CN105444762A - Rapid inertial navigation error correction method for onboard satellite communication in motion - Google Patents

Abstract

The invention discloses a rapid inertial navigation error correction method for onboard satellite communication in motion. The rapid inertial navigation error correction method is a method for correcting MEMS inertial navigation information of antennas for onboard satellite communication in motion by virtue of onboard high-precision navigation information. Onboard high-precision navigation data is broadcasted by an onboard bus, is relatively high in precision but is long in time interval, so that the onboard high-precision navigation data cannot meet control requirements of the antennas for the satellite communication in motion. MEMS inertial navigation data is relatively high in velocity and but is relatively low in precision, so that the MEMS inertial navigation data cannot meet control requirements of the satellite communication in motion. According to the method, MEMS inertial navigation errors are corrected at regular intervals by virtue of the onboard inertial navigation information, so that the requirements of the satellite communication in motion on the control precision and the time interval of attitude data can be simultaneously met, and meanwhile, the rapid calibration of the gyro zero position in the MEMS inertial navigation is realized.

Description

A kind of ins error rapid correction method for airborne communication in moving

Technical field

The invention belongs to navigation field, relate to a kind of modification method of ins error of airborne communication in moving.

Background technology

The mobile communication application of synchronous satellite is commonly called as " communication in moving ", is that the present satellites communications field is in great demand, development is applied rapidly.Except " communication in moving " advantage that satellite communication overlay area extensively, not limits by landform region except having, transmission line is reliable and stable, really achieve the object of broadband, mobile communication.

Aboard, be equipped with airborne ins and communication in moving, communication in moving self comprises again a MEMS inertial navigation and controls for the sensing of antenna simultaneously.For two kinds of inertial navigations, airborne inertial navigation precision is high but attitude information upgrades slow, but and the low posture renewal of MEMS inertial navigation precision is fast.Only carry out Antenna pointing control with MEMS inertial navigation, because MEMS inertial navigation precision is poor, cannot the long-time high-precision attitude stabilization of complete independently, must constantly revise its navigation error by external auxiliary information.Only control antenna for satellite communication in motion with airborne inertial navigation to point to, although pointing accuracy increases, due to the renewal rate that airborne ins is slower, antenna cannot complete the dynamic tracking of wide-angle.

Summary of the invention

The technical matters that the present invention solves is: overcome the deficiencies in the prior art, provide a kind of method of MEMS inertial navigation navigation error of airborne ins navigation information correction communication in moving, MEMS inertial navigation and airborne ins advantage are separately taken into account, learn from other's strong points to offset one's weaknesses, preferably resolve the low and problem of MEMS inertial navigation low precision of airborne ins data rate, MEMS gyro zero can be estimated fast inclined, and revise the attitude error of MEMS inertial navigation, make MEMS inertial navigation can meet airborne communication in moving control overflow for a long time, be more applicable to controlling accurately antenna direction.

Technical solution of the present invention is: a kind of ins error rapid correction method for airborne communication in moving, comprises the steps:

(1) when airborne ins navigation information is effective, obtain the navigation information of airborne ins, comprise the position angle ψ of carrier _p, pitching angle theta _pwith roll angle γ _p, according to described position angle ψ _p, pitching angle theta _pwith roll angle γ _pobtain the attitude matrix of airborne ins with attitude quaternion Q ₀;

(2) by the attitude matrix of airborne ins with attitude quaternion Q ₀under being transformed into MEMS inertial navigation body series by airborne ins body series, obtain the attitude matrix after converting with attitude quaternion Q ₁; The initial point of described airborne ins body series is positioned at the barycenter of carrier, and X-axis, Y-axis, Z axis point to the right of carrier, front and top respectively, and described MEMS inertial navigation body coordinate system is the coordinate system that three sensitive axes of MEMS inertial navigation are formed;

