CN103674034A - Robust navigation method capable of realizing multi-beam velocity and distance measurement correction - Google Patents

Abstract

The invention relates to a robust navigation method capable of realizing multi-beam velocity and distance measurement correction. The robust navigation method comprises the steps of firstly performing inertial navigation computation, then performing distance measurement correction and finally performing velocity correction, wherein a signal source used for distance measurement correction is a distance measuring sensor; signal sources for velocity correction are divided into two kinds, wherein one is the kind of distance measurement signals and the other is the kind of velocity measurement signals; when the distance measurement signals are effective and flight height is within a velocity measurement correction introduction range, height solved through two continuous distance measurement signals is differentiated to obtain velocity in the direction of gravity; when the velocity measurement signals are effective and the flight height is within the velocity measurement correction introduction range, the velocity along the direction of wave beams of the velocity measuring sensor is directly obtained. Lander inertial velocity at a moment corresponding to the distance measurement signal or the velocity measurement signal is calculated by using velocity of two cycles before inertial navigation, then velocity relative to lunar surface is obtained through conversion and the velocity is projected into the direction of height or the direction of velocity measuring wave beams. The robust navigation method capable of realizing multi-beam velocity and distance measurement correction has the advantages that the method is simple and practical, the reliability is high and the high accuracy of the whole navigation process is guaranteed.

Description

The robust air navigation aid that multi-beam tests the speed and finds range and revise

Technical field

The present invention relates to the robust air navigation aid that a kind of multi-beam tests the speed and finds range and revise, belong to the independent navigation field of the moon and survey of deep space soft landing task.

Background technology

Objects outside Earth soft landing is a brand-new task to China.For GNC system, obtaining navigation data is accurately the prerequisite that guarantees the safe soft landing of lander.Because soft landing process time is short, independence is high, therefore external existing similar tasks, comprises that Lunar Landing Mission that Lunar Landing Mission that Apollo, the Surveyor of the U.S. and the Luna series of USSR (Union of Soviet Socialist Republics) etc. have been implemented and European Space Agency and Japan's plan carry out is all using inertial navigation as main navigate mode.

Inertial navigation technology relies on the direct measured angular speed of gyro integration obtain detector attitude angle or set up stable platform and obtain attitude by frame corners with gyro; Rely on accelerometer measures specific force, and integration obtains speed and position.This air navigation aid has the advantages that operational process does not rely on external information.Because inertial navigation is a kind of extrapolation algorithm, so the precision of inertial navigation depends on the error of navigation initial value and inertia device, and it presents ever-increasing feature in time.

For soft landing task, the initial value of inertial navigation is provided by ground observing and controlling and star sensor.And for objects outside Earth, the precision of ground observing and controlling is limited, detector position error reaches km magnitude.Using this as navigation initial value to soft landing, be dangerous.The topographic relief of adding celestial body surface also has uncertainty.Therefore it is infeasible relying on merely inertial navigation to complete soft landing navigation task.

For this reason, external objects outside Earth soft landing task is all equipped with microwave or laser ranging or the sensor that tests the speed, so that inertial navigation is revised.But due to the difference of the difference of various countries' sensor technical merit, the difference of detector scale and soft landing guidance algorithm, the concrete grammar that utilizes range finding or the information that tests the speed to revise inertial navigation is not identical yet.For example, the Apollo lunar module of the U.S. has adopted Microwave Velocity, range finding sensor, and this sensor has three test the speed wave beam and distance measurement beams.Three wave beams that test the speed are worked respectively, and then geometric solution is calculated the speed of lunar module to lunar surface.This sensor has two different gears in addition, can guarantee that distance measurement beam points to lunar surface as far as possible according to the attitude of lunar module, thereby obtain the elevation information of lunar module.Refer to " Bernard A.Kriegsman; Radar-updated inertial navigation of a continuously-powered space vehicle; IEEE transaction on Aerospace and Electronic Systems, Vol.AES-2, No.4; July; 1966,549-565 " and " Floyd V.Bennett, Apollo lunar descent and ascent trajectories; NASA TMX58040, March1970 ".The method of filter correction has adopted the kalman filter method of simplifying.The shortcoming that said method exists is: (1) needs three wave beams effectively simultaneously, and fault-tolerance is bad; (2) only have a distance measurement beam, in order to adapt to the adjustment of large attitude, need movable part, reduced the reliability of system.

