CN106248065B - A kind of method and system of time vehicle launch after effect period and range measurement - Google Patents
A kind of method and system of time vehicle launch after effect period and range measurement Download PDFInfo
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Abstract
The present invention relates to a kind of time vehicle launch after effect period and distance measurement method and system, wherein method includes：The measurement data of aircraft is acquired by micro-inertia sensor and magnetic resistance electronic compass and is handled；According to the situation of change of the measurement data, it is broken down into static data and dynamic data；The initial alignment that attitude of flight vehicle angle is carried out to static data calculates；To the aircraft initial attitude angle of dynamic data and acquisition, attitude of flight vehicle angle information is calculated, and then calculates aircraft speed information and Aircraft position information；Calculating aircraft leaves emitter moment and transmitting gas effect finish time corresponding Aircraft position information, and obtains the operating distance of vehicle launch after effect period accordingly；The operating distance of output vehicle launch after effect period and before collected vehicle launch after effect period time span.The problems such as present invention solves information collection heavy workload existing for high-speed digital photography method, image processing algorithm is complicated.
Description
Technical field
The present invention relates to electroporation field more particularly to a kind of sides of time vehicle launch after effect period and range measurement
Method and system.
Background technology
Due to the vehicle launch after effect period belong to high temperature, high pressure, high speed ternary nonstationary flow, air motion phenomenon is non-
It is often complicated, and the time is very of short duration, only several milliseconds, therefore whether theoretical research or verification in terms of testing measurement all
There is prodigious difficulty.
So far, for vehicle launch after effect period time span and operating distance this 2 important parameters not yet at
Measurement method ripe, conveniently, practical.As although common shadowgraph method can observe some physical phenomenons of after effect period,
Discrete several pictures are can only obtain, it is difficult to the finish time of vehicle launch after effect period accurately be photographed, to be difficult to be flown
Row device emits the time span and operating distance of after effect period；And high-speed digital photography rule existence information collecting work amount is big, figure
As the problems such as Processing Algorithm is complicated, data processing is difficult, hardware device is of high cost and measurement accuracy is not high, therefore, after existing
Time effect phase and the limitation of distance measurement method due to various reasons, cannot be satisfied vehicle launch test measurement demand.
Invention content
In view of above-mentioned analysis, the present invention is intended to provide a kind of method of time vehicle launch after effect period and range measurement
And system, to solve information collection heavy workload existing for existing solution high-speed digital photography method, image processing algorithm it is complicated,
The problems such as data processing is difficult, hardware device is of high cost and measurement accuracy is not high.
The purpose of the present invention is mainly achieved through the following technical solutions：
The present invention provides a kind of time vehicle launch after effect period and distance measurement methods, including：
The measurement data of aircraft is acquired by micro-inertia sensor and magnetic resistance electronic compass and is handled；
According to the situation of change of the measurement data, it is broken down into static data and dynamic data；
The initial alignment that attitude of flight vehicle angle is carried out to static data calculates；It is initial to dynamic data and the aircraft of acquisition
Attitude angle calculates attitude of flight vehicle angle information, and then calculates aircraft speed information and Aircraft position information；
Calculating aircraft leaves emitter moment and transmitting gas effect finish time corresponding position of aircraft letter
Breath, and the operating distance of vehicle launch after effect period is obtained accordingly；
It exports the operating distance of vehicle launch after effect period and collected time vehicle launch after effect period is long before
Degree, and shown.
Further, the micro-inertia sensor specifically includes：Three axis micro-inertia sensors and three axis gyroscopes, then institute
The method of stating further includes：
The error model of the micro- inertial acceleration meter of three axis, the error model of three axis gyroscopes and magnetic resistance electricity are established respectively
The error model of sub- compass；
Error compensation is carried out to measurement data according to above-mentioned error model.
Further, the error model for establishing the micro- inertial acceleration meter of three axis is：
A=A0+KA·Fij·a+δ
In formula, A is the micro- inertial acceleration meter output valve of three axis；A0For the micro- inertial acceleration meter zero bias of three axis；KAIt is micro- for three axis
Inertial acceleration meter scale factor；Fij(i, j=x, y, z) is quadrature error system of the micro- inertial acceleration meter i axis of three axis to j axis
Number；A is that aircraft moves input acceleration；δ is the micro- inertial acceleration meter random error of three axis；
The error model for establishing three axis gyroscopes is：
G=G0+KG(Eij·ω)+Dij·a+ε
In formula, G is three axis gyroscope output valves；G0For the zero bias of three axis gyroscopes；KGFor the mark of three axis gyroscopes
Spend coefficient；ω is that aircraft moves input angular velocity；ε is three axis gyroscope random errors；Eij(i, j=x, y, z) is three axis
Installation error coefficient of the gyroscope i axis to j axis；Dij(i, j=x, y, z) is first order error coefficient related with acceleration；
The error model for establishing magnetic resistance electronic compass is：
ψ=ψc+σ1+σ2sinψ+σ3cosψ+σ4sin(2ψ)+σ5cos(2ψ)
In formula, ψ is the output valve of magnetic resistance electronic compass；ψcFor the aircraft course angle pre-entered；σ1、σ2、σ3、σ4、σ5
For the penalty coefficient of magnetic resistance electronic compass.