(3) angular speed of three-axis gyroscope output is obtained from MEMS inertial navigation from carrier avionics system acquisition number according to propagation time delay Δ t, obtain rotating vector thus utilize rotating vector R to attitude quaternion Q ₁carry out posture renewal, obtain the attitude quaternion Q after upgrading ₄;

(4) with the attitude quaternion Q after upgrading ₄as the navigation hypercomplex number Q of MEMS inertial navigation ₂with sky line traffic control hypercomplex number Q ₃initial value and bind;

(5) again perform step (1) ~ (3), obtain the attitude quaternion Q after upgrading ₄, and and Q ₄corresponding attitude matrix with attitude matrix corresponding carrier position angle ψ _m, pitching angle theta _mwith roll angle γ _m, use the attitude quaternion Q after again upgrading ₄replace the navigation hypercomplex number Q of MEMS inertial navigation ₂; The three axle gyro output quantities simultaneously collected according to MEMS inertial navigation carry out attitude algorithm, obtain the position angle ψ that MEMS inertial navigation navigation hypercomplex number attitude is corresponding _n, pitching angle theta _nwith roll angle γ _n, thus obtain the azimuthal error Δ ψ=ψ of MEMS inertial navigation _n-ψ _m, pitch error Δ θ=θ _n-θ _mwith roll error delta γ=γ _n-γ _m;

(6) according to described azimuthal error Δ ψ, pitch error Δ θ and roll error delta γ, and the propagation time delay Δ t of airborne ins, the gyro zero obtaining MEMS inertial navigation is worth ε partially _x=Δ θ/Δ t, ε _y=Δ γ/Δ t, ε _z=Δ ψ/Δ t, utilizes least squares filtering to estimate ε _x, ε _y, ε _zestimated value l _x, l _y, l _z, the MEMS inertial navigation three axle gyro output quantity after being compensated thus ${\left[\begin{array}{ccc}{\mathrm{\ω}}_{ib}^{bx}& {\mathrm{\ω}}_{ib}^{by}& {\mathrm{\ω}}_{ib}^{bz}\end{array}\right]}^T,{\mathrm{\ω}}_{ib}^{bx}={\mathrm{\ω}}_{0ib}^{bx}-l_x,$ ${\mathrm{\ω}}_{ib}^{by}={\mathrm{\ω}}_{0ib}^{by}-l_y,{\mathrm{\ω}}_{ib}^{bz}={\mathrm{\ω}}_{0ib}^{bz}-l_z;$

(7) by up-to-date sky line traffic control hypercomplex number Q ₃navigation hypercomplex number Q after upgrading with step (5) ₂corresponding attitude angle is compared, and produces the instruction angular speed [ω of sky line traffic control hypercomplex number _0xω _0yω _0z] ^t, and obtain with step (6) ${\left[\begin{array}{ccc}{\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bx}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{by}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bz}}\end{array}\right]}^T$ Be added and obtain final instruction angular speed to sky line traffic control hypercomplex number Q ₃upgrade;

(8) step (5) ~ (7) are repeated, to navigation hypercomplex number Q ₂with sky line traffic control hypercomplex number Q ₃carry out continuous updating.

The present invention's advantage is compared with prior art:

(1), in the inventive method, adopt the MEMS inertial navigation of low cost as communication in moving control module, solve airborne ins data rate lower, the problem of communication in moving realtime control can not be met;

(2), in the inventive method, adopt the navigation error of the navigation information correction MEMS inertial navigation of airborne ins, ensure that the long-time navigation accuracy of MEMS inertial navigation, also ensure that the control accuracy of MEMS inertial navigation to antenna simultaneously;

(3) in the inventive method, adopt the navigation error of the navigation information correction MEMS inertial navigation of airborne ins, give a kind of method that MEMS gyro zero is inclined of estimating, using the inclined feedback quantity as control system of gyro zero estimated, effectively reduce attitude drift.

Accompanying drawing explanation

Fig. 1 is the FB(flow block) of the inventive method.

