CN105841699A - Radar altimeter assistance method aiming to inertial navigation - Google Patents

Abstract

The invention provides a radar altimeter assistance method aiming to inertial navigation. In the method, a measurement quantity is constructed in the manner of calculating earth core altitude of an air vehicle according to position vectors outputted by the inertial navigation, calculating earth core altitude minus earth ellipsoidal radius to obtain the calculated altitude height of the air vehicle, obtaining observed altitude height by means of a measurement value from the radar altimeter with other algorithms, and calculating altitude height minus the observed altitude height to obtain a difference as the measurement quantity, and finally deriving a total differential equation of the measurement quantity. The invention provides a particular embodiment, wherein a strapdown inertial navigation/radar altimeter combined navigation system is formed on the basis of the measurement quantity. Estimation on position error status quantity is achieved, diffusion of position error is inhibited and positioning precision is increased through an indirect method Kalman filtering with feedback compensation.

Description

Radar altimeter householder method for inertial navigation

(1) technical field

The present invention provides a kind of radar altimeter householder method for inertial navigation, belongs to navigation field, leads for improving aircraft Boat positioning precision provides a kind of effective ways.

(2) background technology

Inertial navigation is as a kind of entirely autonomous navigation system, and its ultimate principle is by gyroscope and accelerometer sensitive carrier Rotate and translation information, complete the attitude of aircraft, speed and position through formula layout and resolve, there is output continuously, reliably Property remarkable advantages such as precision is high by force, in short-term, but there is also the shortcoming that error dissipates in time.

Radar altimeter is a kind of active range sensor, and radar has different systems, such as: laser radar or radio thunder Reach, in order to measure aircraft and the distance of a certain target object.Along with guidance navigation is navigated in survey of deep space, space flight with controlling technology Empty fast development, radar range finding technology oneself through be successfully applied to orbit measurement, projectile correction, satellite fix, spacecrafts rendezvous, The aspects such as cruise guidance, space flight landing, geodesic survey, microwave remote sensing.

How improving navigation accuracy during autonomous flight is the most all hot research problem, inertial navigation and star sensor Combination be a kind of conventional independent navigation solution, program mechanism be by merge star sensor high-precision fixed appearance result Thus improve the precision of the inertial navigation system frame of reference, and reduce the cross-coupling effect of accelerometer, can subtract to a certain extent Speed that little integrated acceleration method is brought and site error, but can not fundamentally solve the divergence problem of speed and site error. Inertial navigation can also combine with GPS, and GPS positioning result does not has drift error, but is not usually considered that GPS is entirely autonomous Navigation mode, both combinations are not the most entirely autonomous navigation modes.The present invention is directed to inertial navigation system and a kind of radar altitude is provided Table and the used Kalman Filtering Data Fusion method organized.

(3) summary of the invention

It is an object of the invention to provide a kind of measurement model based on radar altimeter, utilize Kalman filtering aided inertial navigation System obtains higher speed, positional precision.

Aircraft, in flight course, utilizes radar to record vertical high to reflecting surface (sea level or Earth surface plane) of aircraft Angle value.Aircraft observation height above sea level is obtained further, as shown in Figure 1 in conjunction with other method.Export according to inertial navigation system Position vector be calculated aircraft geocentric altitude, deduct earth ellipsoid radius obtain aircraft calculate height above sea level, calculate sea Degree of lifting deducts the observation difference that obtains of height above sea level as measurement.

The position of aircraft that inertial navigation system resolves has the drift error continued to increase in time, the measurement that the inventive method provides Amount and position excursion error have functional relationship, carry out Kalman filtering in conjunction with inertial navigation system state equation, can play benefit Repay the effect of position excursion error, improve the positioning precision of aircraft.A combination thereof navigation system schematic diagram is as shown in Figure 2.

(4) accompanying drawing explanation

Fig. 1 is that under launching inertial system, position of aircraft P obtains height above sea level observed altitude H with radar range finding_oGeometrical relationship.

Fig. 2 is the integrated navigation system schematic diagram of radar altimeter aided inertial navigation system.

(5) detailed description of the invention

For strapdown inertial navigation system (SINS), the present invention is utilized to obtain based on radar altimeter (RA, Radar Altimeter) The system quantities measurement obtained establishes SINS/RA integrated navigation system, uses indirect feedback corrector Kalman filter, filtering knot On the one hand fruit feeds back to inertial navigation system correction deviation, on the other hand correction current time State-output amount, it is achieved that to aircraft position The estimation of the amount of putting, reduces position excursion error.