The navigation sensor that China lunar exploration second phase soft landing adopts is different from external inter-related task, except the IMU for inertial navigation, is also equipped with tellurometer survey, test the speed sensor and laser ranging sensor be for revising inertial navigation.For fear of sensor rotating mechanism fault, landing mission is worked the mischief, the tellurometer survey of China, the sensor that tests the speed are different from Apolline way, are fixed installation.But in order to guarantee can both to test the speed and ranging information in 90 ° of attitude slewing areas in landing mission, the Microwave Velocity of outfit, range finding sensor have three different test the speed wave beam and two different distance measurement beams that point to that point to.Add laser ranging sensor and also there are two different distance measurement beams that point to, so the number of beams revised of can be used in China's soft landing task testing the speed, find range is much larger than Apollo.These wave beams are measured and are asynchronous, and precision is different, utilizing these information and taking into account the needs of System Fault Tolerance effectively and reasonably how, and this becomes the difficult problem that navigation algorithm need to solve.

Summary of the invention

Technology of the present invention is dealt with problems: overcome the deficiencies in the prior art, and the robust air navigation aid that provides a kind of simple and practical multi-beam to test the speed range finding correction, and reliability is high, has guaranteed the omnidistance precision of navigating.

The technical solution adopted in the present invention is: the robust air navigation aid that multi-beam tests the speed and finds range and revise, performing step is as follows

(1) inertial reference calculation

Utilize the inertia attitude of the angle step calculating current time lander of gyro to measure in the previous sampling period, then utilize the speed of the speed increment accumulation calculating current time lander that accelerometer measures obtains, and integration obtains the position of lander under inertial coordinates system;

(2) range finding is revised

If receive the valid data of range finding sensor, and flying height meets the introducing scope that range finding is revised, by these data by matrix being installed and the conversion of lander attitude matrix solves altitude information; Utilize the inertial reference calculation in the first two cycle simultaneously, calculate range finding sensor data lander height on corresponding time point, it is poor that the height that inertial navigation and stadimeter are resolved is respectively done, and according to the site error of altitude correction factor corrected altitude direction; This altitude correction factor changes with lander highly linear, and lander height is higher, and described altitude correction factor is less, and highly more low clearance correction factor is larger;

(3) speed correction

The signal source of speed correction is divided into two kinds, and a kind of is distance measuring signal, and another kind is tachometer signal, when distance measuring signal effective, and flying height meets while testing the speed correction introducing scope, the height that double distance measuring signal is solved carries out difference, obtain the speed of gravity direction, utilize the corresponding lander velocity inertial constantly of speed calculation distance measuring signal in the first two cycle in inertial reference calculation simultaneously, and be converted to the speed of relative lunar surface, it is poor that the speed that the speed that distance measuring signal is obtained and inertial navigation obtain is done, and according to the speed of speed correction factor corrected altitude direction, this speed correction factor is with lander speed linear change, lander speed is faster, speed correction factor is less, speed is less, speed correction factor is larger, effective for tachometer signal, utilize the corresponding lander velocity inertial constantly of speed calculation tachometer signal in the first two cycle in inertial reference calculation, and be converted to the speed of relative lunar surface, it is poor that the speed that the speed that tachometer signal is obtained and inertial navigation obtain is done, and according to the speed of speed correction factor correction beam direction, this speed correction factor is with lander speed linear change, lander speed is faster, speed correction factor is less, and speed is less, and speed correction factor is larger,

(4) each control cycle is repeated to the step of (1)～(3), complete navigation calculation.

Described altitude correction factor and speed correction factor generally change between 0～1.

