CN106990426A - A kind of air navigation aid and guider - Google Patents
A kind of air navigation aid and guider Download PDFInfo
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- CN106990426A CN106990426A CN201710158060.XA CN201710158060A CN106990426A CN 106990426 A CN106990426 A CN 106990426A CN 201710158060 A CN201710158060 A CN 201710158060A CN 106990426 A CN106990426 A CN 106990426A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
- G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/42—Determining position
- G01S19/48—Determining position by combining or switching between position solutions derived from the satellite radio beacon positioning system and position solutions derived from a further system
- G01S19/49—Determining position by combining or switching between position solutions derived from the satellite radio beacon positioning system and position solutions derived from a further system whereby the further system is an inertial position system, e.g. loosely-coupled
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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
- G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
- G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
- G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
- G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
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Abstract
This application discloses a kind of air navigation aid and device, method comprises the following steps：DVB gathers position and speed；Micro inertial measurement unit is by measurement period collection acceleration, gyro instantaneous value；Position, speed and the attitude parameter calculated value of each measurement period are drawn with strap inertial navigation algorithm；According to filtering cycle, DVB outgoing position and speed calibration value；Pitching angle error, roll angle error, course angle error, east orientation speed error, north orientation speed error, trueness error, latitude error, x, y, z direction gyro error, x, y directional acceleration error totally 12 parameters are calculated with Kalman Filtering for Discrete method.The device of the application includes micro inertial measurement unit, DVB, processor；Processor further includes navigational parameter initialization unit, strap-down inertial computing unit, estimation error wave filter.Cost of the present invention is low, small volume, precision high.
Description
Technical field
The application unmanned plane field, more particularly to a kind of use satellite navigation and the machine inertia measurement means navigation method and
Device.
Background technology
Every kind of single navigation system has respective special performance and limitation.Miniature fiber or micro-mechanical inertia navigation dress
Be equipped with low cost, small volume, lightweight advantage, can be independently provide posture, position, velocity information for carrier；Have the disadvantage
Navigation error increases fast with the time, particularly the index error such as the drift of MEMS inertia devices, bias repeatability, it is difficult to by marking in advance
School reaches the target for improving inertial navigation precision.
Beidou satellite navigation system is the satellite navigation system that China voluntarily develops, and it can be in states such as GPS, GLONASS
Outer navigation system provides navigation, positioning, time service service in the case of closing for China user.Current dipper system has been formd
5GEO+5IGSO+5MEO constellation framework, it is possible to achieve positioning, navigation and the time service of the Asian-Pacific area, positioning precision 10m, tests the speed
Precision 0.2m/s, time service precision 50ns.Satellite navigation has the high advantage of round-the-clock, navigation accuracy, but is restricted by use environment,
The poor availability under interference and obstruction conditions, and attitude of carrier information can not be provided for carrier flight control.
Accordingly, it would be desirable to a kind of integrated navigation system method and system be designed, by mini inertia measurement unit (MIMU)
Composition is combined with Beidou satellite receiver a kind of can provide complete carrier movement parameter, navigation accuracy higher system.
The content of the invention
The present invention discloses a kind of air navigation aid and device, and solution utilizes micro inertial measurement unit (MIMU) under the conditions of not demarcating
The problem of navigation.
The embodiment of the present application provides a kind of air navigation aid, for unmanned plane, the unmanned plane comprising micro inertial measurement unit,
DVB, comprises the following steps：
In initial time, unmanned plane position and speed are gathered with the DVB, respectively as position initial value and
Velocity original value；
From the initial time to the calibration moment, accelerated with the micro inertial measurement unit by measurement period collection unmanned plane
Spend instantaneous value, gyro instantaneous value；Attitude parameter instantaneous value is calculated according to acceleration instantaneous value, gyro instantaneous value；During the calibration
It is filtering cycle to carve between the initial time, includes multiple measurement periods；
According to the position initial value, velocity original value, attitude parameter instantaneous value, nothing is drawn with strap inertial navigation algorithm
Man-machine position calculated value, speed calculated value and attitude parameter calculated value in each measurement period；
At the calibration moment, unmanned plane position and speed are gathered with the DVB, respectively as position correction value and
Speed calibration value；
According to the position calculated value at calibration moment, speed calculated value, position correction value, speed calibration value, discrete karr is used
Graceful filtering method calculates attitude parameter error, velocity error, site error, gyro error, acceleration error；The discrete karr
The error variance that graceful filtering method is used is included：Pitching angle error, roll angle error, course angle error, east orientation speed error, north
To velocity error, trueness error, latitude error, x, y, z direction gyro error, x, y directional acceleration error totally 12 parameters；Its
In, unmanned plane right side side is pointed in x, y, z form right angle coordinate system, x directions, before y directions are pointed to along unmanned plane y direction.
