CN105371853A - Mars power descending section navigation method based on TDS and orbiter - Google Patents

Abstract

The present invention relates to a Mars power descending section navigation method based on TDS and an orbiter, and belongs to the technical field of deep space exploration. According to the present invention, a plurality of Doppler radars capable of emitting different wave beam directions are arranged on the TDS, the distances between a lander and different wave beam directions of the Mars surface and the relative velocity are directly obtained, and a radio receiver is arranged on the lander so as to be adopted as the sensor for carrying out radio communication with the orbiter to accurately measure the relative distance between the lander and the orbiter; the lander state equation of the Mars power descending section is established, the lander autonomous navigation measurement model of the Mars power descending section is established, the nonlinear filtering algorithm is used, the autonomous navigation filtering is calculated, and finally the state information of the lander is output; and the calculation time of the measurement data of the TDS and the radio receiver is short so as to meet the real-time requirement of the autonomous navigation.

Description

Based on the Mars power dropping section air navigation aid of TDS and orbiter, orbital vehicle

Technical field

The present invention relates to a kind of based on the Mars power dropping section air navigation aid of decline sensor (TDS) with orbiter, orbital vehicle, belong to field of deep space exploration.

Background technology

Following mars exploration needs lander to realize fixed point soft landing.Power dropping section is the main controlled stages of lander, can control lander and to fly to predetermined target point, realize accurate soft landing by certain guidance algorithm, and premised on this exact position that must provide by navigational system and velocity information.Therefore, structure accurately can determine that the navigation scheme of lander position and speed is the basis of accuracy Mission Success.

The Mars landing task in the past under power section of falling all adopts radar altimeter to measure the elevation information of lander, is then not quite similar to the measurement of velocity information.Mars exploration ramber (MER) utilizes decline Image estimation system (DIMES) to estimate horizontal velocity, and " phoenix number " task, " Mars Pathfinder " task and " pirate number " task more early all have employed radar Doppler and estimate horizontal velocity.MSL (MSL) has carried decline sensor (TDS), is formed by the radar Doppler set in different beams direction, can determine height and the speed of lander.But current navigation sensor all cannot record the horizontal level of lander relative target point so that lander under power the section of falling have larger horizontal position error, guidance system also just uncontrollable lander realize accuracy.

The lander under power section of falling can carry out wireless communication with Mars orbiter, orbital vehicle, measures both relative distance.Because orbiter, orbital vehicle is known relative to the position of landing point, the horizontal level of radio distance-measuring information to lander therefore can be utilized to estimate, then in conjunction with TDS, just can realize the accurate estimation to lander position and speed.

Summary of the invention

The object of the invention is for obtaining power dropping section lander position and velocity information accurately, in conjunction with the navigation problem of Mars power dropping section, adopt decline sensor, radio receiver, inertance element three kinds of sensors, a kind of navigation scheme based on TDS and Mars orbiter, orbital vehicle is proposed, in order to determine the horizontal position information of the height of lander, speed and relative target point, for following Mars accurate soft landing task navigation conceptual design provides technical support and reference.

The section of falling under power, the inertance element of lander navigational system, for a recursion of navigating, is basic sensor; Described TDS there is multiple radar Doppler launching different beams direction, direct acquisition lander, apart from the Distance geometry relative velocity in the multiple different beams direction of martian surface, then calculates the height of lander and to be connected the speed on coordinate system three direction of principal axis in landing point; Described radio receiver is arranged on lander, the sensor of wireless communication is carried out as with orbiter, orbital vehicle, accurately can measure the relative distance of lander and orbiter, orbital vehicle, this relative distance contains the horizontal position information of the relative landing point of lander, greatly can improve the estimated accuracy of lander horizontal level.

Technical scheme of the present invention specifically comprises the steps:

Step 1: the lander state equation setting up Mars power dropping section.

