CN103245963A - Double-antenna GNSS/INS deeply integrated navigation method and device - Google Patents

Double-antenna GNSS/INS deeply integrated navigation method and device Download PDF

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CN103245963A
CN103245963A CN2013101696133A CN201310169613A CN103245963A CN 103245963 A CN103245963 A CN 103245963A CN 2013101696133 A CN2013101696133 A CN 2013101696133A CN 201310169613 A CN201310169613 A CN 201310169613A CN 103245963 A CN103245963 A CN 103245963A
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China
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gps
global
navigation
error
satellite
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CN2013101696133A
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Chinese (zh)
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郭美凤
刘刚
张嵘
包超
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清华大学
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Abstract

The invention provides a double-antenna GNSS (Global Navigation Satellite System)/INS (Inertial Navigation System) deeply integrated navigation method and device. The device comprises a power supply module, an MEMS (Micro Electro Mechanical System) inertial measurement unit, a radio frequency module, an FPGA (Field Programmable Gate Array) module, a DSP (Digital Signal Processor) module and at least two antennas. According to the device, with the assistance of an INS, the probability of cycle slip of the GNSS can be reduced effectively, the integer ambiguity searching space of the GNSS is reduced, attitude measurement information is guaranteed to be continuously outputted, and the tracking loop robustness of the GNSS is enhanced; and with the assistance of a double-antenna GNSS, quick initial alignment of the INS can be realized, the attitude observability of the INS is improved, and the attitude divergence of the INS is inhibited.

Description

Double antenna GNSS/INS deep integrated navigation method and device
Technical field
The invention belongs to the satellite mobile communication technical field, be specifically related to a kind of double antenna GNSS/INS deep integrated navigation method and device.
Background technology
INS (Inertial Navigation System, inertial navigation system) has characteristics such as independence, strong anti-interference and good concealment, and can the high-speed and continuous outgoing position, navigational parameter such as speed, attitude, shortcoming is that navigation error is accumulated in time, can't satisfy the requirement that works long hours.GNSS (Global Navigation Satellite System, GPS (Global Position System)) can the whole world, round-the-clock high precision outgoing position, speed, temporal information, navigation results is accumulation in time not, but the navigation data turnover rate is generally less than 20Hz, and at height dynamically or can lose satellite-signal under the strong interference environment and navigation was lost efficacy.As seen, INS and GNSS have complementary advantages, and the two is used in combination, and can obtain that dynamic is good, the system of strong anti-interference, can high-speed and continuous output navigational parameter and accumulation in time of navigation error.
GNSS/MINS(Global Navigation Satellite System, GPS (Global Position System)/Micro-Electro-Mechanical Systems Inertial Navigation System, the micro electro mechanical inertia navigational system) integrated navigation system refers to the Unit with MIMU(MEMS Inertial Measurement, the micro electro mechanical inertia measuring unit) as sensor, proofread and correct the MINS error by reference informations such as GNSS receiver outgoing position/speed, by the navigational system of MINS outgoing position/speed assisted GNSS receiver acquisition tracking.The GNSS/MINS integrated navigation system is low because of its cost, volume is little, error advantages such as accumulation in time, high sensitivity, anti-high dynamic stress not, and good prospect is arranged in following aviation, vehicular applications.The integrated navigation algorithm has pine combination, tight combination, makes up three kinds deeply at present, and position-based/speed, pseudorange/pseudorange rates, I/Q accumulated value aspect make up respectively.
Because MIMU low precision, can't realize initial alignment by revolutions sensitively, therefore adopt the auxiliary aligning of GNSS receiver outgoing position/velocity information, but because the unobservable property of non-linear and attitude of inertial navigation system under the big misalignment situation causes aiming at and can't realize or aim at overlong time.
On the other hand, because integrated navigation system attitude observability is poor, under the long-time inorganic emotionally condition of carrier, attitude angle can be dispersed, if travel in the expressway such as automobile, long-time nothing is turned and acceleration and deceleration, with the Kalman filtering algorithm of prior art, attitude angle can be dispersed.
A kind of double antenna GPS/INS Integrated Navigation Instrument is disclosed among the patent CN201497509, introduce the attitude observed quantity by double antenna and attempt to solve the problem that attitude angle is dispersed, in this device, be blocked owing to satellite-signal, carrier-to-noise ratio is crossed the low cycle slip that causes easily, and owing to more easily lose satellite-signal, and will carry out long integer ambiguity search after losing satellite-signal, cause not having for a long time problems such as attitude amount output.
Summary of the invention
The present invention one of is intended to solve the problems of the technologies described above at least to a certain extent or provides a kind of useful commerce to select at least.For this reason, first purpose of the present invention is to propose a kind of double antenna GNSS/INS deep integrated navigation method, and second purpose of the present invention is to propose a kind of double antenna GNSS/INS deep integrated navigation device.