Embodiment

As shown in Figure 1, be the FB(flow block) of the inventive method.The inventive method adopts the navigation error of the precise navigation information correction MEMS inertial navigation of airborne ins.

The key step of the inventive method is as follows:

(1) when airborne ins navigation information is effective, obtain the navigation information of airborne ins, comprise the position angle ψ of carrier _p, pitching angle theta _p, roll angle γ _p, according to position angle ψ _p, pitching angle theta _p, roll angle γ _pobtain the attitude matrix of airborne ins attitude quaternion Q ₀.

The navigation information significant instant of airborne ins here, refers to the moment that the frame navigation attitude data transmission being delivered to communication in moving system by airborne-bus completes.

How by position angle ψ _p, pitching angle theta _p, roll angle γ _pobtain the attitude matrix of airborne ins attitude quaternion Q ₀, specifically can see " inertial navigation " (Science Press, Qin Yongyuan writes, in May, 2006 first published) book.

The general expression of attitude of carrier matrix is:

$C_b^T={\left[\begin{array}{ccc}\cos \left(\mathrm{\γ}\right)\cos \left(\mathrm{\ψ}\right)+\sin \left(\mathrm{\γ}\right)\sin \left(\mathrm{\ψ}\right)\sin \left(\mathrm{\θ}\right)& -\cos \left(\mathrm{\γ}\right)\sin \left(\mathrm{\ψ}\right)+\sin \left(\mathrm{\γ}\right)\cos \left(\mathrm{\ψ}\right)\sin \left(\mathrm{\θ}\right)& -\sin \left(\mathrm{\γ}\right)\cos \left(\mathrm{\θ}\right)\\ \sin \left(\mathrm{\ψ}\right)\cos \left(\mathrm{\θ}\right)& \cos \left(\mathrm{\ψ}\right)\cos \left(\mathrm{\θ}\right)& \sin \left(\mathrm{\θ}\right)\\ \sin \left(\mathrm{\γ}\right)\cos \left(\mathrm{\ψ}\right)-\cos \left(\mathrm{\γ}\right)\sin \left(\mathrm{\ψ}\right)\sin \left(\mathrm{\θ}\right)& -\sin \left(\mathrm{\γ}\right)\sin \left(\mathrm{\ψ}\right)-\cos \left(\mathrm{\γ}\right)\cos \left(\mathrm{\ψ}\right)\sin \left(\mathrm{\θ}\right)& \cos \left(\mathrm{\γ}\right)\cos \left(\mathrm{\θ}\right)\end{array}\right]}^T$

Wherein ψ, θ, γ are respectively position angle, the angle of pitch, roll angle.

Attitude of carrier hypercomplex number expression formula is:

$\left\{\begin{array}{c}\left|q_0\right|=0.5\·\sqrt{1+C_T^b(1,1)+C_T^b(2,2)+C_T^b(3,3)}\\ \left|q_1\right|=0.5\·\sqrt{1+C_T^b(1,1)-C_T^b(2,2)-C_T^b(3,3)}\\ \left|q_2\right|=0.5\·\sqrt{1-C_T^b(1,1)+C_T^b(2,2)-C_T^b(3,3)}\\ \left|q_3\right|=0.5\·\sqrt{1-C_T^b(1,1)-C_T^b(2,2)+C_T^b(3,3)}\end{array}\right.$

To q ₀, q ₁, q ₂, q ₃symbol, can be defined as by following formula

$\left\{\begin{array}{c}sign\left(q_0\right)=1\\ sign\left(q_1\right)=sign(C_T^b\left(3,2\right)-C_T^b\left(2,3\right))\\ sign\left(q_2\right)=sign(C_T^b\left(1,3\right)-C_T^b\left(3,1\right))\\ sign\left(q_3\right)=sign(C_T^b\left(2,1\right)-C_T^b\left(1,2\right))\end{array}\right.$

In formula for transposition.