The present invention's advantage is compared with prior art: the present invention does not require range finding, the wave beam that tests the speed has movable part, has increased the reliability of system; In navigation procedure, any one wave beam effectively can both the result to inertial reference calculation be revised on certain orientation, is conducive to improve system survivability, reduces the complicacy of fault diagnosis algorithm; Filtering has adopted the correction factor of linear change, and comparison with standard Kalman filtering is calculated simple, can to filter correction parameter, automatically adjust for the different accuracy of different phase sensor again, to guarantee the omnidistance high precision of navigating.

Accompanying drawing explanation

Fig. 1 is the test the speed theory diagram of the robust air navigation aid that range finding revises of multi-beam of the present invention;

Fig. 2 is the definition of reference frame;

Fig. 3 is the curve of actual lander change in location in soft lunar landing process;

Fig. 4 is the change in location curve that in soft lunar landing process, navigational system provides;

Fig. 5 is the curve of actual lander velocity variations in soft lunar landing process;

The speed change curves that in Fig. 6 soft lunar landing process, navigational system provides.

Embodiment

As shown in Figure 1, the multi-beam of the present invention concrete computation process of robust air navigation aid that range finding revises that tests the speed is as follows:

The first step, carries out inertial reference calculation

(1) foundation of reference frame and navigational system initial value (initialization)

Body series (O _b-X _by _bz _b) connect firmly at lander, initial point overlaps with barycenter, and three axles are parallel to the lander principal axis of inertia.

Navigational system is selected landing inertial coordinates system (O _m-X _iy _iz _i) as with reference to coordinate system: initial point be a moon heart O _m, in nominal, land constantly, X _iaxle points to nominal landing point, Z _iaxle points to direction of motion, Y in the landing orbit plane of planning _iaxle forms right hand orthogonal coordinate system, as shown in Figure 2.If lander landed in nominal landing point with nominal attitude in the nominal time, three axles of land so inertial coordinates system and body coordinate system are parallel.

The initial value of navigational system comprises position, speed and attitude.Wherein position, speed are provided by ground observing and controlling, and attitude is obtained by star sensor and gyro to measure.It should be noted that, the coordinate system that the quick measurement of ground observing and controlling and star adopts may not be landing inertial coordinates system, for example month heart J2000 coordinate system.But because nominal landing point and nominal time of landing is known, therefore according to moon ephemeris just can be easily by these state transitions in landing inertial coordinates system.If position vector is r ⁱ, velocity is v ⁱ, relatively the land attitude matrix of inertial system of body is C _bi(corresponding attitude quaternion is q).

(2) attitude is determined algorithm

Attitude determines that the object of algorithm is to upgrade attitude matrix C _bi, it can obtain according to gyro to measure recursion.If the angular velocity of gyro to measure is ω ^b, Attitude kinematic function can be with quaternion representation:

$\stackrel{\·}{q}=\frac{1}{2}\mathrm{Eq}\left(q\right){\mathrm{\ω}}^b---\left(1\right)$

Wherein

$\mathrm{Eq}\left(q\right)=\left[\begin{array}{ccc}{q}_4& -q_3& q_2\\ q_3& q_4& -q_1\\ -q_2& q_1& q_4\\ -q_1& -q_2& -q_3\end{array}\right]---\left(2\right)$

(1) formula is carried out to the renewal that integration can complete hypercomplex number, further can calculate attitude matrix:

$C_{\mathrm{bi}}=\left[\begin{array}{ccc}{q}_1^2-{\mathrm{<figure-callout\; id="2"\; label="q"\; filenames="HDA0000447199320000021.png,HDA0000447199320000032.png"\; state="\{\{state\}\}">q</figure-callout>}}_2^2-q_3^2+{\mathrm{<figure-callout\; id="4"\; label="q"\; filenames="HDA0000447199320000021.png"\; state="\{\{state\}\}">q</figure-callout>}}_4^2& 2(q_1{\mathrm{<figure-callout\; id="2"\; label="q"\; filenames="HDA0000447199320000021.png,HDA0000447199320000032.png"\; state="\{\{state\}\}">q</figure-callout>}}_2+q_3+q_4)& 2(q_1{q}_3-q_2{q}_4)\\ 2(q_1{q}_2-q_3{q}_4)& -q_1^2+{\mathrm{<figure-callout\; id="2"\; label="q"\; filenames="HDA0000447199320000021.png,HDA0000447199320000032.png"\; state="\{\{state\}\}">q</figure-callout>}}_2^2-q_3^2+{\mathrm{<figure-callout\; id="4"\; label="q"\; filenames="HDA0000447199320000021.png"\; state="\{\{state\}\}">q</figure-callout>}}_4^2& 2(q_2{q}_3+q_1{q}_4)\\ 2(q_1{q}_3+q_2{q}_4)& 2(q_2{q}_3-q_1{q}_4)& -q_1^2-{\mathrm{<figure-callout\; id="2"\; label="q"\; filenames="HDA0000447199320000021.png,HDA0000447199320000032.png"\; state="\{\{state\}\}">q</figure-callout>}}_2^2+q_3^2+{\mathrm{<figure-callout\; id="4"\; label="q"\; filenames="HDA0000447199320000021.png"\; state="\{\{state\}\}">q</figure-callout>}}_4^2\end{array}\right]---\left(3\right)$

(3) position and speed recursion, utilize the speed of the speed increment accumulation calculating current time lander that accelerometer measures obtains, and integration obtains the position of lander under inertial coordinates system

The inertial navigation fundamental equation of describing under landing inertial coordinates system is

${\stackrel{\&CenterDot;\&CenterDot;}{r}}^I=g^I+f^I---\left(4\right)$

Wherein, g ⁱfor moon's gravity acceleration, in the situation that ignoring the non-ball Gravitational perturbation of the moon, have

$g^I={\mathrm{\&mu;}}_m\frac{r^I}{r^3}---\left(5\right)$

Wherein, μ _mfor lunar gravitation constant.F ⁱbe the non-gravitational acceleration (specific force) representing under inertial system, it can obtain according to accelerometer measures.But because accelerometer measures means under body series, i.e. f ^b, therefore need to change according to attitude matrix.

$f^I=C_{\mathrm{ib}}{f}^b=C_{\mathrm{bi}}^T{f}^b---\left(6\right)$

Due to

therefore (4) formula is carried out to two repeated integrals just can obtain respectively the velocity v of lander ⁱwith position vector r ⁱ.

After obtaining position, can also calculate the height of relative lunar surface

h=||r ^I||-r _M (7)

R _mfor r ⁱcorresponding reference lunar surface height.

Second step, the correction of finding range and test the speed (independent navigation filtering)

Utilization is tested the speed, range radar can be revised inertial navigation recursion position and speed.Range finding sensor provides line-of-sight distance, therefore utilizes attitude information and surveys sensor beam direction (direction of visual lines), line-of-sight distance can be converted into the height of the relative lunar surface of lander.The sensor that tests the speed obtains be the speed of the relative lunar surface of lander at the component of beam direction, therefore utilize attitude information and the sensor that tests the speed is installed, and according to detector position the compensation geostrophic ground velocity of landing on the moon, just can obtain the velocity inertial of lander.The information of range finding and the sensor acquisition of testing the speed is compared with position, the velocity information of inertial reference calculation and obtained new breath, just can to the position of inertial navigation, velocity error, revise by Kalman filtering.

When reality is used, owing to testing the speed, finding range sensor Measuring Time inconsistent, be subject to software on star again and calculate quantitative limitation, the EKF in navigation algorithm realizes by matching and optimization filter correction coefficient, and tests the speed, finds range to revise and separately carry out.

(1) range finding is revised

According to measuring the height of the relative lunar surface reference surface of lander constantly and the height that range finding sensor data is converted to, computation and measurement residual error; According to optimizing correction factor, carry out position correction, and utilize the current gravitational acceleration of position calculation of revising.Specific algorithm is:

A. navigation data extrapolation

Because the time of ranging information is often asynchronous with the time of inertial reference calculation, therefore, need to utilize navigation data to extrapolate synchronizing time point.If time corresponding to ranging information is t _u, the time that the most approaching first two steps inertial navigation is calculated is with it t _n-2and t _n-1, corresponding position calculation value is with

utilize the t of inertial navigation prediction _ulander position is constantly

$r_u^I=r_{n-1}^I+\frac{r_{n-1}^I-r_{n-2}^I}{t_{n-1}-t_{n-2}}(t_u-t_{n-1})---\left(\text{8}\right)$

B. error is calculated

The positional information of being calculated by inertial navigation can obtain the height of relative lunar surface

$h_{u,\mathrm{ins}}=\left|\right|r_u^I\left|\right|-r_M---\left(9\right)$

R _mfor

corresponding reference lunar surface height.