Double antenna GNSS/INS deep integrated navigation method according to the embodiment of the invention may further comprise the steps: S1. carries out initialization and parameter setting to GPS (Global Position System) and inertial navigation system respectively; S2. obtain the initial angle of pitch and the initial roll angle of described inertial navigation system; S3. obtain satellite navigation information by described GPS (Global Position System), and first pseudorange, first pseudorange rates, first angle of pitch and first course angle of the described GPS (Global Position System) that will obtain according to described satellite navigation information are as the observed quantity of Kalman filter; S4. according to described first pseudorange, described first pseudorange rates, described first angle of pitch and described first course angle described inertial navigation system is carried out big misalignment initial alignment; And the described inertial navigation system of S5. is resolved derive output inertial navigation system three-dimensional position, inertial navigation system three-dimensional velocity, second angle of pitch, second course angle and second roll angle by inertial navigation, and described inertial navigation system three-dimensional velocity fed back to described GPS (Global Position System), so that described GPS (Global Position System) is carried out the clocking error correction, make described inertial navigation system carry out the systematic error correction, and after correction, satellite-signal is followed the tracks of.
In an embodiment of the present invention, further comprise: the described GPS (Global Position System) of S6. is according to described satellite-signal, carry out navigation calculation and difference attitude algorithm, upgrade first pseudorange, first pseudorange rates, first angle of pitch and first course angle of the described GPS (Global Position System) of output, when described GPS (Global Position System) has output, enter step S7, enter step S5 during no-output; S7. the output result according to described inertial navigation system and described GPS (Global Position System) sets up Kalman filter error equation and Kalman filter observation equation, and described output result is carried out the integrated kalman filter algorithm, enters step S5.
In an embodiment of the present invention, the described GPS (Global Position System) of described step S6 is in described satellite-signal tracing process, monitor the state of track loop in real time and estimate carrier-to-noise ratio, judge whether losing lock of described track loop, if then under described inertial navigation system is auxiliary, described satellite-signal is carried out reacquisition.In an embodiment of the present invention, the state vector of described Kalman filter error equation is:
Wherein, geographic coordinate system is elected east northeast ground coordinate system as, subscript N represent north orientation, E represent east orientation, D represent to, δ L is longitude error, the rad of unit, δ λ are latitude error, the rad of unit, δ h are height error, the m of unit, δ V NBe the north orientation velocity error, the m/s of unit, δ V EBe the east orientation velocity error, the m/s of unit, δ V DFor ground to velocity error, the m/s of unit, Be north orientation attitude error angle, the rad of unit, Be east orientation attitude error angle, the rad of unit, For ground to the attitude error angle, the rad of unit, ε BxBe north gyro zero inclined to one side error, the rad/s of unit, ε ByBe east orientation gyro zero inclined to one side error, the rad/s of unit, ε BzFor ground to gyro zero error partially, the rad/s of unit, Be north orientation accelerometer bias error, the m/s of unit 2, Be east orientation accelerometer bias error, the m/s of unit 2, For ground to the accelerometer bias error, the m/s of unit 2, δ b is clock jitter, Be clock drift, the differential equation computing formula of Kalman filter error equation is:
δ L · = δV N R M + h - δh V N ( R M + h ) 2 δ h · = - δ V D
δ λ · = δV E R N + h sec L + δL V E R N + h tan L sec L - δh V E sec L ( R N + h ) 2
In an embodiment of the present invention, among the described step S7, adopt described GPS (Global Position System) to resolve the difference of described second pseudorange, described second pseudorange rates, described second angle of pitch and described second course angle that output described first pseudorange, described first pseudorange rates, described first angle of pitch and described first course angle and described inertial navigation system derive as the observed quantity of described Kalman filter.
In an embodiment of the present invention, the observed quantity of described Kalman filter is:
Z = δρ 1 · · · δρ n δ ρ · 1 · · · δ ρ · n δθ δγ T
Wherein, δ ρ iINSGPS, I=1,2 ..., n, n are observable number of satellite, ρ INSBe described second pseudorange of deriving according to described inertial navigation system three-dimensional position and satellite three-dimensional position, ρ GPSBe described first pseudorange that described GPS (Global Position System) is resolved output, Be described second pseudorange rates of deriving according to described inertial navigation system three-dimensional position, described inertial navigation system three-dimensional velocity and described satellite three-dimensional position, satellite three-dimensional velocity, Be described first pseudorange rates that described GPS (Global Position System) is resolved output, δ θ is angle of pitch error, and δ γ is the course angle error, and the observation equation of described Kalman filter is Z=HX+V, and wherein V is the observation noise sequence, and observing matrix is:
H = a n × 3 0 n × 3 0 n × 3 0 n × 3 0 n × 5 0 n × 3 a n × 3 0 n × 3 0 n × 3 0 n × 5 0 2 × 3 0 2 × 3 I 2 × 3 0 2 × 3 0 2 × 5
Wherein, a N * 3Be the satellite projection vector, the fundamental renewal equation of Kalman Filtering for Discrete device is:
X ^ k / k - 1 = Φ k , k - 1 X ^ k - 1
P k / k - 1 = Φ k , k - 1 P k - 1 Φ k , k - 1 T + Γ k - 1 Q k - 1 Γ k - 1 T
K k = P k / k - 1 H k T ( H k P k / k - 1 H k T + R k ) - 1
P k = ( I - K k H k ) P k / k - 1 ( I - K k H k ) T + K k P k K k T
X ^ k = X ^ k / k - 1 + K k ( Z k - H k X ^ k / k - 1 )
In an embodiment of the present invention, described GPS (Global Position System) is GPS navigation system, Beidou satellite navigation system, GLONASS navigational system or Galileo navigational system.
