q of the attitude of the star relative to the J2000 inertial system_biCombining J2000 inertia system to orbital system attitude quaternion q_oiObtaining the quaternion q of the attitude of the body relative to the orbital system_boThe following were used:

let q be_bo＝[q₀ q₁ q₂ q₃]^TThe transformation matrix expressed in quaternion form is as follows:

the transformation matrix expressed in Euler angle 3-1-2 rotation order is as follows:

wherein:is a rolling attitude angle, theta is a pitching attitude angle, psi is a yawing attitude angle;

the attitude of the star relative to the orbital system is obtained as follows:

approximately resolving the earth center pointing deviation angle to obtain the earth center pointing rolling deviation angle of the relative track systemThe pitch deviation angle d θ is as follows:

dθ＝θ-θ_H

wherein,roll angle, theta, output for horizon finder_HThe pitch angle output by the horizon sensor.

2. A spacecraft integrated navigation method according to claim 1, wherein the process S1 of performing inertial navigation solution based on J2000 inertial system includes:

let the position r of the satellite under the inertia of J2000_i＝[x y z]^TThen, the three-axis component of the gravitational acceleration of the satellite under the J2000 inertial system is as follows:

wherein: position scalarMu is the gravitational constant, R, corresponding to the distance of the satellite to the Earth's center_eIs the radius of the earth, J₂Is the global non-spherical perturbation coefficient;

setting the attitude transformation matrix from the adding table coordinate system to the J2000 inertial system as A_iaTo obtain the non-inertial acceleration a of the satellite under the J2000 inertial system_a,iThe following were used:

a_a,i＝A_ia·a_a,a

wherein: a is_a,aAcceleration under a summator coordinate system measured by a summator;

let a_g,i＝[a_gx,i a_gy,i a_gz,i]^TThe acceleration a of the satellite under the J2000 inertial system is obtained_iThe following were used:

a_i＝a_g,i+a_a,i

performing inertial navigation calculation in a J2000 inertial system to obtain a position r_iVelocity v_iThe following were used:

v_i＝∫a_i·dt

r_i＝∫v_i·dt。

3. the spacecraft integrated navigation method of claim 2, further comprising determining the J2000 inertial system to orbital system attitude quaternion q from the position/velocity of the inertial navigation output_oiProcedure S2:

let position r under J2000 inertial system_i＝[x y z]^TSpeed, velocityDetermining a position scalar r, a velocity scalar v and a track plane normal vectorThe following were used:

order toObtaining the ascension angle omega of the ascending intersection point, the track inclination angle i and the latitude argument u:

i＝cos^-1(σ₃/σ)

wherein: k- σ₁/σ，h＝-σ₂/σ，

Obtaining quaternion q from J2000 inertia system to orbit system attitude_oiThe following were used:

wherein:

4. a spacecraft integrated navigation method according to claim 3, wherein the process S4 of estimating the position error based on the earth-centered pointing error angle comprises:

pointing the roll deviation angle from the centroid of the opposite orbital systemThe pitch deviation angle d theta and the distance from the satellite to the earth center, namely the position mark r, which are output by inertial navigation resolving, are used for solving the position errors of the X direction and the Y direction under the orbit system as follows:

dx＝r·dθ

by using the coupling relationship of dz and dx, the ratio of the Z-position error dz based on dx is estimated as follows: k is dz ═ k_pxz·dx；

Wherein: k is a radical of_pxzA scale estimation coefficient;

let the rail under position error Deltar_o＝[dx dy dz]^TConversion to J2000 position error Δ r under inertial system_iThe following were used:

Δr_i＝A_oi ^T·Δr_o

wherein: a. the_oiA quaternion q from J2000 inertial system to orbital system attitude for the inertial system to orbital system transformation matrix_oiAnd (6) obtaining.

5. A spacecraft integrated navigation method according to claim 4, wherein the process S5 of estimating the position/velocity error modifier based on PI filtering comprises:

let the k-th step position error under the J2000 inertial system estimated based on the star sensor and the horizon instrument be delta r_i,kLet is given by_i,kPosition/velocity error after clipping ofK step inertial navigation solution position error correction dr_i,kThe following were used:

wherein: k is a radical of_p,rEstimating a scaling factor for the position error correction;

let dr_i,k＝[dx_i,k dy_i,k dz_i,k]^TAndthe triaxial independent estimation position error correction is in the form of:

wherein: k is a radical of_p,x、k_p,y、k_p,zScaling factors are estimated for the three axis position error correction.

6. A spacecraft integrated navigation method according to claim 5, wherein the process S5 of estimating the position/velocity error correction amount based on PI filtering further includes estimating the position/velocity error correction amount based on the position error correction amount dr_i,kProcess of estimating speed error correction amount:

let pair dr_i,kThe position error correction amount after the amplitude limiting isK-th step inertial navigation resolving speed error correction dv_i,kThe following were used:

wherein: k is a radical of_p,vEstimating a scaling factor for the speed error correction;

let dv_i,k＝[dvx_i,k dvy_i,k dvz_i,k]^TAndthe triaxial independent estimation speed error correction is in the form as follows:

wherein: k is a radical of_p,dx、k_p,dy、k_p,dzA scaling factor is estimated for the three axis velocity error correction.

7. A spacecraft integrated navigation method according to claim 6, wherein the integrated navigation process based on position/velocity error correction amount S6 includes:

and (3) deducting the position/speed error correction quantity in the J2000 inertial system, and applying a simplified integral algorithm to carry out inertial navigation solution as follows:

v_i,k＝v_i,k-1+[a_i,k-1+(a_i,k-a_i,k-1)/2]·T-dv_i,k-1

r_i,k＝r_i,k-1+[v_i,k-1+(v_i,k-v_i,k-1)/2]·T-dr_i,k-1

wherein:

a_i,k-1acceleration of the k-1 step;

a_i,kacceleration of the kth step;

dv_i,k-1the speed error correction amount estimated for the k-1 step;

v_i,k-1deducting the speed of the speed error correction quantity for the k-1 step;

v_i,kdeducting the speed of the speed error correction quantity for the k step;

dr_i,k-1the position error correction amount estimated for the (k-1) th step;

r_i,k-1deducting the position of the position error correction quantity for the k-1 step;

r_i,kdeducting the position of the position error correction quantity for the k step;

t is the navigation period.