If range finding sensor wave beam is oriented in the installation of body series

utilize attitude information it can be forwarded in landing inertial coordinates system

$u_R^I=C_{\mathrm{ib}}{u}_R^b---\left(10\right)$

The ρ that is measured as of sensor if find range _u, can calculate highly and be

$h_{u,\mathrm{radar}}={\mathrm{\&rho;}}_u\frac{<u_R^I\&CenterDot;-r_u^I>}{\left|\right|r_u^I\left|\right|}---\left(11\right)$

The error that thus, can obtain high computational is

δh _u=h _u,radar-h _u,ins (12)

C. filter correction

Get altitude correction factor w _hwith highly declining and increase gradually,

$w_h=w_h^{*}(1-h/h_h^{*})---\left(13\right)$

with

for given parameters, can adjust generally according to different task features

value between 0～1, therefore $0\&le;w_h\&le;w_h^{*}.$

Then inertial navigation is carried out to height correction

r ^I=r ^I+w _h·δh _u·r ^I/||r ^I|| (14)

(2) correction of testing the speed

I. utilize the algorithm of elevation information to the correction of gravity direction speed

Utilize the height of the relative lunar surface of oblique distance information clearing lander that twice distance measurement beam does not measure respectively in the same time, and the speed component of the gravity direction of difference acquisition lander, computing velocity is measured residual error, and carries out speed correction according to optimizing correction factor.Specific algorithm is:

A. navigation data extrapolation

The same while processing with ranging data, need to measure t constantly according to the Data Extrapolation of the first two steps inertial navigation sensor that tests the speed _uspeed.If t _n-2and t _n-1speed calculated value is constantly

with

$v_u^I=v_{n-1}^I+\frac{v_{n-1}^I-v_{n-2}^I}{t_{n-1}-t_{n-2}}(t_u-t_{n-1})---\left(15\right)$

B. error is calculated

The speed that inertial navigation calculates is that relative inertness is speed, is then converted into the speed of relative lunar surface.

$v_{g,\mathrm{ins}}^I=v_u^I-{\mathrm{\&omega;}}_m^I\&times;r_u^I---\left(16\right)$

Its component at gravity direction is

$v_{h,\mathrm{ins}}\left(t_k\right)=\frac{<r^I\&CenterDot;v_{g,\mathrm{ins}}^I>}{\left|\right|r^I\left|\right|}---\left(17\right)$

If t _k-1the moment and t _kthe oblique distance information that distance measurement beam records respectively is constantly ρ (t _k-1) and ρ (t _k), then in conjunction with the sensing of distance measurement beam and attitude and the positional information in the corresponding moment, calculate height.With t _kbe example constantly, if the installation of distance measurement beam under body series is oriented to p ^b, it is at the sensing p of inertial system ⁱ(t _k) be

p ^I(t _k)=C _ib(t _k)p ^b (18)

According to range finding sensor distance measurement information, can calculate height so, be designated as

$h_{\mathrm{radar}}\left(t_k\right)=\frac{<p^I\left(t_k\right),r^I\left(t_k\right)>}{\left|\right|r^I\left(t_k\right)}\mathrm{\&rho;}\left(t_k\right)---\left(19\right)$

Equally can be according to t _k-1ranging information computed altitude h constantly _radar(t _k-1).Difference acquisition lander in the speed of gravity direction is so

$v_{h,\mathrm{radar}}\left(t_k\right)=\frac{h_{\mathrm{\&rho;}}\left(t_k\right)-h_{\mathrm{\&rho;}}\left(t_{k-1}\right)}{t_k-t_{k-1}}---\left(20\right)$

