CN109596127A - A kind of air navigation aid of radio auxiliary dead reckoning - Google Patents
A kind of air navigation aid of radio auxiliary dead reckoning Download PDFInfo
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Abstract
The present invention relates to a kind of air navigation aids of radio auxiliary dead reckoning.The present invention includes: the parameter of the airborne each navigation equipment output of acquisition；Obtain externally input initial position message；According to aircraft position, magnetic heading information is added to obtain true course with magnetic variation information；Using dead reckoning method, the location information of current time reckoning is obtained；State equation is established according to dead reckoning error model and oblique distance error model；Measurement equation is established using the location information of dead reckoning system, guidance station location information, oblique distance information；Bonding state equation and measurement equation estimate navigational parameter error using kalman filter method；Navigational parameter is corrected according to the navigational parameter margin of error of estimation, completes the accurate estimation of navigational parameter；The position at obtained current time and wind speed are fed back into dead-reckoning module, for completing the reckoning of subsequent time position, the real-time accurate estimation of navigational parameter is realized by way of Recursive Solution.
Description
Technical field
The invention belongs to Multi-sensor Navigation technologies, are related to a kind of air navigation aid of radio auxiliary dead reckoning.
Background technique
For the navigation accuracy requirement for improving aircraft, it is commonly equipped with a variety of navigation equipments, it is example inertial navigation system (INS), complete
Ball global position system (GPS), air data computer (ADC), radio navigation system etc..INS can be autonomous offer it is complete
Face navigation information, but its expensive error is with time integral.GPS round-the-clock offer position and speed information and can be not present
Cumulative errors, but there are problems that dependence external information, vulnerable to environmental disturbances.Radionavigation is by guidance station switching and signal quality
It influences, navigation results have jump.ADC is capable of providing flight atmospheric parameter, but can not determine aircraft position and posture information.
For the positioning accuracy for improving navigation system, integrated navigation mode is generallyd use both at home and abroad and realizes advantage and disadvantage complementation.
Currently, generalling use inertial navigation system as prime navaid system, and it is whole to use GPS to be combined amendment raising
The precision of a navigation system.But when requiring in face of low costs such as navigation flivver, unmanned planes, generally use more cheap boat
Appearance system (AHRS) replaces INS, merges realization navigation by carrying out with GPS, but entirely leads when GPS failure not can guarantee constantly
The precision property and continuity of boat system；Although being capable of providing aircraft position, navigation knot using DME radio-positioning mode
There is jump in fruit, precision and continuity are all difficult to be guaranteed；AHRS and atmospheric engine are only capable of providing posture and atmospheric parameter, can not
Realize the normal calculating of the comprehensive navigational parameter such as position, ground velocity, wind speed.
Summary of the invention
The technical problems to be solved by the present invention are: in order to solve under inexpensive navigation configuration and still be able to after GPS failure
High-precision, continuous reliable, the comprehensive navigation data of data are provided, make full use of airborne AHRS, ADC and DME parameter, the present invention mentions
The air navigation aid for having gone out a kind of low-cost wireless electricity auxiliary dead reckoning can be flown using this method for inexpensive navigation configuration
Machine provides high-precision continuous comprehensive navigation data, not only without using expensive inertial navigation system, but also still is able to after GPS failure
Ensure that full navigation parameter accurately calculates.
The technical scheme is that a kind of air navigation aid of radio auxiliary dead reckoning comprising the steps of:
Step 1, the parameter of the airborne each navigation equipment output of acquisition, the parameter include:
(1) the magnetic heading information of aviation attitude system output；
(2) true air speed, the pressure altitude information of air data system output；
(3) the oblique distance information of rangefinder output；
Step 2 obtains externally input initial position message, and location information includes longitude, latitude；
Magnetic heading information is added by step 3 according to aircraft position by inquiry database acquisition magnetic variation with magnetic variation information
Obtain true course；
Step 4, according to last moment aircraft position and wind speed, the true course at current time, true air speed and pressure altitude, adopt
With dead reckoning method, the location information of current time reckoning is obtained；
Step 5 establishes state equation according to dead reckoning error model and oblique distance error model；
Step 6 establishes measurement equation using the location information, guidance station location information, oblique distance information of dead reckoning system；
Step 7, bonding state equation and measurement equation estimate navigational parameter error using kalman filter method；
Step 8 is corrected navigational parameter according to the navigational parameter margin of error of estimation, completes accurately estimating for navigational parameter
Meter；
The position at obtained current time and wind speed are fed back to dead-reckoning module by step 9, for completing subsequent time
The real-time accurate estimation of navigational parameter is realized in the reckoning of position by way of Recursive Solution.
