CN110487300A - Vibration absorber influences test method to the performance of inertial navigation system - Google Patents
Vibration absorber influences test method to the performance of inertial navigation system Download PDFInfo
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- 239000006096 absorbing agent Substances 0.000 title claims abstract description 74
- 238000010998 test method Methods 0.000 title claims abstract description 14
- 238000000034 method Methods 0.000 claims abstract description 19
- 238000004422 calculation algorithm Methods 0.000 claims abstract description 8
- 238000002955 isolation Methods 0.000 claims abstract description 4
- 239000011159 matrix material Substances 0.000 claims description 26
- 230000008878 coupling Effects 0.000 claims description 24
- 238000010168 coupling process Methods 0.000 claims description 24
- 238000005859 coupling reaction Methods 0.000 claims description 24
- 230000000631 nonopiate Effects 0.000 claims description 24
- 238000009434 installation Methods 0.000 claims description 16
- 230000035945 sensitivity Effects 0.000 claims description 16
- 230000001133 acceleration Effects 0.000 claims description 14
- 238000012360 testing method Methods 0.000 claims description 11
- 230000009466 transformation Effects 0.000 claims description 7
- 238000005070 sampling Methods 0.000 claims description 3
- 238000002474 experimental method Methods 0.000 abstract description 4
- 238000004088 simulation Methods 0.000 description 15
- 230000003068 static effect Effects 0.000 description 9
- 238000004364 calculation method Methods 0.000 description 4
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- 239000000969 carrier Substances 0.000 description 1
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- 230000000694 effects Effects 0.000 description 1
- 230000010355 oscillation Effects 0.000 description 1
- 230000001235 sensitizing effect Effects 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
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Abstract
The present invention relates to a kind of vibration absorbers to influence test method to the performance of inertial navigation system, belongs to inertial navigation inertial sensor errors calibration technique field.This method comprises the following steps: step 1, establishing conventional gyroscope error model, the gyro error includes gyro misalignment, scale factor error and Random Drift Error；Step 2, the error of the three classes error considers to introduce the influence generated after vibration absorber, establishes new error model；Step 3, it is tested by calibration, using two kinds of error models, provides error information, the validity of verifying gyroscope emulation data for gyro performance is quantitatively evaluated；Step 4, according to obtained error information, to antenna attitude angle error impact analysis, influence of the increase of isolation mounting for attitude algorithm is further analyzed.The present invention establishes the inertial device error model after increasing vibration absorber, analyzes influence of the vibration absorber to navigation system precision, and verified to emulation to corresponding conclusion by experiment.
Description
Technical field
The present invention relates to a kind of vibration absorbers to influence test method to the performance of inertial navigation system, and it is used to belong to inertial navigation
Property sensor error calibration technique field.
Background technique
Inertial navigation system is the prime navaid equipment of numerous carriers such as aviation, land, waterborne.Inertia device is that inertia is led
The chief component of boat system, inertial device error are an important factor for influencing inertial navigation system performance, and precision is direct
Determine the performance of inertial navigation system.When carrier operation, the vibration environment of body can cause the decline of inertia device performance,
To influence navigation accuracy.Vibration environment has uncertain and not timing, very likely leads under aircraft high-speed flight
Causing inertia device installation error and scale factor error, great changes will take place compared with Laboratory Calibration value, causes greatly to navigate and miss
Difference.Therefore, vibration influence caused by inertia system precision how is reduced during aircraft flight, it is real for improving system
It is of great significance with precision.
In view of this, the mode for generalling use physics vibration damping is isolated by inertia device with carrier.Vibration absorber reduces
Carrier angular oscillation, linearly coupled are to inertia device bring adverse effect.However, it is again such that inertia device is no longer completely quick
Feel the motion information of carrier, change the kinetic characteristic of inertia device in a dynamic environment, to influence the measurement essence of inertia device
Degree.Therefore, analysis and test need to be carried out to the inertial guidance index before and after the external vibrating isolation system of increase.