The difference that thus, can obtain the gravity direction speed being calculated by two kinds of means of different is

δv _h=v _h,radar-v _h,ins (21)

C. filter correction

Get speed correction factor w _vwith Speed Reduction, increase gradually,

$w_v=w_v^{\&CenterDot;}(1-|\left|v^I\right||/v_v^{*})---\left(19\right)$

with

for given parameters, can adjust according to different task features, general

value between 0～1, therefore

Then the speed of inertial navigation short transverse is revised

v ^I=v ^I+w _v·δv _h·r ^I/||r ^I|| (20)

II. utilize the algorithm of sensor to the correction of beam direction speed that test the speed

Utilize the speed of the relative lunar surface of lander of knotmeter measurement and the speed component of the beam direction that tests the speed of inertial reference calculation, computing velocity is measured residual error, and carries out speed correction according to optimizing correction factor.Specific algorithm is:

A. navigation data extrapolation

Utilize t _n-2and t _n-1the speed calculated value of inertial navigation is constantly

with

according to (15) formula extrapolation corresponding lander velocity inertial constantly of sensor measurement data that tests the speed.

B. error is calculated

The speed that need to be first relative lunar surface by the rate conversion of inertial reference calculation according to (16) formula, is then projected to beam direction, and specific algorithm is as follows.

If velocity radar wave beam is oriented to body series

utilize attitude information it can be forwarded in landing inertial coordinates system

$u_v^I=C_{\mathrm{ib}}{u}_v^b---\left(21\right)$

The speed of inertial reference calculation at the component of radar beam direction is

$v_{u,\mathrm{ins}}=<v_{g,\mathrm{ins}}^I\&CenterDot;u_v^I>---\left(22\right)$

The velocity survey of the actual acquisition of velocity radar is v _{u, radar}, therefore, both differences are

δv _u=v _u,radar-v _u,ins (23)

C. filter correction

According to (19) formula, get speed correction factor and increase gradually with Speed Reduction, inertial navigation is carried out to speed correction

$v^I=v^I+w_v\&CenterDot;\mathrm{\&delta;}{v}_u\&CenterDot;u_v^I---\left(20\right)$

(3) correction strategy

Above-mentioned range finding and the correction algorithm that tests the speed are all to carry out for single beam, and current time is as long as any one wave beam effectively just can be for the correction of navigating.In makeover process, range finding revise mainly for be altitude channel site error, test the speed revise mainly for be the velocity error of beam direction.For soft landing task, elevation information is the most important, the safety that direct relation lands; And the accuracy requirement of horizontal position information is not high especially, the position correction that therefore adopts the present invention to carry out is suitable.For the correction of testing the speed, horizontal and vertical speed is all quite important, and it is directly connected to the inclined to one side extent of landing speed, affects task success or failure.Although speed correction of the present invention, mainly along beam direction, as long as there is the wave beam of three different directions to measure, just can complete the correction of three-dimensional velocity.Therefore, this algorithm is also simple and effective.

Simulation analysis.No. three moon landing devices of the goddess in the moon of take carry out simulation analysis as example.Simulated conditions is as follows:

● preliminary orbit: perilune position and the speed of 100km * 15km track are highly 15km, horizontal velocity is that 1692m/s(perilune and apocynthion height all exist 1km stochastic error).

● initial attitude: the attitude angle of the inertial system of relatively landing is 0 ° of 110 ° of the angles of pitch, driftage and roll angle.

● initial navigation error: orbit error comprises site error 2km(R direction 0.45km, T direction 1.8km, N direction 0.6km), velocity error 2m/s(R direction 1.65m/s, T direction 1m/s, N direction 0.6m/s), 0.03 ° of attitude error.

● the tellurometer survey sensor totally five groups of antennas that test the speed, wherein find range two groups, test the speed three groups, be arranged on lander-Z face.Specifically be oriented to: the expression of the unit vector of each beam position of three groups of rate antennas (containing receiving antenna and emitting antenna) in lander body series is respectively R1=[cos133 °, cos90 °, cos137 °], R2=[cos154.3265 °, cos75.3330 °, cos110.5823 °], R3=[cos154.3265 °, cos104.6670 °, cos110.5823 °]; Two groups of telemetry antennas (containing receiving antenna and emitting antenna) beam position R4 and R5 respectively along lander body series-Z axis and-X-axis.