Further, the process that the initial alignment calculating at attitude of flight vehicle angle is carried out to static data specifically includes：
The pitching angle theta of initial time aircraft is determined by the acceleration static information of the micro- inertial acceleration meter output of three axis0
With roll angle γ0, formula is：
γ0=arctan (- ax/az) (8)
The course angle ψ of initial time aircraft is determined by magnetic resistance electronic compass0：
ψ0=ψc (9)
The initial attitude angle under aircraft initial rest state is obtained by formula (7), formula (8), formula (9)：Pitching angle theta0, it is horizontal
Roll angle γ0, course angle ψ0。
Further, to the aircraft initial attitude angle of dynamic data and acquisition, using strap inertial navigation algorithm to knot
Fruit carries out continuous integral processing, and is transformed into navigational coordinate system, first calculates the attitude angle information of aircraft：Pitching angle theta, boat
To angle ψ and roll angle γ；The velocity information of aircraft is calculated again：Lateral velocity vx, forward speed vyWith longitudinal velocity vz；Most
Location information lateral displacement x, forward direction displacement y and the length travel z of aircraft are calculated afterwards.
Further, the vehicle launch after effect period time span t=t2-t1, t1When leaving emitter for aircraft
It carves, t2Finish time is acted on for transmitting gas；
The operating distance of the vehicle launch after effect period is calculated according to following formula：
Wherein, x1、y1、z1The emitter moment pair is left for aircraft
The position of aircraft answered；x2、y2、z2Finish time corresponding position of aircraft is acted on for transmitting gas；L is after vehicle launch
Effect phase operating distance.
The present invention also provides a kind of time vehicle launch after effect period and Range Measurement Systems, including：Micro- inertia sensing
Device and magnetic resistance electronic compass, Information Collecting ＆ Processing device, after effect period resolve processor and display control device；Micro-inertia sensor with
Magnetic resistance electronic compass is packaged in one, and is connect with Information Collecting ＆ Processing device by communication bus；Information Collecting ＆ Processing device
Processor is resolved by communication bus with the after effect period to connect；After effect period resolves processor and is connected by communication bus and display control device
It connects；
The micro-inertia sensor, magnetic resistance electronic compass, measurement data for acquiring aircraft simultaneously pass through communication bus
It is sent to described information acquisition processing device；
Described information acquisition processing device, for will the measurement data that received carry out signal conversion after be sent to it is described after
The effect phase resolves processor；
The after effect period resolves processor, for the situation of change according to the measurement data, is broken down into static number
According to dynamic data；The initial alignment that attitude of flight vehicle angle is carried out to static data calculates；Flight to dynamic data and acquisition
Device initial attitude angle calculates attitude of flight vehicle angle information, and then calculates aircraft speed information and position of aircraft letter
Breath；
Calculating aircraft leaves emitter moment and transmitting gas effect finish time corresponding position of aircraft letter
Breath, and the operating distance of vehicle launch after effect period is obtained accordingly；Export the vehicle launch after effect period operating distance and it
Preceding collected vehicle launch after effect period time span is shown to the display control device.
