CN101074881A - Inertial navigation method for moon detector in flexible landing stage - Google Patents

Abstract

An inertia-navigating method of moon craft at soft landing stage includes setting up moon navigation equation to obtain moon navigation orbit, using speed measurer to obtain speed of craft according to attitude information and using distance measurer to obtain position of craft according attitude information, integrating speed and position of craft with inertia navigation orbit to obtain final navigation orbit data after filter-revision.

Description

A kind of inertial navigation method for moon detector in flexible landing stage

Technical field

The invention belongs to moon detector in flexible landing stage inertial navigation technology field.

Background technology

Moon detector in flexible landing stage, inertial navigation is a kind of simple and reliable air navigation aid, and the Lunar series lunar orbiter of USSR (Union of Soviet Socialist Republics), the Surveyor series lunar orbiter of the U.S. and Apollo series lunar orbiter system have all adopted the method for inertial navigation.The principle of inertial navigation is, according to the bookbinding data of preliminary orbit, integration carried out in the output of Inertial Measurement Unit, obtains position of detector and speed.But in the moon detector in flexible landing process, there is error inevitably in the preliminary orbit bookbinding data that provided by ground observing and controlling system; Also there are measuring error in gyro in the Inertial Measurement Unit and accelerometer.The existence of these errors has influenced navigation accuracy, just is being embodied in this based on the weak point of the inertial navigation method of integral principle.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome prior art can't reduce the measuring error of preliminary orbit sum of errors Inertial Measurement Unit at soft lunar landing stage inertial navigation method weak point, a kind of inertial navigation method based on range finding, the correction of testing the speed is provided, and this method can weaken the influence of these errors to navigation accuracy.

The technical solution of method of the present invention is: a kind of inertial navigation method for moon detector in flexible landing stage comprises the following steps:

(1) set up moon navigation equation, obtain the attitude information of lunar orbiter by gyro, comprehensive track initial value of accelerometer and attitude information obtain the inertial navigation track;

(2) knotmeter obtains detector speed according to attitude information; Stadimeter obtains detector position according to attitude information;

(3) inertial navigation track comprehensive survey device speed and position obtain final navigation orbital data through the filtering correction.

The present invention's beneficial effect compared with prior art is: the present invention utilizes range finding, the information that tests the speed is revised the inertial navigation result, has effectively improved the independent navigation precision in soft lunar landing stage, has alleviated the pressure of ground observing and controlling system to a certain extent.

Description of drawings

Fig. 1 is the inertial navigation schematic diagram based on range finding, the correction of testing the speed of the present invention;

Fig. 2 is the site error of inertial navigation under no correction situation;

Fig. 3 is the velocity error of inertial navigation under no correction situation;

Fig. 4 is that inertial navigation of the present invention is in the site error that has under the correction situation;

Fig. 5 is that inertial navigation of the present invention is in the velocity error that has under the correction situation.

Embodiment

The inertial navigation principle of revising based on finding range, testing the speed as shown in Figure 1.

(1) sets up navigation equation.According to the navigation fundamental equation in the inertial coordinates system, obtain the attitude information of lunar orbiter by gyro, comprehensive track initial value of accelerometer and attitude information obtain the inertial navigation track.Inertial coordinates system F _IInterior navigation fundamental equation is

$\underset{\‾}{\overset{\·\·}{r}}=\underset{\‾}{g}+\underset{\‾}{a}$

Wherein, rFor the moon heart to the vector of lander barycenter, gBe the gravitational acceleration vector, aFor acting on the apparent acceleration vector on the lander.Accelerometer measures be that the apparent acceleration of lander relative inertness system is at F _BIn projection $\underset{\‾}{a}=F_B^T{\left[\begin{array}{ccc}{a}_x& a_y& a_z\end{array}\right]}^T.$ Navigation equation is at inertial coordinates system F _IIn be expressed as

$\left[\begin{array}{c}\stackrel{\·\·}{x}\\ \stackrel{\·\·}{y}\\ \stackrel{\·\·}{z}\end{array}\right]=\left[\begin{array}{c}{g}_x\\ g_y\\ g_z\end{array}\right]+C_{\mathrm{IB}}\left[\begin{array}{c}{a}_x\\ a_y\\ a_z\end{array}\right]$

Wherein, $\left[\begin{array}{c}{g}_x\\ g_y\\ g_z\end{array}\right]=-\frac{\mathrm{\μ}}{r^3}\left[\begin{array}{c}x\\ y\\ z\end{array}\right],$ R=(x ²+ y ²+ z ²) ^1/2, C _IBFor body is tied to the attitude transition matrix of inertial system, obtain by gyro to measure.

