CN103591948B - Initial value synchronization method for improving landing navigation accuracy - Google Patents

Initial value synchronization method for improving landing navigation accuracy Download PDF

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Publication number
CN103591948B
CN103591948B CN201310485300.9A CN201310485300A CN103591948B CN 103591948 B CN103591948 B CN 103591948B CN 201310485300 A CN201310485300 A CN 201310485300A CN 103591948 B CN103591948 B CN 103591948B
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0new
landing
initial value
detector
moment
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CN103591948A (en
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张洪华
于萍
关轶峰
赵宇
梁俊
程铭
何健
王大轶
王鹏基
黄翔宇
李骥
杨巍
于洁
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Beijing Institute of Control Engineering
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Beijing Institute of Control Engineering
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass

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  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Manufacturing & Machinery (AREA)
  • Automation & Control Theory (AREA)
  • Position Fixing By Use Of Radio Waves (AREA)

Abstract

The invention discloses an initial value synchronization method for improving the landing navigation accuracy. According to the method, a starting moment of a landing and declining program is adjusted to be consistent with an initial moment of a control period of a satellite-borne computer, a navigation initial value is updated according to the initial moment of the control period of the satellite-borne computer, the synchronization among the navigation initial value, the starting moment of the landing and declining program and the initial moment of the control period is ensured, and the navigation accuracy in the landing and declining process is improved. A moment alignment algorithm of the method is simple, reliable and easy to realize, a large amount of operation of timing, calculation and the like is not required to be performed on the ground, and only the track extrapolation calculation is required to be autonomously performed on a satellite.

Description

A kind of initial value synchronous method that improves landing navigation precision
Technical field
The present invention relates to a kind of initial value synchronous method that improves landing navigation precision, be applicable to the navigation task demand of space probe objects outside Earth landing mission.
Background technology
Landing navigation precision is most important to guaranteeing landing safety.The accuracy of navigation initial value is to guarantee the condition precedent of landing navigation precision.
Due to space probe, all dispose the inertial attitude sensors such as the quick and gyro of star, therefore attitude information can independently be obtained by detector, and detector is with respect to the translation information of specific celestial body, as position and speed etc., detector does not possess the ability of these translation information of active obtaining (position and speed), need to measure be determined by ground observing and controlling system, and in the up spaceborne computer that is injected into detector.
Detector, before starting landing procedure, first needs to receive the position and speed information of the landing procedure Startup time being injected by land station, as navigation initial value.At landing procedure Startup time, detector spaceborne computer gathers the measurement data of star quick and gyro (obtaining the inertia attitude information of detector), accelerometer (obtaining the specific force acceleration information of detector), find range sensor (obtaining the elevation information of the relative lunar surface of detector), the sensor that tests the speed (obtaining the speed of the relative lunar surface of detector), the position and speed information of the detector that ground injects of take is navigation initial value, the calculating of navigating.
The accuracy of navigation initial value, had both comprised accuracy, the accuracy in the corresponding moment of position and speed information of position and speed information, also comprised the accuracy of landing letdown procedure Startup time.Guarantee that both synchronously start, very important to improving landing navigation precision.
The spaceborne computer of detector carries out discrete control according to control cycle, under this mechanism, if simple, adopting and judge that the method whether current time is greater than landing letdown procedure Startup time starts landing letdown procedure, is that the letdown procedure Startup time that can not guarantee to land strictly alignd with the navigation initial value moment of ground injection.Due to detector, speed is larger in orbit, as calculated, and when control cycle is 128ms (representative value in Spacecraft Control cycle), as calculated only because the mismatch of time will cause landing latter stage vertical height error just has nearly hundred meters.
Summary of the invention
The technical matters that the present invention solves is: overcome the deficiencies in the prior art, a kind of initial value synchronous method that improves landing navigation precision is provided, solved due to the navigation initial value moment navigation error problem that cause asynchronous with landing letdown procedure Startup time.
Technical scheme of the present invention is: a kind of initial value synchronous method that improves landing navigation precision, and step is as follows:
1), before detector starts landing letdown procedure, by ground control station, to detector, inject landing letdown procedure Startup time t 0and t 0the positional information r of moment detector 0with velocity information v 0;
2) the landing letdown procedure Startup time t injecting according to ground 0, and current control cycle Startup time t, obtain new landing letdown procedure Startup time t 0new;
3) the positional information r that the spaceborne computer on detector obtains with step 1) 0with velocity information v 0for initial value, with t 0new-t 0for orbit integration duration, adopt the method for track extrapolation, obtain t 0newthe positional information r of moment detector 0newwith velocity information v 0new;
4) spaceborne computer on detector is with t 0newfor the control cycle of initial time starts landing letdown procedure temporarily, spaceborne computer is with t 0newthe positional information r of moment detector 0newwith velocity information v 0newfor initial value, the navigation that starts the process of declining is calculated.
Described step 2) in, obtain new landing procedure Startup time t 0newconcrete grammar be: (t 0new-t)=n * T, and t 0new>t 0, (t 0new-T) <t 0; Wherein n is positive integer, the control cycle that T is spaceborne computer.
The present invention's advantage is compared with prior art:
1) the present invention has guaranteed " landing procedure " Startup time and the accurate corresponding of the initial value moment of navigating, and has avoided due to navigation initial value and the asynchronous navigation error causing of the moment.
2) moment alignment algorithm of the present invention, simple and reliable, is easy to realize.
3) when method of the present invention does not need ground to carry out a large amount of schools again, calculating does not wait work, only needs independently to carry out track extrapolation on star and calculates.
Accompanying drawing explanation
Fig. 1 is initial value asynchronous error delta t schematic diagram constantly;
Fig. 2 is the velocity error schematic diagram that the constantly asynchronous error of initial value causes.
Fig. 3 is the inventive method process flow diagram.
Embodiment
The concrete steps of the inventive method are as shown in Figure 3:
1), before detector starts landing letdown procedure, by ground control station, to detector, inject landing letdown procedure Startup time t 0and t 0the positional information r of moment detector 0with velocity information v 0;
2) the landing letdown procedure Startup time t injecting according to ground 0, and current control cycle Startup time t, calculate and obtain new landing letdown procedure Startup time t 0new, concrete grammar is: (t 0new-t)=n * T, and t 0new>t 0, (t 0new-T) <t 0; Wherein n is positive integer, the control cycle that T is spaceborne computer
3) the positional information r that the spaceborne computer on detector obtains with step 1) 0with velocity information v 0for initial value, with t 0new-t 0for orbit integration duration, adopt the method for track extrapolation, obtain t 0newthe positional information r of moment detector 0newwith velocity information v 0new; Concrete acquisition methods can be referring to: < < satellite orbit and attitude dynamics with control > >, Zhang Renwei, BJ University of Aeronautics & Astronautics, 1998.
4) spaceborne computer on detector is with t 0newfor the control cycle of initial time starts landing letdown procedure temporarily, spaceborne computer is with t 0newthe positional information r of moment detector 0newwith velocity information v 0newfor initial value, the navigation that starts the process of declining is calculated.
Give an example this method be specifically described below:
Take Lunar Landing Mission as example.Lunar orbiter when lunar orbit moves around the moon speed be about 1.7km/s(representative value).If it is synchronous not carry out initial value, t 0newwith t 0mistiming Δ t limiting case under be likely 1 control cycle.Referring to Fig. 1.
Due to the existence of Δ t, when detector starts landing letdown procedure, its speed v 1 is with respect to t 0there is variation in speed v 0 constantly.The vertical velocity error delta v that can obtain as calculated initial navigation is about 0.2m/s.Referring to Fig. 2.
Because the representative value of landing mission required time is greater than 600s, due to approximately 100 meters of the initial value vertical errors that asynchronous error delta t causes constantly.
The present invention not detailed description is known to the skilled person technology.