Further, the after effect period resolves processor and specifically includes：
Error compensation module, by establishing error model based on the micro- inertial acceleration of three axis, three axis gyroscopes respectively
The error model of error model and magnetic resistance electronic compass, and error compensation is carried out to measurement data according to above-mentioned error model；
Quiet dynamic measuring data recognizes module, for the situation of change according to the measurement data, is broken down into static state
Data and dynamic data；
Initial alignment modules, the initial alignment for carrying out attitude of flight vehicle angle to static data calculates, and will calculate
To initial attitude angle be sent to parameter calculating module；
Parameter calculating module, for the aircraft initial attitude angle to dynamic data and acquisition, using strap-down inertial
Algorithm carries out continuous integral processing to result, and is transformed into navigational coordinate system, first calculates the attitude angle information of aircraft, into
And calculate aircraft speed information and Aircraft position information；
Time gap computing module, at the end of leaving emitter moment and transmitting gas effect for calculating aircraft
Corresponding Aircraft position information is carved, and obtains the operating distance of vehicle launch after effect period accordingly, and will output aircraft hair
Penetrate the after effect period operating distance and before collected vehicle launch after effect period time span give the display control device
Further, the error model that the error compensation module establishes the micro- inertial acceleration meter of three axis is：
A=A0+KA·Fij·a+δ
In formula, A is the micro- inertial acceleration meter output valve of three axis；A0For the micro- inertial acceleration meter zero bias of three axis；KAIt is micro- for three axis
Inertial acceleration meter scale factor；Fij(i, j=x, y, z) is quadrature error system of the micro- inertial acceleration meter i axis of three axis to j axis
Number；A is that aircraft moves input acceleration；δ is the micro- inertial acceleration meter random error of three axis；
The error model of three axis gyroscopes established is：
G=G0+KG(Eij·ω)+Dij·a+ε
In formula, G is three axis gyroscope output valves；G0For the zero bias of three axis gyroscopes (10)；KGFor three axis gyroscopes
(10) scale factor；ω is that aircraft moves input angular velocity；ε is three axis gyroscope random errors；Eij(i, j=x, y,
Z) it is installation error coefficient of the three axis gyroscope i axis to j axis；Dij(i, j=x, y, z) misses for first order related with acceleration
Poor coefficient.
The error model of the magnetic resistance electronic compass of foundation is：
ψ=ψc+σ1+σ2sinψ+σ3cosψ+σ4sin(2ψ)+σ5cos(2ψ)
In formula, ψ is the output valve of magnetic resistance electronic compass；ψcFor the aircraft course angle pre-entered；σ1、σ2、σ3、σ4、σ5
For the penalty coefficient of magnetic resistance electronic compass.
Further, the time gap computing module is specifically used for,
The vehicle launch after effect period time span t=t2-t1, t1Emitter moment, t are left for aircraft2For hair
Body of emanating acts on finish time；
The operating distance of the vehicle launch after effect period is calculated according to following formula：
Wherein, x1、y1、z1The emitter moment pair is left for aircraft
The position of aircraft answered；x2、y2、z2Finish time corresponding position of aircraft is acted on for transmitting gas；L is after vehicle launch
Effect phase operating distance.
The present invention has the beneficial effect that：
The present invention solves information collection heavy workload existing for high-speed digital photography method, image processing algorithm complexity, data
The problems such as processing is difficult, hardware device is of high cost and measurement accuracy is not high.
Other features and advantages of the present invention will illustrate in the following description, also, partial become from specification
It obtains it is clear that understand through the implementation of the invention.The purpose of the present invention and other advantages can be by the explanations write
Specifically noted structure is realized and is obtained in book, claims and attached drawing.
Description of the drawings
Attached drawing is only used for showing the purpose of specific embodiment, and is not considered as limitation of the present invention, in entire attached drawing
In, identical reference mark indicates identical component.
Fig. 1 is the structural schematic diagram of system described in the embodiment of the present invention；
Fig. 2 is the flow diagram of the method for the embodiment of the present invention.
Specific implementation mode
Specifically describing the preferred embodiment of the present invention below in conjunction with the accompanying drawings, wherein attached drawing constitutes the application part, and
It is used to illustrate the principle of the present invention together with embodiments of the present invention.
As shown in FIG. 1, FIG. 1 is the structural schematic diagrams of system described in the embodiment of the present invention, may include mainly：Micro- inertia passes
Sensor, magnetic resistance electronic compass, Information Collecting ＆ Processing device, after effect period resolve processor, display control device, communication bus, communication always
Line, communication bus.Wherein micro-inertia sensor includes：The micro- inertial acceleration meter of three axis and the micro- inertial gyroscope of three axis；After effect period
Resolving processor includes：Error compensation module, quiet dynamic data identification module, initial alignment modules and parameter calculating module.It is micro-
Inertial sensor is packaged in one with magnetic resistance electronic compass, and is connect with Information Collecting ＆ Processing device by communication bus；Information
Acquisition processing device resolves processor with the after effect period by communication bus and connect；After effect period resolve processor by communication bus with
Display control device connects；Display control device is used for human-computer interaction.
It should be noted that the specific implementation process due to the system various pieces will carry out in the description of following methods
It is described in detail, therefore details are not described herein again.
