CN108534783A - A kind of aircraft navigation method based on Beidou navigation technology - Google Patents
A kind of aircraft navigation method based on Beidou navigation technology Download PDFInfo
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- CN108534783A CN108534783A CN201810449055.9A CN201810449055A CN108534783A CN 108534783 A CN108534783 A CN 108534783A CN 201810449055 A CN201810449055 A CN 201810449055A CN 108534783 A CN108534783 A CN 108534783A
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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
- G01C21/20—Instruments for performing navigational calculations
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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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- 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
- G01S19/45—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement
- G01S19/47—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement the supplementary measurement being an inertial measurement, e.g. tightly coupled inertial
Abstract
The invention discloses a kind of aircraft navigation methods based on Beidou navigation technology, include the following steps：S100, the target information extraction based on SINS/BDS systems will carry out information extraction by the Federated Filters of no reconfiguration structure to SINS/BDS integrated navigation systems；S200, the target information extraction based on SINS/CNS systems；S300, setting local filter, carry out conventional Kalman filtering to existing integrated navigation information, calculate two groups of local optimum estimated values for obtaining system mode；S400, setting fault detection module, examine the validity of each partial estimation；S500, primary filter structure is established, effective partial estimation value is sent into senior filter and carries out global optimum's information fusion, obtain the global best estimates of system common condition, this method had both absorbed SINS/BDS integrated navigations and had tested the speed the advantage high with positioning accuracy, the high advantage of SINS/CNS integrated navigation accuracy of attitude determination is absorbed again, to obtain it is very high determine appearance, test the speed and positioning accuracy, realize comprehensive optimization of each navigational parameter.
Description
Technical field
The present invention relates to aircraft navigation field, specially a kind of aircraft navigation method based on Beidou navigation technology.
Background technology
In modern war, the informationization of height so that no matter attacker or defender can quick obtaining battlefield
Information, but launch weapon system platform and weapon it is relatively lower speed, can not quickly arrive at target area, and destroy mesh
Mark, opportunity of combat are transient.Speed is slow, be easy it is intercepted be current cruise vehicle disadvantage, it is superb for these problems
Velocity of sound aircraft comes into being.
Eyes as hypersonic aircraft：Navigation system, it is undoubtedly crucial in key.Navigation system is broad sense
The most basic link of flight control system (Navigation, Guidance and Control system), it is the basis of guidance and control loop, is its number
It is one of the component of aircraft key according to source, but there is also following deficiencies for the existing method for aircraft navigation
Place：
For example, application No. is 201611032966.9, patent name be it is a kind of based on the microminiature of triones navigation system without
The patent of invention of people's Aerial vehicle position and air navigation aid：
, using the velocity information for obtaining unmanned plane mounted on the light stream sensor of quadrotor drone bottom, utilization is airborne for it
Inertial navigation unit obtains acceleration information, using airborne vision system acquisition speed information, in conjunction with the original of dipper system position
Beginning measured value, fused filtering obtain the estimation for position and speed；And then it by nonlinear positional control algorithm, realizes
Position of aircraft controls.Present invention is mainly applied to unmanned vehicles and flight control occasion.
But the existing aircraft navigation method based on Beidou navigation technology has the following defects：
(1) currently, China has certain research and application for inertia and satellite, inertia with astronomical combined guidance, still
Theory is still in the research based on inertia combined of multi-sensor information, satellite, astronomy and other system in combination airmanships
Research and experimental exploring stage；
(2) currently, the domestic combination that inertia/astronomy is much used in the navigation system of the strategic arms such as intercontinental missile
Navigation pattern is asked although efficiently solving the protrusion that stabilized platform constantly deviates reference position as the flight time increases
Topic, slows down the speed that inertial navigation system position error accumulates at any time to a certain extent, but in long endurance, high-speed flight etc.
