CN101256080A - Midair aligning method for satellite/inertia combined navigation system - Google Patents

Abstract

The invention provides a GNSS/INS integrated navigation system in-flight alignment method, which belongs to the inertial navigation system. The alignment method includes the following steps: satellite navigation receiver locator acquisition; satellite navigation locator data caching; inertia measurement cell data acquisition; aircraft motion acceleration calculation; three dimension proportion calculation in navigation coordinate system; fast data solution of pitch angle and roll angle trigonometric equation; navigation parameter of strap-down inertial navigation system initialization and so on. The invention can realize the initial alignment in the air of strap-down inertial navigation system in carrier in flight, which solves the problem of fast alignment of strap-down inertial navigation system in flight in airplane, missile and other moving carriers.

Description

The aerial alignment methods of satellite/inertia combined navigation system

One, technical field

The invention belongs to the initial alignment technical field of inertial navigation system, especially initial attitude autoregistration under the engine-bed condition and simulating, verifying technology thereof.

Two, background technology

Initial alignment is distinctive duty of inertial navigation system and process, the given initial velocity of its function, initial position, initial attitude.The given ratio of initial velocity, initial position is easier to, and obtains key and difficult point that the high precision initial attitude is an initial alignment fast.Usually, initial alignment mainly refers to the attitude aligning.According to the difference of alignment so, attitude is aimed to be divided into autoregistration and to be based upon main/sub-inertial navigation information and is transmitted the Transfer Alignment on the basis.The attitude autoregistration mainly is to carry out under the situation of system quiescence, as the horizontal attitude benchmark, utilizes rotational-angular velocity of the earth as azimuth reference with the terrestrial gravitation direction.

From prior art, aerial aligning in the past mainly is a Transfer Alignment, and application mainly is an air weapon.Realize the initial attitude autoregistration aloft in-flight, the initial alignment under the conation attitude condition especially in office then rarely has research, and this mainly is because the difficulty demand bigger and actual engineering system that aerial dynamic condition is aimed at down is less.

The domestic and international open source literature of aerial alignment applications and research is less, mainly comprises:

In August, (1) 1998 US Airways cosmonautics navigation, guidance and control meeting (AIAA Guidance, Navigation, and Control Conference) is numbered the method that inertia/GPS integrated navigation system has adopted aerial statue to aim in the paper " Assesment of integrated GPS/INS for the EX-171 Extended RangeGuided Munition " of AIAA-98-4416 and paper " the AGPS/INS Guidance System for Navy 5 Projectiles " U.S.'s guided cartridge of promoting to a higher rank that U.S.'s Draper laboratory periodical in 1997 is delivered, but, all do not announce know-why and details in the relevant document about this gordian technique;

(2) domestic aspect, in October, 2007, the 15th volume the 52nd phase " Chinese inertial technology " (568-572 page or leaf) publication paper " the airborne SINS/GPS based on GPS observed quantity and model prediction error aims in the air ", in November, 2007, the 28th volume the 6th phase " aviation journal " (1395-1400 page or leaf) publication paper " the aerial rapid alignment technology of MIMU under the big misalignment ".The main filtering problem that aerial aligning is discussed in these two pieces of papers, this filtering problem is the fine alignment stage behind the coarse alignment, attitude for initial time is obtained, it is aerial coarse alignment, then require the dynamic process of aircraft has been proposed extra demand, the principle that adopts static accelerometer sensitive acceleration of gravity down to determine attitude realizes the coarse alignment in the airflight.

(3) the existing disclosed patent status inquiry of the U.S. and China is shown not aerial aligning the, the patent of especially aerial coarse alignment aspect.

The situation of comprehensive above patent both at home and abroad and open source literature, relate in the air on time, general disposal route is by the flight dynamic process is retrained, make aircraft flight as far as possible steadily, aerial engine-bed alignment issues is simplified, suppose the only responsive acceleration of gravity of accelerometer, adopt utilization under the quiet support directly to utilize the ratio between accelerometer output and the acceleration of gravity to determine rough initial level attitude, compensate in the follow-up method of filtering that adopts in-flight again.In the coarse alignment constraint of the dynamic many places of flying and approximate staticize processing are limited the applicability of method greatly, and can not be used for the aerial aligning under the big maneuvering condition, and the aerial exactly alignment issues of big maneuvering condition must satisfy.