Summary of the invention
The present invention proposes a kind of vibration absorber and leads to inertia for inertial attitude tracking system common in guided missile carrier
The performance of boat system influences test method, establishes the inertial device error model after increasing vibration absorber, analyzes vibration damping dress
The influence to navigation system precision is set, and corresponding conclusion is verified to emulation by experiment.
The present invention is to solve its technical problem to adopt the following technical scheme that
A kind of vibration absorber influences test method to the performance of inertial navigation system, includes the following steps:
Step 1, conventional gyroscope error model is established, the gyro error includes gyro misalignment, scale factor error
And Random Drift Error；
Step 2, on the basis of step 1 pair conventional gyro error modeling, the error of three classes error described in step 1 is considered
The influence generated after vibration absorber is introduced, new error model is established；
Step 3, it is tested by calibration, the two kinds of error models obtained using steps 1 and 2, is mentioned for gyro performance is quantitatively evaluated
For error information, the validity of gyroscope emulation data is verified；
Step 4, the error information obtained according to step 3 further analyzes vibration isolation to antenna attitude angle error impact analysis
Influence of the increase of device for attitude algorithm.
Routine gyroscope error model described in step 1 are as follows:
Gyro reality outputω is exported by gyro idealbAnd gyro items error composition, error model are as follows:
Wherein,For gyro reality output, ωbFor the output of error free gyro, εbFor gyro zero bias, ωsFor constant multiplier mistake
Difference, ωθFor nonopiate coupling error, ωGFor acceleration sensitive error, η is random noise item；
1) gyro zero bias
In formula,For component of the gyro zero bias on three axis of gyroscope；
2) gyro scale factor error
In formula,For gyro scale factor error coefficient matrix；
3) gyro accelerometer sensitivity error
In formula,For the coefficient matrix of gyro accelerometer sensitivity error,For
Acceleration on gyroscope three axis of ideal；
4) the nonopiate coupling error of gyro
In formula,For the coefficient matrix of the nonopiate coupling error of gyro；
5) Gyro Random noise
η=εr+ωg
In formula, ωgFor white noise, εrFor single order markoff process random noise.
New gyroscope error model described in step 2 are as follows:
Gyro reality output after vibration absorber is addedω is exported by gyro idealbWith the gyro after addition vibration absorber
Every error composition, are as follows:
Wherein, ε 'bTo increase the gyro zero bias after vibration absorber, ω 'sTo increase the scale factor error after vibration absorber,
ω′θTo increase the nonopiate coupling error after vibration absorber, ω 'GTo increase the acceleration sensitive error after vibration absorber, η ' is
Random noise item after increasing vibration absorber, GATo increase the installation error coefficient matrix after vibration absorber；
1) the gyro zero bias after adding and subtracting vibrating device indicate are as follows:
Wherein:For component of the gyro zero bias on three axis of gyroscope；
2) the gyro scale factor error after adding and subtracting vibrating device indicates are as follows:
Wherein,For the gyro scale factor error coefficient matrix after plus-minus vibrating device；
3) the gyro accelerometer sensitivity error after adding and subtracting vibrating device is expressed as:
Wherein,For the coefficient matrix for adding and subtracting vibrating device gyro accelerometer sensitivity error；
4) the nonopiate coupling error of gyro after adding and subtracting vibrating device indicates are as follows:
Wherein,For the coefficient matrix of the nonopiate coupling error of gyro after plus-minus vibrating device.
Test process is demarcated described in step 3:
(1) random error is tested
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into 1h on turntable and carries out data acquisition, be repeated 3 times；
(2) zero bias are tested
1) gyro is separately fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into progress zero bias sampling, every group of acquisition 1min on double axle table；
(3) acceleration sensitive error testing
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on double axle table and is sampled, every group of acquisition 1min；
(4) scale factor error is tested
1) gyroscope is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) operation turntable make inertia component rotating around three axis accelerometer with ± 10 °/s, ± 30 °/s, ± 50 °/s, ± 70 °/s,
± 90 °/s, ± 110 °/s, ± 130 °/s, 18 ± 150 °/s, ± 180 °/s angular speed rotations；
(5) nonopiate coupling error test
1) gyroscope is fixed on angle vibration table by fixture, connects equipment power supply line and data line；
2) give gyroscope to be powered, start double axle table after 5s, operation turntable make inertia component rotating around three axis accelerometer with ±
10 °/s, ± 30 °/s, ± 50 °/s, ± 70 °/s, ± 90 °/s, ± 110 °/s, ± 130 °/s, ± 150 °/s, ± 180 °/s 18
A angular speed rotation.Wherein, ωiThe angular speed rotated for turntable around gyro i axis.