Navigation accuracy under flight environment of vehicle is not still high, but also there is complicated, volume is big, weight is heavy, poor reliability, fault-tolerant
Property difference and equal protrusion disadvantage difficult in maintenance；
(3) Platform Inertial Navigation System, star tracker and global positioning system (GPS) are combined by domestic have studied at present
Navigation system, be configured with two mutually orthogonal star trackers of sight line in systems, and star tracker is installed
In INS Platform, this allows for coupling too tightly between each subsystem, is affected to the estimation effect of systematic error.
Invention content
In order to overcome the shortcomings of that prior art, the present invention provide a kind of aircraft navigation based on Beidou navigation technology
Method can effectively solve the problem that the problem of background technology proposes.
The technical solution adopted by the present invention to solve the technical problems is：
A kind of aircraft navigation method based on Beidou navigation technology, includes the following steps：
S100, the target information extraction based on SINS/BDS systems, will be right by the Federated Filters of no reconfiguration structure
SINS/BDS integrated navigation systems carry out information extraction；
S200, the target information extraction based on SINS/CNS systems, will be right by the Federated Filters of no reconfiguration structure
SINS/CNS integrated navigation systems carry out information extraction；
S300, setting local filter carry out conventional Kalman filtering to existing integrated navigation information, and calculating is
Two groups of local optimum estimated values of system state；
Two groups of local optimum estimated values are passed through respective fault detect by S400, setting fault detection module respectively respectively
Module examines the validity of each partial estimation；
S500, primary filter structure is established, effective partial estimation value, which is sent into senior filter, carries out global optimum's information
Fusion, obtains the global best estimates of system common condition.
Further, in the step s 100, the specific algorithm of the target information extraction based on SINS/BDS systems is：
S101, the attitude algorithm differential equation, if the quaternary number of description attitude of carrier isThen it meets the differential equation：Wherein, Referred to as posture rate；
S102, posture rate equation is obtained, utilizedWherein,For
The carrier angular speed that gyroscope exports in inertial navigation,For posture battle array determined by posture renewal in step S101,For
The instruction angular speed of mathematical platform；
S103, the differential equation is solved, joint step S101, S102 can obtain attitude quaternion immediatelyInstantaneous value,
It can determine that attitude of carrier matrix
For
To obtain course angle, pitch angle and the roll angle of carrier.
Further, the speed of Federated Filters more new algorithm is as follows in step S100, S200：
S201, acquisition speed renewal equation can obtain the speed of Strapdown Inertial Navigation System more according to the specific force equation of inertial navigation system
The new differential equation
S202, speed in a certain moment navigational coordinate system is determined, if the speed update cycle of inertial navigation is Tv=tm-
tm-1, to above formula in [tm-1, tm] period interior integral, it is collated to obtain tmSpeed of the moment carrier in navigational coordinate system is
S203, determine speed update in rotation effect compensation term,Wherein Δ θmFor
Angular speed is in [tm-1, tm] the period interior angle increment generated；
S204, determine speed update in sculling compensation term.
Further, in step S400, specific bearing calibration is as follows in fault detection module：
S401, setting carrier transition matrix, if the transition matrix of carrier coordinate system (b systems) to navigational coordinate system (n systems) is
S402, calculating matrix transformed error angle, if it is n that navigation, which calculates the navigational coordinate system actually obtained,1System, then it is corresponding
Coordinate conversion matrix isWherein n1There are error angles for opposite n systems of system
S403, coefficient transition matrix is determined, the mathematical platform attitude error angular estimation value exported according to Kalman filterIt calculates from n1It is the transition matrix to n systemsI.e.：
To true strapdown attitude matrix
Further, in step S500, the integrated navigation information fusion algorithm of senior filter is as follows：
S501, the system common condition local optimum estimated value of SINS/BDS integrated navigations is set as XB, corresponding estimation
Mean square deviation is PB；And if the system common condition local optimum estimated value of SINS/CNS integrated navigations is Xc, corresponding estimation is
Variance is PC；
S502, X is setcFor the common condition of integrated navigation system, Xci(i=1,2 ... N) are local filter i to public shape
The local optimum of state estimates that the covariance matrix of the estimation is PCi, δ XciFor each local optimum estimation evaluated error, i.e.,：δXci=
Xci-Xc；
S503, according to the global information blending algorithm without resetting Federated Filters, it is public to obtain SINS/BDS/CNS integrated navigation systems
The global best estimates value X of state and its estimation mean square deviation P altogether, i.e.,：
Further, in step S503, according to the system state estimation value X of previous moment part Kalman filterK-1
And its mean square error PK-1, standard karr is carried out using the output of system state equation, Strapdown Inertial Navigation System and other auxiliary systems
Graceful filtering.