, civil aviation aircraft military along with a new generation is to the further raising of reliability requirement, making strapdown inertial navigation system possess the over-the-air rekey kinetic force is the key of guaranteeing that aircraft resumes operation in some fault/state of necessity, in addition for some guidance chemical weapons device, after emission, just begin to work on power, the first step of its course of work just needs to determine the initial parameter of navigational system under the high at a high speed dynamically flying condition aloft.

Three, summary of the invention

Fundamental purpose of the present invention is, solve the aloft initial alignment problem of strapdown inertial navigation system in the motion carriers such as aircraft, guided missile aloft, in that flight dynamic process and having a smooth flight property are not added under the situation of any restrictive condition, obtain initial position, speed and attitude information.The present invention is specially adapted to in-flight, and aircraft breaks down, when restarting, satellite/inertia combined navigation system obtains initial navigation information, also be specially adapted to some special aircraft, the system that must after taking off, could begin to work on power as guided cartridge etc.

One, aerial coarse alignment method

Airborne satellite/inertia combined navigation system comprehensive utilization satellite and inertia measurement information, by the exact relationship expression formula of inertial navigation system specific force equation foundation about attitude, utilize the numerical solution algorithm, realize aiming in the air, its characteristics are that this method realizes by following steps:

(1) the satellite navigation receiver locator data is gathered: integrated navigation computer is with period T _GContinue to read the navigation information of the specific format of satellite navigation receiver output from the interface that links to each other with satellite navigation receiver, and understand, obtain real-time three-dimensional position, three-dimensional velocity and the flight path azimuthangle of aircraft, wherein, three-dimensional position comprises longitude λ, latitude L and height H, the east orientation speed V under the local geographic coordinate system _E, east orientation speed V _NWith east orientation speed V _U, flight path azimuthangle ψ;

(2) satellite navigation locator data buffer memory: in the memory headroom variable that three-dimensional position, three-dimensional velocity and the flight path azimuthangle of the satellite navigation receiver output of reading in integrated navigation computer in the step (1) and temporal information are kept at navigational computer in the lump, the variable behind the buffer memory is successively with sign of lambda _K-1Expression longitude, L _K-1The expression latitude, H _K-1The expression height, V _{E, k-1}, V _{N, k-1}, V _{U, k-1}Represent respectively east orientation, north orientation and day to speed, ψ _K-1The expression flight path azimuthangle, t _K-1Express time information;

(3) Inertial Measurement Unit data acquisition: immediately following step (2), by analog to digital conversion or serial port, with period T _IRead the three dimensional angular speed and the three-dimensional specific force of gyro and accelerometer measures in the Inertial Measurement Unit.The information that the quick-connecting inertia measurement unit records is aircraft angular motion and the projection of line motion under body axis system b, and three dimensional angular speed and three-dimensional specific force be ω respectively _Ib ^Bx, ω _Ib ^By, ω _Ib ^BzAnd f _x ^b, f _y ^b, f _z ^b, T wherein _GBe T _IIntegral multiple, x, y, z represent three coordinate axis of body system.

(4) aircraft movements acceleration calculation: obtain satellite navigation data first, and through at least 2 period T _GTime after, understand t constantly finishing GPS information _k, longitude λ is arranged in the integrated navigation computer _k, latitude L _k, height H _k, east orientation direction speed V _{E, k}, north orientation speed V _{N, k}, the sky is to speed V _{U, k}, course angle ψ _k, the acceleration of motion of pressing following front and back method of difference calculating aircraft:

$a_E=\frac{V_{E,k}-V_{E,k-1}}{T_G}$

$a_N=\frac{V_{N,k}-V_{N,k-1}}{T_G};$

$a_U=\frac{V_{U,k}-V_{U,k-1}}{T_G}$

(5) real time position, speed, the acceleration of motion that step (1), step (3) and step (4) are obtained, substitution strapdown inertial navigation system specific force equation calculates the three-dimensional specific force under the navigation system, and its computing method are as follows:

$\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]=\left[\begin{array}{c}{a}_E\\ a_N\\ a_U\end{array}\right]+\left[\begin{array}{ccc}0& -(2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)\\ (2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& 0& -(2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)\\ -(2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)& 0\end{array}\right]\left[\begin{array}{c}{V}_{E,k}\\ V_{N,k}\\ V_{U,k}\end{array}\right]-\left[\begin{array}{c}0\\ 0\\ g_0\end{array}\right]$

In the formula, navigation system rotational-angular velocity of the earth down can be calculated by rotational-angular velocity of the earth and local latitude in the projection of navigation system:

${\mathrm{\ω}}_{\mathrm{ie}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{ieE}}^n\\ {\mathrm{\ω}}_{\mathrm{ieN}}^n\\ {\mathrm{\ω}}_{\mathrm{ieU}}^n\end{array}\right]=\left[\begin{array}{c}0\\ {\mathrm{\ω}}_{\mathrm{ie}}\cos L\\ {\mathrm{\ω}}_{\mathrm{ie}}\sin L\end{array}\right]$

Navigation system can be calculated by flying speed and local latitude, the earth radius of aircraft with respect to the rotational angular velocity of the earth:

${\mathrm{\ω}}_{\mathrm{en}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{enE}}^n\\ {\mathrm{\ω}}_{\mathrm{enN}}^n\\ {\mathrm{\ω}}_{\mathrm{enU}}^n\end{array}\right]=\left[\begin{array}{c}-\frac{V_{N,k}}{R}\\ \frac{V_{E,k}}{R}\\ \frac{V_{E,k}}{R}\mathrm{tgL}\end{array}\right]$

Calculate three-dimensional specific force [f under the navigation system _Ef _Nf _U] ^TAnd the specific force [f under the body system that obtains of step (3) _x ^bf _y ^bf _z ^b] ^TBetween concern just like down conversion:

$\left[\begin{array}{c}{f}_x^b\\ f_y^b\\ f_z^b\end{array}\right]=C_n^b\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]$

$C_n^b=\left[\begin{array}{ccc}\cos \mathrm{\γ}\cos \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\cos \mathrm{\γ}\sin \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& -\sin \mathrm{\γ}\cos \mathrm{\θ}\\ \cos \mathrm{\θ}\sin \mathrm{\ψ}& \cos \mathrm{\θ}\cos \mathrm{\ψ}& \sin \mathrm{\θ}\\ \sin \mathrm{\γ}\cos \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\sin \mathrm{\γ}\sin \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& \cos \mathrm{\γ}\cos \mathrm{\θ}\end{array}\right]$

ψ, θ, γ are respectively course angle, the angle of pitch and roll angle.

According to the specific force transformation relation, can set up the trigonometric function equation under the body axis system about the angle of pitch and roll angle:

$f_y^b=(f_E\sin \mathrm{\ψ}+f_N\cos \mathrm{\ψ})\cos \mathrm{\θ}+f_U\sin \mathrm{\θ}---\left(1\right)$

$f_x^b=(f_E\cos \mathrm{\ψ}-f_N\sin \mathrm{\ψ})\cos \mathrm{\γ}+(f_E\sin \mathrm{\θ}\sin \mathrm{\ψ}+f_N\sin \mathrm{\θ}\cos \mathrm{\ψ}-f_U\cos \mathrm{\θ})\sin \mathrm{\γ}---\left(2\right)$

Formula (1) and formula (2) simultaneous are used ψ _kAfter replacing course angle ψ, have only θ, γ is a unknown quantity, and formula (1) only contains unknown quantity of the angle of pitch, and after through type (1) was tried to achieve the angle of pitch, substitution formula (2), then substitution formula (2) also only contained unknown quantity of roll angle.

(6) Fast numerical of the angle of pitch and roll angle trigonometric equation is found the solution: the field of definition of the angle of pitch is for-90 to 90 degree, and in this scope, employing waits the step interval Δ _θ, Δ _θDesirable 0.5～1 degree with pitching field of definition discretize, with the above-mentioned trigonometric equation of angle of pitch numerical value substitution of discretize, compares the optimizing search, satisfies min{|cos θ (f _ESin ψ+f _NCos ψ)+f _USin θ-f _y ^b| θ be the angle of pitch to be asked; Field of definition to roll angle is carried out discretize, and traversal search, satisfies

Min{| (f _ECos ψ-f _NSin ψ) cos γ+(f _ESin θ sin ψ+f _NSin θ cos ψ-f _UCos θ) sin γ-f _x ^b| γ be roll angle to be asked.

(7) pitching angle theta and the roll angle γ that step (6) is tried to achieve is with the longitude λ that has obtained _k, latitude L _k, height H _k, east orientation speed V _{E, k}, north orientation speed V _{N, k}, day to speed V _{U, k}, course angle ψ _kForm original state, the navigational parameter of initialization strapdown inertial navigation system is realized the aerial initial alignment of the dynamic aloft strapdown inertial navigation system of carrier.