In an embodiment of the present invention, the antenna base length of described GPS (Global Position System) is at least 0.5 meter.
Double antenna GNSS/INS deep integrated navigation method according to the embodiment of the invention is introduced the GPS (Global Position System) track loop with inertial navigation system speed, can effectively reduce the probability that cycle slip occurs with the carrier tracking loop bandwidth reduction below 5Hz on the one hand.Secondly, if satellite-signal is lost, can catch satellite-signal fast down the auxiliary of inertial navigation system, guarantee that attitude exports enhancing GPS (Global Position System) track loop robustness continuously.Simultaneously, down auxiliary in described inertial navigation system attitude, can compress GPS (Global Position System) integer ambiguity search volume, accelerate GPS (Global Position System) attitude algorithm speed.
Double antenna GNSS/INS deep integrated navigation device according to the double antenna GNSS/INS deep integrated navigation method that is applied to first purpose of the present invention of the embodiment of the invention, comprise: power module, micro electro mechanical inertia measuring unit, radio-frequency module, FPGA module, DSP module and at least two antennas, wherein: described FPGA module comprises at least two correlator bank; Described DSP module is by the described GPS (Global Position System) satellite-signal of correlator bank demodulation of the described FPGA module of the total line traffic control of EMIF, finish track loop control, navigation calculation, the difference attitude algorithm of described GPS (Global Position System) in described DSP inside modules, and the inertial navigation that carries out described inertial navigation system resolves and the integrated kalman filter algorithm; Described radio-frequency module comprises at least two antennas, place along the carrier longitudinal axis, after its described satellite-signal is sampled through down coversion, intermediate frequency data is imported described FPGA module and is finished Base-Band Processing, in described DSP inside modules described inertial navigation system inertial navigation is resolved the described three-dimensional velocity that obtains and feed back to described GPS (Global Position System) track loop, described GPS (Global Position System) is calculated the carrier doppler frequency with the described satellite-signal of auxiliary tracking.
In an embodiment of the present invention, in the described satellite-signal tracing process of described GPS (Global Position System), monitor the state of described track loop in real time and estimate carrier-to-noise ratio, judge whether losing lock of described track loop, if then under described inertial navigation system is auxiliary, described satellite-signal is carried out reacquisition.
In an embodiment of the present invention, described GPS (Global Position System) is GPS navigation system, Beidou satellite navigation system, GLONASS navigational system or Galileo navigational system.
In an embodiment of the present invention, the antenna base length of described GPS (Global Position System) is at least 0.5 meter.
Double antenna GNSS/INS deep integrated navigation device according to the embodiment of the invention, use double antenna GPS (Global Position System) difference attitude algorithm, utilize this attitude to carry out big misalignment initial alignment and retrain attitude (second course angle, second roll angle and second angle of pitch) and disperse, thereby solve the problem that big misalignment initial alignment problem and attitude (second course angle, second roll angle and second angle of pitch) are dispersed.Simultaneously the inertial navigation system three-dimensional velocity is introduced the GPS (Global Position System) track loop, on the one hand can be with the carrier tracking loop bandwidth reduction below 5Hz, effectively reduce the probability that cycle slip occurs, secondly, if satellite-signal is lost, can catch satellite-signal fast down the auxiliary of inertial navigation system, guarantee that attitude exports the robustness of enhancing GPS (Global Position System) track loop continuously.Simultaneously, down auxiliary in described inertial navigation system attitude, can compress GPS (Global Position System) integer ambiguity search volume, accelerate GPS (Global Position System) attitude algorithm speed.
Additional aspect of the present invention and advantage part in the following description provide, and part will become obviously from the following description, or recognize by practice of the present invention.
Description of drawings
Above-mentioned and/or additional aspect of the present invention and advantage are from obviously and easily understanding becoming the description of embodiment in conjunction with following accompanying drawing, wherein:
Fig. 1 is the algorithm principle figure of the double antenna GNSS/INS deep integrated navigation method of the embodiment of the invention;
Fig. 2 is a kind of algorithm flow chart of double antenna GNSS/INS deep integrated navigation method;
Fig. 3 is the process flow diagram of the double antenna GNSS/INS deep integrated navigation method of the embodiment of the invention;
Fig. 4 is the configuration diagram of double antenna GNSS/INS deep integrated navigation device;
Fig. 5 is the embodiment synoptic diagram of the double antenna GNSS/INS deep integrated navigation device of the embodiment of the invention;
Fig. 6 is the simulation result figure of embodiment of the double antenna GNSS/INS deep integrated navigation device of the embodiment of the invention;
Fig. 7 is another simulation result figure of embodiment of the double antenna GNSS/INS deep integrated navigation device of the embodiment of the invention.