The detailed process of the step 4 are as follows:
Relationship according to quaternary number and Eulerian angles obtains following relational expression, carrier coordinate system be it is right front upper, then by geography
It is the transformation quaternary number to carrier system are as follows:
In formula, ψ, θ, γ are respectively course angle, pitch angle and the roll angle of carrier, q0For real part, q1For imaginary part unit i,
q2For imaginary part unit j, q3For imaginary part unit k；
Quaternary number is obtained by quaternion differential equation and iterates to calculate formula are as follows:
Wherein
Δθ0To take approximate definition value, Δ θ to exponential integralxFor
Δθ0X-component, Δ θyFor Δ θ0Y-component, Δ θzFor Δ θ0Z-component,
Δθ0 2=Δ θx 2+Δθy 2+Δθz 2,
[] indicates the matrix-expand of vector in formula, and
Wherein,The projection x-axis output that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope
The relatively geographical projection y-axis output for tying up to body and fastening of body system,Body system is tied up to for gyroscope body system geography relatively
On projection z-axis output,The projection that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope body system
Relative inertness ties up to the projection output that body is fastened,It is defeated that the projection in Department of Geography is tied up to for gyroscope earth system relative inertness
Out,The output of the projection in Department of Geography is tied up to respect to the earth for gyroscope Department of Geography, is solvedFormula in again need use
Utilize last momentIt is iterated, therefore the solution of Eulerian angles is a loop iteration process；
The transformation relation between fixed vector indicated according to quaternary number, obtains the pass of quaternionic matrix and direction cosines
System:
Finally according to the following formula, attitude angle is found out
The output of target seeker gyro is transformed into antenna attitude angle by above calculating, so that the gyro output with error be reflected
It penetrates as antenna attitude angle error.The gyroscope error model that steps 1 and 2 are established is substituted into, and combines step 3 vibration absorber to gyro shadow
Loud analysis, the influence by emulation preliminary analysis gyro error to antenna attitude angle error.
Beneficial effects of the present invention are as follows:
1, the present invention to vibration absorber installation front and back inertial device error model, including installation error, scale because
Number error and Random Drift Error model, and by calibration test, influence of the vibration absorber installation front and back to basic accuracy is detected,
Using nominal data, analysis obtains the performance under its vibration environment, ensure that the effective use of inertia device precise information.
2, the present invention is by demarcating gyro error, for missile-borne inertial attitude calculation method, more true mould
The course of work of gyro in quasi- tracking platform generates track, gyro output simulation and attitude algorithm have carried out simulation analysis.
It is brought into missile attitude respectively by emulating obtained " former gyro " and " antivibration gyro " simulation output under dynamic and vibration environment
The attitude parameter simulation result resolved by these two types of gyros can be obtained in parameter calculation algorithm, analyze antivibration on this basis
Influence of the gyro product to Guidance Parameter.
Detailed description of the invention
Fig. 1 is the architecture diagram that vibration absorber of the present invention influences test method on the performance of inertial navigation system.
Fig. 2 is the scheme of installation of present invention installation vibration absorber inertia device.
Fig. 3 is aircraft's flight track schematic diagram of the present invention.
