CN107065566A - Each link error distribution method of missile control system - Google Patents
Each link error distribution method of missile control system Download PDFInfo
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- CN107065566A CN107065566A CN201710360800.8A CN201710360800A CN107065566A CN 107065566 A CN107065566 A CN 107065566A CN 201710360800 A CN201710360800 A CN 201710360800A CN 107065566 A CN107065566 A CN 107065566A
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- G—PHYSICS
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- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
- G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
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Abstract
The present invention proposes a kind of each link error distribution method of missile control system, it is possible to achieve distribution from missile control system accuracy at target error to each link actual physics systematic error inside control system.Step 1: the accuracy at target error by having distributed to control system, according to setting weight, accuracy at target error caused by control system as each link is distributed to using the accuracy at target error of control system；Step 2: respectively using seeker error, missile-borne computer error, steering wheel error as input, missile position, which becomes, turns to output, derives itself error to the transmission function of missile position；Step 3: according to obtaining by the corresponding transmission function to missile position of each link, respectively obtaining the steady-state gain that itself error causes missile position to change；Step 4: by accuracy at target error caused by each link of control system, respectively divided by missile position change caused by each link itself error steady-state gain, the as corresponding error of missile control system own physical system.
Description
Technical field：
It is a kind of each link error of missile control system specifically the present invention relates to guided missile system design method field
Distribution method.
Background technology：
The accuracy at target error of guided missile is to weigh the important indicator that guided missile hits target accuracy, and it is real that it is characterized as guided missile
Border drop point is to the air line distance of target location, and dimension is long measure.When designer misses to accuracy at target caused by control system
When difference link each into system is allocated, it can only accomplish to be assigned to " accuracy at target error caused by so-and-so link " this layer
It is secondary, and to the error of physical system of each link itself, the but clear and definite design method of neither one.
The reason for such as steering wheel output angle of rudder reflection error is accuracy at target error caused by steering wheel, but both time
Corresponding relation, and how to go out corresponding angle of rudder reflection error from accuracy at target tolerance design caused by steering wheel, at this stage can only
Determined by experience or test, the design method of neither one parsing solves this problem.
The content of the invention：
Missed to solve missile control system overall hits trueness error each link own physical system into control system
The problem of difference is matched somebody with somebody, the present invention proposes a kind of each link error distribution method of missile control system, it is possible to achieve STT missile system
Distribution from accuracy at target error of uniting to each link actual physics systematic error inside control system.
The present invention is achieved through the following technical solutions：
A kind of each link error distribution method of missile control system, itself error of target seeker link is that target seeker measures guided missile
To the range deviation of laser beam axis, dimension is long measure；The control that itself error of missile-borne computer link exports for it to steering wheel
Command error processed, dimension is chronomere；Itself error of steering wheel link is the angle of rudder reflection of its output, and its dimension is angular unit；
Specifically include following steps：
Step 1: the accuracy at target error by having distributed to control system, according to setting weight, by the hit of control system
Trueness error distributes to steering wheel, missile-borne computer and the target seeker or receiver of control system, is hit as caused by each link
Trueness error；
Step 2: being used as input, guided missile using target seeker or receiver error, missile-borne computer error, steering wheel error respectively
Change in location is output, derives target seeker, missile-borne computer, the transmission function of steering wheel itself error to missile position；
Step 3: according to obtaining by the corresponding transmission function to missile position of each link, respectively obtain target seeker or
Itself error of receiver, missile-borne computer, steering wheel causes the steady-state gain that missile position changes；
Step 4: by accuracy at target error caused by each link of control system, respectively divided by each link itself error causes
Missile position change steady-state gain, as missile control system target seeker, missile-borne computer, steering wheel own physical system pair
The error answered.
The beneficial effects of the invention are as follows realize from trueness error in caused by missile control system life to each ring of control system
Save own physical systematic error distribution, for designer to control system carry out error distribution when provide theoretical direction and according to
According to.
Brief description of the drawings：
Fig. 1 is a kind of each link error distribution method flow chart of missile control system of the invention.
Missile control system block diagram under Fig. 2 laser-beam riding guidance systems.