Embodiment
Describe embodiments of the invention below in detail, the example of described embodiment is shown in the drawings, and wherein identical or similar label is represented identical or similar elements or the element with identical or similar functions from start to finish.Be exemplary below by the embodiment that is described with reference to the drawings, be intended to for explaining the present invention, and can not be interpreted as limitation of the present invention.
In description of the invention, it will be appreciated that, term " " center "; " vertically "; " laterally "; " length "; " width "; " thickness ", " on ", D score, " preceding ", " back ", " left side ", " right side ", " vertically ", " level ", " top ", " end " " interior ", " outward ", " clockwise ", close the orientation of indications such as " counterclockwise " or position is based on orientation shown in the drawings or position relation, only be that the present invention for convenience of description and simplification are described, rather than device or the element of indication or hint indication must have specific orientation, with specific orientation structure and operation, therefore can not be interpreted as limitation of the present invention.
In addition, term " first ", " second " only are used for describing purpose, and can not be interpreted as indication or hint relative importance or the implicit quantity that indicates indicated technical characterictic.Thus, one or more these features can be expressed or impliedly be comprised to the feature that is limited with " first ", " second ".In description of the invention, the implication of " a plurality of " is two or more, unless clear and definite concrete restriction is arranged in addition.
In the present invention, unless clear and definite regulation and restriction are arranged in addition, broad understanding should be done in terms such as term " installation ", " linking to each other ", " connection ", " fixing ", for example, can be fixedly connected, also can be to removably connect, or connect integratedly; Can be mechanical connection, also can be to be electrically connected; Can be directly to link to each other, also can link to each other indirectly by intermediary, can be the connection of two element internals.For the ordinary skill in the art, can understand above-mentioned term concrete implication in the present invention as the case may be.
In the present invention, unless clear and definite regulation and restriction are arranged in addition, first feature second feature it " on " or D score can comprise that first and second features directly contact, can comprise that also first and second features are not directly contacts but by the contact of the additional features between them.And, first feature second feature " on ", " top " and " above " comprise first feature directly over second feature and oblique upper, or only represent that the first characteristic level height is higher than second feature.First feature second feature " under ", " below " and " below " comprise first feature under second feature and tiltedly, or only represent that the first characteristic level height is less than second feature.
As shown in Figure 1, be the algorithm principle figure according to the double antenna GNSS/INS deep integrated navigation method of the embodiment of the invention, wherein, GNSS is the abbreviated form of English Global Navigation Satellite System, concrete Chinese lexical or textual analysis is GPS (Global Position System), INS is the abbreviated form of English Inertial Navigation System, and concrete Chinese lexical or textual analysis is inertial navigation system.Inertial navigation system utilizes micro electro mechanical inertia measuring unit output specific force, angular speed to carry out inertial navigation and resolve, obtain inertial navigation system three-dimensional position, three-dimensional velocity, attitude (comprising second course angle, second angle of pitch and second roll angle), wherein three-dimensional velocity feeds back to the GPS (Global Position System) track loop, calculates the carrier doppler frequency with assisted global satellite navigation system tracking satellite signal.Output first pseudorange, first pseudorange rates, first angle of pitch and first course angle behind GPS (Global Position System) navigation calculation and the difference attitude algorithm, be used for the observed quantity of computer card Thalmann filter, the Kalman filter estimate of error is used for revising INS errors and GPS (Global Position System) clocking error.
As shown in Figure 2, be a kind of algorithm flow chart of double antenna GNSS/INS deep integrated navigation method, and in conjunction with the process flow diagram of double antenna GNSS/INS deep integrated navigation method according to the embodiment of the invention shown in Figure 3, may further comprise the steps:
S1. respectively GPS (Global Position System) and inertial navigation system are carried out initialization and parameter setting.
S2. obtain the initial angle of pitch and the initial roll angle of inertial navigation system.Comprise the estimation of inertial navigation system leveling and zero-bit.
S3. obtain satellite navigation information by GPS (Global Position System), and first pseudorange, first pseudorange rates, first angle of pitch and first course angle of the GPS (Global Position System) that will obtain according to satellite navigation information are as the observed quantity of Kalman filter.
According to the double antenna GNSS/INS deep integrated navigation method of the embodiment of the invention, GPS (Global Position System) is GPS navigation system, Beidou satellite navigation system, GLONASS navigational system or Galileo navigational system.