According to features described above, step 4 Air China position projectional technique are as follows:
According to the true air speed V (t) for the current time t that step 1 obtains, the true course φ (t) of the t moment that step 3 obtains with
And the north orientation wind speed of last moment t-1East orientation wind speedCalculate the north orientation ground velocity V of current time tn(t)
With east orientation ground velocity Ve(t) are as follows:
Wherein north orientation wind speed, east orientation wind speed initial value be taken as zero.
According to the north orientation ground velocity V of the latitude L (t-1) of last moment t-1 and longitude λ (t-1), current time tn(t) and it is eastern
To ground velocity Ve(t), the pressure altitude h (t) for the current time t that step 1 obtains, the latitude of current time t is calculated using reckoning mode
Spend L (t) and longitude λ (t):
L (t)=L (t-1)+Δ tVn(t)/(RM+h(t))
λ (t)=λ (t-1)+Δ tVe(t)/(RN+h(t))cosL(t-1)
Wherein, Δ t is that front and back time is poor, RN=Re(1+fsin2L), RM=Re(1-2f+3fsin2L), f=1/
298.257223563 being the compression of the Earth, Re=6378137.0 meters are earth radius.
According to features described above, state equation in the step 5 specifically:
According to dead reckoning error model and oblique distance error model, aircraft altitude error delta L, longitude error δ λ, height are chosen
Error delta h, north orientation air speed errorEast orientation air speed errorTrue course error delta φ, true air speed error delta V, DME1 oblique distance
Error delta R1, DME2 oblique distance error delta R2As quantity of state X, it may be assumed thatInto
And establish the state equation of Kalman filter:
Wherein state-transition matrix are as follows:
Wherein, τh、τnw、τew、τφ、τV、τRRespectively height, north orientation wind speed, east orientation wind speed, true course, true air speed, DME
The single order Markov correlation time of oblique distance.
Noise matrix is
System noise vector are as follows: w=[nφ,wh,wnw,wew,wφ,wV,wR,wR]T。
Wherein, nφIt is true course white Gaussian noise, wh、wnw、wew、wφ、wV、wRRespectively height, north orientation wind speed, east orientation wind
The single order Markov noise of speed, true course, true air speed, DME oblique distance.
According to features described above, measurement equation in the step 6 specifically:
The oblique distance R measured according to i-th of DMEi, guidance station height hiAnd aircraft altitude h, it utilizesCalculate DDMEiAs i-th of DME measurement oblique distance in horizontal plane projector distance, pushed away using boat position
The latitude L and longitude λ of calculation and the latitude L of i-th of DME guidance stationiWith longitude λiOblique distance is calculated in the projection D of horizontal planeDRi, i.e.,By DDMEiWith DDRiSubtract each other as measurement Z=[DDME1-
DDR1DDME2-DDR2]T, and then establish Kalman filter measurement equation:
Z (t)=H (t) X (t)+v (t)
H=[H in formula11 H12]
V=[v1,v2]T, v1And v2The white Gaussian noise of the zero-mean of respectively DME1 and DME2, i=1,2.
According to features described above, navigational parameter error includes aircraft altitude error delta L, longitude error δ λ, height in the step 7
Spend error delta h, north orientation air speed errorEast orientation air speed errorTrue course error delta φ, true air speed error delta V.
According to features described above, navigational parameter calculation method in the step 8 are as follows:
By step 4 obtain by the dead reckoning of current time t calculate aircraft altitude L (t), longitude λ (t) respectively with step
Latitude error δ L, the longitude error δ λ of 7 Kalman Filter Estimations are added to obtain the current time t accurate latitude L'(t of aircraft), warp
It spends λ ' (t).
T true air speed V (t), the pressure altitude h (t) of the current time measurement that step 1 obtains and step 3 are calculated
True course φ (t) is respectively the same as the true air speed error delta V of step 7 Kalman Filter Estimation, height error δ h, true course error delta φ phase
Add to obtain the current time t accurate true air speed V'(t of aircraft), pressure altitude h'(t) and true course φ ' (t).