(2) by the attitude matrix of airborne ins attitude quaternion Q ₀under being transformed into MEMS inertial navigation body series by airborne ins body series, obtain the attitude matrix after converting with attitude quaternion Q ₁.

Airborne ins body series is the frame of reference of airborne ins, and itself and aircraft carrier coordinate system (round dot at barycenter, X-axis, Y-axis, Z axis point to the right of aircraft, front and top respectively) overlap.MEMS inertial navigation body coordinate system is the frame of reference of MEMS inertial navigation, and three coordinate axis of this coordinate system are three sensitive axes of MEMS inertial navigation, and this coordinate system is determined when dispatching from the factory.

By attitude matrix the concrete grammar be transformed under MEMS inertial navigation body series is

$C{1}_n^m=C_b^m\·C_n^b$

Wherein be the transition matrix between airborne ins body coordinate system and MEMS inertial navigation body coordinate system, by aircraft general assembly, department provides.

By attitude matrix obtain attitude quaternion Q ₁method specifically can see " inertial navigation " (Science Press, Qin Yongyuan writes, in May, 2006 first published) book.

3) angular speed of each axle gyroscope output is obtained from MEMS inertial navigation from Aircraft electric system acquisition number according to propagation time delay Δ t, obtain rotating vector thus utilize rotating vector R to attitude quaternion Q ₁carry out posture renewal, obtain the attitude quaternion Q after upgrading ₄.

the form being is ${\mathrm{\ω}}_{\mathrm{iT}}^T={\left[\begin{array}{ccc}{\mathrm{\ω}}_{\mathrm{iT}}^T\left(x\right)& {\mathrm{\ω}}_{\mathrm{iT}}^T\left(y\right)& {\mathrm{\ω}}_{\mathrm{iT}}^T\left(z\right)\end{array}\right]}^T.$ The attitude motion of carrier in the time period of delay is approximately the rigid body Free-rolling without External Force Acting in inertial space, and the angular speed that each axle gyroscope of MEMS inertial navigation exports is approximately steady state value the three axle angular velocity of rotations of the MEMS inertial navigation body coordinate system that records of MEMS inertial navigation relative to inertial system.

The time delay Δ t of transmission is the parameter that Aircraft electric system provides, and is the amount of knowing, is provided by aircraft design unit.The time delay Δ t of transmission characterizes navigation information and is delivered to time needed for communication in moving system from airborne ins by airborne databus.

Rotating vector is the vector describing finite rotation of rigid body, and in the present invention, the rotating vector obtained with angle increment approximate treatment is

Utilize rotating vector R to attitude quaternion Q ₁carry out posture renewal, obtain the attitude quaternion Q after upgrading ₄method specifically can see " inertial navigation " (Science Press, Qin Yongyuan writes, in May, 2006 first published) book.

(4) with the attitude quaternion Q after upgrading ₄as the navigation hypercomplex number Q of MEMS inertial navigation ₂with sky line traffic control hypercomplex number Q ₃initial value, even also Q ₂=Q ₃=Q ₄.

(5) airborne ins navigation information again effectively time, again perform step (1) (2) (3), obtain the attitude quaternion Q after upgrading ₄, and and Q ₄corresponding attitude matrix with attitude matrix corresponding carrier position angle ψ _m, pitching angle theta _mwith roll angle γ _m, use the attitude quaternion Q after again upgrading ₄directly replace the navigation hypercomplex number Q of MEMS inertial navigation ₂.Due to navigation hypercomplex number Q ₂be only the navigation attitude value of MEMS system, do not participate in the control of antenna directly, so in order to allow Q ₂reach precision high as far as possible, use the attitude quaternion Q after again upgrading ₄directly substituted for the navigation hypercomplex number Q of MEMS inertial navigation ₂.