Fig. 4 .1- Fig. 4 .8 is gyro error calibration comparing result before and after present invention installation vibration absorber, wherein Fig. 4 .1 (a)
It is that gyroscope emulates static x-axis output；Fig. 4 .1 (b) is that gyroscope emulates static y-axis output；Fig. 4 .2 (a) is that gyroscope is true
Static x-axis output；Fig. 4 .2 (b) is the truly static y-axis output of gyroscope；Fig. 4 .3 (a) is that the static x-axis of antivibration gyroscope emulation is defeated
Out；Fig. 4 .3 (b) is that antivibration gyroscope emulates static y-axis output；Fig. 4 .4 (a) is the truly static x-axis output of antivibration gyroscope；Figure
4.4 (b) be the truly static y-axis output of antivibration gyroscope；Fig. 4 .5 (a) is gyroscope emulation dynamic x-axis output；Fig. 4 .5 (b) is
Gyroscope emulates the output of dynamic y-axis；Fig. 4 .6 (a) is the true dynamic x-axis output of gyroscope；Fig. 4 .6 (b) is that gyroscope really moves
The output of state y-axis；Fig. 4 .7 (a) is the emulation dynamic x-axis output of antivibration gyroscope；Fig. 4 .7 (b) is antivibration gyroscope emulation dynamic y-axis
Output, Fig. 4 .8 (a) are the true dynamic x-axis output of antivibration gyroscope；Fig. 4 .8 (b) is the true dynamic y-axis output of antivibration gyroscope.
Fig. 5 .1- Fig. 5 .4 is to inertial attitude calculation result curve in guided missile.Fig. 5 .1 (a) is single emulation attitude angle without subtracting
Vibrating device roll angle error change curve；Fig. 5 .1 (b) is that single emulation attitude angle is bent without vibration absorber pitch angle error change
Line；Fig. 5 .1 (c) is single emulation attitude angle without vibration absorber course angle error change curve；Fig. 5 .2 (a) is single emulation appearance
There is vibration absorber roll angle error change curve at state angle；Fig. 5 .2 (b) is that single emulation attitude angle has vibration absorber pitching angle error
Change curve；Fig. 5 .2 (c) is that single emulation attitude angle has vibration absorber course angle error change curve；Fig. 5 .3 (a) is hundred times
Attitude angle is emulated without vibration absorber roll angle error change curve；Fig. 5 .3 (b) is hundred emulation attitude angles without vibration absorber pitching
Angle error change curve；Fig. 5 .3 (c) is hundred emulation attitude angles without vibration absorber course angle error change curve；Fig. 5 .4 (a)
It is that hundred emulation attitude angles have vibration absorber roll angle error change curve；Fig. 5 .4 (b) is that hundred emulation attitude angles have vibration damping dress
Set pitch angle error change curve；Fig. 5 .4 (c) is that hundred emulation attitude angles have vibration absorber course angle error change curve.
Specific embodiment
Embodiments of the present invention are described below in detail, the example of the embodiment is shown in the accompanying drawings.Below by
The embodiment being described with reference to the drawings is exemplary, and for explaining only the invention, and is not construed as limiting the claims.
As shown in Figure 1, vibration absorber of the present invention influences the principle of test method on the performance of inertial navigation system
As follows, Fig. 2 is gyroscope structure signal:
(1) it is ideal gyroscope ideal output simulation: to contain the gyroscope directly generated by Missile Terminal Guidance trace simulation
The gyroscope ideal output that output and vibrated environment influence.
(2) gyro error model modeling is analyzed: the foundation of error model is mainly carried out according to Gyroscope error parameter,
Wherein error parameter is divided into gyro error parametric nominal value that producer provides, by the analysis testing requirement report of target seeker antivibration gyro
The error parameter value obtained and the measurement error parameter value obtained based on calibration experiment.
(3) gyroscope simulation output: the gyroscope ideal that the error parameter of (2) part is added to (1) part is defeated
Gyroscope simulation output can be obtained in out.Be divided into: the gyroscope simulation of the device without friction based on nominal error parameter is defeated
Out, join comprising the theoretical error parameter under the influence of vibration and the gyroscope simulation output with vibration absorber and based on measurement error
Number and the gyroscope simulation output with vibration absorber.