Under landing point is connected coordinate system, lander state x comprises position vector r=[x, y, z] ^t, velocity v=[v _x, v _y, v _z] ^t, attitude quaternion q=[q ₀, q ₁, q ₂, q ₃], accelerometer drift b _a=[b _ax, b _ayb _az] ^t, gyroscope constant value drift b _ω=[b _{ω x}, b _{ω y}, b _{ω z}] ^t.The state equation of power dropping section lander for:

$\begin{array}{c}\stackrel{\·}{r}=v\\ \stackrel{\·}{v}=T_b^l(a_m-b_a-{\mathrm{\ξ}}_m)+g\\ \stackrel{\·}{q}=\frac{1}{2}\mathrm{\Ωq}\\ {\stackrel{\·}{b}}_a=0\\ {\stackrel{\·}{b}}_{\mathrm{\ω}}=0\end{array}---\left(1\right)$

Wherein to be connected the coordinate conversion matrix of coordinate system for lander body is tied to landing point, to be expressed as follows by hypercomplex number:

$T_b^l=\left[\begin{array}{ccc}{q}_0^2+q_1^2-q_2^2-q_3^2& 2(q_0{q}_3+q_1{q}_2)& 2(q_1{q}_3-q_0{q}_2)\\ 2(-q_0{q}_3+q_1{q}_2)& q_0^2-q_1^2+q_2^2-q_3^2& 2(q_2{q}_3+q_0{q}_1)\\ 2(q_0{q}_2+q_1{q}_3)& 2(q_2{q}_3-q_0{q}_1)& q_0^2-q_1^2-q_2^2+q_3^2\end{array}\right]---\left(2\right)$

Ω is the Jacobian matrix of angular velocity:

$\mathrm{\Ω}=\mathrm{\Ω}\left(\mathrm{\ω}\right)=\mathrm{\Ω}({\mathrm{\ω}}_m-b_{\mathrm{\ω}}-{\mathrm{\ξ}}_{\mathrm{\ω}})=\left[\begin{array}{cccc}0& -{\mathrm{\ω}}_x& -{\mathrm{\ω}}_y& -{\mathrm{\ω}}_z\\ {\mathrm{\ω}}_x& 0& {\mathrm{\ω}}_z& -{\mathrm{\ω}}_y\\ {\mathrm{\ω}}_y& -{\mathrm{\ω}}_z& 0& {\mathrm{\ω}}_1\\ {\mathrm{\ω}}_z& {\mathrm{\ω}}_y& -{\mathrm{\ω}}_x& 0\end{array}\right]---\left(3\right)$

A _mand ω _mbe respectively the output valve of accelerometer and gyro, ξ _mand ξ _ωbe the measuring error of accelerometer and gyro respectively, g is acceleration of gravity.ω _x, ω _y, ω _zfor the angular velocity of rotation of lander is along the component of body series three axles.

The described landing point coordinate that is connected is with target landing point for initial point O, and take east orientation as x-axis, north orientation is y-axis, the coordinate system set up with right hand rule.

Described lander body is with lander barycenter for initial point O, and x-axis points to lander head along the longitudinal axis of lander, and y-axis is vertical with x-axis and point to TDS, the vertical Oxy plane of z-axis, and formation right-handed coordinate system.

Step 2: the lander independent navigation measurement model setting up Mars power dropping section.

If the unit vector of different beams direction under landing point is connected coordinate system of TDS is expressed as d _j(j=1,2 ..., n), n is number of beams (in order to improve navigation accuracy, n>=3), has:

${[d_1{d}_2\·\·\·d_n]}_{3\×n}=T_b^l\·S_{3\×n}---\left(4\right)$

Wherein S _{3 × n}for the unit vector matrix of n beam direction under landing point body series.If lander is r along the different beams direction of TDS apart from areographic distance _j(j=1,2 ... n), then have:

r _j＝z/(d _j·[001] ^T)(5)

If lander is v along the speed in different beams direction _j(j=1,2 ... n), then have:

$\left[\begin{array}{c}{v}_1\\ v_2\\ \·\\ \·\\ \·\\ v_n\end{array}\right]=S_{n\×3}\·{T_l}^b{\left[\begin{array}{c}{v}_x\\ v_y\\ v_z\end{array}\right]}^L---\left(6\right)$

S _{n × 3}with be respectively S _{3 × n}with transposition.