Further, in step S500, computer mathematics emulation is carried out to the quick backoff algorithm of integrated navigation information.
Further, in step S500, senior filter output system error optimization estimates master control system, the master control
System includes attitude algorithm quaternion algebra amount module, pitch angle control module and pid control module；
The attitude algorithm quaternion algebra amount module receives the acceleration signal of three axis accelerometer, the attitude algorithm four
The signal end of first number quantity module is interconnected with three-axis gyroscope；
The output end of the attitude algorithm quaternion algebra amount module connects pitch angle control module, and the pitch angle controls mould
The control terminal of block is connected with pid control module；
The pid control module output end PWM wave adjusts empennage posture.
Compared with prior art, the beneficial effects of the invention are as follows：
(1) aircraft navigation method of the invention had both absorbed SINS/BDS integrated navigations and had tested the speed and high excellent of positioning accuracy
Point, and absorb the high advantage of SINS/CNS integrated navigation accuracy of attitude determination very high determine appearance, tests the speed and positioning accurate to obtain
Degree, realizes comprehensive optimization of each navigational parameter；
(2) use integrated navigation system and its information fusion algorithm of the invention, using Strapdown Inertial Navigation System as public ginseng
Test system, the Big Dipper, celestial navigation system take commonwealth filter technique to design the inertial navigation/Big Dipper/day as secondary navigation system
Literary high-precision integrated navigation system is not only determined appearance, positioning and rate accuracy with very high, but also can effectively be estimated used
The error of property device, realizes the optimization of each navigational parameter, and carry comprehensively at the advantages of having fully absorbed each navigation subsystem
The high comprehensive performance of navigation system.
Description of the drawings
Fig. 1 is the overall flow figure of the present invention；
Fig. 2 is the master control system structural schematic diagram of the present invention.
Specific implementation mode
Following will be combined with the drawings in the embodiments of the present invention, and technical solution in the embodiment of the present invention carries out clear, complete
Site preparation describes, it is clear that described embodiments are only a part of the embodiments of the present invention, instead of all the embodiments.It is based on
Embodiment in the present invention, it is obtained by those of ordinary skill in the art without making creative efforts every other
Embodiment shall fall within the protection scope of the present invention.
As shown in Figure 1, the present invention provides a kind of aircraft navigation methods based on Beidou navigation technology, including walk as follows
Suddenly：
S100, the target information extraction based on SINS/BDS systems, will be right by the Federated Filters of no reconfiguration structure
SINS/BDS integrated navigation systems carry out information extraction；
S200, the target information extraction based on SINS/CNS systems, will be right by the Federated Filters of no reconfiguration structure
SINS/CNS integrated navigation systems carry out information extraction；
S300, setting local filter carry out conventional Kalman filtering to existing integrated navigation information, and calculating is
Two groups of local optimum estimated values of system state；
Two groups of local optimum estimated values are passed through respective fault detect by S400, setting fault detection module respectively respectively
Module examines the validity of each partial estimation；
S500, primary filter structure is established, effective partial estimation value, which is sent into senior filter, carries out global optimum's information
Fusion, obtains the global best estimates of system common condition (SINS error states).
In the present embodiment, SINS is combined two-by-two first with BDS, SINS and CNS, constitutes SINS/BDS, SINS/
CNS integrated navigations part Kalman filter, two local filters calculate the two groups of local optimums estimation for obtaining system mode
XB, XC；Then, two groups of local optimum estimations are examined into the effective of each partial estimation respectively by respective fault detection module
Property；Then, effective partial estimation value is sent into senior filter and carries out global optimum's information fusion, obtain system common condition
The global best estimates of (SINS error states)；Finally, SINS is carried out immediately using obtained SINS error optimizations estimated value
Error correction, and the SINS after correction is exported into the output as SINS/BDS/CNS integrated navigation systems.