(4) missile attitude parameter calculation: the gyroscope analog output value under emulation obtains in (3) different condition is brought into
Three classes attitude parameter simulation result can be obtained into missile attitude parameter calculation algorithm.It will include measurement error parameter and band
The attitude parameter that the gyroscope of vibration absorber resolves is with no vibration absorber and using the resolving of the gyroscope of nominal error parameter
Obtained attitude parameter compares, and can analyze the influence after vibration absorber is added to inertial guidance parameter.And will include
The attitude parameter for vibrating the theoretical error parameter influenced and the gyroscope resolving with vibration absorber is opposite with ideal guided missile track
Than, so that it may analysis obtains the influence to inertial guidance parameter calculation result under vibration of antivibration gyro product.
A specific embodiment of the invention is as follows:
1, the gyroscope error model before and after installation vibration damping is established
The precision of gyro is the determinant of inertial navigation system precision.For inertia sensing device assembly, main mistake
Poor source includes nonopiate coupling error, scale factor error, zero bias and random error etc., error model are as follows:
In formula:For gyro reality output, ωbFor the output of error free gyro, εbFor gyro zero bias, ωsFor constant multiplier
Error, ωθFor nonopiate coupling error, ωGFor acceleration sensitive error, η is random noise item.
1) gyro zero bias
In formula,For component of the gyro zero bias on three axis of gyroscope.
2) gyro scale factor error
In formula,For gyro scale factor error coefficient matrix.
3) gyro accelerometer sensitivity error
In formula,For the coefficient matrix of gyro accelerometer sensitivity error,For
Acceleration on gyroscope three axis of ideal.
4) the nonopiate coupling error of gyro
In formula,For the coefficient matrix of the nonopiate coupling error of gyro.
5) Gyro Random noise
η=εr+ωg
Wherein: εrFor single order markoff process random noise, ωgFor white noise.
After vibration absorber is added, the kinetic characteristic of gyro in a dynamic environment can be changed, to produce to above-mentioned error model
It is raw to influence.
1) the gyro zero bias after vibrating device are added and subtracted
The introducing of vibration absorber, may change the sensitive axis of gyro, to introduce additional zero bias, and vibration absorber
It will affect the output characteristics of gyro under vibration condition, so as to cause the variation of zero bias.Therefore, the gyro zero bias after vibrating device are added and subtracted
It may be expressed as:
Wherein:For component of the gyro zero bias on three axis of gyroscope；
Gyro scale factor error after adding and subtracting vibrating device
The constant multiplier of gyro will affect by the variation of vibration absorber bring gyro sensitivity axis, to change its scale
Factor error.Therefore, the scale factor error after adding and subtracting vibrating device may be expressed as:
In formula,For the gyro scale factor error coefficient matrix after plus-minus vibrating device.
2) the gyro accelerometer sensitivity error after vibrating device is added and subtracted
The introducing of vibration absorber will affect the acceleration characteristic of gyro under vibration condition, to affect acceleration sensitive mistake
Difference.Therefore, the gyro accelerometer sensitivity error after adding and subtracting vibrating device may be expressed as:
In formula,For the coefficient matrix for adding and subtracting vibrating device gyro accelerometer sensitivity error.
3) the nonopiate coupling error of gyro after vibrating device is added and subtracted
The torsion of vibration absorber will affect the orthogonality between each gyro under vibration, make the sensitivity of each gyro
Axis no longer meets orthogonal installation configuration, so as to cause nonopiate coupling error.Therefore, the nonopiate coupling of gyro after vibrating device is added and subtracted
Closing error may be expressed as:
In formula,For the coefficient matrix of the nonopiate coupling error of gyro after plus-minus vibrating device.
Gyro misalignment coefficient matrix after adding and subtracting vibrating device
Gyro can have fix error angle during installing vibration absorber, so that the three of three sensitive axes of gyro and carrier
A quadrature axis cannot be completely coincident, to generate installation error.
Therefore, after considering vibration absorber, inertial sensor component erroi model is represented by
In formula: ε 'bTo increase the gyro zero bias after vibration absorber, ω 'sTo increase the scale factor error after vibration absorber,
ω′θTo increase the nonopiate coupling error after vibration absorber, ω 'GTo increase the acceleration sensitive error after vibration absorber, η ' is
Random noise item after increasing vibration absorber, GATo increase the installation error coefficient matrix after vibration absorber.