As shown in Fig. 2, Fig. 2 is the flow diagram of the method for the embodiment of the present invention, following steps are can specifically include：
Step 201：Measurement data acquisition
After after effect period computing system being installed on aircraft flight device to be measured, micro-inertia sensor and magnetic resistance electronics
Compass starts to acquire aircraft data and be passed after data acquisition processing device handles (voltage signal is converted into physical signal)
It is defeated by after effect period resolving processor.
Step 202：Error compensation module compensates error of measured data
After after effect period resolves the measurement data that processor receives micro-inertia sensor and magnetic resistance electronic compass, mended by error
It repays module and carries out error compensation, wherein the process of error compensation includes mainly：The mistake of the micro- inertial acceleration meter of three axis is established respectively
The error model of differential mode type, the error model of three axis gyroscopes and magnetic resistance electronic compass；According to above-mentioned error model to surveying
It measures data and carries out error compensation
Above-mentioned each error compensation model is described as follows respectively：
Consider the zero bias of the micro- inertial acceleration meter of three axis, installation error, Random Drift Error item, ignores second order or more dynamic
Small errors, the error model for establishing the micro- inertial acceleration meter of three axis are：
A=A0+KA·Fij·a+δ (1)
In formula, A is the micro- inertial acceleration meter output valve of three axis；A0For the micro- inertial acceleration meter zero bias of three axis；KAIt is micro- for three axis
Inertial acceleration meter scale factor；Fij(i, j=x, y, z) is quadrature error system of the micro- inertial acceleration meter i axis of three axis to j axis
Number；A is that aircraft moves input acceleration；δ is the micro- inertial acceleration meter random error of three axis.
The expression formula of each matrix is：
A=[Ax Ay Az]T；A0=[Ax0 Ay0 Az0]T；
δ=[δx δy δz]T；A=[ax ay az]T。
In formula, Ax、Ay、AzFor the output of micro- three axis of inertial acceleration meter x, y, z；Ax0、Ay0、Az0For micro- inertial acceleration meter
The zero bias of three axis of x, y, z；KAx、KAy、KAzFor the scale factor of micro- three axis of inertial acceleration meter x, y, z；Fxy、Fxz、Fyx、Fyz、
Fzx、FzyIt is the corresponding i axis of micro- inertial acceleration meter to the quadrature error coefficient (i, j=x, y, z) of j axis；δx、δy、δzMicro- inertia
The random error of three axis of accelerometer x, y, z；ax、ay、azFor aircraft x, y, z three-axis moving input acceleration, i.e. aircraft is transported
Dynamic true acceleration.
Similarly, zero bias, installation error, quadrature error and the Random Drift Error for considering three axis gyroscopes, ignore two
The above dynamic Small errors of rank, the error model for establishing three axis gyroscopes are：
G=G0+KG(Eij·ω)+Dij·a+ε (2)
In formula, G is three axis gyroscope output valves；G0For the zero bias of three axis gyroscopes (10)；KGFor the micro- inertia top of three axis
The scale factor of spiral shell instrument (10)；ω is that aircraft moves input angular velocity；ε is the micro- inertial gyroscope random error of three axis；Eij(i,
J=x, y, z) it is installation error coefficient of the micro- inertial gyroscope i axis of three axis to j axis；Dij(i, j=x, y, z) is micro- inertia gyroscope
Instrument first order error coefficient related with acceleration.
The expression formula of each matrix is：
G=[Gx Gy Gz]T；G0=[Gx0 Gy0 Gz0]T；
ω=[ωx ωy ωz]T；ε=[εx εy εz]T。
In formula, Gx、Gy、GzFor the output of micro- three axis of inertial gyroscope x, y, z of three axis；Gx0、Gy0、Gz0For micro- inertial gyroscope
The zero bias of three axis of x, y, z；Exy、Exz、Eyx、Eyz、Ezx、EzyIt is the corresponding i axis of the micro- inertial gyroscope of three axis to the installation error of j axis
Coefficient (i, j=x, y, z)；Dxx、Dxy、Dxz、Dyx、Dyy、Dyz、Dzx、Dzy、DzzIt is the corresponding i axis of the micro- inertial gyroscope of three axis to j
(i, j=x, y, z) first order error coefficient related with acceleration of axis；KGx、KGy、KGzFor the micro- inertial gyroscope x, y, z of three axis
The scale factor of three axis；ωx、ωy、ωzFor the angular speed of aircraft x, y, z three-axis moving input；εx、εy、εzIt is micro- used for three axis
The random error of three axis of property gyroscope x, y, z.