Further, using the calibration moment as new initial time, circulation performs the step.
The most preferred embodiment of the application is also comprised the steps of：Acceleration with the acceleration error to next filtering cycle
Degree instantaneous value is modified；The gyro instantaneous value of next filtering cycle is modified with the gyro error.
The embodiment of the present application also provides a kind of guider, for unmanned plane, includes micro inertial measurement unit, satellite reception
Machine, processor；The micro inertial measurement unit, for since initial time, producing gyro instantaneous value according to measurement period, adding
Instantaneous state of velocity；The DVB, for receiving satellite positioning information, in initial time outgoing position initial value and speed
Initial value；According to filtering cycle outgoing position calibration value and speed calibration value from initial time；The processor, for basis
The position initial value, velocity original value, gyro instantaneous value, acceleration instantaneous value, nobody is drawn with strap inertial navigation algorithm
Position calculated value, speed calculated value and attitude parameter calculated value of the machine in each measurement period；According to the position at calibration moment
Calculated value, speed calculated value, position correction value, speed calibration value, with Kalman Filtering for Discrete method calculate attitude parameter error,
Velocity error, site error, gyro error, acceleration error；The error variance bag that the Kalman Filtering for Discrete method is used
Contain：Pitching angle error, roll angle error, course angle error, east orientation speed error, north orientation speed error, trueness error, latitude are missed
Difference, x, y, z direction gyro error, x, y directional acceleration error totally 12 parameters；Wherein, x, y, z form right angle coordinate system, x side
To sensing unmanned plane right side side, before y directions are pointed to along unmanned plane y direction.
Further, the processor further comprising navigational parameter initialization unit, strap-down inertial computing unit,
Estimation error wave filter；The navigational parameter initialization unit, adds for receiving acceleration instantaneous value, gyro instantaneous value, basis
Instantaneous state of velocity, gyro instantaneous value calculate attitude parameter instantaneous value；The strap-down inertial navigation elements, it is described for receiving
Position initial value, velocity original value, attitude parameter instantaneous value, show that unmanned plane is measured at each with strap inertial navigation algorithm
Position calculated value, speed calculated value and the attitude parameter calculated value in cycle；The estimation error wave filter, is filtered with discrete Kalman
Wave method calculates attitude parameter error, velocity error, site error, gyro error, acceleration error.
Preferably, the guider further includes output calibration unit；The output calibration unit, uses the posture
Parameter error is modified to the attitude parameter calculated value.
In any one embodiment of the present invention, it is preferable that the guider further comprising input correction unit, is used
The acceleration error is modified to the acceleration instantaneous value of next filtering cycle, filters week to next with the gyro error
The gyro instantaneous value of phase is modified, the input of the navigational parameter initialization unit.
At least one above-mentioned technical scheme that the embodiment of the present application is used can reach following beneficial effect：
Other integrated navigation system implementation methods are contrasted, the design has complete data handling procedure, using the present invention
The integrated navigation based on micro inertial measurement unit (MIMU), Beidou satellite receiver can be realized well.
Mierotubule-associated proteins error of the present invention need not be demarcated in advance, on the one hand periodically defeated using Beidou satellite receiver
Position, speed and the speed course initialization inertial navigation calculating parameter gone out, another aspect estimation error wave filter is estimated in real time
Meter inertial navigation error is simultaneously modified, and may finally obtain the available integrated navigation result of precision.
The present invention is using miniature fiber or micro-mechanical inertia measuring unit compared with traditional laser or fiber-optic inertial unit
Cost is lower, volume is smaller, weight is lighter；The present invention can make integrated navigation system in other satellites using Beidou satellite receiver
Still normal work in the case that navigation system (GPS, GLONASS etc.) is closed.