S4. first pseudorange, first pseudorange rates, first angle of pitch and first course angle are carried out big misalignment initial alignment to inertial navigation system.And
S5. inertial navigation system is resolved derive output inertial navigation system three-dimensional position, inertial navigation system three-dimensional velocity, second angle of pitch, second course angle and second roll angle by inertial navigation, and the inertial navigation system three-dimensional velocity fed back to GPS (Global Position System), so that GPS (Global Position System) is carried out the clocking error correction, make inertial navigation system carry out the systematic error correction, and after correction, satellite-signal is followed the tracks of.
S6. GPS (Global Position System) is according to satellite-signal, carry out navigation calculation and difference attitude algorithm, upgrade first pseudorange, first pseudorange rates, first angle of pitch and first course angle of output GPS (Global Position System), when GPS (Global Position System) has output, enter step S7, enter step S5 during no-output.
GPS (Global Position System) is monitored the state of track loop in real time and is estimated carrier-to-noise ratio in the satellite-signal tracing process, judges whether losing lock of track loop, if then down satellite-signal is carried out reacquisition in that inertial navigation system is auxiliary.
S7. the output result according to inertial navigation system and GPS (Global Position System) sets up Kalman filter error equation and Kalman filter observation equation, and the result carries out the integrated kalman filter algorithm to output, enters step S5.
Adopt GPS (Global Position System) to resolve the difference of first pseudorange, first pseudorange rates, first angle of pitch and first course angle of output and second pseudorange, second pseudorange rates, second angle of pitch and second course angle that inertial navigation system is derived as the observed quantity of Kalman filter.
The state vector of Kalman filter error equation is:
Wherein, geographic coordinate system is elected east northeast ground coordinate system as, subscript N represent north orientation, E represent east orientation, D represent to, δ L is longitude error, the rad of unit, δ λ are latitude error, the rad of unit, δ h are height error, the m of unit, δ V NBe the north orientation velocity error, the m/s of unit, δ V EBe the east orientation velocity error, the m/s of unit, δ V DFor ground to velocity error, the m/s of unit, Be north orientation attitude error angle, the rad of unit, Be east orientation attitude error angle, the rad of unit, For ground to the attitude error angle, the rad of unit, ε BxBe north gyro zero inclined to one side error, the rad/s of unit, ε ByBe east orientation gyro zero inclined to one side error, the rad/s of unit, ε BzFor ground to gyro zero error partially, the rad/s of unit, Be north orientation accelerometer bias error, the m/s of unit 2, Be east orientation accelerometer bias error, the m/s of unit 2, For ground to the accelerometer bias error, the m/s of unit 2, δ b is clock jitter, Be clock drift,
The differential equation computing formula of Kalman filter error equation is:
δ L · = δV N R M + h - δh V N ( R M + h ) 2 δ h · = - δ V D
δ λ · = δV E R N + h sec L + δL V E R N + h tan L sec L - δh V E sec L ( R N + h ) 2
The observed quantity of Kalman filter is:
Z = δρ 1 · · · δρ n δ ρ · 1 · · · δ ρ · n δθ δγ T
Wherein, δ ρ iINSGPS, I=1,2 ..., n, n are observable number of satellite, ρ INSBe second pseudorange of deriving according to three-dimensional position and the satellite three-dimensional position of inertial navigation system, ρ GPSBe first pseudorange that GPS (Global Position System) is resolved output, Be three-dimensional position, three-dimensional velocity and the satellite three-dimensional position according to inertial navigation system, second pseudorange rates that the satellite three-dimensional velocity is derived, Be first pseudorange rates that GPS (Global Position System) is resolved output, δ θ is angle of pitch error, and namely first angle of pitch and second angle of pitch is poor, and δ γ is the course angle error, and namely first course angle and second course angle is poor
The observation equation of Kalman filter is Z=HX+V, and wherein V is the observation noise sequence, and observing matrix is:
H = a n × 3 0 n × 3 0 n × 3 0 n × 3 0 n × 5 0 n × 3 a n × 3 0 n × 3 0 n × 3 0 n × 5 0 2 × 3 0 2 × 3 I 2 × 3 0 2 × 3 0 2 × 5
Wherein, a N * 3Be the satellite projection vector,
The fundamental renewal equation of Kalman Filtering for Discrete device is:
X ^ k / k - 1 = Φ k , k - 1 X ^ k - 1
P k / k - 1 = Φ k , k - 1 P k - 1 Φ k , k - 1 T + Γ k - 1 Q k - 1 Γ k - 1 T
K k = P k / k - 1 H k T ( H k P k / k - 1 H k T + R k ) - 1
P k = ( I - K k H k ) P k / k - 1 ( I - K k H k ) T + K k P k K k T
X ^ k = X ^ k / k - 1 + K k ( Z k - H k X ^ k / k - 1 )
In above-mentioned Kalman filter error equation, remove influences in a small amount such as earth rotation, kalman filter state transition matrix F and observing matrix H are respectively:
F = 0 I 0 0 0 0 0 0 0 f n × C b n 0 0 0 0 0 0 0 - C b n 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 0 0 0 0 0 0 0 , H = a n × 3 0 n × 3 0 n × 3 0 n × 3 0 n × 5 0 n × 3 a n × 3 0 n × 3 0 n × 3 0 n × 5 0 2 × 3 0 2 × 3 I 2 × 3 0 2 × 3 0 2 × 5
I wherein 2 * 3The expression order is a matrix of 2.Observability matrix is:
Q j = H HF HF 2 HF 3 . . . HF 14 = a n × 3 0 0 0 0 0 a n × 3 0 0 0 0 0 I 2 × 3 0 0 0 a n × 3 0 0 0 0 0 f n × C b n 0 0 0 0 0 - I 2 × 3 * C b n 0 0 f n × C b n 0 0 0 0 0 - f n × C b n 0 0 0 0 0 . . . . . . . . . . . . . . .