2, gyro error scaling method
(1) random error is tested
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into 1h on turntable and carries out data acquisition, be repeated 3 times.
(2) zero bias are tested
1) gyro is separately fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into progress zero bias sampling, every group of acquisition 1min on double axle table
(3) acceleration sensitive error testing
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on double axle table and is sampled, every group of acquisition 1min.
(4) scale factor error is tested
1) gyroscope is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) operation turntable make inertia component rotating around three axis accelerometer with ± 10 °/s, ± 30 °/s, ± 50 °/s, ± 70 °/s,
± 90 °/s, ± 110 °/s, ± 130 °/s, 18 ± 150 °/s, ± 180 °/s angular speed rotations.Rate self-calibration transfer mesa-shaped
State and the input of each axis gyro standard are shown in Table 1.Wherein: ωiThe angular speed rotated for turntable around gyro i axis.
1 gyro dynamic calibration turning table control information of table and the corresponding sensitizing input of gyro
(5) nonopiate coupling error test
1) gyroscope is fixed on angle vibration table by fixture, connects equipment power supply line and data line；
2) give gyroscope to be powered, start double axle table after 5s, operation turntable make inertia component rotating around three axis accelerometer with ±
10 °/s, ± 30 °/s, ± 50 °/s, ± 70 °/s, ± 90 °/s, ± 110 °/s, ± 130 °/s, ± 150 °/s, ± 180 °/s 18
A angular speed rotation.Wherein: the angular speed rotated for turntable around gyroaxis.
3, the computation of missile-borne inertial attitude tracking system
Inertial attitude tracking system is one of core sensor part of guided missile, when target seeker tracks target,
Attitude reference can be provided.The system is usually made of the gyro of three orthogonal installations, passes through the movement angle of gyro sensitivity guided missile
Speed, and then the posture of target seeker is solved.Quaternion Method is the algorithms most in use of gyro attitude algorithm, process are as follows: first
First gyro is exported and carries out error compensation, then carries out quaternary number equation solution, then solves pose transformation matrix and posture
Angle.
Relationship according to quaternary number and Eulerian angles can obtain Department of Geography to the transformation quaternary number of carrier system:
In formula, ψ, θ, γ are respectively course angle, pitch angle and the roll angle of carrier, q0For real part, q1For imaginary part unit i,
q2For imaginary part unit j, q3For imaginary part unit k.
Quaternary number can be obtained by quaternion differential equation and iterate to calculate formula
In formulaΔθ0To take approximate definition value, Δ to exponential integral
θxFor Δ θ0X-component, Δ θyFor Δ θ0Y-component, Δ θzFor Δ θ0Z-component, Δ θ0 2=Δ θx 2+Δθy 2+Δθz 2,
[] indicates the matrix-expand of vector in formula.And
Wherein,The projection x-axis output that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope
The relatively geographical projection y-axis output for tying up to body and fastening of body system,Body is tied up to for gyroscope body system geography relatively to fasten
Projection z-axis output,The projection that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope body system phase
Inertial system is exported in the projection that body is fastened,The output of the projection in Department of Geography is tied up to for gyroscope earth system relative inertness,The output of the projection in Department of Geography is tied up to respect to the earth for gyroscope Department of Geography, is solvedFormula in again need useIt can
To utilize last momentIt is iterated, therefore the solution of Eulerian angles is a loop iteration process.
The transformation relation between fixed vector indicated according to quaternary number, can obtain quaternionic matrix and direction cosines
Relational expression (11).
Finally according to formula (12), attitude angle can be found out.
The output of target seeker gyro is transformed into antenna attitude angle by above calculating, so as to which the gyro with error is defeated
It is mapped as antenna attitude angle error out
It invents proposed vibration absorber in order to verify and influences the correctness and effectively of test method on inertia device performance
Property, model is established using the method for the present invention, carries out error calibration using double axle table, Matlab emulation is carried out according to data basis
Verifying.Because gyro misalignment and scale factor error need certain attitude maneuver to be motivated, design aircraft calibration
Track is as shown in figure 3, the comparison of inertial device error calibration result is as shown in Fig. 4 .1- Fig. 4 .8 before and after installation vibration absorber.