Claims (2)

1. improve an initial value synchronous method for landing navigation precision, it is characterized in that step is as follows:
1), before detector starts landing letdown procedure, by ground control station, to detector, inject landing letdown procedure Startup time t 0and t 0the positional information r of moment detector 0with velocity information v 0;
2) the landing letdown procedure Startup time t injecting according to ground 0, and current control cycle Startup time t, obtain new landing letdown procedure Startup time t 0new;
3) the positional information r that the spaceborne computer on detector obtains with step 1) 0with velocity information v 0for initial value, with t 0new-t 0for orbit integration duration, adopt the method for track extrapolation, obtain t 0newthe positional information r of moment detector 0newwith velocity information v 0new;
4) spaceborne computer on detector is with t 0newfor the control cycle of initial time starts landing letdown procedure temporarily, spaceborne computer is with t 0newthe positional information r of moment detector 0newwith velocity information v 0newfor initial value, the navigation that starts the process of declining is calculated.
2. a kind of initial value synchronous method that improves landing navigation precision according to claim 1, is characterized in that: described step 2), obtain new landing procedure Startup time t 0newconcrete grammar be: (t 0new-t)=n * T, and t 0new>t 0, (t 0new-T) <t 0; Wherein n is positive integer, the control cycle that T is spaceborne computer.
CN201310485300.9A 2013-10-16 2013-10-16 Initial value synchronization method for improving landing navigation accuracy Active CN103591948B (en)

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Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101074881A (en) * 2007-07-24 2007-11-21 北京控制工程研究所 Inertial navigation method for moon detector in flexible landing stage
WO2008054482A2 (en) * 2006-03-24 2008-05-08 Dula Arthur M Solar system positioning system
CN101219713A (en) * 2007-12-26 2008-07-16 北京控制工程研究所 Satellitic self-determination orbital transfer method
CN102116628A (en) * 2009-12-31 2011-07-06 北京控制工程研究所 High-precision navigation method for landed or attached deep sky celestial body detector
CN103335654A (en) * 2013-06-19 2013-10-02 北京理工大学 Self-navigation method for planetary power descending branch

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP5329409B2 (en) * 2006-08-11 2013-10-30 シエラ・ネバダ・コーポレイション A method for fusing multiple GPS measurement types into a weighted least squares solution

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2008054482A2 (en) * 2006-03-24 2008-05-08 Dula Arthur M Solar system positioning system
CN101074881A (en) * 2007-07-24 2007-11-21 北京控制工程研究所 Inertial navigation method for moon detector in flexible landing stage
CN101219713A (en) * 2007-12-26 2008-07-16 北京控制工程研究所 Satellitic self-determination orbital transfer method
CN102116628A (en) * 2009-12-31 2011-07-06 北京控制工程研究所 High-precision navigation method for landed or attached deep sky celestial body detector
CN103335654A (en) * 2013-06-19 2013-10-02 北京理工大学 Self-navigation method for planetary power descending branch

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