Radio receiver receives the radio signal from m orbiter, orbital vehicle, and therefrom obtain signal propagates into lander time from orbiter, orbital vehicle, calculate the relative distance between lander and m orbiter, orbital vehicle, the relative distance between lander and i-th orbiter, orbital vehicle is expressed as

D _i＝c·t _i(7)

In formula, c is the light velocity, t _iit is the travel-time of the radio signal arrival lander that i-th orbiter, orbital vehicle is launched.

Navigation measurement model based on TDS and orbiter, orbital vehicle is:

$y=h\left(x\right)+\mathrm{\υ}=\left[\begin{array}{c}R\\ V\\ D\end{array}\right]+\left[\begin{array}{c}{R}_{\mathrm{\υ}}\\ V_{\mathrm{\υ}}\\ D_{\mathrm{\υ}}\end{array}\right]---\left(8\right)$

R=[r in formula ₁..., r _n], V=[v ₁..., v _n], D=[D ₁..., D _m], R _υ, V _υ, D _υbe respectively the measuring error of radio receiver.

Step 3: independent navigation filtering is resolved.

According to the Mars power dropping section state equation that step 1 obtains and measurement model y=h (the x)+υ that step 2 obtains, the state of lander is estimated by Navigation algorithm.Because state equation and measurement equation are non-linear, adopt nonlinear filtering algorithm, the final status information exporting lander.

The present invention adopts EKF (ExtendKalmanFilter, EKF), and Unscented kalman filtering (UnscentedKalmanFilter, UKF) algorithm is to improve Navigation precision and speed of convergence.

Beneficial effect

(1) merge the metrical information of decline sensor and radio receiver, the position complete to lander and velocity information are estimated.

(2) the measurement data resolving time of TDS and radio receiver is short, meets the demand of independent navigation real-time.

Accompanying drawing explanation

Fig. 1 is the Mars power dropping section air navigation aid process flow diagram based on TDS and orbiter, orbital vehicle;

Fig. 2 is that in embodiment, the be connected real trace under being and filtering track of landing point contrasts, wherein (a) is for lander is apart from the estimated distance in impact point x direction and the contrast of actual distance, (b) estimating speed for lander x direction and the contrast of true velocity, c () is for lander is apart from the estimated distance in impact point y direction and the contrast of actual distance, (d) estimating speed for lander y direction and the contrast of true velocity, e () estimates height and the contrast of true altitude for lander, (f) estimating speed for lander z direction and the contrast of true velocity.

Embodiment

In order to better objects and advantages of the present invention are described, below in conjunction with accompanying drawing and example, summary of the invention is described further.

This example is for the navigation scheme of Mars power dropping segment base in TDS and orbiter, orbital vehicle, in conjunction with the ranging and range rate information of six wave beam TDS, the radio distance-measuring information of lander and single orbiter, orbital vehicle, adopts extended Kalman filter to carry out filtering and resolves, and realizes the high precision navigation of power dropping section.The specific implementation method of this example is as follows:

Step 1: Mars power dropping section state equation is set up

Lander state equation is set up under landing point is connected coordinate system.Lander state x comprises position vector r=[x, y, z] ^t, velocity v=[v _x, v _y, v _z] ^t, attitude quaternion q=[q ₀, q ₁, q ₂, q ₃], acceleration takes into account gyroscope constant value drift b _a=[b _ax, b _ay, b _az] ^t, $b_{\mathrm{\ω}}={[b_{\mathrm{\ωx}},b_{\mathrm{\ωy}},b_{{\mathrm{\ω}}_z}]}^T.$ The state equation of power dropping section write as the form of (1) formula.