According to the track in the terminal guidance stage of Fig. 3 design, the reality output of the gyro in guided missile operational process is simulated, and is led to
Crossing inertial device error influences principle to antenna attitude angle error, has carried out 100 emulation.For ascertainment error, with items
For the calibration value of error parameter as reference, being generated in 100 emulation experiments using calibration value is the random number of maximum value as top
Spiral shell instrument error parameter, as a result such as Fig. 5 .1- Fig. 5 .4.
Comparing result can be seen that one to gyroscope simulation data and the true output data of gyroscope in Fig. 4 .1- Fig. 4 .8
Cause property is verified, and gyroscope, which emulates data, has validity.It can be seen that the gyroscope output of emulation from Fig. 5 .1- Fig. 5 .4
Data can be in the static dynamic output characteristic with simulation gyroscope preferable under current intelligence, gyroscope simulation data and true output
Consistency be verified, for inertia device vibration absorber installation have certain guidance value.
The above examples only illustrate the technical idea of the present invention, and this does not limit the scope of protection of the present invention, all
According to the technical idea provided by the invention, any changes made on the basis of the technical scheme each falls within the scope of the present invention
Within.
Claims (5)
1. a kind of vibration absorber influences test method to the performance of inertial navigation system, which comprises the steps of:
Step 1, establish conventional gyroscope error model, the gyro error include gyro misalignment, scale factor error and with
Machine drift error；
Step 2, on the basis of step 1 pair conventional gyro error modeling, the error of three classes error described in step 1 is considered to introduce
New error model is established in the influence generated after vibration absorber；
Step 3, it is tested by calibration, the two kinds of error models obtained using steps 1 and 2, provides mistake for gyro performance is quantitatively evaluated
Difference data, the validity of verifying gyroscope emulation data；
Step 4, the error information obtained according to step 3 further analyzes isolation mounting to antenna attitude angle error impact analysis
Influence of the increase for attitude algorithm.
2. vibration absorber according to claim 1 influences test method to the performance of inertial navigation system, which is characterized in that
Routine gyroscope error model described in step 1 are as follows:
Gyro reality outputω is exported by gyro idealbAnd gyro items error composition, error model are as follows:
Wherein,For gyro reality output, ωbFor the output of error free gyro, εbFor gyro zero bias, ωsFor scale factor error,
ωθFor nonopiate coupling error, ωGFor acceleration sensitive error, η is random noise item；
1) gyro zero bias
In formula,For component of the gyro zero bias on three axis of gyroscope；
2) gyro scale factor error
In formula,For gyro scale factor error coefficient matrix；
3) gyro accelerometer sensitivity error
In formula,For the coefficient matrix of gyro accelerometer sensitivity error,For gyro
Acceleration on instrument three axis of ideal；
4) the nonopiate coupling error of gyro
In formula,For the coefficient matrix of the nonopiate coupling error of gyro；
5) Gyro Random noise
η=εr+ωg
In formula, ωgFor white noise, εrFor single order markoff process random noise.
3. vibration absorber according to claim 2 influences test method to the performance of inertial navigation system, which is characterized in that
New gyroscope error model described in step 2 are as follows:
Gyro reality output after vibration absorber is addedω is exported by gyro idealbIt is every with the gyro after addition vibration absorber
Error composition, are as follows:
Wherein, ε 'bTo increase the gyro zero bias after vibration absorber, ω 'sTo increase the scale factor error after vibration absorber, ω 'θ
To increase the nonopiate coupling error after vibration absorber, ω 'GTo increase the acceleration sensitive error after vibration absorber, η ' is to increase
Random noise item after adding and subtracting vibrating device, GATo increase the installation error coefficient matrix after vibration absorber；
1) the gyro zero bias after adding and subtracting vibrating device indicate are as follows:
Wherein:For component of the gyro zero bias on three axis of gyroscope；
2) the gyro scale factor error after adding and subtracting vibrating device indicates are as follows:
Wherein,For the gyro scale factor error coefficient matrix after plus-minus vibrating device；
3) the gyro accelerometer sensitivity error after adding and subtracting vibrating device is expressed as:
Wherein,For the coefficient matrix for adding and subtracting vibrating device gyro accelerometer sensitivity error；
4) the nonopiate coupling error of gyro after adding and subtracting vibrating device indicates are as follows:
Wherein,For the coefficient matrix of the nonopiate coupling error of gyro after plus-minus vibrating device.