Two, the simulating, verifying of aerial alignment methods

The simulation checking system of aerial alignment methods is characterized in that:

Comprise flight path generator (1), Inertial Measurement Unit simulator (2), satellite navigation receiver simulator (3), brother's formula acceleration are produced device (4), centripetal acceleration maker (5), acceleration of motion maker (6), navigation system specific force maker (7) and the aerial attitude error comparer (8) of aiming at down;

Described flight path generator is variable with time, recursion in time, produce successively that aircraft quickens, takes off, straight and level flight, turning, climb, flat flying, slow down to glide, spiral, contourly data such as exact position, speed, attitude, acceleration and attitude rate in the process such as fly nonstop to, and above-mentioned information is passed to Inertial Measurement Unit simulator and satellite navigation receiver simulator;

The rotational angular velocity of the data computation body coordinate system relative inertness coordinate system that described Inertial Measurement Unit simulator transmits according to the flight path generator and body system specific force down, and add noise signal, simulate the output signal of Inertial Measurement Unit;

Flight-path angle, three-dimensional velocity and the three-dimensional position signal of the output of described satellite navigation receiver simulator analog satellite navigation neceiver;

The three-dimensional velocity information of the base area revolutions of described taxi driver brother's formula acceleration maker, the output of satellite navigation receiver simulator generates brother's formula acceleration signal in the aircraft movements;

Described centripetal acceleration maker generates centripetal acceleration signal in the aircraft movements according to the three-dimensional velocity of satellite navigation receiver simulator output;

Described acceleration of motion maker generates acceleration of motion signal in the aircraft movements according to the adjacent three-dimensional velocity of output constantly of measuring before and after the satellite navigation receiver simulator;

Described navigation system down the specific force maker with the output signal of brother's formula acceleration maker, centripetal acceleration maker, acceleration of motion maker and the ratio force signal under the synthetic navigation of the acceleration of gravity system;

The ratio force signal of specific force maker and Inertial Measurement Unit simulator under the trigonometric function equation solution program reception navigation system of the described angle of pitch and roll angle, find the solution the angle of pitch and roll angle in the aerial aligning, flight-path angle, three-dimensional position and speed initialization strapdown inertial navigation system parameter that the angle of pitch of trying to achieve and roll angle produce with the satellite navigation receiver simulator;

The accurate attitude angle that attitude angle that described aerial aligning attitude error comparer will be tried to achieve by aerial alignment methods and flight path generator produce compares, when verifying that aerial alignment algorithm is applied in various state of flight, and the initial alignment precision that is reached.

Four, Figure of description

Fig. 1 is aerial alignment methods simulation checking system synoptic diagram

Fig. 2 is aerial alignment methods simulating, verifying flight path figure

Fig. 3 is the omnidistance aerial attitude error curve map of aiming at

Five, embodiment

In order to realize aircraft under various state of flights, the aerial aligning of satellite/inertia combined navigation system, need finish the work:

(1) star navigation neceiver locator data is gathered: integrated navigation computer is with period T _GContinue to read the navigation information of the specific format of satellite navigation receiver output from the interface that links to each other with satellite navigation receiver, and understand, obtain real-time three-dimensional position, three-dimensional velocity and the flight path azimuthangle of aircraft, wherein, three-dimensional position comprises longitude λ, latitude L and height H, the east orientation speed V under the local geographic coordinate system _E, east orientation speed V _NWith east orientation speed V _U, flight path azimuthangle ψ;

(2) satellite navigation locator data buffer memory: in the memory headroom variable that three-dimensional position, three-dimensional velocity and the flight path azimuthangle of the satellite navigation receiver output of reading in integrated navigation computer in the step (1) and temporal information are kept at navigational computer in the lump, the variable behind the buffer memory is successively with sign of lambda _K-1Expression longitude, L _K-1The expression latitude, H _K-1The expression height, V _{E, k-1}, V _{N, k-1}, V _{U, k-1}Represent respectively east orientation, north orientation and day to speed, ψ _K-1The expression flight path azimuthangle, t _K-1Express time information;

(3) Inertial Measurement Unit data acquisition: immediately following step (2),, read the three dimensional angular speed and the three-dimensional specific force of gyro and accelerometer measures in the Inertial Measurement Unit with period T I by analog to digital conversion or serial port.The information that the quick-connecting inertia measurement unit records is aircraft angular motion and the projection of line motion under body axis system b, and three dimensional angular speed and three-dimensional specific force be ω respectively _Ib ^Bx, ω _Ib ^By, ω _Ib ^BzAnd f _x ^b, f _y ^b, f _z ^b, T wherein _GBe T _IIntegral multiple, x, y, z represent three coordinate axis of body system.