It is carried out at normal temperatures when in view of ejection test, the calibration of three axis micro- inertial acceleration meter and the micro- inertial gyroscope of three axis
Experiment carries out under normal temperature condition, ignores the influence that temperature exports sensor.Micro- inertia is determined using " six-position testing method "
The zero bias of accelerometer, scale factor, orthogonal axle misalignment coefficient, micro- inertial gyroscope are to acceleration sensitive item；Using " speed
Rate indexing method of testing " determines the zero bias of micro- inertial gyroscope, scale factor, orthogonal axle misalignment coefficient.
Formula (1), formula (2) become and got in return
In formula, a=[ax ay az]TFor the output of micro- three axis of inertial acceleration meter of three axis after compensation, i.e. aircraft moves
The acceleration actually entered；ω=[ωx ωy ωz]TFor the output of micro- three axis of inertial gyroscope of three axis after compensation, as fly
Row device moves the angular speed actually entered.
The error model for establishing magnetic resistance electronic compass is：
ψ=ψc+σ1+σ2sinψ+σ3cosψ+σ4sin(2ψ)+σ5cos(2ψ) (5)
In formula, ψ is the output valve of magnetic resistance electronic compass；ψcFor the aircraft course angle pre-entered, i.e., after error compensation
Actual heading angle；σ1、σ2、σ3、σ4、σ5For the penalty coefficient of magnetic resistance electronic compass.
Using least square method, that is, it is based on " error sum of squares minimum ", between 0 °~360 °, each 15 ° totally 24 realities
A progress error testing is tested, 24 groups of data is obtained, remembers error equation：
U Ω=H (6)
In formula,Through meter
The penalty coefficient σ of magnetic resistance electronic compass can be obtained by calculating1、σ2、σ3、σ4、σ5。
Error compensation module is transferred to quiet dynamic measuring data identification module after carrying out error compensation to measurement data.
Step 203：Quiet dynamic measuring data identification module identification aircraft static state and dynamic data
Quiet dynamic measuring data identification module receives after the measurement data of error compensation, is surveyed according to micro-inertia sensor
The situation of change of amount data picks out the static state and dynamic of aircraft, and measurement data is decomposed into static data and dynamic number
According to.Wherein, static data is used for the initial alignment at attitude of flight vehicle angle, and dynamic data is used for the posture renewal of aircraft, speed
Update and location updating.
Step 204：Initial alignment modules calculating aircraft initial attitude angle
The static data that initial alignment modules recognize module according to quiet dynamic measuring data carries out the first of attitude of flight vehicle angle
Begin to be aligned and calculate.
The pitching angle theta of initial time aircraft is determined by the acceleration static information of the micro- inertial acceleration meter output of three axis0
With roll angle γ0, formula is：
γ0=arctan (- ax/az) (8)
The course angle ψ of initial time aircraft is determined by magnetic resistance electronic compass0：
ψ0=ψc (9)
The initial attitude angle under aircraft initial rest state is obtained by formula (7), formula (8), formula (9)：Pitching angle theta0, it is horizontal
Roll angle γ0, course angle ψ0。
The initial attitude angle information of aircraft is sent to the parameter meter in after effect period resolving processor by initial alignment modules
Calculate module.
Step 205：Parameter calculating module calculating aircraft movable information, and then calculate aircraft speed information and fly
Row device location information；
(acquisition flies for the dynamic data that parameter calculating module goes out according to quiet Dynamic Identification module identification and initial alignment modules
Row device initial attitude angle carries out continuous integral processing to result using strap inertial navigation algorithm, and is transformed into navigational coordinate system
In, first calculate the attitude angle information of aircraft：Pitching angle theta, course angle ψ and roll angle γ；The speed of aircraft is calculated again
Information：Lateral velocity vx, forward speed vyWith longitudinal velocity vz；Finally calculate the location information lateral displacement x, preceding of aircraft
To displacement y and length travel z.
The attitude angle information of aircraft calculates：
The attitude matrix T of aircraft coordinate system b to navigational coordinate system n is：
Attitude matrix T and quaternary number q=[q0 q1 q2 q3]TRelationship it is as follows：
The initial attitude angle obtained by initial alignment modules obtains original state matrix T by formula (10)0, it is aircraft
Posture renewal provide initial value.According to formula (11), by original state matrix T0Quaternary number q=[q can be found out0 q1 q2 q3]T
Initial value q0, by initial value q0As the input of following formula (16), continuous integral processing is carried out.
Since the micro- inertial gyroscope angular speed of three axis is measured in aircraft coordinate system, need to convert it to navigation seat
In mark system, have：
In formula,For the angular speed of aircraft in navigational coordinate system,For aircraft coordinate system
The angular speed of interior aircraft.