4. vibration absorber according to claim 1 influences test method to the performance of inertial navigation system, which is characterized in that
Test process is demarcated described in step 3:
(1) random error is tested
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into 1h on turntable and carries out data acquisition, be repeated 3 times；
(2) zero bias are tested
1) gyro is separately fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on into progress zero bias sampling, every group of acquisition 1min on double axle table；
(3) acceleration sensitive error testing
1) gyro is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) gyro is rested on double axle table and is sampled, every group of acquisition 1min；
(4) scale factor error is tested
1) gyroscope is fixed on double axle table by fixture, connects equipment power supply line and data line；
2) operation turntable make inertia component rotating around three axis accelerometer with ± 10 °/s, ± 30 °/s, ± 50 °/s, ± 70 °/s, ±
90 °/s, ± 110 °/s, ± 130 °/s, 18 ± 150 °/s, ± 180 °/s angular speed rotations；
(5) nonopiate coupling error test
1) gyroscope is fixed on angle vibration table by fixture, connects equipment power supply line and data line；
2) give gyroscope to be powered, start double axle table after 5s, operation turntable make inertia component rotating around three axis accelerometer with ± 10 °/
S, ± 30 °/s, ± 50 °/s, ± 70 °/s, ± 90 °/s, ± 110 °/s, ± 130 °/s, ± 150 °/s, 18 angles ± 180 °/s
Rate rotation.Wherein, ωiThe angular speed rotated for turntable around gyro i axis.
5. vibration absorber according to claim 1 influences test method to the performance of inertial navigation system, which is characterized in that
The detailed process of the step 4 are as follows:
Relationship according to quaternary number and Eulerian angles obtains following relational expression, carrier coordinate system be it is right front upper, then by Department of Geography to
The transformation quaternary number of carrier system are as follows:
In formula, ψ, θ, γ are respectively course angle, pitch angle and the roll angle of carrier, q0For real part, q1For imaginary part unit i, q2For
Imaginary part unit j, q3For imaginary part unit k；
Quaternary number is obtained by quaternion differential equation and iterates to calculate formula are as follows:
Wherein
Δθ0To take approximate definition value, Δ θ to exponential integralxFor Δ θ0
X-component, Δ θyFor Δ θ0Y-component, Δ θzFor Δ θ0Z-component,
Δθ0 2=Δ θx 2+Δθy 2+Δθz 2,
[] indicates the matrix-expand of vector in formula, and
Wherein,The projection x-axis output that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope body
The relatively geographical projection y-axis output for tying up to body and fastening of system,Tie up to what body was fastened for gyroscope body system is relatively geographical
Z-axis output is projected,The projection that body is fastened is tied up to for gyroscope body system geography relatively,For gyroscope body, system is opposite
The projection output that inertial system is fastened in body,The output of the projection in Department of Geography is tied up to for gyroscope earth system relative inertness,The output of the projection in Department of Geography is tied up to respect to the earth for gyroscope Department of Geography, is solvedFormula in again need useBenefit
With last momentIt is iterated, therefore the solution of Eulerian angles is a loop iteration process；
The transformation relation between fixed vector indicated according to quaternary number, obtains the relationship of quaternionic matrix and direction cosines:
Finally according to the following formula, attitude angle is found out
The output of target seeker gyro is transformed into antenna attitude angle by above calculating, so that the gyro output with error is mapped as
Antenna attitude angle error substitutes into the gyroscope error model that steps 1 and 2 are established, and gyro is influenced in conjunction with step 3 vibration absorber
Analysis, the influence by emulation preliminary analysis gyro error to antenna attitude angle error.
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