(4) aircraft movements acceleration calculation: obtain satellite navigation data first, and through at least 2 period T _GTime after, understand t constantly finishing GPS information _k, longitude λ is arranged in the integrated navigation computer _k, latitude L _k, height H _k, east orientation direction speed V _{E, k}, north orientation speed V _{N, k}, the sky is to speed V _{U, k}, course angle ψ _k, the acceleration of motion of pressing following front and back method of difference calculating aircraft:

$a_E=\frac{V_{E,k}-V_{E,k-1}}{T_G}$

$a_N=\frac{V_{N,k}-V_{N,k-1}}{T_G};$

$a_U=\frac{V_{U,k}-V_{U,k-1}}{T_G}$

(5) real time position, speed, the acceleration of motion that step (1), step (3) and step (4) are obtained, substitution strapdown inertial navigation system specific force equation calculates the three-dimensional specific force under the navigation system, and its computing method are as follows:

$\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]=\left[\begin{array}{c}{a}_E\\ a_N\\ a_U\end{array}\right]+\left[\begin{array}{ccc}0& -(2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)\\ (2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& 0& -(2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)\\ -(2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)& 0\end{array}\right]\left[\begin{array}{c}{V}_{E,k}\\ V_{N,k}\\ V_{U,k}\end{array}\right]-\left[\begin{array}{c}0\\ 0\\ g_0\end{array}\right]$

In the formula, navigation system rotational-angular velocity of the earth down can be calculated by rotational-angular velocity of the earth and local latitude in the projection of navigation system:

${\mathrm{\ω}}_{\mathrm{ie}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{ieE}}^n\\ {\mathrm{\ω}}_{\mathrm{ieN}}^n\\ {\mathrm{\ω}}_{\mathrm{ieU}}^n\end{array}\right]=\left[\begin{array}{c}0\\ {\mathrm{\ω}}_{\mathrm{ie}}\cos L\\ {\mathrm{\ω}}_{\mathrm{ie}}\sin L\end{array}\right]$

Navigation system can be calculated by flying speed and local latitude, the earth radius of aircraft with respect to the rotational angular velocity of the earth:

${\mathrm{\ω}}_{\mathrm{en}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{enE}}^n\\ {\mathrm{\ω}}_{\mathrm{enN}}^n\\ {\mathrm{\ω}}_{\mathrm{enU}}^n\end{array}\right]=\left[\begin{array}{c}-\frac{V_{N,k}}{R}\\ \frac{V_{E,k}}{R}\\ \frac{V_{E,k}}{R}\mathrm{tgL}\end{array}\right]$

Calculate three-dimensional specific force [f under the navigation system _Ef _Nf _U] ^TAnd the specific force [f under the body system that obtains of step (3) _x ^bf _y ^bf _z ^b] ^TBetween concern just like down conversion:

$\left[\begin{array}{c}{f}_x^b\\ f_y^b\\ f_z^b\end{array}\right]=C_n^b\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]$

$C_n^b=\left[\begin{array}{ccc}\cos \mathrm{\γ}\cos \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\cos \mathrm{\γ}\sin \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& -\sin \mathrm{\γ}\cos \mathrm{\θ}\\ \cos \mathrm{\θ}\sin \mathrm{\ψ}& \cos \mathrm{\θ}\cos \mathrm{\ψ}& \sin \mathrm{\θ}\\ \sin \mathrm{\γ}\cos \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\sin \mathrm{\γ}\sin \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& \cos \mathrm{\γ}\cos \mathrm{\θ}\end{array}\right]$

ψ, θ, γ are respectively course angle, the angle of pitch and roll angle.

According to the specific force transformation relation, can set up the trigonometric function equation under the body axis system about the angle of pitch and roll angle:

$f_y^b=(f_E\sin \mathrm{\ψ}+f_N\cos \mathrm{\ψ})\cos \mathrm{\θ}+f_U\sin \mathrm{\θ}---\left(1\right)$

$f_x^b=(f_E\cos \mathrm{\ψ}-f_N\sin \mathrm{\ψ})\cos \mathrm{\γ}+(f_E\sin \mathrm{\θ}\sin \mathrm{\ψ}+f_N\sin \mathrm{\θ}\cos \mathrm{\ψ}-f_U\cos \mathrm{\θ})\sin \mathrm{\γ}---\left(2\right)$

Formula (1) and formula (2) simultaneous are used ψ _kAfter replacing course angle ψ, have only θ, γ is a unknown quantity, and formula (1) only contains unknown quantity of the angle of pitch, and after through type (1) was tried to achieve the angle of pitch, substitution formula (2), then substitution formula (2) also only contained unknown quantity of roll angle.