Equivalent control system block diagram when Fig. 3 seeker errors are inputted as system.
Embodiment：
Each step to the present invention by taking laser-beam riding guidance missile control system as an example is specifically described below.
Missile control system block diagram under laser-beam riding guidance system is as shown in Fig. 2 in the control system of guided missile, using super
Preceding correction link, is not used automatic pilot, steering wheel adoption rate link.
In control system, the transmission function of corrective network is：
Wherein KjFor corrective network gain, a is indexing coefficient, and T is time constant.
RmFor missile air range, KsFor steering wheel gain.Angle of rudder reflection is to the transmission function of guided missile normal acceleration：
It is that seeker error, missile-borne computer error and steering wheel are missed respectively if the error source one of control system has three
Difference.Here error is defined as making between real output value and idea output the difference of difference.
Seeker error refers to target seeker after laser beam information is received, and its guided missile exported is missed apart from optical axis position
The true value and difference of the target seeker reality output to the position of missile-borne computer of difference, i.e. guided missile apart from optical axis position.It is one
Length physical quantity.
Missile-borne computer error refers to after the missile position information that missile-borne computer have received target seeker output, by it
The rudder command error that portion calculates processing and exported.Under laser-beam riding guidance system, guided missile is generally middle-size and small-size guided missile, the rudder of use
Machine is electrodynamic type steering wheel or vapour-pressure type steering wheel, therefore the control signal of missile-borne computer output is voltage signal, missile-borne computer
Output error be missile-borne computer reality output control voltage and the difference of desired output voltage, it is a voltage physical
Amount.
Steering wheel error shows as the angle of rudder reflection error of steering wheel output, i.e., after the rudder instruction that missile-borne computer is provided is received,
The difference of steering wheel reality output angle of rudder reflection and theoretical output angle of rudder reflection, it is a physical quantity in units of angle or radian.
If the error of three error sources carried is represented under conditions of input is certain, error source reality output and ideal
The difference of output, with it is defeated to error source enter it is unrelated, it can thus be assumed that the error of three kinds of error sources is uncertainty system error, and
And be separate.Therefore, it is possible to use separate random error distribution method is carried out to each link in control system
The distribution of accuracy at target error.
Step one：If the accuracy at target error criterion that overall system distributes to control system is | σcs|, control system distribution
Accuracy at target error criterion to target seeker, missile-borne computer and steering wheel is respectively | σc1|、|σc2| and | σc3|.According to difference
Weight is allocated to the corresponding accuracy at target error of each link in control system, if target seeker, missile-borne computer and steering wheel
Accuracy at target error accounts for the weight respectively λ of the total accuracy at target error of control systemc1, λc2And λc3, wherein λc1+λc2+λc3=
1.According to separate random error distribution method, then have：
So, then σ is met2 c1+σ2 c2+σ2 c3=σ2 cs, obey separate random error allocation criteria.
Step 2：Below by taking target seeker as an example, the biography of the corresponding accuracy at target error of its physical system error is derived
Pass relation.Inputted by control system of seeker error, missile position is output, can when assuming that error source only has target seeker
Control system block diagram after being deformed is as shown in Figure 3.
By Fig. 3, it can derive that missile position change is on the transmission function of seeker error：
Similarly, missile accuracy error is on the transmission function of missile-borne computer error：
Missile accuracy error is on the transmission function of steering wheel error：
Step 3：In figure 3, when seeker error input is 0, missile position output is also 0.Because guided missile hits essence
Degree error represents the distance between guided missile physical location and ideal position, therefore when error source only has target seeker, the position of guided missile
The physical location deviation that output is guided missile caused by seeker error is put, is exactly accuracy at target error caused by target seeker, formula
(2) it is exactly target seeker physical system error to the transmission function between its caused accuracy at target error.Target seeker physical system
Error can be represented to the carry-over factor between its caused accuracy at target error with the steady-state gain of formula (2).When time convergence
When infinite, the value of formula (2) levels off to 1, therefore can obtain target seeker (or receiver) itself error and cause what missile position changed
Steady-state gain is 1.Missile-borne computer itself error, which can similarly be obtained, causes the steady-state gain that missile position changes to be 1/Kj, steering wheel is certainly
The steady-state gain that body error causes missile position to change is 1/ (KjKs)。
Step 4：Target seeker physical system error is to the transitive relation between its caused accuracy at target error：
|σc1|=| σseeker|
Wherein, | σseeker| represent that target seeker detects guided missile to laser beam optical axis distance error, unit is rice.