As seen, double antenna GNSS/INS deep integrated navigation system only needs to accelerate or deceleration along y direction, and observability matrix has reduced the requirement motor-driven to carrier with regard to full rank.Need to prove that the antenna base length of GPS (Global Position System) is at least 0.5 meter.
The observed quantity of the Kalman filter of present embodiment has added attitude (first course angle and first angle of pitch), has effectively improved the observability of system state.Owing to there is attitude auxiliary, can obtain second course angle of system fast, finish aligning.Under the long-term inorganic emotionally condition of carrier, assurance system attitude is not dispersed.
The double antenna GNSS/INS deep integrated navigation method of the embodiment of the invention is introduced the GPS (Global Position System) track loop with inertial navigation system speed, can effectively reduce the probability that cycle slip occurs with the carrier tracking loop bandwidth reduction below 5Hz on the one hand.Secondly, if satellite-signal is lost, can catch satellite-signal fast down the auxiliary of inertial navigation system, guarantee that attitude exports enhancing GPS (Global Position System) track loop robustness continuously.Simultaneously, down auxiliary in described inertial navigation system attitude, can compress GPS (Global Position System) integer ambiguity search volume, accelerate GPS (Global Position System) attitude algorithm speed.
As shown in Figure 4, the configuration diagram for double antenna GNSS/INS deep integrated navigation device comprises:
Power module 100, micro electro mechanical inertia measuring unit 200, radio-frequency module 300, FPGA module 400, DSP module 500, at least two antennas 600 and 700.
The double antenna 600 and 700 of GPS (Global Position System), antenna 600 and 700 base length are at least 0.5 meter, place along the carrier longitudinal axis, can measure first angle of pitch and first course angle of carrier like this.Two antennas 600 and 700 with satellite-signal through the down coversion sampling after, intermediate frequency data input FPGA module 400 is finished Base-Band Processing, DSP module 500, be two correlator bank 410 and the 420 correct demodulation GPS (Global Position System) satellite-signals that digital signal processing module passes through the total line traffic control FPGA module 400 of EMIF, finish track loop control, navigation calculation, the difference attitude algorithm of GPS (Global Position System) in DSP module 500 inside, and the inertial navigation that carries out inertial navigation system resolves and the integrated kalman filter algorithm.In DSP module 500 inside the inertial navigation system inertial navigation is resolved the three-dimensional velocity that obtains and feed back to the GPS (Global Position System) track loop, GPS (Global Position System) is calculated the carrier doppler frequency with auxiliary tracking satellite signal.In the satellite-signal tracing process of GPS (Global Position System), monitor the state of track loop in real time and estimate carrier-to-noise ratio, judge whether losing lock of track loop, if then down satellite-signal is carried out reacquisition in that inertial navigation system is auxiliary.Need to prove that present embodiment can compatible all GPS (Global Position System), comprise GPS navigation system, big-dipper satellite positioning system, glonass system or Galileo navigational system.
As shown in Figure 5, be the embodiment synoptic diagram according to the double antenna GNSS/INS deep integrated navigation device of the embodiment of the invention, the aerial flight path of turning of the aircraft of its design, the linear velocity of aircraft is 200m/s, gyroscope zero stability partially is 30 °/h, and accelerometer bias stability is 2mg.Use the intermediate frequency satellite signal simulator to generate the GPS (Global Position System) intermediate frequency data, 7 satellites in view are set.Adopt Shuangzi sample inertial navigation algorithm, GPS (Global Position System) pre-detection integral time is 20ms, and carrier tracking loop uses 2 rank FLL(Frequency Lock Loop, FLL) auxiliary 3 rank PLL(Phase Lock Loop, phaselocked loop) structure, and use carrier wave ring auxiliary code ring to follow the tracks of., GPS (Global Position System) carrier wave ring FLL bandwidth 10Hz, PLL bandwidth 5Hz, sign indicating number ring DLL(Delay Lock Loop, delay lock loop auxiliary down in inertial navigation system) bandwidth is 1Hz, loop stability is followed the tracks of.Simulation result as shown in Figure 6 and Figure 7.As seen, integrated navigation system three-dimensional position error of the present invention is less than 3m, and the three-dimensional velocity error is less than 0.5m/s, and attitude error is less than 0.5 degree.Wherein as shown in Figure 6, pitch angle deviation is the difference between second angle of pitch of second angle of pitch derived of inertial navigation system and emulation, in like manner the roll angle deviation is the difference between second roll angle of second roll angle derived of inertial navigation system and emulation, and the course angle deviation is the difference between second course angle of second course angle derived of inertial navigation system and emulation.