The north orientation wind speed of last moment t-1With east orientation wind speedEstimate respectively with step 7 Kalman filtering
Count obtained north orientation air speed errorEast orientation air speed errorAddition obtains the north orientation wind speed at current time at current time tWith east orientation wind speed
Utilize the accurate true air speed V'(t of current time t aircraft), true course φ ' (t), north orientation wind speedWith east orientation wind
SpeedCalculate the north orientation ground velocity V of current time aircraftn' (t), east orientation ground velocity Ve' (t), specifically:
The beneficial effects of the present invention are:
One, the method for the present invention is proposed without using expensive inertial navigation system, after GPS failure, is sufficiently used
The output parameter of existing AHRS, DME and ADC system realize the estimation of aircraft navigation parameter using the method for integrated navigation, at
This reduces by 80% or more relative to inertial navigation/GPS system, and method is simply easily achieved, especially suitable for the flivver, nothing of opening the navigation or air flight
Man-machine equal inexpensive aircraft platform.
Two, the present invention is fed back by dead reckoning method, and by revised position and wind speed, realization aircraft position,
The navigational parameters such as speed, wind speed accurately calculate, and can be carried out to comprehensive navigational parameter is enriched by navigational parameter calculation method
It calculates, the comprehensive of data, continuity and inertial navigation are suitable.
Three, the present invention is by radio auxiliary filter estimation method to aircraft position, ground velocity, wind speed, height, true course etc.
Parameter is accurately estimated that the navigational parameter after solving the problems, such as GPS failure accurately calculates, and navigation accuracy and data flatness are than single
Pure radionavigation is more excellent, suitable for the precision navigation demand under the conditions of no GPS.
Detailed description of the invention
Fig. 1 is the structure principle chart of the air navigation aid of radio auxiliary dead reckoning of the present invention.
Specific embodiment
The following further describes the specific embodiments of the present invention with reference to the drawings.
Referring to Fig. 1, the present invention under the premise of not using inertial navigation system and GPS, make full use of existing AHRS,
The output parameter of ADC, DME system realizes accurately calculating for aircraft navigation parameter comprising the steps of:
Step 1, the parameter of the airborne each navigation equipment output of acquisition, the parameter include:
(1) the magnetic heading information of aviation attitude system output；
(2) true air speed, the pressure altitude information of air data system output；
(3) the oblique distance information of rangefinder output.
Step 2 obtains externally input initial position message, and location information includes longitude, latitude.
Magnetic heading information is added by step 3 according to aircraft position by inquiry database acquisition magnetic variation with magnetic variation information
Obtain true course.
Step 4, according to last moment aircraft position and wind speed, the true course at current time, true air speed and pressure altitude, adopt
With dead reckoning method, the location information of current time reckoning is obtained.
Dead reckoning method are as follows:
According to the true air speed V (t) for the current time t that step 1 obtains, the true course φ (t) of the t moment that step 3 obtains with
And the north orientation wind speed of last moment t-1East orientation wind speedCalculate the north orientation ground velocity V of current time tn(t)
With east orientation ground velocity Ve(t) are as follows:
Wherein north orientation wind speed, east orientation wind speed initial value be taken as zero.
According to the north orientation ground velocity V of the latitude L (t-1) of last moment t-1 and longitude λ (t-1), current time tn(t) and it is eastern
To ground velocity Ve(t), the pressure altitude h (t) for the current time t that step 1 obtains, the latitude of current time t is calculated using reckoning mode
Spend L (t) and longitude λ (t):
L (t)=L (t-1)+Δ tVn(t)/(RM+h(t))
λ (t)=λ (t-1)+Δ tVe(t)/(RN+h(t))cosL(t-1)
Wherein, Δ t is that front and back time is poor, RN=Re(1+fsin2L), RM=Re(1-2f+3fsin2L), f=1/
298.257223563 being the compression of the Earth, Re=6378137.0 meters are earth radius.
Step 5 establishes state equation according to dead reckoning error model and oblique distance error model.
State equation specifically:
According to dead reckoning error model and oblique distance error model, aircraft altitude error delta L, longitude error δ λ, height are chosen
Error delta h, north orientation air speed errorEast orientation air speed errorTrue course error delta φ, true air speed error delta V, DME1 oblique distance
Error delta R1, DME2 oblique distance error delta R2As quantity of state X, it may be assumed thatInto
And establish the state equation of Kalman filter:
Wherein state-transition matrix are as follows:
Wherein, τh、τnw、τew、τφ、τV、τRRespectively height, north orientation wind speed, east orientation wind speed, true course, true air speed, DME
The single order Markov correlation time of oblique distance.
Noise matrix is
System noise vector are as follows: w=[nφ,wh,wnw,wew,wφ,wV,wR,wR]T。
Wherein, nφIt is true course white Gaussian noise, wh、wnw、wew、wφ、wV、wRRespectively height, north orientation wind speed, east orientation wind
The single order Markov noise of speed, true course, true air speed, DME oblique distance.
Step 6 establishes measurement equation using the location information, guidance station location information, oblique distance information of dead reckoning system.