In the step s 100, the specific algorithm of the target information extraction based on SINS/BDS systems is：
S101, the attitude algorithm differential equation, if the quaternary number of description attitude of carrier isThen it meets the differential equation：Wherein, Referred to as posture rate；
S102, posture rate equation is obtained, utilizedWherein,For
The carrier angular speed that gyroscope exports in inertial navigation,For posture battle array determined by posture renewal in step S101,For
The instruction angular speed of mathematical platform；
S103, the differential equation is solved, joint step S101, S102 can obtain attitude quaternion immediatelyInstantaneous value,
It can determine that attitude of carrier matrix
For
To obtain course angle, pitch angle and the roll angle of carrier.
The speed of Federated Filters more new algorithm is as follows in step S100, S200：
S201, acquisition speed renewal equation can obtain the speed of Strapdown Inertial Navigation System more according to the specific force equation of inertial navigation system
The new differential equation
S202, speed in a certain moment navigational coordinate system is determined, if the speed update cycle of inertial navigation is Tv=tm-
tm-1, to above formula in [tm-1, tm] period interior integral, it is collated to obtain tmSpeed of the moment carrier in navigational coordinate system is
S203, determine speed update in rotation effect compensation term,Wherein Δ θmFor
Angular speed is in [tm-1, tm] the period interior angle increment generated；
S204, determine speed update in sculling compensation term.
In step S400, specific bearing calibration is as follows in fault detection module：
S401, setting carrier transition matrix, if the transition matrix of carrier coordinate system (b systems) to navigational coordinate system (n systems) is
S402, calculating matrix transformed error angle, if it is n that navigation, which calculates the navigational coordinate system actually obtained,1System, then it is corresponding
Coordinate conversion matrix isWherein n1There are error angles for opposite n systems of system
S403, coefficient transition matrix is determined, the mathematical platform attitude error angular estimation value exported according to Kalman filterIt calculates from n1It is the transition matrix to n systemsI.e.：
To true strapdown attitude matrix
In step S500, the integrated navigation information fusion algorithm of senior filter is as follows：
S501, the system common condition local optimum estimated value of SINS/BDS integrated navigations is set as XB, corresponding estimation
Mean square deviation is PB；And if the system common condition local optimum estimated value of SINS/CNS integrated navigations is Xc, corresponding estimation is
Variance is PC；
S502, X is setcFor the common condition of integrated navigation system, Xci(i=1,2 ... N) are local filter i to public shape
The local optimum of state estimates that the covariance matrix of the estimation is PCi, δ XciFor each local optimum estimation evaluated error, i.e.,：δXci=
Xci-Xc；
S503, according to the global information blending algorithm without resetting Federated Filters, it is public to obtain SINS/BDS/CNS integrated navigation systems
The global best estimates value X of state and its estimation mean square deviation P altogether, i.e.,：
In the present embodiment, after the global best estimates value X for obtaining Strapdown Inertial Navigation System error state, need according to this most
Excellent estimated value carries out error correction to Strapdown Inertial Navigation System in time, finally, will be by the Strapdown Inertial Navigation System of Systematic Error Correction
Output of the output as SINS/BDS/CNS high-precision integrated navigation systems, specifically include the posture, speed, position of carrier
It sets, the navigation informations such as angular speed and acceleration.
In the present embodiment, each local filter is not interfere with each other, independent parallel is filtered in no resetting commonwealth filter technique,
There is no what reset and feedback brought to influence each other, when a certain subsystem breaks down, do not interfere with other navigation subsystems and
The normal work of entire integrated navigation system, to while ensureing integrated navigation system precision, have better fault-tolerance
And reliability.Therefore matched based on system design, the redundancy of a variety of different performance navigation subsystems without resetting commonwealth filter technique
It is set to reliability and fault-tolerance of the integrated navigation system under complicated adverse circumstances and provides effective guarantee.