Brief description of the drawings
Accompanying drawing described herein is used for providing further understanding of the present application, constitutes the part of the application, this Shen
Schematic description and description please is used to explain the application, does not constitute the improper restriction to the application.In the accompanying drawings：
Fig. 1 is the embodiment flow chart of air navigation aid of the present invention；
Fig. 2 is the embodiment flow chart that air navigation aid of the present invention includes circulation calibration；
Fig. 3 is the embodiment flow chart that estimation error is carried out with Kalman Filtering for Discrete method；
Fig. 4 is the embodiment block diagram of guider of the present invention；
Fig. 5 is the embodiment block diagram that guider of the present invention further optimizes；
Fig. 6 is the embodiment schematic diagram that guider of the present invention was inputted, exported calibration.
Embodiment
To make the purpose, technical scheme and advantage of the application clearer, below in conjunction with the application specific embodiment and
Technical scheme is clearly and completely described corresponding accompanying drawing.Obviously, described embodiment is only the application one
Section Example, rather than whole embodiments.Based on the embodiment in the application, those of ordinary skill in the art are not doing
Go out the every other embodiment obtained under the premise of creative work, belong to the scope of the application protection.
In the application whole embodiments, unmanned plane right side side, y directions are pointed in x, y, z form right angle coordinate system, x directions
Before being pointed to along unmanned plane y direction.
Below in conjunction with accompanying drawing, the technical scheme that each embodiment of the application is provided is described in detail.
Fig. 1 is the embodiment flow chart of air navigation aid of the present invention.Specifically comprise the steps of：
Step 11, in initial time, unmanned plane position and speed are gathered with the DVB, respectively as at the beginning of position
Initial value and velocity original value；
Step 12, from the initial time to calibration the moment, with the micro inertial measurement unit by measurement period gather nothing
Man-machine acceleration instantaneous value, gyro instantaneous value；Attitude parameter instantaneous value is calculated according to acceleration instantaneous value, gyro instantaneous value；Institute
The when a length of filtering cycle between calibration moment and the initial time is stated, multiple measurement periods are included；
Wherein, in first initial time, the computational methods of the attitude parameter instantaneous value are：
pitch0=arcsin (ay/g0) formula 1.1
roll0=-arcsin (ax/g0) formula 1.2
Pitch in formula0、roll0For initial pitch angle, roll angle；ax、ayFor accelerometer x, y directions output valve；g0Attach most importance to
Power acceleration.
For example, gathering gyro, the accelerometer information of micro inertial measurement unit (MIMU) with measurement period T1 (20ms).
Controlled it should be noted that inertial navigation exports attitude of carrier for attitude of carrier, it is desirable to which real-time is high, therefore sets
Meter inertial navigation calculates for short cycle T 1 (being 20ms)；Increase inertial navigation output result (posture, position, speed with the time of calculating
Degree) error increase is, it is necessary to which periodicity is aided in using the information of Beidou satellite receiver to improve precision, and this method embodiment is used
Filtering cycle T2 (being 1s).
In step 12, for example can with the longitude of Beidou satellite receiver, latitude, height, ground velocity course and it is described just
The beginning angle of pitch, initial roll angle, calculate UAV Attitude, speed, position initial value；
Step 13, according to the position initial value, velocity original value, attitude parameter instantaneous value, calculated with strap-down inertial
Method draws position calculated value, speed calculated value and attitude parameter calculated value of the unmanned plane in each measurement period；
Location parameter includes longitude, latitude, height, and speed parameter includes east orientation speed, north orientation speed, sky orientation speed, appearance
State parameter includes course, the angle of pitch, roll angle；
Step 14, by the filtering cycle, at the calibration moment, unmanned plane position and speed are gathered with the DVB
Degree, respectively as position correction value and speed calibration value；
For example, gathering information (longitude, latitude, height, the east orientation speed of Beidou satellite receiver with filtering cycle T2 (1s)
Degree, north orientation speed, sky orientation speed, ground velocity course).
Step 15, according to the calibration position calculated value at moment, speed calculated value, position correction value, speed calibration value, with from
Dissipate kalman filter method and calculate attitude parameter error, velocity error, site error, gyro error, acceleration error, it is described from
The error variance that scattered kalman filter method is used is included：Pitching angle error, roll angle error, course angle error, east orientation speed
Error, north orientation speed error, trueness error, latitude error, x, y, z direction gyro error, x, y directional acceleration error totally 12
Parameter.
Step 16, using the error variance of 12 parameters to the position calculated value, speed calculated value and attitude parameter meter
Calculation value is calibrated；
For example when DVB, to participate in position location satellite number more, when geometric dilution of precision is less than threshold value, integrated navigation day
To speed, height directly output satellite receiver sky orientation speed, height；Otherwise integrated navigation exports micro inertial measurement unit
(MIMU) navigation sky orientation speed, height.