(2) state equation.Navigation equation is converted into state equation.If state variable $X={\left[\begin{array}{cccccc}x& y& z& \stackrel{\·}{x}& \stackrel{\·}{y}& \stackrel{\·}{z}\end{array}\right]}^T,$ State equation

$\stackrel{\·}{X}=f\left(X\right)+U$

Wherein

$f\left(X\right)=\left[\begin{array}{c}\stackrel{\·}{x}\\ \stackrel{\·}{y}\\ \stackrel{\·}{z}\\ -\frac{\mathrm{\μ}}{r^3}x\\ -\frac{\mathrm{\μ}}{r^3}y\\ -\frac{\mathrm{\μ}}{r^3}z\end{array}\right]$ $U=\left[\begin{array}{c}0\\ 0\\ 0\\ C_{\mathrm{IB}}\left[\begin{array}{c}{a}_x\\ a_y\\ a_z\end{array}\right]\end{array}\right]$

If X _RefBe inertial navigation mode, the error state δ X=X-X of define system _RefBy state equation can the guiding system error state the linearization state equation

$\frac{d}{\mathrm{dt}}\left(\mathrm{\δX}\right)=\mathrm{A\δX}$

Wherein

$A={\frac{\&PartialD;f\left(X\right)}{\&PartialD;X}|}_{X=X_{\mathrm{ref}}}=\left[\begin{array}{cc}{0}_{3\&times;3}& I_3\\ \mathrm{<figure-callout\; id="0"\; label="M"\; filenames="A20071013018800091.png,A20071013018800092.png"\; state="\{\{state\}\}">M</figure-callout>}& 0_{3\&times;3}\end{array}\right]$

$M=-\frac{\mathrm{\&mu;}}{{(x_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+y_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+z_{\mathrm{ref}}^2)}^{\frac{5}{2}}}\left[\begin{array}{ccc}{-2x}_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+y_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+z_{\mathrm{ref}}^2& -{3x}_{\mathrm{ref}}{y}_{\mathrm{ref}}& -{3x}_{\mathrm{ref}}{z}_{\mathrm{ref}}\\ {-3x}_{\mathrm{ref}}{y}_{\mathrm{ref}}& x_{\mathrm{ref}}^2-{2y}_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+z_{\mathrm{ref}}^2& -3z_{\mathrm{ref}}{y}_{\mathrm{ref}}\\ {-3x}_{\mathrm{ref}}{z}_{\mathrm{ref}}& -{3z}_{\mathrm{ref}}{y}_{\mathrm{ref}}& x_{\mathrm{<figure-callout\; id="2"\; label="ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">ref</figure-callout>}}^2+y_{\mathrm{ref}}^2-2{\mathrm{<figure-callout\; id="2"\; label="z\; ref"\; filenames="A20071013018800091.png"\; state="\{\{state\}\}">z</figure-callout>}}_{\mathrm{ref}}^{}\end{array}\right]$

(3) observation equation.Stadimeter, knotmeter obtain position of detector and speed according to attitude information.The speed v that knotmeter is measured _FBe body coordinate system F _BCoordinate system F relatively is connected _FSpeed at F _BIn expression.Body F _BRelative inertness F _ISpeed

v _I＝ v _F+ ω× r _I

Wherein, ωBe moon spin velocity vector, size is ω _LFollowing formula is at F _IIn be expressed as

$\left[\begin{array}{c}\stackrel{\&CenterDot;}{x}\\ \stackrel{\&CenterDot;}{y}\\ \stackrel{\&CenterDot;}{z}\end{array}\right]=C_{\mathrm{IB}}{v}_F+{\mathrm{\&omega;}}^{\&times;}\left[\begin{array}{c}x\\ y\\ z\end{array}\right]$