Double antenna GNSS/INS deep integrated navigation device according to the embodiment of the invention, use double antenna GPS (Global Position System) difference attitude algorithm, utilize this attitude to carry out big misalignment initial alignment and retrain attitude (second course angle, second roll angle and second angle of pitch) and disperse, thereby solve the problem that big misalignment initial alignment problem and attitude (second course angle, second roll angle and second angle of pitch) are dispersed.Simultaneously the inertial navigation system three-dimensional velocity is introduced the GPS (Global Position System) track loop, on the one hand can be with the carrier tracking loop bandwidth reduction below 5Hz, effectively reduce the probability that cycle slip occurs, secondly, if satellite-signal is lost, can catch satellite-signal fast down the auxiliary of inertial navigation system, guarantee that attitude exports the robustness of enhancing GPS (Global Position System) track loop continuously.Simultaneously, down auxiliary in described inertial navigation system attitude, can compress GPS (Global Position System) integer ambiguity search volume, accelerate GPS (Global Position System) attitude algorithm speed.
Describe and to be understood that in the process flow diagram or in this any process of otherwise describing or method, expression comprises module, fragment or the part of code of the executable instruction of the step that one or more is used to realize specific logical function or process, and the scope of preferred implementation of the present invention comprises other realization, wherein can be not according to order shown or that discuss, comprise according to related function by the mode of basic while or by opposite order, carry out function, this should be understood by the embodiments of the invention person of ordinary skill in the field.
In the description of this instructions, concrete feature, structure, material or characteristics that the description of reference term " embodiment ", " some embodiment ", " example ", " concrete example " or " some examples " etc. means in conjunction with this embodiment or example description are contained at least one embodiment of the present invention or the example.In this manual, the schematic statement to above-mentioned term not necessarily refers to identical embodiment or example.And concrete feature, structure, material or the characteristics of description can be with the suitable manner combination in any one or more embodiment or example.
Although illustrated and described embodiments of the invention above, be understandable that, above-described embodiment is exemplary, can not be interpreted as limitation of the present invention, those of ordinary skill in the art can change above-described embodiment under the situation that does not break away from principle of the present invention and aim within the scope of the invention, modification, replacement and modification.

Claims (12)

1. a double antenna GNSS/INS deep integrated navigation method is characterized in that, may further comprise the steps:
S1. respectively GPS (Global Position System) and inertial navigation system are carried out initialization and parameter setting;
S2. obtain the initial angle of pitch and the initial roll angle of described inertial navigation system;
S3. obtain satellite navigation information by described GPS (Global Position System), and first pseudorange, first pseudorange rates, first angle of pitch and first course angle of the described GPS (Global Position System) that will obtain according to described satellite navigation information are as the observed quantity of Kalman filter;
S4. according to described first pseudorange, described first pseudorange rates, described first angle of pitch and described first course angle described inertial navigation system is carried out big misalignment initial alignment; And
S5. described inertial navigation system is resolved derive output inertial navigation system three-dimensional position, inertial navigation system three-dimensional velocity, second angle of pitch, second course angle and second roll angle by inertial navigation, and described inertial navigation system three-dimensional velocity fed back to described GPS (Global Position System), so that described GPS (Global Position System) is carried out the clocking error correction, make described inertial navigation system carry out the systematic error correction, and after correction, satellite-signal is followed the tracks of.
2. double antenna GNSS/INS deep integrated navigation method as claimed in claim 1 is characterized in that, further comprises:
S6. described GPS (Global Position System) is according to described satellite-signal, carry out navigation calculation and difference attitude algorithm, upgrade first pseudorange, first pseudorange rates, first angle of pitch and first course angle of the described GPS (Global Position System) of output, when described GPS (Global Position System) has output, enter step S7, enter step S5 during no-output;
S7. the output result according to described inertial navigation system and described GPS (Global Position System) sets up Kalman filter error equation and Kalman filter observation equation, and described output result is carried out the integrated kalman filter algorithm, enters step S5.
3. as claim 1 and 2 described double antenna GNSS/INS deep integrated navigation methods, it is characterized in that, the described GPS (Global Position System) of described step S6 is in described satellite-signal tracing process, monitor the state of track loop in real time and estimate carrier-to-noise ratio, judge whether losing lock of described track loop, if then under described inertial navigation system is auxiliary, described satellite-signal is carried out reacquisition.