Measurement equation specifically:
The oblique distance R measured according to i-th of DMEi, guidance station height hiAnd aircraft altitude h, it utilizesCalculate DDMEiAs i-th of DME measurement oblique distance in horizontal plane projector distance, dead reckoning is utilized
Latitude L and longitude λ and i-th of DME guidance station latitude LiWith longitude λiOblique distance is calculated in the projection D of horizontal planeDRi, i.e.,By DDMEiWith DDRiSubtract each other as measurement Z=[DDME1-
DDR1DDME2-DDR2]T, and then establish Kalman filter measurement equation:
Z (t)=H (t) X (t)+v (t)
H=[H in formula11H12]
V=[v1,v2]T, v1And v2The white Gaussian noise of the zero-mean of respectively DME1 and DME2, i=1,2.
Step 7, bonding state equation and measurement equation estimate navigational parameter error using kalman filter method.
Navigational parameter error includes aircraft altitude error delta L, longitude error δ λ, height error δ h, north orientation air speed errorEast orientation air speed errorTrue course error delta φ, true air speed error delta V.
Step 8 is corrected navigational parameter according to the navigational parameter margin of error of estimation, completes accurately estimating for navigational parameter
Meter.
Navigational parameter calculation method are as follows:
By step 4 obtain by the dead reckoning of current time t calculate aircraft altitude L (t), longitude λ (t) respectively with step
Latitude error δ L, the longitude error δ λ of 7 Kalman Filter Estimations are added to obtain the current time t accurate latitude L'(t of aircraft), warp
It spends λ ' (t).
T true air speed V (t), the pressure altitude h (t) of the current time measurement that step 1 obtains and step 3 are calculated
True course φ (t) is respectively the same as the true air speed error delta V of step 7 Kalman Filter Estimation, height error δ h, true course error delta φ phase
Add to obtain the current time t accurate true air speed V'(t of aircraft), pressure altitude h'(t) and true course φ ' (t).
The north orientation wind speed of last moment t-1With east orientation wind speedEstimate respectively with step 7 Kalman filtering
Count obtained north orientation air speed errorEast orientation air speed errorAddition obtains the north orientation wind speed at current time at current time tWith east orientation wind speed
Utilize the accurate true air speed V'(t of current time t aircraft), true course φ ' (t), north orientation wind speedWith east orientation wind
SpeedCalculate the north orientation ground velocity V of current time aircraftn' (t), east orientation ground velocity Ve' (t), specifically:
The position at obtained current time and wind speed are fed back to dead-reckoning module by step 9, for completing subsequent time
The real-time accurate estimation of navigational parameter is realized in the reckoning of position by way of Recursive Solution.
Claims (6)
1. a kind of air navigation aid of radio auxiliary dead reckoning, it is characterized in that the method comprises the steps of:
Step 1: the parameter of the airborne each navigation equipment output of acquisition, the parameter include: the magnetic heading letter of aviation attitude system output
Breath；True air speed, the pressure altitude information of air data system output；The oblique distance information of rangefinder output；
Step 2: obtaining externally input initial position message, location information includes longitude, latitude；
Step 3: obtaining magnetic variation according to aircraft position by inquiry database, magnetic heading information being added to obtain with magnetic variation information
True course；
Step 4: being used according to last moment aircraft position and wind speed, the true course at current time, true air speed and pressure altitude
Dead reckoning method obtains the location information of current time reckoning；
Step 5: establishing state equation according to dead reckoning error model and oblique distance error model；
Step 6: establishing measurement equation using the location information of dead reckoning system, guidance station location information, oblique distance information；
Step 7: bonding state equation and measurement equation estimate navigational parameter error using kalman filter method；
Step 8: the navigational parameter margin of error according to estimation is corrected navigational parameter, the accurate estimation of navigational parameter is completed；
Step 9: the position at obtained current time and wind speed are fed back to dead-reckoning module, for completing subsequent time position
The real-time accurate estimation of navigational parameter is realized in the reckoning set by way of Recursive Solution.