In step S503, according to the system state estimation value X of previous moment part Kalman filterK-1And its it is square
Error PK-1, standard Kalman filtering is carried out using the output of system state equation, Strapdown Inertial Navigation System and other auxiliary systems.
In the present embodiment, under normal circumstances, it can be obtained according to step S503：
XK=XK/K-1+KK*(ZK-HK*XK/K-1)
PK=(I-KK*HK)PK/K-1(I-KK*HK)T+KKRKKK T
According to above formula as it can be seen that when navigation system works in normal circumstances, the defeated of each navigation subsystem can be passed through
Go out just-in-time construction and goes out measurement information ZK, to the standard Kalman filtering algorithm just measurement information Z by being constructedKConstantly disappear
Influence except factors such as initial value error, system noise and system modelling errors to state recurrence estimation value, and then make system mode
Recurrence estimation value increasingly approaching to reality value, i.e. the filtering accuracy of system is higher and higher.
In step S500, computer mathematics emulation is carried out to the quick backoff algorithm of integrated navigation information, in the present embodiment,
SINS/BDS/CNS high-precision integrated navigation systems utilize without resetting Federated Kalman Filtering technology, given full play to SINS,
The advantages of each navigation subsystem of BDS, CNS, has and very high determine appearance, test the speed and positioning accuracy, moreover it is possible to each inertia device mistake
Difference is effectively estimated, and improves the comprehensive performance of integrated navigation system comprehensively, realizes the optimization of each navigational parameter of system.
As shown in Fig. 2, in step S500, senior filter output system error optimization estimates master control system, the master
Control system includes attitude algorithm quaternion algebra amount module, pitch angle control module and pid control module；The attitude algorithm quaternary
Number quantity module receives the acceleration signal of three axis accelerometer, the signal end and three of the attitude algorithm quaternion algebra amount module
Axis gyroscope interconnects；The output end of the attitude algorithm quaternion algebra amount module connects pitch angle control module, described to bow
The control terminal of elevation angle control module is connected with pid control module；The pid control module output end PWM wave adjusts empennage appearance
State.
In the present embodiment, pid control module is using complementary filter algorithm to the acceleration transducer with low frequency characteristic
Information has the function of low frequency filtering, so as to get posture information it is more smooth, while reducing high-frequency information to posture information
Interference；The low-frequency noise that gyro sensor obtains has high-pass filtering effect, so as to get posture information reduce gyroscopic drift
Interference, enable the work that gyroscope is steady in a long-term.
In the present embodiment, complementary filter method utilizes the characteristic of the two sensors, them is allowed to be complementary to one another, and improves sensor
Measurement accuracy and system dynamic property.
In the present embodiment, during solving flapping wing aircraft posture using complementary filter algorithm, data are carried out initial
Change is handled, and it is 3-axis acceleration a that accelerometer, which obtains data,x, ay, az, the data that gyroscope obtains are three axis angular rate wx, wy,
wz, initial angular velocity is integrated to obtain the three-axis attitude angle of initial time, the three-axis attitude angle of usual original state is all 0, institute
It is identical in an initial condition, to need the regulation inceptive direction with the coordinate of navigation system, solves four under original state
First number is as follows：
Before flapping wing aircraft is flown, in order to indicate convenient, need to adjust initial position, in an initial condition, θ,It is set as zero degree.It is q=[1000] to obtain initial quaternary number.
It is unitization in order to conveniently can generally be carried out acceleration value in calculating process in the present embodiment.Utilize complementary filter
Method corrects gyro parameter, with the weight that in body indicates that indicates with accelerometer measure of the acceleration of gravity in inertial coodinate system
Power acceleration value does multiplication cross operation, obtains an error amount, this error amount is known as compensation correction error, is then exactly to utilize this
A error correction gyro.