4. as the described double antenna GNSS/INS of claim 1 to 3 deep integrated navigation method, it is characterized in that the state vector of described Kalman filter error equation is:
Wherein, geographic coordinate system is elected east northeast ground coordinate system as, subscript N represent north orientation, E represent east orientation, D represent to, δ L is longitude error, the rad of unit, δ λ are latitude error, the rad of unit, δ h are height error, the m of unit, δ V NBe the north orientation velocity error, the m/s of unit, δ V EBe the east orientation velocity error, the m/s of unit, δ V DFor ground to velocity error, the m/s of unit, Be north orientation attitude error angle, the rad of unit, Be east orientation attitude error angle, the rad of unit, For ground to the attitude error angle, the rad of unit, ε BxBe north gyro zero inclined to one side error, the rad/s of unit, ε ByBe east orientation gyro zero inclined to one side error, the rad/s of unit, ε BzFor ground to gyro zero error partially, the rad/s of unit, Be north orientation accelerometer bias error, the m/s of unit 2, Be east orientation accelerometer bias error, the m/s of unit 2, For ground to the accelerometer bias error, the m/s of unit 2, δ b is clock jitter, Be clock drift,
The differential equation computing formula of Kalman filter error equation is:
5. double antenna GNSS/INS deep integrated navigation method as claimed in claim 2, it is characterized in that, among the described step S7, adopt described GPS (Global Position System) to resolve the difference of described second pseudorange, described second pseudorange rates, described second angle of pitch and described second course angle that output described first pseudorange, described first pseudorange rates, described first angle of pitch and described first course angle and described inertial navigation system derive as the observed quantity of described Kalman filter.
6. as the described double antenna GNSS/INS of claim 1 to 5 deep integrated navigation method, it is characterized in that the observed quantity of described Kalman filter is:
Wherein, δ ρ iINSGPS, I=1,2 ..., n, n are observable number of satellite, ρ INSBe described second pseudorange of deriving according to described inertial navigation system three-dimensional position and satellite three-dimensional position, ρ GPSBe described first pseudorange that described GPS (Global Position System) is resolved output, Be described second pseudorange rates of deriving according to described inertial navigation system three-dimensional position, described inertial navigation system three-dimensional velocity and described satellite three-dimensional position, satellite three-dimensional velocity, Be described first pseudorange rates that described GPS (Global Position System) is resolved output, δ θ is angle of pitch error, and δ γ is the course angle error,
The observation equation of described Kalman filter is Z=HX+V, and wherein V is the observation noise sequence, and observing matrix is:
Wherein, a N * 3Be the satellite projection vector,
The fundamental renewal equation of Kalman Filtering for Discrete device is:
7. double antenna GNSS/INS deep integrated navigation method as claimed in claim 1 is characterized in that, described GPS (Global Position System) is GPS navigation system, Beidou satellite navigation system, GLONASS navigational system or Galileo navigational system.
8. double antenna GNSS/INS deep integrated navigation method as claimed in claim 1 is characterized in that the antenna base length of described GPS (Global Position System) is at least 0.5 meter.
9. a double antenna GNSS/INS deep integrated navigation device that is applied to each described double antenna GNSS/INS deep integrated navigation method among the claim 1-8 is characterized in that, comprising:
Power module, micro electro mechanical inertia measuring unit, radio-frequency module, FPGA module, DSP module and at least two antennas, wherein:
Described FPGA module comprises at least two correlator bank;
Described DSP module is by the described GPS (Global Position System) satellite-signal of correlator bank demodulation of the described FPGA module of the total line traffic control of EMIF, finish track loop control, navigation calculation, the difference attitude algorithm of described GPS (Global Position System) in described DSP inside modules, and the inertial navigation that carries out described inertial navigation system resolves and the integrated kalman filter algorithm;
Described radio-frequency module comprises at least two antennas, place along the carrier longitudinal axis, after its described satellite-signal is sampled through down coversion, intermediate frequency data is imported described FPGA module and is finished Base-Band Processing, in described DSP inside modules described inertial navigation system inertial navigation is resolved the described three-dimensional velocity that obtains and feed back to described GPS (Global Position System) track loop, described GPS (Global Position System) is calculated the carrier doppler frequency with the described satellite-signal of auxiliary tracking.
10. double antenna GNSS/INS deep integrated navigation device as claimed in claim 9, it is characterized in that, in the described satellite-signal tracing process of described GPS (Global Position System), monitor the state of described track loop in real time and estimate carrier-to-noise ratio, judge whether losing lock of described track loop, if then under described inertial navigation system is auxiliary, described satellite-signal is carried out reacquisition.
11., it is characterized in that described GPS (Global Position System) is GPS navigation system, Beidou satellite navigation system, GLONASS navigational system or Galileo navigational system as the described double antenna GNSS/INS of claim 9 to 10 deep integrated navigation device.
12. a kind of double antenna GNSS/INS deep integrated navigation device as claimed in claim 9 is characterized in that the antenna base length of described GPS (Global Position System) is at least 0.5 meter.
CN2013101696133A 2013-05-09 2013-05-09 Double-antenna GNSS/INS deeply integrated navigation method and device CN103245963A (en)

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