● correction factor is taken as $w_h^{*}=0.5,h_h^{*}=8000m,w_v^{*}=0.4,v_v^{*}=300m/s.$

Navigation strategy is: apart from lunar surface 8km, highly introduce ranging information, apart from lunar surface 3.5km, highly introduce the information of testing the speed; Apart from more than lunar surface height 8km and below 30m, only utilize inertial navigation.

Change in location curve under the landing inertial system providing with navigation that Fig. 3 has provided landing mission reality, the error of navigation as shown in Figure 4.Before the information of externally finding range as seen from Figure 4 and test the speed is introduced, inertial navigation location compute is dispersed; Below 8km, height (after about 360s) introducing of ranging information contributes to revise the site error of x direction (short transverse), but there is no capability for correcting for the site error of y, z direction (horizontal direction); Below 3.5km, (after the about 450s) information that tests the speed is introduced, and it can reduce the divergence speed of horizontal level to a certain extent by revising the speed of y, z direction, has improved the precision of navigation.

The change curve of the lander velocity inertial providing with navigation that Fig. 5 has provided power decline process reality, the error of navigation as shown in Figure 6.Data from Fig. 6, before Microwave Velocity data are introduced navigational system, the velocity error of inertial reference calculation increases gradually; After Microwave Velocity data are introduced, velocity error is revised, and has guaranteed the precision of navigation.

Non-elaborated part of the present invention belongs to techniques well known.

Claims (2)

1. the robust air navigation aid that multi-beam tests the speed and finds range and revise, is characterized in that performing step is as follows:

(1) inertial reference calculation

Utilize the inertia attitude of the angular speed calculation current time lander of gyro to measure in the previous sampling period, then utilize specific force that accelerometer measures obtains to calculate the speed of current time lander, and integration obtains the position of lander under inertial coordinates system;

(2) range finding is revised

If receive the valid data of range finding sensor, and flying height meets the introducing scope that range finding is revised, by these data by matrix being installed and the conversion of lander attitude matrix solves altitude information; Utilize the inertial reference calculation in the first two cycle simultaneously, calculate range finding sensor data lander height on corresponding time point, it is poor that the height that inertial navigation and stadimeter are resolved is respectively done, and according to the site error of altitude correction factor corrected altitude direction; This altitude correction factor changes with lander highly linear, and lander height is higher, and described altitude correction factor is less, and highly more low clearance correction factor is larger;

(3) speed correction

The signal source of speed correction is divided into two kinds, and a kind of is distance measuring signal, and another kind is tachometer signal, when distance measuring signal effective, and flying height meets while testing the speed correction introducing scope, the height that double distance measuring signal is solved carries out difference, obtain the speed of gravity direction, utilize the corresponding lander velocity inertial constantly of speed calculation distance measuring signal in the first two cycle in inertial reference calculation simultaneously, and be converted to the speed of relative lunar surface, it is poor that the speed that the speed that distance measuring signal is obtained and inertial navigation obtain is done, and according to the speed of speed correction factor correction gravity direction, this speed correction factor is with lander speed linear change, lander speed is faster, speed correction factor is less, speed is less, speed correction factor is larger, when tachometer signal is effective, utilize the corresponding lander velocity inertial constantly of speed calculation tachometer signal in the first two cycle in inertial reference calculation, and be converted to the speed of relative lunar surface, it is poor that the speed that the speed that tachometer signal is obtained and inertial navigation obtain is done, and according to the speed of speed correction factor correction beam direction, this speed correction factor is with lander speed linear change, lander speed is faster, speed correction factor is less, and speed is less, and speed correction factor is larger,

(4) each control cycle is repeated to the step of (1)～(3), complete navigation calculation.

2. the robust air navigation aid that multi-beam according to claim 1 tests the speed and finds range and revise, is characterized in that: described altitude correction factor and speed correction factor are 0-1.

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