It is obvious to a person skilled in the art that invention is not limited to the details of the above exemplary embodiments, Er Qie
In the case of without departing substantially from spirit or essential attributes of the invention, the present invention can be realized in other specific forms.Therefore, no matter
From the point of view of which point, the present embodiments are to be considered as illustrative and not restrictive, and the scope of the present invention is by appended power
Profit requires rather than above description limits, it is intended that all by what is fallen within the meaning and scope of the equivalent requirements of the claims
Variation is included within the present invention.Any reference signs in the claims should not be construed as limiting the involved claims.
Claims (8)
1. a kind of aircraft navigation method based on Beidou navigation technology, it is characterised in that：Include the following steps：
S100, the target information extraction based on SINS/BDS systems, will be to SINS/ by the Federated Filters of no reconfiguration structure
BDS integrated navigation systems carry out information extraction；
S200, the target information extraction based on SINS/CNS systems, will be to SINS/ by the Federated Filters of no reconfiguration structure
CNS integrated navigation systems carry out information extraction；
S300, setting local filter, carry out conventional Kalman filtering to existing integrated navigation information, calculate acquisition system shape
Two groups of local optimum estimated values of state；
S400, setting fault detection module, respectively by two groups of local optimum estimated values respectively by respective fault detection module,
Examine the validity of each partial estimation；
S500, primary filter structure is established, effective partial estimation value, which is sent into senior filter, carries out global optimum's information fusion,
Obtain the global best estimates of system common condition.
2. a kind of aircraft navigation method based on Beidou navigation technology according to claim 1, it is characterised in that：In step
In rapid S100, the specific algorithm of the target information extraction based on SINS/BDS systems is：
S101, the attitude algorithm differential equation, if the quaternary number of description attitude of carrier isThen it meets the differential equation：Wherein, Referred to as posture rate；
S102, posture rate equation is obtained, utilizedWherein,For strapdown
The carrier angular speed that gyroscope exports in inertial navigation,For posture battle array determined by posture renewal in step S101,For mathematics
The instruction angular speed of platform；
S103, the differential equation is solved, joint step S101, S102 can obtain attitude quaternion immediatelyInstantaneous value, can be true
Make attitude of carrier matrix
For
To obtain course angle, pitch angle and the roll angle of carrier.
3. a kind of aircraft navigation method based on Beidou navigation technology according to claim 1, it is characterised in that：In step
The speed more new algorithm of Federated Filters is as follows in rapid S100, S200：
S201, speed renewal equation is obtained, according to the specific force equation of inertial navigation system, the speed update that can obtain Strapdown Inertial Navigation System is micro-
Divide equation
S202, speed in a certain moment navigational coordinate system is determined, if the speed update cycle of inertial navigation is Tv=tm-tm-1,
To above formula in [tm-1, tm] period interior integral, it is collated to obtain tmSpeed of the moment carrier in navigational coordinate system is
S203, determine speed update in rotation effect compensation term,Wherein Δ θmFor angle speed
Degree is in [tm-1, tm] the period interior angle increment generated；
S204, determine speed update in sculling compensation term.