In the application items embodiment, the DVB receives navigation signal, specifically, applied to big-dipper satellite,
Gps satellite or other navigation satellites.When application big-dipper satellite, the DVB is the literary star receiver of the Big Dipper.
Fig. 2 is the embodiment flow chart that air navigation aid of the present invention includes circulation calibration.
The most preferred embodiment of the application is also comprised the steps of：
Step 17, with the acceleration error acceleration instantaneous value of next filtering cycle is modified；Use the top
Spiral shell error is modified to the gyro instantaneous value of next filtering cycle.Gyro that each Kalman filtering cycle (T2) obtains plus
Speedometer error estimate is to next cycle gyro, accelerometer initial data amendment.
Further, need to reinitialize when recapturing unlocked, repeat step 11~17.
When the continuous tracking satellite positioning signal of the DVB, further, using the calibration moment as new
Initial time, circulation performs step 13~17.
By said process, Mierotubule-associated proteins error is periodically defeated using Beidou satellite receiver without demarcating in advance
Position, speed and the speed course initialization inertial navigation calculating parameter gone out, and design error estimation filter are estimated in real time
Meter inertial navigation error is simultaneously modified, and can finally give the available navigation results of precision.
Fig. 3 is the embodiment flow chart that estimation error is carried out with Kalman Filtering for Discrete method.Filtered using estimation error
Device carries out inertial navigation estimation error, it is considered to micro inertial measurement unit (MIMU) feature and operation time, and wave filter selection 12 is joined
Several error state variable, specifically, error state variable X element is made up of following physical quantity：Angle of pitch error delta φe、
Roll angle error delta φn, course angle error delta φu；East orientation speed error delta ve, north orientation speed error delta vn；Longitude error Δ λ,
Latitude error Δ L；X, y, z direction gyro error εbx, εby, εbz；X, y directional acceleration meter errorI.e.
Step 21, the 12 parameter error state set up according to the relation between inertial navigation system error, basic navigation parameter
Model.Specifically, error state model is
Wherein F is state-transition matrix, and dimension is 12 × 12, and nonzero term is：
F (0,1)=ωiesinL；F (0,2)=- ωiecosL； F (1,0)=- ωiesinL；F (1,5)=- ωiesinL；F (2,0)=ωiecosL； F (3,1)=- g；F
(4,0)=g；
In formula, ωieFor rotational-angular velocity of the earth, L is latitude, and R is earth radius,For attitude matrix (dimension is 3 ×
3), ve、vnFor east orientation speed, north orientation speed, g is acceleration of gravity, τgyroFor gyro correlation time, τaccIt is related for accelerometer
Time, in technical solution of the present invention, gyro, accelerometer correlation time value 5s.
There are a many kinds in inertial navigation system error source, such as site error, velocity error, initial attitude error, gyro bias plus
Discrepancies in speedometers, scale coefficient error etc., every error can be divided into droop and noise item again.It is inertial navigation system error, basic
The change that there is inertial navigation system error in approximate corresponding relation, fixed sample interval between navigational parameter can be expressed as coefficient
It is multiplied by the form that error term adds noise item.The corresponding error term coefficient of the error change amount of each in expression formula is to determine,
It can be drawn by inertial navigation error transfer relationship.Different according to the error state variable of selection, error state model is write as
Expression formula form is different.The error state model of the form of formula 2 can specifically be set up.
In error state model, G is noise matrix, and dimension is 12 × 1, is met
Wherein, Wgyro=[wgx wgy wgz]T, Wacc=[wax way waz]T, wgx、wgy、wgzFor gyroscope measurement noise,
wax、way、wazFor accelerometer measures noise.
Estimation error wave filter is calculated by kalman filter method, and calculating cycle is the filtering cycle T2 (such as 1s),
Each filtering cycle, which is calculated, obtains the step estimate of state one：
In formula,Respectively pitching angle error, roll angle error, course angle error estimate；Point
Wei not east orientation speed error estimate, north orientation speed error estimate；Respectively longitude, latitude error estimate；Respectively x, y, z direction gyroscope error estimation value；Respectively x, y directional acceleration meter estimation error
Value.