When in view of experiment, the variation of aircraft absolute distance is little, therefore,
In formula,The angular speed of aircraft relative inertness spatial rotational is measured in aircraft system for the micro- inertial gyroscope of three axis
In projection, i.e., the output of micro- inertial gyroscope after error compensation
Due to
According to the associative law of multiplication of quaternary number, can be obtained by formula (12)：
The form that formula (15) is write as matrix is had：
In addition,
The angular speed measured with the micro- inertial gyroscope of three axis Jing Guo error compensationTo formula (16) into
Row quadravalence dragon lattice-Ku Tafa is calculated, and does normalized according to formula (17), you can realizes the real-time update of quaternary number.
After formula (16) formula and (17) calculate quaternary number in real time, the update of attitude matrix is completed by formula (11), and can
According to the transformational relation of formula (10) and formula (11) can inverse obtain aircraft attitude angle information：Pitching angle theta, course angle ψ, cross
Roll angle γ.
The velocity information of aircraft calculates：
The acceleration information a that the micro- inertial acceleration meter of three axis through overcompensation measuresx、ay、az, pass through attitude matrix T and hair
The transformation of coordinate system is penetrated to navigational coordinate system, the speed update of aircraft in navigational coordinate system is carried out by once integrating.
In formula, g is terrestrial gravitation acceleration.
In conjunction with the attitude angle information of the calculated aircraft in front, second order Runge-Kutta method calculating is carried out to formula (18), is obtained
To the velocity information of aircraft：
The positional information calculation of aircraft：
Integral and calculating is carried out again to get to the location information of aircraft to formula (19)：x、y、z.
Step 206：After effect period time span calculates
The vehicle launch after effect period refers to that aircraft leaves the emitter moment to transmitting gas effect finish time.It is corresponding
The emitter moment is left in aircraft, is denoted as t1, aircraft will not be subject to the supporting role of emitter；Corresponding to transmitting
Gas acts on finish time, is denoted as t2, aircraft, which will not be subject to, emits the thrust that gas generates.The two moment fly
The acceleration of device will occur significantly to change, therefore t1With t2It can be straight according to the acquisition information of Information Collecting ＆ Processing device (3)
It connects to obtain, i.e.,
T=t2-t1 (20)
In formula, t is vehicle launch after effect period time span, the information collection frequency of precision and after effect period computing system
It is related.
Step 207：After effect period operating distance calculates
T can be solved according to formula (13)1、t2Moment corresponding Aircraft position information, you can after obtaining vehicle launch
The operating distance of effect phase.
In formula, x1、y1、z1Emitter moment t is left for aircraft1Corresponding position of aircraft；x2、y2、z2For transmitting
Gas acts on finish time t2Corresponding position of aircraft；L is vehicle launch after effect period operating distance.
After effect period resolving processor output vehicle launch after effect period time span information t and operating distance information L, and by
Communication bus C is transferred to display control device.
Step 208：Display control device is by the vehicle launch after effect period time span information received and operating distance information
Output display.
So far, the measurement of vehicle launch after effect period time span and operating distance is completed.
In conclusion an embodiment of the present invention provides a kind of method of time vehicle launch after effect period and range measurement and
System, solve information collection heavy workload, image processing algorithm are complicated existing for high-speed digital photography method, data processing is difficult,
The problems such as hardware device is of high cost and measurement accuracy is not high, when can effectively measure vehicle launch the time span of after effect period with
Operating distance.
It will be understood by those skilled in the art that realizing all or part of flow of above-described embodiment method, meter can be passed through
Calculation machine program is completed to instruct relevant hardware, and the program can be stored in computer readable storage medium.Wherein, institute
It is disk, CD, read-only memory or random access memory etc. to state computer readable storage medium.
Although the present invention and its advantage has been described in detail it should be appreciated that without departing from by the attached claims
Defined by can carry out various changes, replacement and transformation in the case of the spirit and scope of the present invention.Moreover, the model of the application
Enclose the specific embodiment for being not limited only to process, equipment, means, method and steps described in specification.In the art is common
Technical staff executes and corresponding reality described herein from the disclosure it will be readily understood that can be used according to the present invention
Apply the essentially identical function of example or obtain the result essentially identical with it, existing and process to be developed in future, equipment,
Means, method or step.Therefore, the attached claims purport includes such process, equipment, hand in the range of them
Section, method or step.
The foregoing is only a preferred embodiment of the present invention, but scope of protection of the present invention is not limited thereto,
Any one skilled in the art in the technical scope disclosed by the present invention, the change or replacement that can be readily occurred in,
It should be covered by the protection scope of the present invention.