Wherein, ${\mathrm{\&omega;}}^{\&times;}=\left[\begin{array}{ccc}0& {-\mathrm{\&omega;}}_{\mathrm{<figure-callout\; id="0"\; label="L"\; filenames="A20071013018800091.png,A20071013018800092.png"\; state="\{\{state\}\}">L</figure-callout>}}& 0\\ {\mathrm{\&omega;}}_{\mathrm{<figure-callout\; id="0"\; label="L"\; filenames="A20071013018800091.png,A20071013018800092.png"\; state="\{\{state\}\}">L</figure-callout>}}& 0& 0\\ 0& 0& 0\end{array}\right].$ Stadimeter can obtain the distance of detector along the sensor direction of visual lines to lunar surface, in conjunction with month radius of a ball and attitude, can obtain position of detector

Then measure equation

$Z=h\left(X\right)=\left[\begin{array}{c}{C_{\mathrm{IB}}}^{-1}[\left[\begin{array}{c}\stackrel{\&CenterDot;}{x}\\ \stackrel{\&CenterDot;}{y}\\ \stackrel{\&CenterDot;}{z}\end{array}\right]-{\mathrm{\&omega;}}^{\&times;}\left[\begin{array}{c}x\\ y\\ z\end{array}\right]]\\ x\\ y\\ z\end{array}\right]$

Linearization gets

$C={\frac{\&PartialD;h\left(X\right)}{\&PartialD;X}|}_{X{=X}_{\mathrm{ref}}}=\left[\begin{array}{c}{C_{\mathrm{IB}}}^{-1}\left[\begin{array}{cc}{-\mathrm{\&omega;}}^{\&times;}& I_3\end{array}\right]\\ \left[\begin{array}{cc}{I}_3& 0_{1\&times;3}\end{array}\right]\end{array}\right]$

Can prove,

Nonsingular, promptly system is considerable.

(4) filtering equations.If

Be the estimated value of δ X,

Estimated value for X.Adopt kalman filter method to try to achieve gain matrix K, then filtering computing formula is as follows

(5) simulation result.Suppose navigation preliminary orbit bookbinding error (577m, 577m, 577m, 0.577m/s, 0.577m/s 0.577m/s), only utilizes IMU to navigate, the site error of simulation result as shown in Figure 2, velocity error as shown in Figure 3.From simulation result as can be seen, only utilize the independent navigation of IMU, can't revise the initial navigation error, position and velocity error increased and increase along with the time, caused the final bigger navigation deviation that exists.In the time of 310 seconds, introduce test the speed and ranging information to IMU output revise, the site error of simulation result as shown in Figure 4, velocity error as shown in Figure 5, table 1, table 2 have provided in 500s, 600s and the 700s moment, site error and the velocity error correlation data before and after revising.As can be seen, position and velocity error are all significantly revised.

Contrast (m) before and after the correction of table 1 site error

	Site error (before revising)			Site error (revising the back)
							X-axis	Y-axis	The Z axle	X-axis	Y-axis	The Z axle
	500s	-1080	-880	-780	-120	-115							-15
600s							-1240	-950	-820	-119	-35	-20
	700s	-1410	-1010	-870	-120	11							-9

Contrast (m/s) before and after the correction of table 2 velocity error

	Velocity error (before revising)			Velocity error (revising the back)
							X-axis	Y-axis	The Z axle	X-axis	Y-axis	The Z axle
	500s	-1.55	-0.62	-0.4	-0.1	-0.15							0.2
600s							-1.83	-0.7	-0.45	0.07	-0.1	0.1
	700s	-2.1	-0.77	-0.45	-0.01	0.05							-0.02

。

Claims (4)

1. an inertial navigation method for moon detector in flexible landing stage is characterized in that comprising the following steps:

(1) set up moon navigation equation, obtain the attitude information of lunar orbiter by gyro, comprehensive track initial value of accelerometer and attitude information obtain the inertial navigation track;

(2) knotmeter obtains detector speed according to attitude information; Stadimeter obtains detector position according to attitude information;

(3) inertial navigation track comprehensive survey device speed and position obtain final navigation orbital data through the filtering correction.

2. inertial navigation method for moon detector in flexible landing stage according to claim 1 is characterized in that: setting up moon navigation equation in the described step (1) is the interior navigation fundamental equation of inertial coordinates system.

3. inertial navigation method for moon detector in flexible landing stage according to claim 1 is characterized in that: the filtering correction in the described step (3) is to have utilized position of detector and velocity information.

4. according to claim 1 or 3 described inertial navigation method for moon detector in flexible landing stage, it is characterized in that: described filtering modification method is a Kalman filtering method.

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