Step 22, with the step estimate of state oneAmendment is missed
Difference, updates attitude matrix.Attitude rectification model is
In formula,For current period attitude matrix,For attitude matrix after amendment, calculated for next cycle S5.
It is used for follow-up strapdown that current period attitude matrix can obtain revised attitude matrix through attitude misalignment amendment
Lead calculating.Attitude rectification is carried out according to the order of course angle error, pitching angle error, roll angle error.Revolved according to above-mentioned angle
Turn attitude matrix after can sequentially being corrected.Attitude matrix and current period attitude matrix, coefficient matrix after specific amendment
Relational expression is as irised out shown in part.
Step 23, according to speed, position correction model, the step estimate of use state one is exported to micro inertial measurement unit
East orientation speed, north orientation speed, latitude, longitude are modified.
Set up estimation error wave filter observation model such as following formula：
The error observation model set up in formula 6 is using velocity location difference as observed quantity, micro inertial measurement unit
(MIMU) the position and speed difference of the speed, position and the Beidou satellite receiver output that export can be expressed as ins error and add survey
Measure the form of noise.
Wherein ve_INS、vn_INS、LINS、λINSTo obtain speed calculated value and position meter according to micro inertial measurement unit output
The east orientation speed of calculation value, north orientation speed, latitude, longitude ve_BD、vn_BD、LBD、λBDThe east orientation speed exported for Beidou satellite receiver
Degree, north orientation speed, latitude, longitude；wve、wvnFor tachometric survey noise；wλ、wLFor position measurement noise.wve、wvnConnect for the Big Dipper
Receipts machine tachometric survey noise, is determined, w by receiver velocity accuracy parameterλ、wLFor Beidou receiver position measurement noise, by receiving
Put precision parameter decision in seat in the plane.
Four ins error parameters can only directly be obtained by being seen from formula 6, but be closed from the state model F matrix of formula 2
System is it can be seen that change and remaining ins error relating to parameters of this four error parameters.It can be solved entirely by Kalman filtering
Portion's error parameter.Including：Attitude parameter (course, the angle of pitch, roll angle), location parameter (longitude λINS, latitude LINS, height),
Speed parameter (east orientation speed ve_INS, north orientation speed vn_INS, sky orientation speed).
The step estimate of utilization state one, operating speed, position correction model
V in formulae、vn, L, λ be revised east orientation speed, north orientation speed, latitude, longitude navigation results.
Fig. 4 is the embodiment block diagram of guider of the present invention.The embodiment of the present application also provides a kind of guider, for nothing
It is man-machine, include micro inertial measurement unit 100, DVB 200, processor 300；The micro inertial measurement unit, for from
Initial time starts, and gyro instantaneous value, acceleration instantaneous value are produced according to measurement period；The DVB, for receiving
Satellite positioning information, in initial time outgoing position initial value and velocity original value；It is defeated according to filtering cycle from initial time
Out position calibration value and speed calibration value；The processor, for according to the position initial value, velocity original value, gyro wink
Duration, acceleration instantaneous value, position calculated value of the unmanned plane in each measurement period, speed are drawn with strap inertial navigation algorithm
Spend calculated value and attitude parameter calculated value；According to the position calculated value at calibration moment, speed calculated value, position correction value, speed
Calibration value, attitude parameter error, velocity error, site error, gyro error, acceleration are calculated with Kalman Filtering for Discrete method
Error；The error variance that the Kalman Filtering for Discrete method is used is included：Pitching angle error, roll angle error, course angle are missed
Difference, east orientation speed error, north orientation speed error, trueness error, latitude error, x, y, z direction gyro error, x, y direction accelerate
Spend error totally 12 parameters；Wherein, unmanned plane right side side is pointed in x, y, z form right angle coordinate system, x directions, and y directions are along unmanned plane
Before y direction is pointed to.