(1) measurement model is set up

As it is shown in figure 1, R_eFor earth radius.At launch inertial coordinate system O_l-X_lY_lZ_l(hereinafter referred to as l system, and sit as navigation Mark system) under, position vector r and launch point horizontal plane angle θ meets

$sin\mathrm{\θ}=\frac{y}{r}=\frac{y}{\sqrt{x^2+y^2+z^2}}---\left(1\right)$

In formula, x, y, z is respectively SINS position calculation result coordinate under l system.Then this moment height above sea level H_cFor:

$\begin{array}{c}{H}_c=\sqrt{r^2+R_e^2+2{\mathrm{rR}}_e\sin \mathrm{\θ}}-R_e\\ =\sqrt{x^2+y^2+{(R_e+y)}^2}-R_e\end{array}---\left(2\right)$

If the height above sea level measured value obtained by radar altimeter is H_o, then SINS solves calculated altitude H_cWith H_oDifference Δ H meet such as Lower relation:

$\mathrm{\Δ}H=H_c-H_o=\[\begin{array}{ccc}\frac{\∂H_c}{\∂x}& \frac{\∂H_c}{\∂y}& \frac{\∂H_c}{\∂z}\end{array}\]\left[\begin{array}{c}\mathrm{\δ}x\\ \mathrm{\δ}y\\ \mathrm{\δ}z\end{array}\right]+w_h---\left(3\right)$

Wherein:

$\begin{array}{c}\frac{\∂H_c}{\∂x}=\frac{x}{\sqrt{x^2+y^2+z^2+R_e^2+2{\mathrm{yR}}_e}}=\frac{x}{\sqrt{x^2+z^2+{(R_e+y)}^2}}\\ \frac{\∂H_c}{\∂y}=\frac{R_e+y}{\sqrt{x^2+y^2+z^2+R_e^2+2{\mathrm{yR}}_e}}=\frac{R_e+y}{\sqrt{x^2+z^2+{(R_e+y)}^2}}\\ \frac{\∂H_c}{\∂z}=\frac{z}{\sqrt{x^2+y^2+z^2+R_e^2+2{\mathrm{yR}}_e}}=\frac{z}{\sqrt{x^2+z^2+{(R_e+y)}^2}}\end{array}---\left(4\right)$

w_hFor the measurement noise of radar altimeter, its size is affected by attitude error and reflecting surface fluctuating.δ x, δ y, δ z are respectively Strapdown resolves the site error with nominal track.

(2) Kalman filter equation is set up

Using the Kalman filter of indirect method band feedback compensation, therefore selecting system navigation error parameter is as state variable, bag The concrete element contained is: SINS mathematical platform misalignment φ_x,φ_y,φ_z, velocity error on three axles of l system δv_x,δv_y,δv_zWith site error δ x, δ y, δ z, gyro drift ε_x,ε_y,ε_zAnd accelerometer bias_x,▽_y,▽_z, That is:

X=[φ_x φ_y φ_z δv_x δv_y δv_z δx δy δz ε_x ε_y ε_z ▽_x ▽_y ▽_z]^T (5)

System state equation is:

$\stackrel{\·}{X}\left(t\right)=F\left(t\right)X\left(t\right)+G\left(t\right)W\left(t\right)---\left(6\right)$

Wherein, F (t) is state-transition matrix, and G (t) is that system noise drives battle array, and W (t) is system noise.

With Δ H as observed quantity, system measurements equation is:

Z (t)=Δ H=HX (t)+V (t) (7)

In formula,

$H=\left[\begin{array}{ccccccccccccccc}0& 0& 0& 0& 0& 0& \frac{\∂H_c}{\∂x}& \frac{\∂H_c}{\∂y}& \frac{\∂H_c}{\∂z}& 0& 0& 0& 0& 0& 0\end{array}\right]---\left(8\right)$

V (t)=w_h, for measurement noise.

Claims (3)

1. for the radar altimeter householder method of inertial navigation, it is characterised in that: calculate according to the position vector of inertial navigation output To aircraft geocentric altitude, deducting earth ellipsoid radius and obtain aircraft calculating height above sea level, radar altimeter measured value combines Other algorithm obtains observation height above sea level, calculates height above sea level and deducts the observation difference that obtains of height above sea level as measurement.

Radar altimeter householder method for inertial navigation the most according to claim 1, it is characterised in that: based on this measurement Set up inertial navigation/radar altimeter Integrated Navigation Algorithm, using this measurement as the one-dimensional observed quantity of Kalman filter.

Radar altimeter householder method for inertial navigation the most according to claim 1, it is characterised in that: based on this measurement Set up inertial navigation/star sensor/radar altimeter Integrated Navigation Algorithm, inertial navigation system and star sensor measure and obtained The misaligned angle of the platform obtained merges with this measurement, as the four-dimensional observed quantity of Kalman filter.