Claims (8)
1. a kind of time vehicle launch after effect period and distance measurement method, which is characterized in that including：
The measurement data of aircraft is acquired by micro-inertia sensor and magnetic resistance electronic compass and is handled；Micro- inertia passes
Sensor specifically includes three axis micro-inertia sensors and three axis gyroscopes, and the method further includes：
The error model of the micro- inertial acceleration meter of three axis, the error model of three axis gyroscopes and magnetic resistance electronics sieve are established respectively
The error model of disk；
The error model for establishing the micro- inertial acceleration meter of three axis is：
A=A0+KA·Fij·a+δ
In formula, A is the micro- inertial acceleration meter output valve of three axis；A0For the micro- inertial acceleration meter zero bias of three axis；KAFor the micro- inertia of three axis
Accelerometer scale factor；Fij(i, j=x, y, z) is quadrature error coefficient of the micro- inertial acceleration meter i axis of three axis to j axis；A is
Aircraft moves input acceleration；δ is the micro- inertial acceleration meter random error of three axis；
The error model for establishing three axis gyroscopes is：
G=G0+KG(Eij·ω)+Dij·a+ε
In formula, G is three axis gyroscope output valves；G0For the zero bias of three axis gyroscopes；KGFor the scale system of three axis gyroscopes
Number；ω is that aircraft moves input angular velocity；ε is three axis gyroscope random errors；Eij(i, j=x, y, z) is the micro- top of three axis
Installation error coefficient of the spiral shell instrument i axis to j axis；Dij(i, j=x, y, z) is first order error coefficient related with acceleration；
The error model for establishing magnetic resistance electronic compass is：
ψ=ψc+σ1+σ2sinψ+σ3cosψ+σ4sin(2ψ)+σ5cos(2ψ)
In formula, ψ is the output valve of magnetic resistance electronic compass；ψcFor the aircraft course angle pre-entered；σ1、σ2、σ3、σ4、σ5For magnetic
Hinder the penalty coefficient of electronic compass；
Error compensation is carried out to measurement data according to above-mentioned error model；
According to the situation of change of the measurement data, it is broken down into static data and dynamic data；
The initial alignment that attitude of flight vehicle angle is carried out to static data calculates；To the aircraft initial attitude of dynamic data and acquisition
Angle calculates attitude of flight vehicle angle information, and then calculates aircraft speed information and Aircraft position information；
Calculating aircraft leaves emitter moment and transmitting gas effect finish time corresponding Aircraft position information, and
The operating distance of vehicle launch after effect period is obtained accordingly；
The operating distance of output vehicle launch after effect period and before collected vehicle launch after effect period time span, and
It is shown.
2. according to the method described in claim 1, it is characterized in that, carrying out the initial alignment at attitude of flight vehicle angle to static data
The process of calculating specifically includes：
The pitching angle theta of initial time aircraft is determined by the acceleration static information of the micro- inertial acceleration meter output of three axis0And roll
Angle γ0, formula is：
γ0=arctan (- ax/az) (8)
The course angle ψ of initial time aircraft is determined by magnetic resistance electronic compass0：
ψ0=ψc (9)
The initial attitude angle under aircraft initial rest state is obtained by formula (7), formula (8), formula (9)：Pitching angle theta0, roll angle
γ0, course angle ψ0。
3. according to the method described in claim 1, it is characterized in that, aircraft initial attitude angle to dynamic data and acquisition,
Continuous integral processing is carried out to result using strap inertial navigation algorithm, and is transformed into navigational coordinate system, flight is first calculated
The attitude angle information of device：Pitching angle theta, course angle ψ and roll angle γ；The velocity information of aircraft is calculated again：Lateral velocity vx、
Forward speed vyWith longitudinal velocity vz；Finally calculate location information lateral displacement x, forward direction displacement y and the length travel of aircraft
z。
4. the method according to any one of Claim 1-3, which is characterized in that when the vehicle launch after effect period
Between length t=t2-t1, t1Emitter moment, t are left for aircraft2Finish time is acted on for transmitting gas；
The operating distance of the vehicle launch after effect period is calculated according to following formula：
Wherein, x1、y1、z1It is corresponding that the emitter moment is left for aircraft
Position of aircraft；x2、y2、z2Finish time corresponding position of aircraft is acted on for transmitting gas；L is the vehicle launch after effect period
Operating distance.