(6) Fast numerical of the angle of pitch and roll angle trigonometric equation is found the solution: the field of definition of the angle of pitch is for-90 to 90 degree, and in this scope, employing waits the step interval Δ _θ, Δ _θDesirable 0.5～1 degree with pitching field of definition discretize, with the above-mentioned trigonometric equation of angle of pitch numerical value substitution of discretize, compares the optimizing search, satisfies min{|cos θ (f _ESin ψ+f _NCos ψ)+f _USin θ-f _y ^b| θ be the angle of pitch to be asked; Field of definition to roll angle is carried out discretize, and traversal search, satisfies

Min{| (f _ECos ψ-f _NSin ψ) cos γ+(f _ESin θ sin ψ+f _NSin θ cos ψ-f _UCos θ) sin γ-f _x ^b| γ be roll angle to be asked.

(7) pitching angle theta and the roll angle γ that step (6) is tried to achieve is with the longitude λ that has obtained _k, latitude L _k, height H _k, east orientation speed V _{E, k}, north orientation speed V _{N, k}, day to speed V _{U, k}, course angle ψ _kForm original state, the navigational parameter of initialization strapdown inertial navigation system is realized the aerial initial alignment of the dynamic aloft strapdown inertial navigation system of carrier.

For the correctness and the precision of aerial alignment methods, need to make up simulation checking system.The structural representation of simulation checking system is seen Fig. 1, comprise flight path generator 1, Inertial Measurement Unit simulator 2, satellite navigation receiver simulator 3, brother's formula acceleration are produced device 4, centripetal acceleration maker 5, acceleration of motion maker 6, the following specific force maker 7 of navigation system and are aimed at attitude error comparer 8 in the air; Wherein the output of flight path generator 1 is connected in the input of Inertial Measurement Unit simulator 2, the input of satellite navigation receiver simulator 3 respectively, and the data computation body coordinate system that Inertial Measurement Unit simulator 2 transmits according to flight path generator 1 is with respect to the rotational angular velocity of inertial coordinates system and the specific force under the body system; The centripetal acceleration maker 5 of three velocity informations that the form acceleration maker 4 of three velocity informations that regional independently information of reception and miniature navigation neceiver simulator are exported and reception satellite navigation receiver simulator are exported is imported the specific force maker 7 that navigates under being respectively with the acceleration of motion maker 6 and the acceleration of gravity information of the three-dimensional velocity information that the reception satellite navigation receiver simulator front and back adjacent measurement moment exports.Find the solution the roll angle and the aerial attitude error comparer 8 and the strap-down inertial parameter initialization aimed at of angle of pitch input of the aerial aligning that obtains; The flight-path angle signal of satellite navigation receiver simulator 3 analog satellite conference receivers, three-dimensional velocity signal and three-dimensional position signal input strapdown inertial navigation system parameter initialization.

In the simulating, verifying, the flight path generator is variable with time, promotion analogue system operation, the flight path of generation is seen Fig. 2, flight path comprises that aircraft quickens, takes off, straight and level flight, turning, climb, flat flying slowed down and glided, spirals, and process such as contourly flies nonstop to, the high-frequency vibration is jolted, data such as acceleration and attitude rate, and above-mentioned information passed to Inertial Measurement Unit simulator and satellite navigation receiver simulator.

Six, the effect of invention

The present invention is based on the auxiliary of satellite navigation information, for various mobile process, realized strap-down inertial The aerial coarse alignment of system has also been built simulation checking system, has verified the aerial alignment methods correctness of the present invention With in acceleration and deceleration, turn, spiral, validity and applicability under the various dynamic flying conditions such as high frequency jolts.

Utilize the aerial simulation checking system of aiming at, constantly by aerial alignment algorithm, obtain the attitude value at each, with Standard flight path attitude compares, and is aimed at initial error in the air and sees Fig. 3, aerial alignment methods of the present invention Coarse alignment method can be applied to different mission phases and the state of flight of aircraft. Accelerometer bias 10^-3Meter per second², gyroscopic drift 1 degree/hour, 30 meters of satellite navigation system position errors, range rate error 0.2 meter per second Under the simulated conditions, about 1.5 degree of horizontal coarse alignment precision, enter navigational state after, the navigation calculation attitude can be In short time, reach navigation accuracy behind the fine alignment.