$\begin{array}{c}\stackrel{\·}{r}=v\\ \stackrel{\·}{v}=T_b^l(a_m-b_a-{\mathrm{\ξ}}_m)+g\\ \stackrel{\·}{q}=\frac{1}{2}\mathrm{\Ωq}\\ {\stackrel{\·}{b}}_a=0\\ {\stackrel{\·}{b}}_{\mathrm{\ω}}=0\end{array}---\left(1\right)$

Wherein to be connected the coordinate transformation matrix being for body is tied to landing point:

$T_b^l=\left[\begin{array}{ccc}{q}_0^2+q_1^2-q_2^2-q_3^2& 2(q_0{q}_3+q_1{q}_2)& 2(q_1{q}_3-q_0{q}_2)\\ 2(-q_0{q}_3+q_1{q}_2)& q_0^2-q_1^2+q_2^2-q_3^2& 2(q_2{q}_3+q_0{q}_1)\\ 2(q_0{q}_2+q_1{q}_3)& 2(q_2{q}_3-q_0{q}_1)& q_0^2-q_1^2-q_2^2+q_3^2\end{array}\right]---\left(2\right)$

Ω is the Jacobian matrix of angular velocity, and expression formula is as shown in (3) formula.

$\mathrm{\Ω}=\mathrm{\Ω}\left(\mathrm{\ω}\right)=\mathrm{\Ω}({\mathrm{\ω}}_m-b_{\mathrm{\ω}}-{\mathrm{\ξ}}_{\mathrm{\ω}})=\left[\begin{array}{cccc}0& -{\mathrm{\ω}}_x& -{\mathrm{\ω}}_y& -{\mathrm{\ω}}_z\\ {\mathrm{\ω}}_x& 0& {\mathrm{\ω}}_z& -{\mathrm{\ω}}_y\\ {\mathrm{\ω}}_y& -{\mathrm{\ω}}_z& 0& {\mathrm{\ω}}_1\\ {\mathrm{\ω}}_z& {\mathrm{\ω}}_y& -{\mathrm{\ω}}_x& 0\end{array}\right]---\left(3\right)$

A _mand ω _mbe respectively the output valve of accelerometer and gyro, ξ _mand ξ _ωbe the measuring error of accelerometer and gyro respectively, g is acceleration of gravity.

Step 2: Mars power dropping section independent navigation measurement model is set up

Decline sensor TDS can record lander along six beam directions apart from areographic distance and speed, if the unit vector of six beam directions under landing point is connected coordinate system is expressed as d _j(j=1,2 ..., 6), then have:

${[d_1{d}_2\·\·\·d_6]}_{3\×6}=T_B^L\·S_{3\×6}---\left(4\right)$

Wherein S _{3 × 6}be the unit vector matrix of six beam directions under body series.If lander along different beams direction apart from areographic distance be r _j(j=1,2 ..., 6), then have:

r _j＝z/(d _j·[001] ^T)(5)

If lander is v along the speed of three beam directions _j(j=1,2 ..., 6), then have:

$\left[\begin{array}{c}{v}_1\\ v_2\\ v_3\end{array}\right]=S_{3\×6}^T\·T_L^B{\left[\begin{array}{c}{v}_x\\ v_y\\ v_z\end{array}\right]}^L---\left(6\right)$

In formula, for S _{3 × 6}transposed matrix.

Radio receiver receives the radio signal from an orbiter, orbital vehicle, and therefrom obtains signal propagates into lander time from orbiter, orbital vehicle, and then can calculate the relative distance between lander and orbiter, orbital vehicle, is expressed as

D ₁＝c·t ₁(7)

In formula, c is the light velocity, t ₁for the travel-time of the radio signal arrival lander that orbiter, orbital vehicle is launched.

Navigation measurement model based on TDS and orbiter, orbital vehicle is:

$y=h\left(x\right)+\mathrm{\υ}=\left[\begin{array}{c}R\\ V\\ D_1\end{array}\right]+\left[\begin{array}{c}{R}_{\mathrm{\υ}}\\ V_{\mathrm{\υ}}\\ D_{\mathrm{\υ}}\end{array}\right]---\left(8\right)$

R=[r in formula ₁, r ₂..., r ₆], V=[v ₁, v ₂..., v _n], R _υ, V _υ, D _υbe respectively measuring error.