4. a kind of aircraft navigation method based on Beidou navigation technology according to claim 1, it is characterised in that：In step
In rapid S400, specific bearing calibration is as follows in fault detection module：
S401, setting carrier transition matrix, if the transition matrix of carrier coordinate system (b systems) to navigational coordinate system (n systems) is
S402, calculating matrix transformed error angle, if it is n that navigation, which calculates the navigational coordinate system actually obtained,1It is, then corresponding coordinate
Transition matrix isWherein n1There are error angles for opposite n systems of system
S403, coefficient transition matrix is determined, the mathematical platform attitude error angular estimation value exported according to Kalman filterIt calculates from n1It is the transition matrix to n systemsI.e.：
To true strapdown attitude matrix
5. a kind of aircraft navigation method based on Beidou navigation technology according to claim 1, it is characterised in that：In step
In rapid S500, the integrated navigation information fusion algorithm of senior filter is as follows：
S501, the system common condition local optimum estimated value of SINS/BDS integrated navigations is set as XB, corresponding estimation mean square deviation
For PB；And if the system common condition local optimum estimated value of SINS/CNS integrated navigations is Xc, corresponding estimation mean square deviation is
PC；
S502, X is setcFor the common condition of integrated navigation system, Xci(i=1,2 ... N) are local filter i to common condition
Local optimum estimates that the covariance matrix of the estimation is PCi, δ XciFor each local optimum estimation evaluated error, i.e.,：δXci=Xci-
Xc；
S503, according to without resetting Federated Filters global information blending algorithm, obtain the public shape of SINS/BDS/CNS integrated navigation systems
The global best estimates value X and its estimation mean square deviation P of state, i.e.,：
6. a kind of aircraft navigation method based on Beidou navigation technology according to claim 5, it is characterised in that：In step
In rapid S503, according to the system state estimation value X of previous moment part Kalman filterK-1And its mean square error PK-1, utilize
The output of system state equation, Strapdown Inertial Navigation System and other auxiliary systems carries out standard Kalman filtering.
7. a kind of aircraft navigation method based on Beidou navigation technology according to claim 1, it is characterised in that：In step
In rapid S500, computer mathematics emulation is carried out to the quick backoff algorithm of integrated navigation information.
8. a kind of aircraft navigation method based on Beidou navigation technology, it is characterised in that：In step S500, senior filter is defeated
Go out systematic error optimal estimation to master control system, the master control system includes attitude algorithm quaternion algebra amount module, pitch angle control
Molding block and pid control module；
The attitude algorithm quaternion algebra amount module receives the acceleration signal of three axis accelerometer, the attitude algorithm quaternary number
The signal end of quantity module is interconnected with three-axis gyroscope；
The output end of the attitude algorithm quaternion algebra amount module connects pitch angle control module, the pitch angle control module
Control terminal is connected with pid control module；
The pid control module output end PWM wave adjusts empennage posture.
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109813299A (en) * | 2019-03-06 | 2019-05-28 | 南京理工大学 | A kind of integrated navigation information fusion method based on Interactive Multiple-Model |
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| CN114689054A (en) * | 2022-02-24 | 2022-07-01 | 中国电子科技集团公司第十研究所 | High-precision navigation method and device for Takang system, flight equipment and storage medium |
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Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109813299A (en) * | 2019-03-06 | 2019-05-28 | 南京理工大学 | A kind of integrated navigation information fusion method based on Interactive Multiple-Model |
| CN110057382A (en) * | 2019-04-23 | 2019-07-26 | 西北工业大学 | A kind of inertial navigation numerical value update method based on launching coordinate system |
| CN110057382B (en) * | 2019-04-23 | 2021-07-09 | 西北工业大学 | Strapdown inertial navigation value updating method based on emission coordinate system |
| CN111649744A (en) * | 2020-05-15 | 2020-09-11 | 北京自动化控制设备研究所 | Combined navigation positioning method based on assistance of dynamic model |
| CN111649744B (en) * | 2020-05-15 | 2023-08-15 | 北京自动化控制设备研究所 | Combined navigation positioning method based on dynamic model assistance |
| CN111947654A (en) * | 2020-08-13 | 2020-11-17 | 杭州北斗东芯科技有限公司 | Navigation and control integrated chip and control method thereof |
| CN112629538A (en) * | 2020-12-11 | 2021-04-09 | 哈尔滨工程大学 | Ship horizontal attitude measurement method based on fusion complementary filtering and Kalman filtering |
| CN113048987A (en) * | 2021-03-12 | 2021-06-29 | 湘潭大学 | Vehicle navigation system positioning method |
| CN114689054A (en) * | 2022-02-24 | 2022-07-01 | 中国电子科技集团公司第十研究所 | High-precision navigation method and device for Takang system, flight equipment and storage medium |
| CN114689054B (en) * | 2022-02-24 | 2023-06-20 | 中国电子科技集团公司第十研究所 | Takang system high-precision navigation method and device, flight equipment and storage medium |
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