Fig. 5 is the embodiment block diagram that guider of the present invention further optimizes.Further, the processor 300 enters one
Step includes navigational parameter initialization unit 301, strap-down inertial computing unit 302, estimation error wave filter 303；It is described to lead
Navigate parameter initialization unit, for receiving acceleration instantaneous value, gyro instantaneous value, according to acceleration instantaneous value, gyro instantaneous value
Calculate attitude parameter instantaneous value (such as formula 1.1~1.2)；The strap-down inertial computing unit, for receiving the position
Initial value, velocity original value, attitude parameter instantaneous value, draw unmanned plane in each measurement period with strap inertial navigation algorithm
Position calculated value, speed calculated value and attitude parameter calculated value；The estimation error wave filter, uses Kalman Filtering for Discrete side
Method calculate attitude parameter error, velocity error, site error, gyro error, acceleration error totally 12 variables (such as formula 2~
5)。
Fig. 6 is the embodiment schematic diagram that guider of the present invention was inputted, exported calibration.The guider is further
Include output calibration unit 400；The output calibration unit, is entered with the attitude parameter error to the attitude parameter calculated value
Row amendment, obtains revised position, speed, attitude parameter.In any one embodiment of the present invention, it is preferable that described to lead
Device navigate further comprising input correction unit 500, with acceleration instantaneous value of the acceleration error to next filtering cycle
It is modified, the gyro instantaneous value of next filtering cycle is modified with the gyro error, the navigational parameter initialization
The input of unit.
It should be understood by those skilled in the art that, embodiments of the invention can be provided as method, system or computer program
Product.Therefore, the present invention can be using the reality in terms of complete hardware embodiment, complete software embodiment or combination software and hardware
Apply the form of example.Moreover, the present invention can be used in one or more computers for wherein including computer usable program code
The computer program production that usable storage medium is implemented on (including but is not limited to magnetic disk storage, CD-ROM, optical memory etc.)
The form of product.
The present invention is the flow with reference to method according to embodiments of the present invention, equipment (system) and computer program product
Figure and/or block diagram are described.It should be understood that can be by every first-class in computer program instructions implementation process figure and/or block diagram
Journey and/or the flow in square frame and flow chart and/or block diagram and/or the combination of square frame.These computer programs can be provided
The processor of all-purpose computer, special-purpose computer, Embedded Processor or other programmable data processing devices is instructed to produce
A raw machine so that produced by the instruction of computer or the computing device of other programmable data processing devices for real
The device for the function of being specified in present one flow of flow chart or one square frame of multiple flows and/or block diagram or multiple square frames.
It should also be noted that, term " comprising ", "comprising" or its any other variant are intended to nonexcludability
Comprising so that process, method, commodity or equipment including a series of key elements are not only including those key elements, but also wrap
Include other key elements being not expressly set out, or also include for this process, method, commodity or equipment intrinsic want
Element.In the absence of more restrictions, the key element limited by sentence "including a ...", it is not excluded that wanted including described
Also there is other identical element in process, method, commodity or the equipment of element.
Embodiments herein is the foregoing is only, the application is not limited to.For those skilled in the art
For, the application can have various modifications and variations.It is all any modifications made within spirit herein and principle, equivalent
Replace, improve etc., it should be included within the scope of claims hereof.
Claims (10)
1. a kind of air navigation aid, for unmanned plane, the unmanned plane includes micro inertial measurement unit, DVB, its feature
It is, comprises the following steps：
In initial time, unmanned plane position and speed are gathered with the DVB, respectively as position initial value and speed
Initial value；
From the initial time to the calibration moment, unmanned plane acceleration wink is gathered by measurement period with the micro inertial measurement unit
Duration, gyro instantaneous value；Attitude parameter instantaneous value is calculated according to acceleration instantaneous value, gyro instantaneous value；It is described calibration the moment and
When a length of filtering cycle between the initial time, includes multiple measurement periods；
According to the position initial value, velocity original value, attitude parameter instantaneous value, unmanned plane is drawn with strap inertial navigation algorithm
In the position calculated value, speed calculated value and attitude parameter calculated value of each measurement period；
At the calibration moment, unmanned plane position and speed are gathered with the DVB, respectively as position correction value and speed
Calibration value；
According to the position calculated value at calibration moment, speed calculated value, position correction value, speed calibration value, filtered with discrete Kalman
Wave method calculates attitude parameter error, velocity error, site error, gyro error, acceleration error；Discrete Kalman's filter
The error variance that wave method is used is included：Pitching angle error, roll angle error, course angle error, east orientation speed error, north orientation speed
Spend error, trueness error, latitude error, x, y, z direction gyro error, x, y directional acceleration error totally 12 parameters；Wherein,
Unmanned plane right side side is pointed in x, y, z form right angle coordinate system, x directions, before y directions are pointed to along unmanned plane y direction.
2. air navigation aid as claimed in claim 1, it is characterised in that
Using the calibration moment as new initial time, circulation performs the step.