Similarly, missile-borne computer physical system error is to the transitive relation between its caused accuracy at target error：
|σ2|=| σcomputer|/Kj
Wherein, | σcomputer| the control voltage error of missile-borne computer output is represented, unit is volt.
Steering wheel physical system error is to the transitive relation between its caused accuracy at target error：
|σc3|=| σservor|/(KjKs)
Wherein, | σservor| the angle of rudder reflection error of steering wheel output is represented, unit is radian.
To sum up, can obtain each link own physical systematic error of missile control system is：
Claims (5)
1. a kind of each link error distribution method of missile control system, itself error of target seeker link measures guided missile for target seeker and arrived
The range deviation of laser beam axis, dimension is long measure；The control that itself error of missile-borne computer link exports for it to steering wheel
Command error, dimension is chronomere；Itself error of steering wheel link is the angle of rudder reflection of its output, and its dimension is angular unit；Its
It is characterised by, specifically includes following steps：
Step 1: the accuracy at target error by having distributed to control system, according to setting weight, by the accuracy at target of control system
Error distributes to steering wheel, missile-borne computer and the target seeker or receiver of control system, is used as accuracy at target caused by each link
Error；
Step 2: being used as input, missile position using target seeker or receiver error, missile-borne computer error, steering wheel error respectively
Change turns to output, derives target seeker, missile-borne computer, the transmission function of steering wheel itself error to missile position；
Step 3: according to obtaining, by the corresponding transmission function to missile position of each link, respectively obtaining target seeker or reception
Itself error of machine, missile-borne computer, steering wheel causes the steady-state gain that missile position changes；
Step 4: by accuracy at target error caused by each link of control system, lead respectively divided by caused by each link itself error
The steady-state gain of change in location is played, as missile control system target seeker, missile-borne computer, steering wheel own physical system is corresponding
Error.
2. a kind of each link error distribution method of missile control system as claimed in claim 1, it is characterised in that described leads
Take the lead error be target seeker receive laser beam information after, its export guided missile apart from optical axis position error, i.e. guided missile distance
The true value of optical axis position and difference of the target seeker reality output to the position of missile-borne computer.
3. a kind of each link error distribution method of missile control system as claimed in claim 1, it is characterised in that described bullet
Upper computer error is that missile-borne computer have received after the missile position information that target seeker is exported, by the processing of its internal calculation
The rudder command error of output.
4. a kind of each link error distribution method of missile control system as claimed in claim 1, it is characterised in that described rudder
The angle of rudder reflection error that chance error difference exports for steering wheel, i.e., after the rudder instruction that missile-borne computer is provided is received, steering wheel reality output rudder
Drift angle and the difference of theoretical output angle of rudder reflection.
5. each link error distribution method of a kind of missile control system as described in any one in claim 1-4, its feature
It is, according to separate random error distribution method, the accuracy at target error meets following allocation criterias：
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So, then σ is met2 c1+σ2 c2+σ2 c3=σ2 cs；Wherein | σcs| the accuracy at target for distributing to control system for overall system is missed
Poor index, | σc1|、|σc2|、|σc3| it is respectively the accuracy at target mistake that control system distributes to target seeker, missile-borne computer and steering wheel
Poor index, λc1, λc2And λc3The hit essence of control system always is accounted for for the accuracy at target error of target seeker, missile-borne computer and steering wheel
Spend the weight of error.
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| CN109270840A (en) * | 2018-09-28 | 2019-01-25 | 四川航天系统工程研究所 | A kind of missile control system time-variable correction network discretization method |
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| CN109270840B (en) * | 2018-09-28 | 2024-05-17 | 四川航天系统工程研究所 | Time-varying correction network discretization method for missile control system |
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