The three axle gyro output quantities simultaneously collected according to MEMS inertial navigation carry out attitude algorithm, obtain the position angle ψ that MEMS inertial navigation navigation hypercomplex number is corresponding _n, pitching angle theta _nwith roll angle γ _n, thus obtain azimuthal error Δ ψ, pitch error Δ θ and the roll error delta γ of MEMS inertial navigation, wherein

Δψ＝ψ _n-ψ _m，Δγ＝γ _n-γ _m，Δθ＝θ _n-θ _m

How to carry out attitude algorithm according to three axle gyro output quantities, obtain the position angle ψ that MEMS inertial navigation navigation hypercomplex number is corresponding _n, pitching angle theta _nwith roll angle γ _n, specifically can see " inertial navigation " (Science Press, Qin Yongyuan writes, in May, 2006 first published) book.

(6) according to azimuthal error Δ ψ, pitch error Δ θ and the roll error delta γ of MEMS inertial navigation, and the propagation time delay Δ t of airborne ins, because gyro zero characterizes the zero drift of gyro partially, be reacted in reality be exactly the drift value of gyro in section at any time divided by this time span, the gyro zero obtaining MEMS inertial navigation partially value is

ε _x＝Δθ/Δt，ε _y＝Δγ/Δt，ε _z＝Δψ/Δt

Then ε is estimated by least squares filtering _x, ε _y, ε _zestimated value l _x, l _y, l _z.

The measurement equation of least squares filtering in, the observed quantity that each filtering calculates $\stackrel{\‾}{Z}={\left[\begin{array}{ccc}{l}_x& l_y& l_z\end{array}\right]}^T,$ X=[ε _xε _yε _z] ^tε _x, ε _y, ε _zestimated value, get three rank unit matrix, it is error in measurement matrix.With the increase of iterations k, X=[ε _xε _yε _z] ^titerative estimate result be:

$X_{k+1}=X_k+\frac{1}{2}\·(Z_k-X_k)$

MEMS inertial navigation three axle gyro output quantity after being compensated thus ${\left[\begin{array}{ccc}{\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bx}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{by}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bz}}\end{array}\right]}^T,$

$x_{ib}^{bx}={\mathrm{\ω}}_{0ib}^{bx}-{\mathrm{\ϵ}}_x,{\mathrm{\ω}}_{ib}^{by}={\mathrm{\ω}}_{0ib}^{by}-{\mathrm{\ϵ}}_y,{\mathrm{\ω}}_{ib}^{bz}={\mathrm{\ω}}_{0ib}^{bz}-{\mathrm{\ϵ}}_z.$

(7) by up-to-date sky line traffic control hypercomplex number Q ₃with the navigation hypercomplex number Q that step (5) obtains ₂corresponding attitude angle is compared, and produces sky line traffic control hypercomplex number instruction angular speed [ω _0xω _0yω _0z] ^t, and obtain with step (6) ${\left[\begin{array}{ccc}{\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bx}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{by}}& {\mathrm{\ω}}_{\mathrm{ib}}^{\mathrm{bz}}\end{array}\right]}^T$ The angular velocity that addition obtains is for upgrading a day line traffic control hypercomplex number Q ₃.

How by sky line traffic control hypercomplex number Q ₃with navigation hypercomplex number Q ₂the corresponding attitude angle method producing sky line traffic control hypercomplex number instruction angular speed of comparing is 201410265808.2 see application number, the patent that name is called " biquaternion antenna for satellite communication in motion control method and system based on MEMS inertial navigation ".Adopt described patented method to upgrade Q ₃, significantly can reduce the noise in the antenna attitude information of day line traffic control hypercomplex number.

(8), when airborne ins navigation information is effective each time, repeated execution of steps (5) ~ (7), can to navigation hypercomplex number Q ₂with sky line traffic control hypercomplex number Q ₃carry out continuous updating.