Step 3: independent navigation filtering is resolved

According to Mars power dropping section state equation and measurement model y=h (x)+υ, the state of lander can be estimated by Navigation algorithm.Because state equation and measurement equation are non-linear, EKF (ExtendKalmanFilter, EKF) is thus adopted to improve Navigation precision and speed of convergence, the final status information exporting lander.Simulation parameter arranges as shown in table 1.

Table 1 simulation parameter is arranged

Parameter name	Average	Mean square deviation
			Lander initial position (m)	(1000，800，1700)	(3500，1400，500)
Lander initial estimation speed (m/s)	(-25，-16，-74)	(5，5，5)
			Orbiter, orbital vehicle initial position (km)	(-476，479，217)	/
Accelerometer zero deviation (m/s 2)	3e-3	/
			Gyroscope constant value drift (°/h)	0.5	/
TDS distance accuracy (m)	0.1	/
			TDS rate accuracy (m/s)	0.01	/
Radio distance-measuring precision (m)	50	/

Navigation accuracy is as shown in table 2

Table 2 is based on the navigation accuracy of TDS and orbiter, orbital vehicle

As can be seen from Fig. 2 and table 2, adopt the vision navigation system based on TDS and image measurement, the estimated value of lander position in three directions and speed all can convergence actual value rapidly, illustrates that this navigation scheme can estimate the complete position of lander and velocity information exactly.

Claims (7)

1., based on the Mars power dropping section air navigation aid of TDS and orbiter, orbital vehicle, it is characterized in that: specifically comprise the steps:

Step 1: the lander state equation setting up Mars power dropping section;

Under landing point is connected coordinate system, lander state x comprises position vector r=[x, y, z] ^t, velocity v=[v _x, v _y, v _z] ^t, attitude quaternion q=[q ₀, q ₁, q ₂, q ₃], accelerometer drift b _a=[b _ax, b _atb _az] ^t, gyroscope constant value drift b _ω=[b _{ω x}, b _{ω y}, b _{ω z}] ^t; The state equation of power dropping section lander for:

$\begin{array}{c}\stackrel{\·}{r}=v\\ \stackrel{\·}{v}=T_b^l(a_m-b_a-{\mathrm{\ξ}}_m)+g\\ \stackrel{\·}{q}=\frac{1}{2}\mathrm{\Ωq}\\ {\stackrel{\·}{b}}_a=0\\ {\stackrel{\·}{b}}_{\mathrm{\ω}}=0\end{array}---\left(1\right)$

Wherein to be connected the coordinate conversion matrix of coordinate system for lander body is tied to landing point, to be expressed as follows by hypercomplex number:

$T_b^l=\left[\begin{array}{ccc}{q}_0^2+q_1^2-q_2^2-q_3^2& 2(q_0{q}_3+q_1{q}_2)& 2(q_1{q}_3-q_0{q}_2)\\ 2(-q_0{q}_3+q_1{q}_2)& q_0^2-q_1^2+q_2^2-q_3^2& 2(q_2{q}_3+q_0{q}_1)\\ 2(q_0{q}_2+q_1{q}_3)& 2(q_2{q}_3-q_0{q}_1)& q_0^2-q_1^2-q_2^2+q_3^2\end{array}\right]---\left(2\right)$

Ω is the Jacobian matrix of angular velocity:

$\mathrm{\Ω}=\mathrm{\Ω}\left(\mathrm{\ω}\right)=\mathrm{\Ω}({\mathrm{\ω}}_m-b_{\mathrm{\ω}}-{\mathrm{\ξ}}_{\mathrm{\ω}})=\left[\begin{array}{cccc}0& -{\mathrm{\ω}}_x& -{\mathrm{\ω}}_y& -{\mathrm{\ω}}_z\\ {\mathrm{\ω}}_x& 0& {\mathrm{\ω}}_z& -{\mathrm{\ω}}_y\\ {\mathrm{\ω}}_y& -{\mathrm{\ω}}_z& 0& {\mathrm{\ω}}_1\\ {\mathrm{\ω}}_z& {\mathrm{\ω}}_y& -{\mathrm{\ω}}_x& 0\end{array}\right]---\left(3\right)$