The present invention breaks down for aloft aircraft, and satellite/inertia combined navigation system restarts, and Some special aircraft, as guided cartridge etc. must strapdown inertial navigation system could begin after taking off on the electrician The situations such as work have the meaning of particular importance.

Claims (1)

1. the aerial alignment methods of a satellite/inertia combined navigation system is characterized in that, may further comprise the steps:

(1) the satellite navigation receiver locator data is gathered: integrated navigation computer is with period T _GContinue to read the navigation information of the specific format of satellite navigation receiver output from the interface that links to each other with satellite navigation receiver, and understand, obtain real-time three-dimensional position, three-dimensional velocity and the flight path azimuthangle of aircraft, wherein, three-dimensional position comprises longitude λ, latitude L and height H, the east orientation speed V under the local geographic coordinate system _E, east orientation speed V _NWith east orientation speed V _U, flight path azimuthangle ψ;

(2) satellite navigation locator data buffer memory: in the memory headroom variable that three-dimensional position, three-dimensional velocity and the flight path azimuthangle of the satellite navigation receiver output of reading in integrated navigation computer in the step (1) and temporal information are kept at navigational computer in the lump, the variable behind the buffer memory is successively with sign of lambda _K-1Expression longitude, L _K-1The expression latitude, H _K-1The expression height, V _{E, k-1}, V _{N, k-1}, V _{U, k-1}Represent respectively east orientation, north orientation and day to speed, ψ _K-1The expression flight path azimuthangle, t _K-1Express time information;

(3) Inertial Measurement Unit data acquisition: immediately following step (2), by analog to digital conversion or serial port, read the three dimensional angular speed and the three-dimensional specific force of gyro and accelerometer measures in the Inertial Measurement Unit with period T I, the information that the quick-connecting inertia measurement unit records is aircraft angular motion and the projection of line motion under body axis system b, and three dimensional angular speed and three-dimensional specific force be ω respectively _Ib ^Bx, ω _Ib ^By, ω _Ib ^BzAnd f _x ^b, f _y ^b, f _z ^b, T wherein _GBe T _IIntegral multiple, x, y, z represent three coordinate axis of body system;

(4) aircraft movements acceleration calculation: obtain satellite navigation data first, and through at least 2 period T _GTime after, understand t constantly finishing GPS information _k, longitude λ is arranged in the integrated navigation computer _k, latitude L _k, height H _k, east orientation direction speed V _{E, k}, north orientation speed V _{N, k}, the sky is to speed V _{U, k}, course angle ψ _k, the acceleration of motion of pressing following front and back method of difference calculating aircraft:

$a_E=\frac{V_{E,k}-V_{E,k-1}}{T_G}$

$a_N=\frac{V_{N,k}-V_{N,k-1}}{T_G};$

$a_U=\frac{V_{U,k}-V_{U,k-1}}{T_G}$

(5) real time position, speed, the acceleration of motion that step (1), step (3) and step (4) are obtained, substitution strapdown inertial navigation system specific force equation calculates the three-dimensional specific force under the navigation system, and its computing method are as follows:

$\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]=\left[\begin{array}{c}{a}_E\\ a_N\\ a_U\end{array}\right]+\left[\begin{array}{ccc}0& -(2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)\\ (2{\mathrm{\ω}}_{\mathrm{ieU}}^n+{\mathrm{\ω}}_{\mathrm{enU}}^n)& 0& -(2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)\\ -(2{\mathrm{\ω}}_{\mathrm{ieN}}^n+{\mathrm{\ω}}_{\mathrm{enN}}^n)& (2{\mathrm{\ω}}_{\mathrm{ieE}}^n+{\mathrm{\ω}}_{\mathrm{enE}}^n)& 0\end{array}\right]\left[\begin{array}{c}{V}_{E,k}\\ V_{N,k}\\ V_{U,k}\end{array}\right]-\left[\begin{array}{c}0\\ 0\\ g_0\end{array}\right]$

In the formula, navigation system rotational-angular velocity of the earth down can be calculated by rotational-angular velocity of the earth and local latitude in the projection of navigation system:

${\mathrm{\ω}}_{\mathrm{ie}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{ieE}}^n\\ {\mathrm{\ω}}_{\mathrm{ieN}}^n\\ {\mathrm{\ω}}_{\mathrm{ieU}}^n\end{array}\right]=\left[\begin{array}{c}0\\ {\mathrm{\ω}}_{\mathrm{ie}}\cos L\\ {\mathrm{\ω}}_{\mathrm{ie}}\sin L\end{array}\right]$

Navigation system can be calculated by flying speed and local latitude, the earth radius of aircraft with respect to the rotational angular velocity of the earth:

${\mathrm{\ω}}_{\mathrm{en}}^n=\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{enE}}^n\\ {\mathrm{\ω}}_{\mathrm{enN}}^n\\ {\mathrm{\ω}}_{\mathrm{enU}}^n\end{array}\right]=\left[\begin{array}{c}-\frac{V_{N,k}}{R}\\ \frac{V_{E,k}}{R}\\ \frac{V_{E,k}}{R}\mathrm{tgL}\end{array}\right]$

Calculate three-dimensional specific force [f under the navigation system _Ef _Nf _U] ^TAnd the specific force [f under the body system that obtains of step (3) _x ^bf _y ^bf _z ^b] ^TBetween concern just like down conversion:

$\left[\begin{array}{c}{f}_x^b\\ f_y^b\\ f_z^b\end{array}\right]=C_n^b\left[\begin{array}{c}{f}_E\\ f_N\\ f_U\end{array}\right]$

$C_n^b=\left[\begin{array}{ccc}\cos \mathrm{\γ}\cos \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\cos \mathrm{\γ}\sin \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& -\sin \mathrm{\γ}\cos \mathrm{\θ}\\ \cos \mathrm{\θ}\sin \mathrm{\ψ}& \cos \mathrm{\θ}\cos \mathrm{\ψ}& \sin \mathrm{\θ}\\ \sin \mathrm{\γ}\cos \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\sin \mathrm{\ψ}& -\sin \mathrm{\γ}\sin \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\cos \mathrm{\ψ}& \cos \mathrm{\γ}\cos \mathrm{\θ}\end{array}\right]$

ψ, θ, γ are respectively course angle, the angle of pitch and roll angle,

According to the specific force transformation relation, can set up the trigonometric function equation under the body axis system about the angle of pitch and roll angle:

$f_y^b=(f_E\sin \mathrm{\ψ}+f_N\cos \mathrm{\ψ})\cos \mathrm{\θ}+f_U\sin \mathrm{\θ}---\left(1\right)$

$f_x^b=(f_E\cos \mathrm{\ψ}-f_N\sin \mathrm{\ψ})\cos \mathrm{\γ}+(f_E\sin \mathrm{\θ}\sin \mathrm{\ψ}+f_N\sin \mathrm{\θ}\cos \mathrm{\ψ}-f_U\cos \mathrm{\θ})\sin \mathrm{\γ}---\left(2\right)$

Formula (1) and formula (2) simultaneous are used ψ _kAfter replacing course angle ψ, have only θ, γ is a unknown quantity, and formula (1) only contains unknown quantity of the angle of pitch, and after through type (1) was tried to achieve the angle of pitch, substitution formula (2), then substitution formula (2) also only contained unknown quantity of roll angle;

(6) Fast numerical of the angle of pitch and roll angle trigonometric equation is found the solution: the field of definition of the angle of pitch is for-90 to 90 degree, and in this scope, employing waits the step interval Δ _θ, Δ _θDesirable 0.5～1 degree with pitching field of definition discretize, with the above-mentioned trigonometric equation of angle of pitch numerical value substitution of discretize, compares the optimizing search, satisfies min{|cos θ (f _ESin ψ+f _NCos ψ)+f _USin θ-f _y ^b| θ be the angle of pitch to be asked; Field of definition to roll angle is carried out discretize, and traversal search, satisfies min{| (f _ECos ψ-f _NSin ψ) cos γ+(f _ESin θ sin ψ+f _NSin θ cos ψ-f _UCos θ) sin γ-f _x ^b| γ be roll angle to be asked;

(7) pitching angle theta and the roll angle γ that step (6) is tried to achieve is with the longitude λ that has obtained _k, latitude L _k, height H _k, east orientation speed V _{E, k}, north orientation speed V _{N, k}, day to speed V _{U, k}, course angle ψ _kForm original state, the navigational parameter of initialization strapdown inertial navigation system is realized the aerial initial alignment of the dynamic aloft strapdown inertial navigation system of carrier.

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