The navigation error of MEMS inertial navigation is the error partially brought of MEMS inertial navigation gyro zero mainly, in order to revise the error that MEMS inertial navigation gyro zero brings partially, according to navigation hypercomplex number Q ₂with sky line traffic control hypercomplex number Q ₃merge and resolve iteration, not only have modified the navigation error of MEMS inertial navigation, also estimate except zero of MEMS gyro is inclined, while making MEMS inertial navigation reach higher navigation accuracy, achieve the level and smooth of antenna for satellite communication in motion and accurately control.

The content be not described in detail in instructions of the present invention belongs to the known technology of those skilled in the art.

Claims (1)

1., for an ins error rapid correction method for airborne communication in moving, it is characterized in that comprising the steps:

(1) when airborne ins navigation information is effective, obtain the navigation information of airborne ins, comprise the position angle ψ of carrier _p, pitching angle theta _pwith roll angle γ _p, according to described position angle ψ _p, pitching angle theta _pwith roll angle γ _pobtain the attitude matrix of airborne ins with attitude quaternion Q ₀;

(2) by the attitude matrix of airborne ins with attitude quaternion Q ₀under being transformed into MEMS inertial navigation body series by airborne ins body series, obtain the attitude matrix after converting with attitude quaternion Q ₁; The initial point of described airborne ins body series is positioned at the barycenter of carrier, and X-axis, Y-axis, Z axis point to the right of carrier, front and top respectively, and described MEMS inertial navigation body coordinate system is the coordinate system that three sensitive axes of MEMS inertial navigation are formed;

(3) angular speed of three-axis gyroscope output is obtained from MEMS inertial navigation from carrier avionics system acquisition number according to propagation time delay Δ t, obtain rotating vector thus utilize rotating vector R to attitude quaternion Q ₁carry out posture renewal, obtain the attitude quaternion Q after upgrading ₄;

(4) with the attitude quaternion Q after upgrading ₄as the navigation hypercomplex number Q of MEMS inertial navigation ₂with sky line traffic control hypercomplex number Q ₃initial value and bind;

(5) again perform step (1) ~ (3), obtain the attitude quaternion Q after upgrading ₄, and and Q ₄corresponding attitude matrix with attitude matrix corresponding carrier position angle ψ _m, pitching angle theta _mwith roll angle γ _m, use the attitude quaternion Q after again upgrading ₄replace the navigation hypercomplex number Q of MEMS inertial navigation ₂; The three axle gyro output quantities simultaneously collected according to MEMS inertial navigation carry out attitude algorithm, obtain the position angle ψ that MEMS inertial navigation navigation hypercomplex number attitude is corresponding _n, pitching angle theta _nwith roll angle γ _n, thus obtain the azimuthal error Δ ψ=ψ of MEMS inertial navigation _n-ψ _m, pitch error Δ θ=θ _n-θ _mwith roll error delta γ=γ _n-γ _m;

(6) according to described azimuthal error Δ ψ, pitch error Δ θ and roll error delta γ, and the propagation time delay Δ t of airborne ins, the gyro zero obtaining MEMS inertial navigation is worth ε partially _x=Δ θ/Δ t, ε _y=Δ γ/Δ t, ε _z=Δ ψ/Δ t, utilizes least squares filtering to estimate ε _x, ε _y, ε _zestimated value l _x, l _y, l _z, the MEMS inertial navigation three axle gyro output quantity after being compensated thus

(7) by up-to-date sky line traffic control hypercomplex number Q ₃navigation hypercomplex number Q after upgrading with step (5) ₂corresponding attitude angle is compared, and produces the instruction angular speed [ω of sky line traffic control hypercomplex number _0xω _0yω _0z] ^t, and obtain with step (6) be added and obtain final instruction angular speed to sky line traffic control hypercomplex number Q ₃upgrade;

(8) step (5) ~ (7) are repeated, to navigation hypercomplex number Q ₂with sky line traffic control hypercomplex number Q ₃carry out continuous updating.