3. air navigation aid as claimed in claim 2, it is characterised in that also comprise the steps of：
The acceleration instantaneous value of next filtering cycle is modified with the acceleration error；
The gyro instantaneous value of next filtering cycle is modified with the gyro error.
4. the air navigation aid as described in claims 1 to 3 any one, it is characterised in that
In the Kalman Filtering for Discrete method, error state variable X element is made up of following physical quantity：Angle of pitch error delta
φe, roll angle error delta φn, course angle error delta φu；East orientation speed error delta ve, north orientation speed error delta vn；Longitude error
Δ λ, latitude error Δ L；X, y, z direction gyro error εbx,εby,εbz；X, y directional acceleration meter errorI.e.
The 12 parameter error state model is set up according to the relation between inertial navigation system error, basic navigation parameter.Specifically,
Error state model isWherein F is state-transition matrix, and dimension is 12 × 12, and nonzero term is：
F (0,1)=ωiesinL；F (0,2)=- ωiecosL； F (1,0)=- ωiesinL；F (1,5)=- ωiesinL；F (2,0)=ωiecosL； F (3,1)=- g；F
(4,0)=g；
Wherein, ωieFor rotational-angular velocity of the earth, L is latitude, and R is earth radius,For attitude matrix, dimension is 3 × 3, ve、
vnRespectively east orientation speed, north orientation speed, g is acceleration of gravity, τgyroFor gyro correlation time, τaccDuring for accelerometer correlation
Between；
G is noise matrix, and dimension is 12 × 1, is met
Wherein, Wgyro=[wgx wgy wgz]T, Wacc=[wax way waz]T, wgx、wgy、wgzFor gyroscope measurement noise, wax、way、
wazFor accelerometer measures noise.
5. the air navigation aid as described in claims 1 to 3 any one, it is characterised in that the measurement period is 20ms.
6. the air navigation aid as described in claims 1 to 3 any one, it is characterised in that the filtering cycle is 1s.
7. a kind of guider, for unmanned plane, comprising micro inertial measurement unit, DVB, processor, its feature exists
In,
The micro inertial measurement unit, for since initial time, gyro instantaneous value, acceleration wink to be produced according to measurement period
Duration；
The DVB, for receiving satellite positioning information, in initial time outgoing position initial value and velocity original value；
According to filtering cycle outgoing position calibration value and speed calibration value from initial time；
The processor, for according to the position initial value, velocity original value, gyro instantaneous value, acceleration instantaneous value, with victory
Inertial navigation algorithm show that unmanned plane is calculated in the position calculated value, speed calculated value and attitude parameter of each measurement period
Value；According to the position calculated value at calibration moment, speed calculated value, position correction value, speed calibration value, Kalman Filtering for Discrete is used
Method calculates attitude parameter error, velocity error, site error, gyro error, acceleration error；The Kalman Filtering for Discrete
The error variance that method is used is included：Pitching angle error, roll angle error, course angle error, east orientation speed error, north orientation speed
Error, trueness error, latitude error, x, y, z direction gyro error, x, y directional acceleration error totally 12 parameters；Wherein, x,
Unmanned plane right side side is pointed in y, z form right angle coordinate system, x directions, before y directions are pointed to along unmanned plane y direction.
8. guider as claimed in claim 7, it is characterised in that the processor further initializes single comprising navigational parameter
Member, strap-down inertial computing unit, estimation error wave filter；
The navigational parameter initialization unit, for receive acceleration instantaneous value, gyro instantaneous value, according to acceleration instantaneous value,
Gyro instantaneous value calculates attitude parameter instantaneous value；
The strap-down inertial navigation elements, it is instantaneous for receiving the position initial value, velocity original value, attitude parameter
Value, position calculated value, speed calculated value and posture of the unmanned plane in each measurement period are drawn with strap inertial navigation algorithm
Parameter calculated value；
The estimation error wave filter, attitude parameter error, velocity error, position mistake are calculated with Kalman Filtering for Discrete method
Difference, gyro error, acceleration error.
9. guider as claimed in claim 8, it is characterised in that further comprising output calibration unit；
The output calibration unit, is modified with the attitude parameter error to the attitude parameter calculated value.
10. the guider as described in claim 7~9 any one, it is characterised in that further correct unit comprising input；
The input correction unit is modified to the acceleration instantaneous value of next filtering cycle with the acceleration error, uses the top
Spiral shell error is modified to the gyro instantaneous value of next filtering cycle, the input of the navigational parameter initialization unit.
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