2. air navigation aid according to claim 1, it is characterized in that: step 4 Air China position projectional technique are as follows:
According to the true air speed V (t) for the current time t that step 1 obtains, the true course φ (t) for the t moment that step 3 obtains and
The north orientation wind speed of last moment t-1East orientation wind speedCalculate the north orientation ground velocity V of current time tn(t) and
East orientation ground velocity Ve(t) are as follows:
Wherein north orientation wind speed, east orientation wind speed initial value be taken as zero；
According to the north orientation ground velocity V of the latitude L (t-1) of last moment t-1 and longitude λ (t-1), current time tn(t) and east orientation ground velocity
Ve(t), the pressure altitude h (t) for the current time t that step 1 obtains, the latitude L (t) of current time t is calculated using reckoning mode
With longitude λ (t):
L (t)=L (t-1)+Δ tVn(t)/(RM+h(t))
λ (t)=λ (t-1)+Δ tVe(t)/(RN+h(t))cosL(t-1)
Wherein, Δ t is that front and back time is poor, RN=Re(1+fsin2L), RM=Re(1-2f+3fsin2L), f=1/
298.257223563 being the compression of the Earth, Re=6378137.0 meters are earth radius.
3. air navigation aid according to claim 1, it is characterized in that: state equation in the step 5 specifically:
According to dead reckoning error model and oblique distance error model, aircraft altitude error delta L, longitude error δ λ, height error are chosen
δ h, north orientation air speed errorEast orientation air speed errorTrue course error delta φ, true air speed error delta V, DME1 oblique distance error delta
R1, DME2 oblique distance error delta R2As quantity of state X, it may be assumed thatAnd then it establishes
The state equation of Kalman filter:
Wherein state-transition matrix are as follows:
Wherein, τh、τnw、τew、τφ、τV、τRRespectively height, north orientation wind speed, east orientation wind speed, true course, true air speed, DME oblique distance
Single order Markov correlation time；
Noise matrix is
System noise vector are as follows: w=[nφ,wh,wnw,wew,wφ,wV,wR,wR]T；
Wherein, nφIt is true course white Gaussian noise, wh、wnw、wew、wφ、wV、wRRespectively height, north orientation wind speed, east orientation wind speed, true
Course, true air speed, DME oblique distance single order Markov noise.
4. air navigation aid according to claim 1, it is characterized in that: measurement equation in the step 6 specifically:
The oblique distance R measured according to i-th of DMEi, guidance station height hiAnd aircraft altitude h, it utilizes
Calculate DDMEiAs i-th DME measurement oblique distance in horizontal plane projector distance, the latitude L and longitude λ of dead reckoning and the are utilized
The latitude L of i DME guidance stationiWith longitude λiOblique distance is calculated in the projection D of horizontal planeDRi, i.e.,By DDMEiWith DDRiSubtract each other as measurement Z=[DDME1-DDR1
DDME2-DDR2]T, and then establish Kalman filter measurement equation:
Z (t)=H (t) X (t)+v (t)
H=[H in formula11 H12]
V=[v1,v2]T, v1And v2The white Gaussian noise of the zero-mean of respectively DME1 and DME2, i=1,2.
5. air navigation aid according to claim 1, it is characterized in that: navigational parameter error includes aircraft latitude in the step 7
Spend error delta L, longitude error δ λ, height error δ h, north orientation air speed errorEast orientation air speed errorTrue course error delta
φ, true air speed error delta V.
6. air navigation aid according to claim 1, it is characterized in that: navigational parameter calculation method in the step 8 are as follows:
By step 4 obtain by the dead reckoning of current time t calculate aircraft altitude L (t), longitude λ (t) respectively with step 7
Latitude error δ L, the longitude error δ λ of Kalman Filter Estimation are added to obtain the current time t accurate latitude L'(t of aircraft), warp
It spends λ ' (t)；
T true air speed V (t), the pressure altitude h (t) of the current time measurement that step 1 obtains and step 3 are calculated true
Course φ (t) is respectively the same as the true air speed error delta V of step 7 Kalman Filter Estimation, height error δ h, true course error delta φ phase
Add to obtain the current time t accurate true air speed V'(t of aircraft), pressure altitude h'(t) and true course φ ' (t)；
The north orientation wind speed of last moment t-1With east orientation wind speedIt is obtained respectively with step 7 Kalman Filter Estimation
The north orientation air speed error arrivedEast orientation air speed errorAddition obtains the north orientation wind speed at current time at current time t
With east orientation wind speed
Utilize the accurate true air speed V'(t of current time t aircraft), true course φ ' (t), north orientation wind speedWith east orientation wind speedCalculate the north orientation ground velocity V of current time aircraftn' (t), east orientation ground velocity Ve' (t), specifically:
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| CN111006659A (en) * | 2019-12-06 | 2020-04-14 | 江西洪都航空工业集团有限责任公司 | Navigation system with multi-navigation-source information fusion function |
| CN111947658A (en) * | 2020-06-30 | 2020-11-17 | 北京航天控制仪器研究所 | Low-cost autonomous navigation device and navigation method for communication-assisted positioning |
| 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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| 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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