A _mand ω _mbe respectively the output valve of accelerometer and gyro, ζ _mand ζ _ωbe the measuring error of accelerometer and gyro respectively, g is acceleration of gravity; ω _x, ω _y, ω _zfor the angular velocity of rotation of lander is along the component of body series three axles;

Step 2: the lander independent navigation measurement model setting up Mars power dropping section;

If the unit vector of different beams direction under landing point is connected coordinate system of TDS is expressed as d _j, j=1,2 ..., n, n are number of beams, have:

${[d_1{d}_2\·\·\·d_n]}_{3\×n}=T_b^l\·S_{3\×n}---\left(4\right)$

Wherein S _{3 × n}for the unit vector matrix of n beam direction under landing point body series; If lander is r along the different beams direction of TDS apart from areographic distance _j, then have:

r _j＝z/(d _j·[001] ^T)(5)

If lander is v along the speed in different beams direction _j, then have:

$\left[\begin{array}{c}{v}_1\\ v_2\\ \·\\ \·\\ \·\\ v_n\end{array}\right]=S_{n\×3}\·{T_l}^b{\left[\begin{array}{c}{v}_x\\ v_y\\ v_z\end{array}\right]}^L---\left(6\right)$

S _{n × 3}with be respectively S _{3 × n}with transposition;

Radio receiver receives the radio signal from m orbiter, orbital vehicle, and therefrom obtain signal propagates into lander time from orbiter, orbital vehicle, calculate the relative distance between lander and m orbiter, orbital vehicle, the relative distance between lander and i-th orbiter, orbital vehicle is expressed as

D _i＝c·t _i(7)

In formula, c is the light velocity, t _iit is the travel-time of the radio signal arrival lander that i-th orbiter, orbital vehicle is launched;

Navigation measurement model based on TDS and orbiter, orbital vehicle is:

$y=h\left(x\right)+\mathrm{\υ}=\left[\begin{array}{c}R\\ V\\ D\end{array}\right]+\left[\begin{array}{c}{R}_{\mathrm{\υ}}\\ V_{\mathrm{\υ}}\\ D_{\mathrm{\υ}}\end{array}\right]---\left(8\right)$

R=[r in formula ₁..., r _n], V=[v ₁..., v _n], D=[D ₁..., D _m], R _υ, V _υ, D _υbe respectively the measuring error of radio receiver;

Step 3: independent navigation filtering is resolved;

According to the Mars power dropping section state equation that step 1 obtains and measurement model y=h (the x)+υ that step 2 obtains, estimated the state of lander by nonlinear navigation filtering algorithm, export the status information of lander.

2. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, is characterized in that: described nonlinear navigation filtering algorithm comprises EKF, Unscented kalman filtering algorithm.

3. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, it is characterized in that: described TDS has multiple radar Doppler launching different beams direction, direct acquisition lander, apart from the Distance geometry relative velocity in the multiple different beams direction of martian surface, then calculates the height of lander and to be connected the speed on coordinate system three direction of principal axis in landing point.

4. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, it is characterized in that: described radio receiver is arranged on lander, carry out the sensor of wireless communication as with orbiter, orbital vehicle, accurately can measure the relative distance of lander and orbiter, orbital vehicle.

5. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, it is characterized in that: the described landing point coordinate that is connected is for initial point O with target landing point, take east orientation as x-axis, north orientation is y-axis, the coordinate system set up with right hand rule.

6. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, it is characterized in that: described lander body is for initial point O with lander barycenter, x-axis points to lander head along the longitudinal axis of lander, y-axis is vertical with x-axis and point to TDS, the vertical Oxy plane of z-axis, and form right-handed coordinate system.

7. the Mars power dropping section air navigation aid based on TDS and orbiter, orbital vehicle according to claim 1, is characterized in that: n >=3.