5. a kind of time vehicle launch after effect period and Range Measurement System, which is characterized in that including：Micro-inertia sensor and magnetic
Hinder electronic compass, Information Collecting ＆ Processing device, after effect period resolving processor and display control device；Micro-inertia sensor and magnetic resistance electricity
Sub- compass is packaged in one, and is connect with Information Collecting ＆ Processing device by communication bus；Information Collecting ＆ Processing device passes through logical
Letter bus resolves processor with the after effect period and connect；After effect period resolves processor and is connect with display control device by communication bus；
The micro-inertia sensor, magnetic resistance electronic compass, measurement data for acquiring aircraft are simultaneously sent by communication bus
Give described information acquisition processing device；
Described information acquisition processing device, for being sent to the after effect period after the measurement data received is carried out signal conversion
Resolve processor；
After effect period resolves processor, for the situation of change according to the measurement data, be broken down into static data with
Dynamic data；The initial alignment that attitude of flight vehicle angle is carried out to static data calculates；To at the beginning of the aircraft of dynamic data and acquisition
Beginning attitude angle calculates attitude of flight vehicle angle information, and then calculates aircraft speed information and Aircraft position information；
Calculating aircraft leaves emitter moment and transmitting gas effect finish time corresponding Aircraft position information, and
The operating distance of vehicle launch after effect period is obtained accordingly；It exports the operating distance of vehicle launch after effect period and acquires before
To vehicle launch after effect period time span shown to the display control device.
6. system according to claim 5, which is characterized in that the after effect period resolves processor and specifically includes：
Error compensation module, the error by establishing error model based on the micro- inertial acceleration of three axis, three axis gyroscopes respectively
The error model of model and magnetic resistance electronic compass, and error compensation is carried out to measurement data according to above-mentioned error model；
Quiet dynamic measuring data recognizes module, for the situation of change according to the measurement data, is broken down into static data
With dynamic data；
Initial alignment modules, initial alignment calculating for carrying out attitude of flight vehicle angle to static data, and will be calculated
Initial attitude angle is sent to parameter calculating module；
Parameter calculating module, for the aircraft initial attitude angle to dynamic data and acquisition, using strap inertial navigation algorithm
Continuous integral processing is carried out to result, and is transformed into navigational coordinate system, the attitude angle information of aircraft, Jin Erji are first calculated
Calculate aircraft speed information and Aircraft position information；
Time gap computing module leaves the emitter moment for calculating aircraft and transmitting gas acts on finish time pair
The Aircraft position information answered, and the operating distance of vehicle launch after effect period is obtained accordingly, and will be after output vehicle launch
The operating distance of effect phase and before collected vehicle launch after effect period time span give the display control device.
7. system according to claim 6, which is characterized in that
The error model for the micro- inertial acceleration meter of three axis that the error compensation module is established is：
A=A0+KA·Fij·a+δ
In formula, A is the micro- inertial acceleration meter output valve of three axis；A0For the micro- inertial acceleration meter zero bias of three axis；KAFor the micro- inertia of three axis
Accelerometer scale factor；Fij(i, j=x, y, z) is quadrature error coefficient of the micro- inertial acceleration meter i axis of three axis to j axis；A is
Aircraft moves input acceleration；δ is the micro- inertial acceleration meter random error of three axis；
The error model of three axis gyroscopes established is：
G=G0+KG(Eij·ω)+Dij·a+ε
In formula, G is three axis gyroscope output valves；G0For the zero bias of three axis gyroscopes (10)；KGFor three axis gyroscopes (10)
Scale factor；ω is that aircraft moves input angular velocity；ε is three axis gyroscope random errors；Eij(i, j=x, y, z) is
Installation error coefficient of the three axis gyroscope i axis to j axis；Dij(i, j=x, y, z) is first order error system related with acceleration
Number；
The error model of the magnetic resistance electronic compass of foundation is：
ψ=ψc+σ1+σ2sinψ+σ3cosψ+σ4sin(2ψ)+σ5cos(2ψ)
In formula, ψ is the output valve of magnetic resistance electronic compass；ψcFor the aircraft course angle pre-entered；σ1、σ2、σ3、σ4、σ5For magnetic
Hinder the penalty coefficient of electronic compass.
8. system according to claim 6, which is characterized in that the time gap computing module is specifically used for,
The vehicle launch after effect period time span t=t2-t1, t1Emitter moment, t are left for aircraft2To emit gas
Body acts on finish time；
The operating distance of the vehicle launch after effect period is calculated according to following formula：
Wherein, x1、y1、z1It is corresponding that the emitter moment is left for aircraft
Position of aircraft；x2、y2、z2Finish time corresponding position of aircraft is acted on for transmitting gas；L is the vehicle launch after effect period
Operating distance.
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